Patent Description:
Modern gas turbine engines are subject to demanding operating conditions involving significant levels of force. Components of the gas turbine engines such as fan blade rotors can be subject to damage that can shorten operational life of the engine or the components thereof, or can require costly repair or remanufacture of the gas turbine engine or the components thereof. In some cases, damage to the component(s) can result in catastrophic failure of the component(s) and accompanying failure of the engine and damage to the aircraft and risk to flight operations. Previous attempts to manage this risk have involved inspection of the gas turbine engine and the components thereof in the hopes of identifying damage such as cracks in metal alloys before they propagate to the point of causing failure of the component(s). However, such inspection regimens may not identify all damage, or may identify damage after a point at which repair is possible, which can lead to costly replacement of the component(s).

<CIT> discloses a method of repairing a turbine engine component. The method includes the steps of preparing a repair surface on the turbine engine component, cold gas-dynamic spray of repair surface, vacuum sintering, and hot isostatic pressing followed by a heat treatment. <CIT> discloses a method of making an integrally bladed rotor composed of a titanium alloy. The method includes thermally processing a titanium alloy rotor disk and friction welding a blade to the rotor disk. <CIT> discloses a control system for determining a maintenance date for restoration of a coating of a component of an engine of an aircraft based on monitored parameters of the aircraft in use.

A method of servicing a gas turbine engine is disclosed. According to the method, a component comprising a titanium alloy is removed from the gas turbine engine after operating the gas turbine engine with the component in service. The removed component is subjected to heat treatment, and the heat-treated component is re-installed into the gas turbine engine or installed into a different gas turbine engine. The removing of the component from the gas turbine engine is performed in response to predetermined criteria of operating the gas turbine engine. The titanium alloy includes soft grains oriented for slip and hard grains not oriented for slip, and prior to the heat treatment the titanium alloy includes dislocations at boundaries between the soft grains and the hard grains. The heat treatment annihilates the dislocations at the boundaries between the soft grains and the hard grains. The heat treatment involves heating the component in a range of <NUM>°F to <NUM>°F (<NUM> to <NUM>) during <NUM> hour to <NUM> hours.

The data of the predetermined criteria of operating the gas turbine engine may be collected by a controller including a microprocessor operatively connected to sensors that monitor the predetermined criteria.

The predetermined criteria may include a cumulative time of operation of the gas turbine engine comprising said component.

The predetermined criteria may include a cumulative number of operation cycles of the gas turbine engine comprising said component.

The predetermined criteria may include a stress level applied to said component during operation of the gas turbine engine comprising said component.

The heat treatment may be performed below a beta transus temperature of the titanium alloy.

The component may be re-installed into the gas turbine engine or installed into a different gas turbine engine without mechanical repair of the component.

The component may be selected from a rotor hub or a bladed rotor hub.

Prior to the heat treatment the titanium alloy may include dislocations between metal grains in the titanium alloy.

The heat treatment may annihilate the dislocations in the metal grains of the titanium alloy.

The titanium alloy may include an alpha phase and a beta phase.

The titanium alloy may be selected from Ti-6Al-4V, Ti-6Al-6V-2Sn, Ti-6Al-2Sn-4Zr-2Mo, Ti <NUM> (<NUM>. 8Al-4Sn-<NUM>. 08C), Ti-<NUM> (6Al-<NUM>. 75Sn-4Zr-<NUM>. 45Si), Ti-<NUM>(8Al-1Mo-1V), Ti-<NUM> (6Al-5Zr-<NUM>. 25Si), or Ti-<NUM> (5Al-2Sn-2Zr-4Mo-4Cr).

As shown in <FIG>, an aircraft includes an aircraft body <NUM>, which can include one or more bays <NUM> beneath a center wing box. The bay <NUM> can contain and/or support one or more components of the aircraft <NUM>. Also shown in <FIG>, the aircraft includes one or more engines <NUM>. The engines <NUM> are typically mounted on the wings <NUM> of the aircraft and are connected to fuel tanks (not shown) in the wings, but may be located at other locations depending on the specific aircraft configuration.

The exemplary engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis. A relative to an engine static structure <NUM> via several bearing systems <NUM>.

In one disclosed embodiment, the engine <NUM> bypass ratio is greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>. 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.

The fan section <NUM> of the engine <NUM> is designed for a particular flight conditiontypically cruise at about <NUM> Mach (<NUM>/s) and about <NUM>,<NUM> feet (<NUM>,<NUM> meters). The flight condition of <NUM> Mach (<NUM>/s) and <NUM>,<NUM> ft (<NUM>,<NUM> 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 lbf of thrust the engine produces at that minimum point.

As mentioned above, components of a gas turbine engine can be removed from service and subj ected to heat treatment. In some aspects, the operation of the gas turbine engine in service can be on an aircraft, as operation of a gas turbine engine on an aircraft can expose engine component(s) to conditions that subject the component(s) to cold dwell fatigue as described in further detail below. In some aspects, the components to be serviced according to this disclosure include gas turbine engine rotor components. An example embodiment of a gas turbine rotor <NUM> is schematically shown in <FIG>. As shown in <FIG>, the example embodiment of a gas turbine rotor <NUM> includes a shaft <NUM>, a bladed hub or rim <NUM> that includes a hub or rim <NUM> and airfoil blades <NUM>. In some aspects, an engine component such as the gas turbine rotor <NUM> can be disassembled into sub-assemblies such as the shaft <NUM> and the bladed hub or rim <NUM>, followed by heat treatment of such sub-assemblies. In other aspects, the entire component (e.g., the entire gas turbine rotor <NUM>) can be subject to heat treatment. In still other aspects, only a portion of the engine component (e.g., areas around a weld joint) can be subjected to localized heat treatment. The component(s) to which heat treatment can be applied as disclosed herein can include various titanium alloys, including but not limited to multiphase titanium alloys such as alpha-beta titanium alloys, and single phase alloys such as alpha titanium alloys, or near-alpha titanium alloys. Examples of titanium alloys to which heat treatment can be applied include but are not limited to Ti-6Al-4V, Ti-6Al-6V-2Sn, Ti-6Al-2Sn-4Zr-2Mo. Ti <NUM> (<NUM>. 8Al-4Sn-<NUM>. 08C), Ti-<NUM> (6Al-<NUM>. 75Sn-4Zr-<NUM>. 45Si), Ti-<NUM>(8Al-1Mo-1V), Ti-<NUM> (6Al-5Zr-<NUM>. 25Si), or Ti-<NUM> (5Al-2Sn-2Zr-4Mo-4Cr).

The application of heat treatment involves heating the component(s) to a temperature in a range having a low end of <NUM>°F (<NUM>) and an upper end of <NUM>°F (<NUM>). The heat treatment temperature is maintained for a duration in a range of <NUM> hour to <NUM> hours.

The application of heat treatment can involve heating the component(s) to a temperature in a range having a low end of <NUM>°F (<NUM>), <NUM>°F (<NUM>), or <NUM>°F (<NUM>), and an upper end of <NUM>°F (<NUM>), <NUM>°F (<NUM>), or <NUM>°F (<NUM>). The above range endpoints can be independently combined to produce a number of different ranges, and every possible range that can be formed by combination of the above endpoints is hereby expressly disclosed. In some aspects, the heat treatment is performed at a temperature that is below the beta transus temperature of the titanium alloy. In some aspects, a heat treatment temperature can be maintained for a duration in a range having a low end of <NUM> hour, <NUM> hours, or <NUM> hours, and an upper end of <NUM> hours, <NUM> hours, or <NUM> hours. The above range endpoints can be independently combined to produce a number of different ranges, and every possible range that can be formed by combination of the above endpoints is hereby expressly disclosed. The time and temperature are interrelated, with greater amounts of time generally required at lower temperatures and lower amounts of time required at higher temperatures. For example, <NUM>-<NUM> hours may be sufficient at temperatures of <NUM>°F-<NUM>°F (<NUM> - <NUM>). Heating can be provided by exposing the component(s) to a heat source such by placing the component(s) in a furnace or by other heating techniques such as conductive heating. The heat treatment temperature is generally provided at a level that should not have a significant impact to alter the basic grain morphology of the titanium alloy, so quenching protocol will generally not be critical for performance, and cooling of the alloy can be accomplished in whatever manner may be efficient for processing operations. For example, cooling can be provided by deactivating whatever heat source was used for the heat treatment, or by removing the heat source or removing the component(s) from the presence of the heat source, and allowing the component(s) to cool naturally and/or by assisting with convective cooling such as contacting the components with moving air from an air source such as a fan.

In some aspects, the heat treatment can be applied at service intervals for the gas turbine engine in which the component(s) are operated. Various criteria can be used to identify an interval for removal of the component(s) from the gas turbine engine and application of heat treatment. In some aspects, removal of the component(s) for application of heat treatment can be performed in response to predetermined criteria based on operation of the gas turbine engine. For example, in some aspects, the predetermined criteria can include a cumulative time of operation of the gas turbine engine. For example, in some aspects, an interval based on a cumulative time of operation can be a regular fixed interval or can be a variable interval that can be based on other variables including but not limited to total lifetime hours of operation of the engine in service (e.g., longer intervals earlier in the life of the engine and shorter intervals later in the life of the engine), number of engine operating cycles (e.g., with a greater number of cycles prompting a shorter interval), and/or operating conditions such as speed or temperature as further described below. In some aspects, the predetermined criteria can include a number of engine operating cycles (e.g., a startup to shutdown cycle). In some aspects, a predetermined criteria based on operating cycles can be in a range having a lower end of <NUM> cycles, <NUM> cycles, or <NUM> cycles, and an upper end of <NUM> cycles, <NUM> cycles, or <NUM> cycles. The above range endpoints can be independently combined to produce a number of different ranges, and every possible range that can be formed by combination of the above endpoints is hereby expressly disclosed. In some aspects, the predetermined criteria, can include a stress and/or a temperature to which the component(s) are exposed during engine operation. Stress can be assessed indirectly, such as by monitoring engine speed or torque, both of which can be directly related to a level of stress on the engine components. For example, as described in more detail below, in some aspects the heat treatment can alleviate accumulated strain that may be influenced by speed or torque and/or temperature to which the components are subjected during engine operation. Any one or more of the above examples of predetermined criteria can be used by themselves or can be combined to identify a point at which the component(s) should be removed from operational service for application of heat treatment. For example, a prospective interval based on a cumulative number of hours of engine operation can be adjusted upward or downward based on whether the accumulated hours of engine operation or number of operating cycles occurred at engine speeds or torque and/or temperatures that were more or less likely to promote accumulation of strain in the titanium alloy. Data on any of the above criteria can be collected by a controller including a microprocessor operatively connected to sensors that can monitor the specified criteria. In some aspects, the microprocessor can be programmed with instructions for an algorithm that calculates an interval based on a variety of criteria, such as an algorithm in which points are accumulated based on operating conditions that can promote accumulation of strain, with higher point values applied to conditions that promote more strain accumulation (e.g., higher applied stress) and lower point values applied to conditions that promote less stress accumulation (e.g., lower applied stress). For example, an hour of operation at a low speed or torque may accumulate a specified point value whereas an hour of operation at a higher speed or torque may accumulate a higher point value, with heat treatment applied when accumulated points reach a predetermined threshold.

In some aspects, heat treatment can provide a technical effect of promoting annihilation of dislocations between grains in a titanium alloy's microstructure. Although this disclosure is not bound by any particular mechanistic theory, it is believed that such dislocations can form when a titanium alloy component is subjected to prolonged and/or cycled periods of stress at temperatures less than about <NUM>°F (<NUM>), leading to a phenomenon of cold dwell fatigue. Unchecked, these dislocations can glide through soft grains in the alloy that are oriented for slip and unload stress onto hard grains that are not oriented for slip. Hard grains can be characterized as grains where the c-axis of the HCP (hexagonal close packed) crystal is oriented <NUM>° to <NUM>° with respect to the stress axis, and soft grains can be characterized as grains where the c-axis of the HCP crystal is oriented <NUM>° to <NUM>° with respect to the stress axis. Accumulation of dislocations at the hard grains can result in localized fields of high stress between the anisotropic distributions of soft and hard grains, resulting in reduced fatigue tolerance and ultimately the formation of cracks that can propagate to the point of a catastrophic failure of the component. Such dislocations can form in titanium alloys including single-phase grain structures such as alpha or near-alpha grain structures, but also in multiphase titanium alloy grain structures such as two-phase alpha-beta grain structures that can be especially susceptible to this phenomenon at least in part because slip can cross boundaries between phases and also because the multiple phases can provide additional grain boundaries along which dislocations can accumulate.

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 claims.

Claim 1:
A method of servicing a gas turbine engine (<NUM>), comprising:
removing a component (<NUM>, <NUM>, <NUM>. <NUM>) comprising a titanium alloy from the gas turbine engine after operating the gas turbine engine with the component in service;
subjecting the component to heat treatment;
re-installing the component into the gas turbine engine or installing the component into a different gas turbine engine, wherein
the removing of the component (<NUM>, <NUM>, <NUM>, <NUM>) from the gas turbine engine (<NUM>) is performed in response to predetermined criteria of operating the gas turbine engine (<NUM>), characterized in that
the titanium alloy includes soft grains oriented for slip and hard grains not oriented for slip,
wherein prior to the heat treatment the titanium alloy includes dislocations at boundaries between the soft grains and the hard grains,
the heat treatment annihilates the dislocations at the boundaries between the soft grains and the hard grains, and
the heat treatment involves heating the component in a range of <NUM>°F to <NUM>°F (<NUM> to <NUM>) during <NUM> hour to <NUM> hours.