Systems and methods of blade tip repair and manufacturing using field assisted sintering

A method of repairing, or manufacturing, a tip for an airfoil can comprise: removing a portion of the airfoil that reduces a radial height of the airfoil and defines a first mating surface of the airfoil; forming a blade tip body for the airfoil, the blade tip body defining a second mating surface; and joining the first mating surface to the second mating surface via a Field Assisted Sintering Technology (“FAST”) process.

FIELD

The present disclosure relates generally to blade tip repair systems and methods, and more particularly to, systems and methods for repairing a tip of an airfoil via field assisted sintering.

BACKGROUND

Gas turbine engines (such as those used in electrical power generation or used in modern aircraft) typically include a compressor, a combustor section, and a turbine. The compressor and the turbine typically include a series of alternating rotors and stators. A rotor generally comprises a rotor disk and a plurality of airfoils. The rotor may be an integrally bladed rotor (“IBR”) or a mechanically bladed rotor.

The rotor disk and airfoils in the IBR are one piece (i.e., monolithic, or nearly monolithic) with the airfoils spaced around the circumference of the rotor disk. Conventional IBRs may be formed using a variety of technical methods including integral casting, machining from a solid billet, or by welding or bonding the airfoils to the rotor disk.

High pressure turbine blade tips experience distress during service. This distress manifest itself in the form of tip rubs, oxidation, erosion, and thermal mechanical fatigue (TMF), impact damage and mechanical rub with the shroud. Often these failure modes are combined and can accelerate distress. Eventually the distress becomes severe enough where the tip condition may drive an engine off wing for repair/overhaul. At overhaul, engine run turbine blades are repaired and restored whenever possible. Repaired turbine blades are much more economical than buying new replacement parts. However, today's turbine blade tip repair techniques do not have the durability of newly manufactured parts and show decreased life the more repairs they have had. Most blade tips are limited to 1-3 repairs depending on their operational environment.

SUMMARY

A method of repairing a tip for an airfoil is disclosed herein. In various embodiments, the method comprises: removing a portion of the airfoil that reduces a radial height of the airfoil and defines a first mating surface of the airfoil; forming a blade tip body for the airfoil, the blade tip body defining a second mating surface; and joining the first mating surface to the second mating surface via a Field Assisted Sintering Technology (“FAST”) process.

In various embodiments, the first mating surface and the second mating surface are substantially complimentary in shape.

In various embodiments, the first mating surface and the second mating surface each comprises ribs and at least partially define cooling passages of a cooling network for the airfoil.

In various embodiments, the airfoil comprises a first metal alloy and the blade tip body comprises a second metal alloy, and wherein the first metal alloy is different from the second metal alloy.

In various embodiments, forming the blade tip body comprises casting the blade tip body from a metal alloy.

In various embodiments, removing the portion of the airfoil includes grinding the portion of the airfoil to remove a defect from the tip of the airfoil.

In various embodiments, the first mating surface and the second mating surface each comprise a flatness between 0.0001 inches and 0.01 inches.

In various embodiments, a surface roughness of the first mating surface and the second mating surface are each less than 16 micro inches.

A method of manufacture is disclosed herein. In various embodiments, the method comprises: forming a main airfoil body of a bladed rotor, the main airfoil body comprising a first metal alloy; forming a blade tip body for the bladed rotor, the blade tip body comprising a second metal alloy, the second metal alloy being different from the first metal alloy; and joining a first mating surface of the main airfoil body to a second mating surface of the blade tip body via a Field Assisted Sintering Technology (“FAST”) process.

In various embodiments, the first mating surface and the second mating surface each comprises ribs and at least partially define cooling passages of a cooling network for the airfoil.

In various embodiments, the method further comprises joining each blade tip body in a plurality of the blade tip body to a corresponding main airfoil body in a plurality of the main airfoil body via the FAST process.

In various embodiments, the first mating surface and the second mating surface are substantially complimentary in shape.

In various embodiments, the first metal alloy and the second metal alloy are both a single crystal alloy.

In various embodiments, forming the blade tip body includes casting a preform of the blade tip body.

In various embodiments, the first mating surface of the main airfoil body is disposed radially outward from a cooling passages network disposed within the main airfoil body.

A bladed rotor is disclosed herein. In various embodiments, the bladed rotor comprises: a hub; a rotor disk extending radially outward from the hub and defining a platform; and an airfoil extending radially outward from the airfoil, the airfoil comprising a main airfoil body made from a first metal alloy and a blade tip body made from a second metal alloy, the second metal alloy being different from the first metal alloy.

In various embodiments, the airfoil comprises a joint region disposed between the main airfoil body and the blade tip body, the joint region disposed radially outward from a cooling passages network in the main airfoil body.

In various embodiments, the bladed rotor comprises a plurality of the airfoil.

In various embodiments, the blade tip body and the main airfoil body each comprise a single crystal alloy.

In various embodiments, the rotor disk and the main airfoil body are mechanically coupled.

DETAILED DESCRIPTION

Disclosed herein is a blade tip that is build up using the Field Assisted Sintering Technology (“FAST”) process. FAST is a solid-state joining process which combines mechanical force, temperature, and electrical current to bond two mating surfaces via diffusion, such as a solid-state diffusion. The process can be used to join similar or dissimilar materials.

In various embodiments, a blade tip can be ground down to remove a defect, or defects, then a second blade tip (e.g., a crystal preform, such as a single crystal or polycrystal preform) can be bonded onto a mating surface of the grinded down blade to replace the blade tip. In various embodiments, the second blade tip is a single crystal preform. In various embodiments, the defect, or defects can be due to distress or oxidization. The single crystal material could be the same composition as the base material or a material with enhanced durability which is desirable at the tip due to environment and high local metal temperatures. The present disclosure is not limited in this regard.

In various embodiments, the systems and methods for blade tip repair disclosed herein can result in longer lasting, more durable blade tip repairs, longer time on wing for engines utilizing repaired blades. In various embodiments, the present disclosure is not limited to repair methods. In this regard, the process disclosed herein could be utilized for a newly manufactured part. For example, structural limitations for a tip of a blade can be limited based on oxidation, erosion, and thermal mechanical fatigue (“TMF”), impact damage and mechanical rub with the shroud in accordance with various embodiments. In contrast, other parts of a blade (e.g., a root of a blade) can be limited based on damage tolerance, high cycle fatigue stress, low cycle fatigue stress, or the like. In this regard, the process disclosed herein can facilitate a blade tip being made of a different material relative to a remainder of the blade, resulting in greater overall durability for the blade, in accordance with various embodiments.

With reference toFIG.1A, a gas turbine engine20is shown according to various embodiments. Gas turbine engine20may be a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26, and a turbine section28. In operation, fan section22can drive air along a path of bypass airflow B while compressor section24can drive air along a core flow path C for compression and communication into combustor section26then expansion through turbine section28. Although depicted as a turbofan gas turbine engine20herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures, single spool architecture or the like.

Gas turbine engine20may generally comprise a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure36or engine case via several bearing systems38,38-1, etc. Engine central longitudinal axis A-A′ is oriented in the Z direction on the provided X-Y-Z axes. It should be understood that various bearing systems38at various locations may alternatively or additionally be provided, including for example, bearing system38, bearing system38-1, etc.

Low speed spool30may generally comprise an inner shaft40that interconnects a fan42, a low pressure compressor44and a low pressure turbine46. Inner shaft40may be connected to fan42through a geared architecture48that can drive fan42at a lower speed than low speed spool30. Geared architecture48may comprise a gear assembly60enclosed within a gear housing62. Gear assembly60couples inner shaft40to a rotating fan structure. High speed spool32may comprise an outer shaft50that interconnects a high pressure compressor52and high pressure turbine54. A combustor56may be located between high pressure compressor52and high pressure turbine54. A mid-turbine frame57of engine static structure36may be located generally between high pressure turbine54and low pressure turbine46. Mid-turbine frame57may support one or more bearing systems38in turbine section28. Inner shaft40and outer shaft50may be concentric and rotate via bearing systems38about 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 may be compressed by low pressure compressor44then high-pressure compressor52, mixed and burned with fuel in combustor56, then expanded over high-pressure turbine54and low-pressure turbine46. Turbines46,54rotationally drive the respective low speed spool30and high-speed spool32in response to the expansion.

Referring now toFIG.1Band still toFIG.1A, according to various embodiments, each of low pressure compressor44, high pressure compressor52, low pressure turbine46, and high pressure turbine54in gas turbine engine20may comprise one or more stages or sets of rotating blades101and one or more stages or sets of stationary vanes102axially interspersed with the associated blade stages but non-rotating about engine central longitudinal axis A-A′. The compressor and turbine sections24,28may include rotor assemblies110. Each compressor stage and turbine stage may comprise multiple interspersed stages of blades101and vanes102. Within the rotor assemblies110of gas turbine engine20are multiple rotor disks, which may include one or more cover plates or minidisks. The blades101rotate about engine central longitudinal axis A-A′, while the vanes102remain stationary with respect to engine central longitudinal axis A-A′. For example,FIG.1Bschematically shows, by example, a portion of an engine section80, which is illustrated as a turbine section28of gas turbine engine20. It will be understood that the repair and/or manufacturing systems and methods in the present disclosure are not limited to the turbine section28and could extend to other sections of the gas turbine engine20, including but not limited to compressor section24.

Engine section80may include alternating rows of blades101and vanes102comprising airfoils100that extend into the core flow path C. The blades101may each include a bladed rotor100. In various embodiments, the bladed rotor100can comprise a mechanically bladed rotor (i.e., each airfoil103mechanically coupled to the rotor disk104). However, the present disclosure is not limited in this regard. For example, the bladed rotor100is an integrally bladed rotor, such that the airfoils103(e.g., blades) and rotor disks102are formed from a single integral component (i.e., a monolithic component formed of a single piece), in accordance with various embodiments.

The airfoils103extend radially outward from the rotor disk104. An outer engine case120is disposed radially outward from a tip of each airfoil103. The outer engine case120comprises an abradable material122disposed radially adjacent to the tip of each airfoil103. In this regard, the tip of each airfoil103comprises a coating, as described further herein, that includes an abrasive material. The abrasive material is configured to interface with the abradable material122of the outer engine case during operation of the gas turbine engine20. Initially, the abrasive material of the coating cuts into the abradable material, forming a trench, a recess, or the like. The coating is configured protect the tips of airfoils103for the bladed rotor100from burning up during operation of the gas turbine engine20.

Referring now toFIG.2, a perspective view of a bladed rotor200is illustrated in accordance with various embodiments. The bladed rotor200can be in accordance with any of the bladed rotor100fromFIG.1A. The present disclosure is not limited in this regard. The bladed rotor200comprises a hub202, a platform205, and a plurality of an airfoils206. Each airfoil206of the bladed rotor200extends radially outward from a respective platform205. For example, the airfoil206extends radially outward from a root212of the airfoil210to a tip214of the airfoil210. The root212can be integral with the platform205or coupled to the platform205as described previously herein. The present disclosure is not limited in this regard.

Referring now toFIG.3, a front view of a portion of a bladed rotor200is illustrated with like numerals depicting like elements, in accordance with various embodiments. In various embodiments, the airfoil210of the bladed rotor200comprises a main airfoil body302, and a blade tip body306. In various embodiments, the main airfoil body302comprises a first material and the blade tip body306comprises a second material. For example, as described previously herein, the structurally limiting features of the main airfoil body302(e.g., stress from high cycle fatigue, stress from low cycle fatigue, damage tolerance, or the like) can be different compared to the structurally limiting feature of the blade tip body306(e.g., oxidation, erosion, TMF, or the like). In this regard, the material for the main airfoil body302can comprise a greater Von Mises yield stress relative to the blade tip body306, in accordance with various embodiments. Thermal fatigue is a dramatically severe form of fatigue. Whereas normal high-cycle fatigue occurs at stresses comfortably in the elastic range (i.e., usually well below the yield point) thermal fatigue is driven by thermal strains that force deformation well into the plastic flow regime. The maximum stresses, as a consequence, are therefore well above the yield point. In this regard, thermomechanical fatigue, which is far more likely to occur at a tip214of an airfoil210, can be limited by tensile strain at low temperature changing to compressive (or lower tensile) strain at high temperatures. Thus, a material that has a lower Von Mises yield stress relative to the material of the blade tip body306can potentially handle thermal fatigue better, resulting in a more durable airfoil210relative to typical blades. In various embodiments, the tip body can be a second alloy that has better rub characteristics, improving sealing capability with the outer air seal rather than improving durability.

In various embodiments, the main airfoil body302can comprise a first alloy (e.g., a superalloy, such as a cobalt or a nickel-based alloy), and the blade tip body306can comprise a second alloy that is different from the first alloy (e.g., a superalloy, such as a cobalt or a nickel-based alloy). In various embodiments, the first alloy and the second alloy can both comprise a single crystal alloy; however, the present disclosure is not limited in this regard. In various embodiments, a single crystal alloy can be more robust relative to polycrystal alloys. In various embodiments, the first alloy can comprise between 56% and 62% nickel, between 4% and 12% chromium, between 1.5% and 6% molybdenum, between 5 and 15% cobalt, between 3 and 7% tungsten, between 4 and 12% tantalum, and between 3% and 7% aluminum (e.g., an alloy in accordance with PWA 1426, PWA 1429, PWA 1432, PWA 1437, PWA 1440, PWA 1449, PWA 1455, PWA 1475, PWA 1480, PWA 1483, PWA 1484, PWA 1487, PWA 663, PWA 1497, PWA 1499, or the like). In various embodiments, the second alloy can comprise between 54% and 72% nickel, between 4% and 20% chromium, between 0.0% and 10% molybdenum, between 7% and 19% cobalt, between 0% and 6% niobium, between 0% and 10% tungsten, between 0% and 10% titanium, between 0% and 8% tantalum, between 0% and 6% rhenium, between 0% and 7% aluminum, and between 0% and 2% hafnium (e.g., Rene 100, Rene 104, Rene 108, Rene 120, Rene 125, Rene 142, Rene 220, Rene 41, Rene 77, Rene 80, Rene 88, Rene 95, Rene N4, Rene N5, Rene N500, Rene 195, or the like). In various embodiments, the first alloy comprises PWA 1429 and the second alloy comprises Rene 195. However, the present disclosure is not limited in this regard. In various embodiments, the tip body can be a second alloy that has better rub characteristics, improving sealing capability with the outer air seal rather than improving durability.

In various embodiments, the main airfoil body302is cast from the first alloy and the blade tip body306is cast from the second alloy. In various embodiments, the main airfoil body302is integrally formed (i.e., formed as a single piece) with the remainder of the bladed rotor (e.g., the rotor disk204, the hub202fromFIG.2, etc.). However, the present disclosure is not limited in this regard.

In various embodiments, the main airfoil body302is coupled to the blade tip body306via the FAST process as described further herein. In various embodiments, the joining process may produce acceptable material properties in and around a joint region (e.g., region304) or threshold location of the first segment. In various embodiments, the processing order depends upon which path produces the desired combination of properties at, for example, both a joint and in the material of the main airfoil body302and the material of the blade tip body306. In various embodiments, a threshold location of the first segment may be determined to be, for example, a location with high structural margins and low stress. Such a location may be determined by performing a series of stress tests of the first segment to determine the location where the first segment may be coupled to the second segment. In various embodiments, the FAST process may be used to join the main airfoil body302to the blade tip body306. In various embodiments of the airfoil210, the heat treatment of the airfoil210occurs after a joining process. In various embodiments of the airfoil210, the heat treatment includes at least one of a solution heat treatment, a precipitation heat treatment, or stress relief. However, the present disclosure is not limited in this regard.

Referring now toFIG.4, a front view of a portion of a bladed rotor200is illustrated with like numerals depicting like elements, in accordance with various embodiments. In various embodiments, the airfoil210further comprises a cooling passage network402disposed through the main airfoil body302. In various embodiments, the region304joining the main airfoil body302and the blade tip body306can be disposed vertically above the cooling passage network402. In this regard, the joining region304can be a flat mating surface of the main airfoil body302being joined, via the FAST process, to a flat mating surface of the blade tip body306. In this regard, a strong joint can be created between the main airfoil body302and the blade tip body306, in accordance with various embodiments.

Referring now toFIG.5, a front view of a portion of a bladed rotor500is illustrated with like numerals depicting like elements, in accordance with various embodiments. In various embodiments, the bladed rotor500is at overhaul (or another maintenance interval). In this regard, the bladed rotor500comprises a airfoil210with a defect504proximate the tip214of the blade. In various embodiments, the defect504can extend radially inward along the airfoil210to a location prior to the cooling passage network402or past the cooling passage network402fromFIG.4. The repair processes disclosed herein are not limited in this regard as described further herein.

Referring now toFIGS.6and7, a flow chart for a repair process600(FIG.6), and illustrations of the repair process600(FIG.7) for repairing a tip of a blade for a bladed rotor is illustrated, in accordance with various embodiments. In various embodiments, the process600comprises removing a portion of a tip of an airfoil (step602). In various embodiments, removing the portion of the tip of the airfoil can comprise grinding the portion of the airfoil off the airfoil. In various embodiments, after step602, a mating surface702of an airfoil body701is created. In various embodiments, the mating surface702can be relatively flat and relatively smooth. For example, a surface roughness of the mating surface can be less than or equal to 16 microinches, or less than or equal to 8 microinches. In various embodiments, a flatness of the mating surface702can be between 0.0001 inches (0.0025 cm) and 0.01 (0.25 cm), or between 0.0001 inches (0.0025 cm) and 0.005 inches (0.013 cm). However, the present disclosure is not limited in this regard. In various embodiments, by having a relatively flat mating surface and a relatively smooth mating surface, a more robust joint can be generated from the solid-state joining process (i.e., the FAST process).

The FAST process is a low voltage, direct current (DC) pulsed current activated, pressure-assisted sintering, and synthesis technique. The FAST process can be used to synthesize adjacent components (e.g., the main airfoil body302and the blade tip body306fromFIG.6) and/or to densify materials in one step. Since the main airfoil body302and blade tip body306are electrically conductive, energy can be dissipated directly within the main airfoil body302and the blade tip body306and the electrically conductive parts of the pressing tool.

The FAST process comprises a mechanical loading system, which acts at the same time as high-power electrical circuit, placed in a controlled atmosphere. The electrical conductivity of the materials used for tooling facilitate low voltages (typically below 10 V applied to the whole set-up) and produce high currents (typically from 1 to 10 kA) leading to efficient heating. Even in the case of electrically non-conductive sintering powder, heat is quickly and efficiently transferred to the main airfoil body302and/or the blade tip body306described previously herein. Depending on the used hardware it is possible to define pulse and pause durations or more specialized pulse patterns. Typical pulse duration is in the order of a few milliseconds. Owing to the compact geometry of the die and punches, sintering cycles with heating rates as high as 1000° C. per minute are thus possible and enable to significantly reduce the total duration of the process and energy costs. Standard cooling rates up to 150° C. per minute are possible; additional active cooling under gas flow enable quenching rates of 400° C. min. At the same time, the simultaneous application of a uniaxial mechanical pressure enhances densification (maximal loads typically between 50 and 250 kN). The process can take place under vacuum or protective gas at atmospheric pressure: all heated parts are kept in a water-cooled chamber. Control of the processing cycle is usually done by temperature measurement (using either thermocouples or axial/radial pyrometers, etc.) but can also be achieved by other methods like power, current, or simply by displacement control. Maximal temperature achieved by using standard graphite tools lies up to 2400° C.

In various embodiments, in response to removing the portion of the airfoil in step602, a span (i.e., a radial length measured from the platform205) of the airfoil body701is reduced. With brief reference toFIGS.8A,8B, and8Ca mating surface812that is defined by a first cross-sectional profile810(FIG.8A), a mating surface822that is defined by a second cross-sectional profile820(FIG.8B), and a mating surface832that is defined by a third cross-sectional profile830(FIG.8C) are illustrated, in accordance with various embodiments. In various embodiments, the first cross-sectional profile810ofFIG.8Acorresponds to a mating surface812of an airfoil body701that is being repaired where the defect504fromFIG.5does not extend radially into the cooling passages network402fromFIG.4. In contrast, the second cross-sectional profile820ofFIG.8Bcorresponds to a mating surface702of an airfoil body701that is being repaired where the defect504fromFIG.5extends radially into the cooling passages network402fromFIG.4. Thus, the process600fromFIG.6is not limited based on a radial length of the defect504in accordance with various embodiments. For example, the FAST process disclosed herein can facilitate the joining of a blade tip body (e.g., blade tip body306fromFIG.3) to the airfoil body701being repaired via process600, in accordance with various embodiments. In various embodiments, the second cross-sectional profile820defines a mating surface822which is defined by cooling passages824, ribs826, and exterior walls828that define the pressure and suction surfaces of the airfoil. Although illustrated inFIG.8Bas including ribs826, the present disclosure is not limited in this regard. For example, a third cross-sectional profile830having a hollow cavity834disposed therethrough is also in accordance with the present disclosure.

Referring back toFIGS.6and7, the process600further comprises forming a blade tip body703(step604). In various embodiments, the blade tip body703is in accordance with the blade tip body306fromFIG.3described previously herein. However, the present disclosure is not limited in this regard. In various embodiments, the blade tip body306defines a mating surface (e.g., mating surface912fromFIG.9Ain response to the airfoil body701having the mating surface812ofFIG.8Aor mating surface922fromFIG.9Bin response to the airfoil body701having the mating surface822ofFIG.8B. In this regard, the mating surface912,922of the blade tip body703can be substantially complimentary to the mating surface812,822of the airfoil body701. “Substantially complimentary” as described herein refers to being within a 0.1 inch profile (0.25 cm profile) of the other mating surface, or being within a 0.05 inch profile (0.127 cm) of the other mating surface.

In various embodiments, step604further comprises casting the blade tip body703. However, the present disclosure is not limited in this regard. For example, the blade tip body703can be additively manufactured and still be within the scope of this disclosure.

In various embodiments, the process600further comprises aligning the mating surface812,822,832of the airfoil body701with the mating surface912,922,932of the blade tip body703(step606) and joining the airfoil body701to the blade tip body via a FAST process (step608). In this regard, as described previously herein, the blade tip body703can comprise a material configured to withstand limiting structural characteristics typical of a blade tip to extend a life of a repaired bladed rotor, in accordance with various embodiments. In this regard, step604can comprise casting the blade tip body from a super alloy that is different from a material of the airfoil body701. For example, the blade tip body703can comprise Rene 195, whereas the airfoil can comprise PWA 1429, in accordance with various embodiments. However, the present disclosure is not limited in this regard, and any alloy being coupled to an airfoil during a repair process as described herein is within the scope of this disclosure. For example, a super alloy that is configured for wear resistance or improved sealing with the shroud could be utilized and still be within the scope of the present disclosure.