Patent Description:
Gas turbines are well-known in the art. It is an ongoing quest within the gas turbine field to increase the thermal efficiency of the gas turbine cycle. One way this has been accomplished is via the development of increasingly temperature-resistant materials, or materials that are able to maintain their structural integrity over time at high temperatures. For this reason, the hot gas path components of gas turbine engines are often formed from superalloy materials. The term 'superalloy' is used herein as it is commonly used in the art to refer to a highly corrosion and oxidation resistant alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, e.g., > <NUM>.

Despite their strength, superalloy components in the hot gas path of a turbine engine are susceptible to damage (defects) due to their long term exposure to significant thermal and mechanical stresses. It is generally known that superalloy materials are among the most difficult materials to repair. Welding of many superalloys, however, is difficult because of the propensity of these materials to develop weld solidification cracking and strain age cracking. Thus, repair processes for superalloy materials which eliminate welding while maintaining the structural integrity of the part are desired.

<CIT> discloses a method for weld-free repair of turbine diaphragms. <CIT> describes a method for weld-free repair of a root section of a turbine blade, in particular of the so called angel wings. The general structure of turbine blades, in particular having a squealer portion at their radially outer end, is described in <CIT> and <CIT>.

Briefly described, a first aspect of the present invention is related to a method for repairing a tip portion of a turbine blade in accordance with claim <NUM>.

The method allows for the use of boron in the braze material to reduce the temperature at which brazing is carried out.

A second aspect of the present invention is related to a pre-sintered preform in accordance with claim <NUM>.

Further preferred embodiments are defined by the dependent claims.

To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are explained hereinafter with reference to implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods.

The components and materials described hereinafter as making up the various embodiments are intended to be illustrative and not restrictive. Many suitable components and materials that would perform the same or a similar function as the materials described herein are intended to be embraced within the scope of embodiments of the present disclosure.

A gas turbine engine may comprise a compressor section, a combustor and a turbine section. The compressor section compresses ambient air. The combustor combines the compressed air with a fuel and ignites the mixture creating combustion products comprising hot gases that form a working fluid. The working fluid travels to the turbine section. Within the turbine section are circumferential alternating rows of vanes and blades, the blades being coupled to a rotor. Each pair of rows of vanes and blades forms a stage in the turbine section. The turbine section comprises a fixed turbine casing, which houses the vanes, blades and rotor.

The turbine blades include a radially inner root and a radially outer tip. The tip of a turbine blade can have a tip feature to reduce the size of the gap between ring segments and blades in the gas path of the turbine to prevent tip flow leakage, which reduces the amount of torque generated by the turbine blades. The tip features can be referred to as squealer tips and incorporated onto the tips of blades to help reduce aerodynamic losses between turbine stages. These features are designed to minimize the leakage between the blade tip and the ring segment.

Currently, structural defects affecting the tip area of a turbine component such as a blade or vane involve grinding and a weld build-up of the squealer tip utilizing a filler material. Additionally, cracks may be removed in the shelf portion of the tip and airfoil by weld repairing the damaged section with the filler material at ambient temperature or at elevated temperature using a hotbox weld repair process. Hot-box weld repairs may take eight hours or more to complete and the requirement for working inside of the hot box to maintain the elevated temperature makes it difficult to perform such welds.

Broadly, the inventor proposes a braze process utilizing a pre-sintered preform (PSP) having a varied composition for repairing a structural defect of a tip portion of a turbine blade. `Braze only' processes may be used to repair the tip damage of a turbine blade without the need for any weld repair process. Since no welding is needed, all the shortcomings of welding processes are eliminated - such as the need for overage heat treatment, the need for skilled welders, and heat affected zone cracks.

Pre-sintered preforms (PSPs) typically contain a powder mixture of base alloy particles and braze alloy particles that is pre-sintered so that the particles establish a metallurgical bond. Additionally, pre-sintered preforms do not include a binder material which creates voids. The pre-sintered material is formed in a net shape that may be used in a repair process such as that proposed. In the case of the proposed repair process, a composite boron base PSP of a turbine blade tip may be utilized.

Referring to <FIG>, a portion of a turbine engine <NUM> is shown. A centerline <NUM> is shown to represent an axial center of the turbine engine <NUM>. A radial direction Ra is shown in a direction that is radially outward. Further, a working fluid Wf direction is shown. A turbine blade <NUM> is formed from a root portion <NUM> coupled to a rotor disc (not shown) and an elongated portion forming an airfoil <NUM> that extends outwardly from a platform <NUM> coupled to the root portion <NUM>. At an opposite end of the turbine blade <NUM>, the blade <NUM> is composed of a tip <NUM> opposite the root section <NUM>, a leading edge <NUM>, and a trailing edge <NUM>. Connecting the leading edge <NUM> and the trailing edge <NUM> is radially extending a pressure side <NUM> and a suction side <NUM> of the airfoil <NUM>. Along the tip end <NUM> of the turbine blade <NUM> is a tip feature in position to reduce the size of the gap between ring segments <NUM> and blades <NUM> in a gas path of a turbine to prevent tip flow leakage, which reduces the amount of torque generated by the turbine blades <NUM>. The tip feature is referred to as a squealer or squealer tip and is incorporated onto the tips of blades to help reduce aerodynamic losses between turbine stages. These features are designed to minimize the leakage between the blade tip <NUM> and the ring segment <NUM>.

<FIG> represents a conventional squealer tip <NUM> location where spaced apart tip walls extend directly up from and extending the length of the pressure side <NUM> and suction side <NUM> of the blade <NUM>. At the tip end <NUM> lying in between the squealer tip walls <NUM> lies a tip shelf <NUM> having a tip shelf surface <NUM>.

As noted above, it is appreciated that during operation the blades, particularly in the early stages of the turbine engine, may be susceptible to significant thermal and mechanical stresses. Accordingly, particularly with some superalloys, it is common to see cracking and other defects develop on the tip of the blade, particularly in the squealer tip walls <NUM> and shelf area <NUM> of the tip <NUM>. <FIG>, for example, illustrates the tip <NUM> comprising cracks (discontinuities) <NUM> extending into the squealer tips <NUM> and in the shelf region <NUM> of the tip. While cracking is shown, other defects such as squealer tip rub or shelf rub may also be considered defects needing repair.

The turbine blade may comprise any suitable metal material. In an embodiment, the turbine blade may comprise a superalloy material. Exemplary superalloys include but are not limited to Hastelloy, Inconel (e.g. IN100, IN600, IN713), Waspaloy, Rene alloys, Haynes alloys, Incoloy, MP98T, TMS alloys, and CMSX (e.g. CMSX-<NUM>) single crystal alloys. In a particular embodiment, the turbine component is formed from an Alloy <NUM> material (a CM247 or MAR-M247 material as is known in the art and commercially available from Praxair Surface Technologies). In an embodiment, the Alloy <NUM> material may have a composition within the following ranges (in wt.

<FIG> illustrates a composite PSP tip coupon <NUM> that will be used to repair a turbine blade. The composite tip coupon <NUM> includes a tip shelf (cap) portion <NUM> and a squealer portion <NUM>. The tip shelf portion <NUM> comprises a first composition and the squealer portion <NUM> comprises a second composition. The first composition and the second composition are different. Each portion of the PSP may comprise a powder mixture comprising braze particles and superalloy particles formed into the tip shelf shape and a squealer tip shape, respectively, and configured to mate to a remaining portion of a turbine blade such as a blade airfoil. In an embodiment, a first thickness of the tip shelf portion <NUM> may be in a range <NUM> in. to <NUM> in. (<NUM>,<NUM> to <NUM>,<NUM>) while a second thickness of the squealer portion <NUM> may be in a range <NUM> in. to <NUM> in. (<NUM>,<NUM> to <NUM>,<NUM>). If the blade shelf is very thin, for example <NUM> in. (<NUM>,<NUM>), its base thickness may be increased by attaching additional shelf PSP material <NUM> to the existing shelf portion before attaching the composite coupon.

The braze material may comprise any suitable material known in the art for brazing which contains at least an amount of boron effective to reduce a melting temperature of the braze material relative to the same braze material without an amount of boron. In an embodiment, the amount of boron may be an amount of boron effective to reduce a melting temperature of the braze material to a desired degree. In a particular embodiment, the braze material comprises an amount of boron plus a first powder material including the same alloy components as that in the damaged area of the blade to be brazed/repaired. Suitable braze material compositions may be found in <CIT>, Brazing of Superalloy Components with Hydrogen Addition for Boron Capture.

Referring now to <FIG>, a method for repairing a tip portion of a turbine blade having a structural defect is presented. The turbine blade <NUM> is prepared for a repair process by first removing the damaged tip portion <NUM> from the remaining turbine blade airfoil <NUM>. Removing the damaged section entails machining and/or cutting at least the squealer portion <NUM> of the turbine blade airfoil <NUM> having the damaged section. Additionally, if the shelf section <NUM> includes wide cracks, e.g. cracks wider than approximately <NUM> inch (<NUM>,<NUM>), the cracks may be filled with paste. When the turbine blade <NUM> comprises Alloy <NUM>, for example, the paste may be a <NUM> paste which may be formed by mixing Alloy <NUM> with a suitable binder. In an embodiment, the upper surface of the remaining blade airfoil <NUM> may then be machined, for example, to produce a smooth, flat surface in order to mate with a surface of the composite PSP tip portion <NUM> which will replace the removed damaged tip section. One difference between this process and other traditional processes is that the paste application need only be utilized when a thru crack exists in the shelf portion. Otherwise, the PSP composition during brazing may take care of the defect, i.e. the braze material will flow into and fill the cracks. This in turn allows for a complete repair and facilitates determining where the paste is needed.

Prior to paste application and brazing, optionally, an area including the structural defect of the blade may be cleaned. In an embodiment, the cleaning step may be carried out using, fluoride ion cleaning (FIC). In a particular embodiment, the damaged area, including the defect may be cleaned via fluoride ion cleaning (FIC) process to ready the damage surface for brazing. In some situations, cracks may need to be physically opened up prior to FIC process. In some embodiments, the FIC process includes cleaning with hydrogen fluoride gas. Use of FIC cleaning advantageously removes unwanted oxides and residual coating remnants (e.g., diffusion coating remnants) within the defects, as well as on a surface of the blade.

In accordance with the invention, a surface of the PSP tip coupon <NUM> is applied to a prepared surface of the remaining blade airfoil <NUM>. The PSP tip coupon <NUM> may be lightly affixed by spot welding to the surface of the blade shelf surface <NUM>, particularly by spot welding in only one or two locations in order to avoid constraining the PSP tip coupon <NUM> or to avoid having it pop off during brazing. A gap may exist between the surface to be brazed and the PSP tip coupon <NUM>. The PSP tip coupon <NUM> will form to the surface being brazed during a brazing heat treatment cycle.

Once the PSP tip coupon <NUM> containing braze material has been applied as desired or necessary, the turbine blade along with the PSP is subjected to a heat treatment (referred to herein as 'brazing' or a `brazing process') in order to at least melt the braze material and allow the molten braze material to flow into the defect. In an embodiment, the brazing heat treatment may be a controlled heat process as described in <CIT>, Brazing of Superalloy Components with Hydrogen Addition for Boron Capture. The main difference between this heat treatment and other brazing heat treatment processes is that the component, namely the turbine blade and the affixed PSP, are subjected to alternating stages within a hydrogen environment (`hydrogen stage") and within a vacuum environment ("vacuum stage") while heating the braze material and at least a portion of the turbine blade. This alternation process <NUM> may be seen in the Braze Heat Treatment Chart seen in <FIG>. Alternating between a hydrogen stage and a vacuum stage helps to eliminate centreline eutectics and improve remelt properties post braze due to deboronization of the low melt portion of the braze mixture. Elimination of centreline eutectics avoids brittle zones in the repair area which tend to crack during service conditions.

<FIG> illustrates a perspective view of a turbine blade <NUM> repaired utilizing a composite PSP coupon <NUM>. <FIG> shows the tip <NUM> after the brazing. In certain embodiments, finishing processes such as blending, machining, cooling hole drilling and coating operations may be performed as applicable depending on the desired final blade configuration.

Claim 1:
A method for repairing a tip portion (<NUM>) of a turbine blade (<NUM>) having a structural defect (<NUM>), comprising:
providing a turbine blade (<NUM>) with a structural defect (<NUM>) in a tip portion (<NUM>) of the turbine blade (<NUM>);
removing the damaged section by machining or cutting the tip portion (<NUM>) of the turbine blade (<NUM>);
providing a pre-sintered preform (<NUM>) configured to mate with an upper surface of a remaining portion of the turbine blade (<NUM>);
applying the pre-sintered preform (<NUM>) to the upper surface, wherein the pre-sintered preform (<NUM>) comprises a superalloy material and a braze material; and
subjecting the pre-sintered preform (<NUM>) and the remaining portion of the turbine blade (<NUM>) to a brazing process to melt the braze material and fill in the structural defect (<NUM>),
wherein the pre-sintered preform (<NUM>) comprises a first portion (<NUM>) having a first composition and a second portion (<NUM>) having a second composition,
and wherein the first portion corresponds to a shelf portion (<NUM>) of the tip and the second portion corresponds to a squealer portion (<NUM>) of the tip,
and wherein the first composition and the second composition are different.