Abstract:
A rotor blade for a turbine engine includes a first side that defines a first contact face with a hardcoat and a second side that defines a second contact face without a hardcoat.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This disclosure was made with Government support under N00019-02-C-3003 awarded by The United States Navy. The Government has certain rights in this invention. 
    
    
     BACKGROUND 
     The present disclosure relates to a gas turbine engine, and more particularly to a blade thereof. 
     Gas turbine engines often include a multiple of rotor assemblies within a fan section, compressor section and turbine section. Each rotor assembly has a multitude of blades attached about a rotor disk. Each blade includes a root section that attaches to the rotor disk, a platform section, and an airfoil section that extends radially outwardly from the platform section. The airfoil section may include a shroud which interfaces with adjacent blades. In some instances, galling may occur on the mating faces of each blade shroud caused by blade deflections due to vibration. 
     SUMMARY 
     A rotor blade for a turbine engine according to an exemplary aspect of the present disclosure includes a first side that defines a first contact face with a hardcoat and a second side that defines a second contact face without a hardcoat. 
     A rotor assembly for a turbine engine according to an exemplary aspect of the present disclosure includes a plurality of adjacent blades, a first of said plurality of adjacent blades having a hardcoat on a first contact face in contact with a second contact face without a hardcoat on a second of the plurality of adjacent blades. 
     A method of manufacturing a rotor blade according to an exemplary aspect of the present disclosure includes hardcoating only one contact face of a rotor blade having a first side that defines a first contact face and a second side that defines a second contact face. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic illustration of a gas turbine engine; 
         FIG. 2  is a general perspective view of a disk assembly form a turbine sectional view of a gas turbine engine; 
         FIG. 3  is a side view of a shrouded turbine blade; 
         FIG. 4  is a suction side perspective view of the shrouded turbine blade; 
         FIG. 5  is a pressure side perspective view of the shrouded turbine blade; and 
         FIG. 6  is a perspective view of the disk assembly and three turbine blade shrouds. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  10  which generally includes a fan section  12 , a compressor section  14 , a combustor section  16 , a turbine section  18 , an augmentor section  20 , and an exhaust duct assembly  22 . The compressor section  14 , combustor section  16 , and turbine section  18  are generally referred to as the core engine. An engine longitudinal axis X is centrally disposed and extends longitudinally through these sections. While a particular gas turbine engine is schematically illustrated in the disclosed non-limiting embodiment, it should be understood that the disclosure is applicable to other gas turbine engine configurations, including, for example, gas turbines for power generation, turbojet engines, high bypass turbofan engines, low bypass turbofan engines, turboshaft engines, etc. 
     The turbine section  18  may include, for example, a High Pressure Turbine (HPT), a Low Pressure Turbine (LPT) and a Power Turbine (PT). It should be understood that various numbers of stages and cooling paths therefore may be provided. 
     Referring to  FIG. 2 , a rotor assembly  30  such as that of a stage of the LPT is illustrated. The rotor assembly  30  includes a plurality of blades  32  circumferentially disposed around a respective rotor disk  34 . The rotor disk  34  generally includes a hub  36 , a rim  38 , and a web  40  which extends therebetween. It should be understood that a multiple of disks may be contained within each engine section and that although one blade from the LPT section is illustrated and described in the disclosed embodiment, other sections will also benefit herefrom. Although a particular rotor assembly  30  is illustrated and described in the disclosed embodiment, other sections which have other blades such as fan blades, low pressure compressor blades, high pressure compressor blades, high pressure turbine blades, low pressure turbine blades, and power turbine blades may also benefit herefrom. 
     With reference to  FIG. 3 , each blade  32  generally includes an attachment section  42 , a platform section  44 , and an airfoil section  46  along a blade axis B. Each of the blades  32  is received within a blade retention slot  48  formed within the rim  38  of the rotor disk  34 . The blade retention slot  48  includes a contour such as a dove-tail, fir-tree or bulb type which corresponds with a contour of the attachment section  42  to provide engagement therewith. The airfoil section  46  defines a pressure side  46 P ( FIG. 5 ) and a suction side  46 S ( FIG. 4 ). 
     A distal end section  46 T includes a tip shroud  50  that may include rails  52  which define knife edge seals which interface with stationary engine structure (not shown). The rails  52  define annular knife seals when assembled to the rotor disk  34  ( FIG. 6 ; with three adjacent blades shown). That is, the tip shroud  50  on one blade  32  interfaces with the tip shroud  50  on an adjacent blade  32  to form an annular turbine ring tip shroud. 
     With reference to  FIGS. 4 and 5 , each tip shroud  50  includes a suction side shroud contact face  54 S and a pressure side shroud contact face  54 P. The suction side shroud contact face  54 S on each blade contacts the pressure side shroud contact face  54 P on an adjacent blade when assembled to the rotor disk  34  to form the annular turbine ring tip shroud ( FIG. 2 ). 
     In one non limiting embodiment, the blade  32  is manufactured of a single crystal superalloy with one of either the suction side shroud contact face  54 S or the pressure side shroud contact face  54 P having a hardface coating such as a laser deposited cobalt based hardcoat. That is, the hardface coating contacts the non-hardface coating in a shroud contact region defined by the suction side shroud contact face  54 S and the corresponding pressure side shroud contact face  54 P between each blade  32  on the rotor disk  34 . The suction side shroud contact face  54 S or the pressure side shroud contact face  54 P to which the hardface coating is applied may be ground prior to application of the hardface deposition or weld to prepare the surface and then finish ground after the application of the hardface to maintain a desired shroud tightness within the annular turbine ring tip shroud. 
     By reducing wear on the mating surfaces of a blade shroud, there is an increase in the functional life of the blade due to consistent blade damping. Applicant has determined that contact of dissimilar metals reduces wear and engine test confirmed less wear as compared to base metal on base metal and hardface coat on hardface coat interfaces. This is in contrast to conventional understanding of shroud contact faces in which each contact face is generally of the same material. 
     It should be understood that although a tip shroud contact interface is illustrated in the disclosed non-limiting embodiment, other contact interfaces such as a partial span shroud will also benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations are possible in light of the above teachings. Non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this disclosure. It is, therefore, to be understood that within the scope of the appended claims, the disclosure may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this disclosure.