Abstract:
An identification feature is used to unmistakably identify internal features present in different generations of turbine blade designs. The identification feature is located on a root portion of the turbine blade and protrudes to provide a visually identifiable feature that is also readable by a coordinate measuring machine, but does not interfere with installation or operation of the turbine blade. The weight of the identification feature is in a specific proportion to the weight of the turbine blade in order to prevent interfere with operation of the turbine blade during high-speed rotation in a gas turbine engine.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines and more particularly to identification of turbine blades having internal features. In gas turbine engines, fuel is combusted in compressed air created by a compressor to produce heated gases. The heated gases are used to turn turbine blades, or airfoils, to produce rotational power for, among other things, operating the compressor. During operation of the gas turbine engine, temperatures inside the combustion chamber can reach 2500° F., resulting in the blades being subject to temperatures in excess of 1700° F. In order to cool the turbine blades, relatively cooler compressed air that bypasses the combustion chamber, or bleed air, is forced through internal passages of the blades. The passages include pathways or channels having various geometries in order to direct the bleed air throughout the interior of the blade. The bleed air flowing through the passages maintains a temperature gradient throughout the entirety of the blade at which the blade can properly function. 
   For performance or manufacturing reasons, it is sometimes necessary to change or modify the interior features of a particular blade model. Meanwhile, the exterior of that blade must be maintained the same in order to meet the design of the specific gas turbine engines in which that model of blade is used. Traditionally, a model number that identifies the interior features of the turbine blade is cast on the exterior of the turbine blade casting. The cast model numbers produce a shallowly indented number on the surface of the turbine blade. The shallow numbers do not create any protrusions or cavities that upset the balance of the blade while it is rotating. Any, even small, disproportion of weight along the length of the turbine blade can produce vibrations during the high-speed rotations produced in gas turbine engines. 
   While the cast model numbers are small enough to prevent any interference with the operation or installation of the blade, the numerals are often illegible and confusingly similar. For example, a cast “9” may look like a “0.” Thus, a turbine blade having second generation internal features would look identical to a turbine blade having first generation internal features, and there would be no positive way to identify which generation of internal features it possesses. Therefore, blades could be improperly introduced into the production stream where they would receive incorrect finishing procedures that are not discovered until a later time. It is desirable for production cost and safety considerations to completely eliminate the possibility of these mistakes. There is, therefore, a need for a turbine blade having an identification feature that unmistakably identifies the internal features of visually identical turbine blades without interfering with the operation of the blade itself. 
   BRIEF SUMMARY OF THE INVENTION 
   The present invention is directed towards a positive identification feature used to identify internal features of turbine blades. The invention comprises a protruding identification that unmistakably identifies the internal features of the turbine blade. The protruding identification feature is visually identifiable and readable by a coordinate measuring machine. The protruding identification feature is located on a root portion of the turbine blade so as to prevent interference with installation of the turbine blade. The protruding identification feature weighs approximately 0.1% or less of the weight of the turbine blade in order to prevent interference with operation of the turbine blade. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  shows a partially cut away perspective view of a gas turbine engine showing a turbine section in which the present invention is used. 
       FIG. 2  shows a partially exploded perspective view of the turbine section of  FIG. 1  showing a turbine blade and rotor disc assembly. 
       FIG. 3  shows the root section of a turbine blade having first generation internal features. 
       FIG. 4  shows the root section of a turbine blade having second generation internal features in which the present invention is used. 
   

   DETAILED DESCRIPTION 
     FIG. 1  shows a partially cut away perspective view of gas turbine engine  10  showing turbine section  12  in which the present invention is used. Gas turbine engine  10  includes turbine section  12 , which is positioned between combustion chamber  14  and nozzle  16 . Casing  18  shrouds turbine section  12 , combustion chamber  14  and nozzle  16 . Turbine section  10  is a multi-stage turbine and includes turbine blades  20 A,  20 B and  20 C; rotor discs  22 A,  22 B and  22 C and turbine shaft  24 . Turbine blades  20 A,  20 B and  20 C are radially aligned around the periphery of rotor discs  22 A,  22 B and  22 C, respectively. Rotor discs  22 A,  22 B and  22 C are co-axially attached to turbine shaft  24 , which extends in an axial direction into gas turbine engine  10 . 
   Fuel is combusted in high-pressure air inside combustion chamber  14  in order to produce heated gases having high density and high pressure. As the heated gases exit combustion chamber  14 , they enter turbine section  12  at a high velocity. The high-density gases impinge on turbine blades  20 A,  20 B and  20 C to produce rotational movement of rotor discs  22 A,  22 B and  22 C, which in turn rotate turbine shaft  24 . Rotation of turbine shaft  24  produces mechanical power for driving the compressor section of gas turbine engine  10 . The heated gases continue through turbine section  12  and are forced through nozzle  16 . Nozzle  16  increases the velocity of the gases as they exit gas turbine engine  10  in order to produce forward thrust for propelling an aircraft. 
     FIG. 2  shows a partially exploded perspective view of cut away turbine section  12  of  FIG. 1  showing the assembly of turbine blades  20 A and rotor disc  22 A. Turbine blades  20 A are radially arranged around the outer circumference of rotor disc  22 A. Turbine blades  20 A include foil  26 , shroud  28 , platform  30  and root  32 . Rotor disc  22 A includes slots  34  aligned along the outer circumference of rotor disc  22 A. Slots  34  receive roots  32  of turbine blades  20 A. Slots  34  include serrations  36 , and roots  32  include tangs  38  having a matching profile with that of serrations  36 . In typical embodiments, roots  32  have a “fir tree” or “dove tail” configuration as is known in the art. Roots  32  are inserted into slots  34  the axial direction so tangs  38  are locked into serrations  36 . Tangs  38  and serrations  36  secure turbine blade  20 A in the radial direction during rotation of rotor disc  22 A and distribute the load produced by the centrifugal momentum of rotating turbine blade  20 A. Serrations  36  and tangs  38  also allow for thermal expansion of roots  30  and rotor disc  22 A in the extreme temperatures reached in gas turbine engine  10 . Additionally, rivets or other fastening mechanisms are used to hold turbine blades  20 A in the axial direction. 
   When turbine blades  20 A are inserted into rotor disc  22 A, shrouds  28  align to form a continuous barrier that assists in preventing gas leakage around the tips of the turbine blade. Shrouds  28  also prevent vibration and bending of foils  26 . In other embodiments, shrouds  28  are not used and the blade tips of foils  26  are cut to a knife-edge tip. Similarly, platforms  30  align to form a continuous boundary between turbine blades  20 A and roots  30 . 
   Typically, bleed air used for cooling turbine blades  20 A is introduced through an opening located on the bottom of root  32 , whereby it enters passages of an interior cooling system. The interior cooling system includes various features and passages in which the bleed air flows. The bleed air travels through the passages on the interior of turbine blade  20 A and whisks heat away from foil  26 . Typically, the heated bleed air exits the interior of turbine blade  20 A through one or more small orifices  40  located on the trailing edge of foil  26  or on the concave suction side (not shown) of foil  26 . 
     FIG. 3  shows the root section of turbine blade  20 A having first generation internal features. For a particular turbine blade design, changes to the interior features may occur mid-production to increase performance of the blade. However, the exterior of every generation of turbine blade  20 A is identical to each other, thereby producing an interchangeable part that will always fit in the gas turbine engines it was designed for use in. 
     FIG. 4  shows root section  32  of turbine blade  20 A′ having second generation, or post-modification, interior features in which the present invention is used. Once a change has been made to the interior design of the model of turbine blade comprising turbine blade  20 A, identification feature  42  is added to root section  32  to produce turbine blade  20 A′. Identification feature  42  provides a mistake proof means for distinguishing turbine blade  20 A from  20 A′. 
   Identification feature  42  provides a positive, raised protuberance that can be recognized by visual inspection. Identification feature  42  also provides a feature that can be measured with a Coordinate Measuring Machine (CMM). During manufacture of turbine blade  20 A′ the blade is inspected for dimensional tolerances before being sent for additional machining procedures. Identification feature  42  provides a positive feature that can be included in the dimensional tolerance checklist and checked for with the CMM. This ensures that the turbine blade being inspected is in fact turbine blade  20 A′ and that it will receive machining procedures intended for blades with second generation internal features. 
   The location of identification feature  42  is selected to not interfere with the operation of turbine blade  20 A′. For example, it is unfeasible to put an identifying mark on foil portion  26  because that would interfere with impingement of the hot air on foil  26  and would cause vibration of foil  26 . For similar reasons, it would be unfeasible to put an identifying feature on shroud  28  or platform  30 . Also, it is impracticable to put an identifying feature in the sides of root portion  32  because that would interfere with alignment of serrations  36  and tangs  38 . Considering these factors, identification feature  42  is placed on front surface  44  of root portion  32 . In other embodiments, identification feature  42  is placed on the rear surface of root portion  32 . In  FIG. 4  identification feature  42  is placed on root portion  32  off-center of front surface  44 . This moves identification feature away from the parting line of the casting of turbine blade  20 A′ and allows the mold for turbine blade  20 A to be adapted for forming turbine blade  20 A′. In other embodiments, identification feature  42  is centered on front surface  44  of root portion  32 . Placing identification feature  42  on root portion  32  also minimizes the vibration effect caused by identification feature  42  on foil  26 . 
   To further prevent identification feature  42  from interfering with operation and installation of turbine blade  20 A′, identification feature  42  is placed in recess  46  located on front surface  44  of root portion  32 . Recess  46  is pre-formed into the casting of turbine blade  20 A′ for weight reduction purposes or other functional purposes. Additionally, recess  46  can be machined into turbine blade  20 A′ for the purposes of receiving identification feature  42 . Thus, in order to minimize the interference of identification feature  42  on the installation and operation of turbine blade  20 A′, identification feature  42  does not extend beyond the forward most portion of the leading edge of root portion  32 . 
   During operation of gas turbine engine  10 , rotor disc  22 A rotates at speeds of approximately 15000 revolutions per minute (RPM). During these high-speed rotations the tangential velocity of the tips of turbine blade  20 A′ can reach speeds up to Mach 2. Thus, placing even a small amount of mass on turbine blade  20 A′ creates a large force that will interfere with true rotation of rotor disc  22 A and foil  26 . The centrifugal force generated by the mass of identification feature  42  has the potential to create vibrations in the rotation of turbine blade  20 A′. When the centrifugal forces exert stresses beyond the stress limits of turbine blade  20 A′, especially compounded with resonance vibration, catastrophic failure of turbine blade  20 A′ will occur. 
   Using standard mechanics formulas, the size and mass of an identification feature  42  that will not cause excessive stresses in turbine blade  20 A′ can be determined. It has been determined that when placing identification feature  42  on root portion  32 , an identification feature weighing approximately 0.1% or less of the total weight of turbine blade  20 A will not produce excessive stresses in turbine blade  20 A′. Therefore, in one embodiment, identification feature  42  weighs 0.1% of turbine blade  20 A′. For example, if turbine blade  20 A′ weighs 0.84 lbs., identification feature  42  weighs approximately 0.0008 lbs. or less. This prevents excessive stresses in and vibration of turbine blade  20 A′ during high-speed rotation of rotor disc  22 A during operation of gas turbine engine  10 . 
   The specific shape of identification feature  42  can have various embodiments. In  FIG. 4 , identification feature  42  is a vertical rib. An additional vertical rib identification feature  42 , or a differently shaped identification feature  42 , can be added to identify each subsequent generation of turbine blade  20 A. In various embodiments, identification feature  42  can be circular, star shaped or triangular. The size and shape of each identification feature, or the plurality of identification features, is limited by being maintained at or below approximately 0.1% of the weight of turbine blade  20 A to prevent perturbation of turbine blade  20 A′ during rotation of rotor disc  22 A. The size and shape of identification feature is also limited because it must not interfere with the installation of turbine blade  20 A′. 
   The present invention has been described as applied to turbine blades used in the turbine section of a gas turbine engine. The protruding identification feature can also be used in rotor blades used in the compressor section of gas turbine engines or in other rotating foils or blades having varying interior features. 
   Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.