Abstract:
A method of designing a turbine blade includes the steps of forming at least two notches on a tip of a turbine blade, each of the at least two notches having a known dimension. The turbine blade has a pressure side and a suction side. The method further includes the step of operating a gas turbine engine including the turbine blade to expand a length of the turbine blade such that the tip of the turbine engages a casing. The method further includes the steps of viewing the tip of the turbine blade after the step of operating of the gas turbine engine, determining an appearance of the notches on the tip and determining a manufacturing length of the turbine blade based on the step of determining the appearance the notches.

Description:
BACKGROUND OF THE INVENTION 
     This application relates generally to a method of measuring tip erosion of a turbine blade during development and testing of the turbine blade. 
     During operation of a gas turbine engine, a turbine blade can tilt or expand due to creep (because of temperature and centrifugal forces). When a tip of the turbine blade rubs against a casing of the gas turbine engine, the tip can erode over time. It is important for the turbine blade to have a proper length to reduce wear at the tip while still providing a seal between the tip and the casing. During development of the gas turbine engine and the turbine blade, the gas turbine engine must be disassembled to access the hardware and the turbine blade to measure and determine any erosion, rub and tilt of the tip of the turbine blade, which is costly. 
     In one prior gas turbine engine, a seal serration part at a tip of a turbine blade includes a single notch. Over time and during normal operation of the gas turbine engine, the seal serration part rubs against a case to wear the seal serration part until the notch is eventually eliminated from the tip. When it is visually determined that the notch is eliminated, this indicates that the turbine blade is approaching fracture due to creep and must be replaced. 
     SUMMARY OF THE INVENTION 
     A method of designing a turbine blade includes the steps of forming at least two notches on a tip of a turbine blade, each of the at least two notches having a known dimension. The turbine blade has a pressure side and a suction side. The method further includes the step of operating a gas turbine engine including the turbine blade to expand a length of the turbine blade such that the tip of the turbine engages a casing. The method further includes the steps of viewing the tip of the turbine blade after the step of operating of the gas turbine engine, determining an appearance of the notches on the tip and determining a manufacturing length of the turbine blade based on the step of determining the appearance the notches. 
     A turbine blade includes a tip and at least two notches formed on the tip. Each of the least two notches have a known dimension. The turbine blade has a pressure side and a suction side. 
     A gas turbine engine assembly includes a casing including a hole and a turbine blade including a tip and at least two notches formed on the tip. Each of the at least two notches have a known dimension, and the turbine blade has a pressure side and a suction side. A borescope is inserted through the hole in the casing to view the notches on the tip. 
     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a simplified cross-sectional view of a standard gas turbine engine; 
         FIG. 2  illustrates a turbine blade with two notches formed on a tip; 
         FIG. 3  illustrates a turbine blade with multiple notches formed on the tip; and 
         FIG. 4  illustrates a turbine blade after operation of the gas turbine engine. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     As shown in  FIG. 1 , a gas turbine engine  10 , such as a turbofan gas turbine engine, is circumferentially disposed about an engine centerline (or axial centerline axis  12 ). The gas turbine engine  10  includes a fan  14 , a low pressure compressor  16 , a high pressure compressor  18 , a combustion section  20 , a high pressure turbine  22  and a low pressure turbine  24 . This application can extend to engines without a fan, and with more or fewer sections. 
     Air is pulled into the gas turbine engine  10  by the fan  14  and flows through a low pressure compressor  16  and a high pressure compressor  18 . Fuel is mixed with the air, and combustion occurs within the combustion section  20 . Exhaust from combustion flows through a high pressure turbine  22  and a low pressure turbine  24  prior to leaving the gas turbine engine  10  through an exhaust nozzle  25 . 
     As is known, air is compressed in the compressors  16  and  18 , mixed with fuel, burned in the combustion section  20 , and expanded in the turbines  22  and  24 . Rotors  26  rotate in response to the expansion, driving the compressors  16  and  18  and the fan  14 . The compressors  16  and  18  include alternating rows of rotating compressor blades  28  and static airfoils or vanes  30 . The turbines  22  and  24  include alternating rows of metal rotating airfoils or turbine blades  32  and static airfoils or vanes  34 . It should be understood that this view is included simply to provide a basic understanding of the sections in a gas turbine engine  10  and not to limit the invention. This invention extends to all types of gas turbines for all types of applications, in addition to other types of turbines, such as vacuum pumps, air of gas compressors, booster pump applications, steam turbines, etc. 
       FIG. 2  illustrates a turbine blade  32 . The turbine blade  32  includes a root  48  received in a rotor disk (not shown), a platform  64 , an airfoil  50 , and a tip  42 . The turbine blade  32  includes a leading edge  52  and a trailing edge  54 . The turbine blade  32  also has a pressure side  56  and a suction side  58 . 
     Prior to operation of the gas turbine engine  10 , there is a gap between the tip  42  of the turbine blade  32  and the casing  36 . During operation of the gas turbine engine  10 , the turbine blades  32  expand due to heat and centrifugal forces such that the tip  42  rubs the casing  36 , creating a seal. However, if the turbine blade  32  expands too much due to creep, the tip  42  can erode and wear. The turbine blade  32  can also tilt, causing a different amount of erosion and wear on either the pressure side  56  or the suction side  58  of the tip  42  of the turbine blade  32 . 
     During the developmental and testing phase of the gas turbine engine  10  and the turbine blade  32 , at least two notches  60  of known depth are formed on the tip  42  of the turbine blade  32 . In one example, one of the at least two notches  60  is formed on the pressure side  56 , and the other of the at least two notches is formed on the suction side  58  (as shown in  FIG. 2 ). In another example, the least two notches  60  are both formed on the pressure side  56  or are both formed on the suction side  58 . Alternately, a plurality of notches  60  can be formed on both the pressure side  56  and the suction side  58  (as shown in  FIG. 3 ). 
     During development and testing of the gas turbine engine  10 , the at least two notches  60  function as wear indicators that indicate how much wear occurs on the tip  42  of the turbine blade  32  during testing. Based on the data obtained from the wear of the at least two notches  60 , the turbine blade  32  can be designed to have a specific length based on expected expansion and wear due to creep and tilt to ensure that there is optimal contact between the turbine blade  32  and the casing  36  during operation of the gas turbine engine  10  to create a seal while reducing wear. 
     In one example, the at least two notches  60  are machined. In one example, the at least two notches  60  are semi-circular in shape. The semi-circular shape minimizes stress concentration. 
     In the example shown in  FIG. 3 , notches  60  having various radii are formed on the tip  42  of the turbine blade  32 . The notches  60  are shown for illustrative purposes only and are not shown to scale. In one example, closest to the leading edge  52 , a set of notches  60   a  and  60   b  is formed on the pressure side  56  and the suction side  58  of the turbine blade  32 , respectively. Another set of notches  60   c  and  60   d  is formed closer to the trailing edge  54  on the pressure side  56  and the suction side  58  of the turbine blade  32 , respectively. Another set of notches  60   e  and  60   f  is formed even closer to the trailing edge  54  than the set of notches  60   c  and  60   d  on the pressure side  56  and the suction side  58  of the turbine blade  32 , respectively. The location and the radius of each of the notches  60   a ,  60   b ,  60   c ,  60   d ,  60   e  and  60   f  on the tip  42  of the turbine blade  32  are a function of design. 
     The turbine blade  32  in the developmental stage has a length L that is slightly longer than that the expected length of the final design of the turbine blade  32 . In one example, the middle notches  60   c  and  60   d  each have a radius that is equal to the amount of wear that is expected when the gas turbine engine  10  is tested. That is, once the gas turbine engine  10  is tested, it is expected that the material above the notches  60   c  and  60   c  will be rubbed away such that the bottom of the notches  60   c  and  60   d  now define the tip  42 . The length L of the turbine blade  32  and the radius of each the notches  60   c  and  60   d  are selected such this will be the expected result. However, as explained below, this might not be the case. 
     In a first example, the notches  60   a  and  60   b  have a radius of 0.005 mils (0.000127 mm), the notches  60   c  and  60   d  have a radius of 0.010 mils (0.000254 mm), and the notches  60   e  and  60   f  have a radius of 0.015 mils (0.000381 mm). However, the tip  42  of the turbine blade  32  can include any number of notches  60  each having any radius and the notches  60  can be placed in any location and configuration on the tip  42  of the turbine blade  32 . The sequence and quantity of the notches  60  will be predetermined based on the needed understanding of the rub phenomenon that occurs during operating of the gas turbine engine  10  during development and testing. 
     In a second example, the turbine blade  32  can include a fourth set of notches  60   g  and  60   h  (shown in dashed lines in  FIG. 3 ) that have a radius of 0.005 mils that is located closer to the trailing edge  54  than the notches  60   e  and  60   f . In this example, from the leading edge  52  to the trailing edge  54 , the notches  60   a  and  60   b  have a radius of 0.005 mils (0.000127 mm), the notches  60   c  and  60   d  have a radius of 0.015 mils (0.000381 mm), the notches  60   e  and  60   f  have a radius of 0.010 mils (0.000254 mm), and the notches  60   g  and  60   h  have a radius of 0.005 mils (0.000127 mm). 
     After the notches  60  are formed in the tip  42  of the turbine blade  32  and the gas turbine engine  10  is assembled, it is operated and tested. As the turbine blades  32  rotate and increase in temperature, they expand in length, and the tips  42  rub against the casing  36 . After operation of the gas turbine engine  10  during the test ends, the turbine blades  32  cool and retract in length. 
     A borescope  62  (shown schematically) is then used to view the notches  60  and determine if any of the notches  60  have be eliminated due to erosion or rub of the tip  42  against the casing  36 . The gas turbine engine  10  includes a pre-existing hole (not shown) that is filled with a plug (not shown). The plug is removed from the pre-existing hole, and the borescope  62  is inserted into a pre-existing hole to view the tip  42  of the turbine blade  32 . 
     The borescope  62  is employed to view and determine how much of the tip  42  has worn away during testing of the gas turbine engine  10 . As each notch  60  has a known radius, it can be determined how much of the tip  42  of the turbine blade  32  has worn away during operation by viewing the tip  42  and determining which notches  60  remain and which notches  60  have been eliminated due to wear or rub against the casing  36 . From this information, the proper length of the turbine blade  32  for manufacture and actual use can be determined, and the turbine blades  32  that will be manufactured for use in actual operating gas turbine engines  10  will have this manufacturing length. 
     For example, as stated above, the middle notches  60   c  and  60   d  each have a radius that is equal to the amount of wear that is expected when the gas turbine engine  10  is tested. Returning to the first example, as shown in  FIG. 4 , if the middle notches  60   c  and  60   d  have been completely eliminated during testing due to rubbing of the tip  42  with the casing  36  (which also means the notches  60   a  and  60   b  with the smaller radii have been eliminated by rubbing), but the notches  60   e  and  60   f  (which have a larger radii) remain, this indicates that 0.010 mils (0.000254 mm) of material has eroded from the airfoil  50  during the test. Based on this knowledge, it can be determined that the turbine blades  32  are to be manufactured with a manufacturing length that is 0.010 mils (0.000254 mm) less than the length L of the turbine blade  32  prior to the test. 
     In another example, if only the notches  60   a  and  60   b  are eliminated during the test due to rubbing of the tip  42  with the casing  36 , this indicates that 0.005 mils (0.000127 mm) of material has eroded from the airfoil  50  during the test. Based on this knowledge, it can be determined that the turbine blades  32  are to be manufactured with a manufacturing length that is 0.005 mils (0.000127 mm) less than the length L of the turbine blade  32  prior to the test. 
     By viewing the notches  60  each having a known radius remaining on the tip  42  of the turbine blade  32  after the test cycle with a borescope  62 , it can be determined how much of the airfoil  50  has eroded because of rub and wear with the casing  36 . The turbine blade  32  can then be manufactured with the determined manufacturing length so that when the turbine blade  32  expands due to creep during use, the tip  42  of the turbine blade  32  contacts the casing  36  to create a proper seal while reducing wear. 
     Alternately, the amount of wear of the notches  60   a ,  60   c  and  60   e  on the pressure side  56  is compared to the amount of wear of the notches  60   b ,  60   d  and  60   f  on the suction side  58  of the turbine blade  32  after testing by viewing with the borescope  62 . If it is viewed based on the visual appearance of the notches  60  that there is more wear on one side  56  or  58  of the turbine blade  32  than the other side  56  or  58  of the turbine blade  32  due to the elimination of more notches  60  on one side  56  or  58  of the turbine blade  32  than the other side  56  or  58  of the turbine blade, this could indicate that tilt is occurring. The turbine blade  32  can then be designed and manufactured to take this into account. 
     By collecting data on erosion and wear of the tip  42  of the turbine blade  32  during testing and determining the amount of erosion and wear to the tip  42  due to creep and/or tilt prior to manufacturing the turbine blade  32  and assembling the gas turbine engine  10  for actual use, the turbine blade  32  can be designed to have a length that prevents erosion and wear during actual use while still providing a seal. By viewing the condition and existence of the notches  60  after testing the gas turbine engine  10  and visually evaluating their condition, presence or absence by the borescope  62  based on the known radii, any creep and tilt can be detected and be taken into consideration when designing and determining the actual length of the turbine blades  32 . 
     By using a borescope  62  to view the condition of the tip  42  of the turbine blade  32 , it is not necessary to disassemble the gas turbine engine  10  during development and engine testing, which provides a cost saving. Evaluation and disposition of several potential distress modes (i.e., creep, erosion, and tilt) is possible without tearing down the gas turbine engine  10  and needing measuring devices. Therefore, the turbine blade  32  can be made with the proper specifications, size and length prior to manufacturing. 
     The foregoing description is only exemplary of the principles of the invention. Many modifications and variations are possible in light of the above teachings. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than using the example embodiments which have been specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.