Abstract:
A method of calculating a throat area of a section of a machinery, according to an exemplary aspect of the present disclosure includes, among other things, outlining a boundary of the throat area of the section, selecting a plurality of inspection points along the boundary of the throat area, dividing the throat area into a plurality of sections, calculating an individual area of each of the plurality of sections and summing the individual areas of each of the plurality of sections to calculate the throat area.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a gas turbine engine, and more particularly to a method of calculating a throat area of a section of a gas turbine engine. 
         [0002]    Gas turbine engines typically include a compressor section, a combustor section and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
         [0003]    The compressor and turbine sections of the gas turbine engine typically include alternating rows of rotating blades and stationary vanes. The rotating blades either create or extract energy from the hot combustion gases that are communicated through the gas turbine engine, and the vanes convert the velocity of the airflow into pressure and prepare the airflow for the next set of blades. 
         [0004]    The vanes are typically arranged in an array such that adjacent vanes of a vane assembly are circumferentially disposed about a longitudinal centerline axis of the gas turbine engine. The hot combustion gases that are communicated through the gas turbine engine flow through a throat area that extends between the adjacent vanes of the vane assembly. The throat area controls the pressure ratio and mass flow rate of the gas turbine engine. 
         [0005]    The throat area associated with a particular section of a gas turbine engine, such as a turbine section, can be calculated using a series of gauges that measure the location of a pressure side and suction side of each airfoil of a vane assembly associated with the turbine section. These measurements can be used to mathematically calculate a throat area associated with the vane assembly. This method assumes axi-symmetric flow path surfaces and ignores any undulations or flow path end wall contouring of the platforms of the vane assembly, which potentially invalidate the calculation. 
       SUMMARY 
       [0006]    A method of calculating a throat area of a section of a machinery, according to an exemplary aspect of the present disclosure includes, among other things, outlining a boundary of the throat area of the section, selecting a plurality of inspection points along the boundary of the throat area, dividing the throat area into a plurality of sections, calculating an individual area of each of the plurality of sections and summing the individual areas of each of the plurality of sections to calculate the throat area. 
         [0007]    In a further non-limiting embodiment of the foregoing method of calculating a throat area of a section of a machinery, the step of outlining includes outlining the boundary of the throat area using a CAD model of the section. 
         [0008]    In a further non-limiting embodiment of either of the foregoing methods of calculating a throat area of a section of a machinery, the step of dividing includes radially dividing the throat area into the plurality of sections. 
         [0009]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the plurality of sections are trapezoidal shaped. 
         [0010]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the plurality of sections are triangular shaped. 
         [0011]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, a first portion of the plurality of sections are trapezoidal shaped and a second portion of the plurality of sections are triangular shaped. 
         [0012]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the method includes calculating an error value is associated with the throat area. 
         [0013]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the step of adjusting the throat area is based on the error value. 
         [0014]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the step of calculating includes measuring a distance between a first portion of the plurality of inspections points located on an outer diameter portion of the boundary and a second portion of the plurality of inspection points located on an inner diameter portion of the boundary. 
         [0015]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the step of measuring a location of each of the plurality of inspection points uses a coordinate measuring machine (CMM). 
         [0016]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the section is a mid-turbine frame of a gas turbine engine. 
         [0017]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a machinery, the section is a turbine section of a gas turbine engine. 
         [0018]    A method of calculating a throat area of a section of a gas turbine engine, according to another exemplary aspect of the present disclosure includes, among other things, radially dividing a throat area of the section into a plurality of sections, calculating an individual area associated with each of the plurality of sections of the throat area and summing the individual areas of each of the plurality of sections to calculate the throat area. 
         [0019]    In a further non-limiting embodiment of the foregoing method of calculating a throat area of a section of a gas turbine engine, the section is a turbine section of the gas turbine engine. 
         [0020]    In a further non-limiting embodiment of either the foregoing methods of calculating a throat area of a section of a gas turbine engine, the method includes outlining a boundary of the throat area of the turbine section in a CAD model prior to the step of radially dividing. 
         [0021]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a gas turbine engine, the method includes selecting a plurality of inspection points along the boundary of the throat area. 
         [0022]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a gas turbine engine, the boundary includes an outer diameter portion, an inner diameter portion, a suction side portion and a pressure side portion and the plurality of inspection points are selected on each of the outer diameter portion, the inner diameter portion, the suction side portion and the pressure side portion. 
         [0023]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a gas turbine engine, the step of radially dividing includes measuring a radial distance between a plurality of outer diameter portions and a plurality of inner diameter portions of a boundary of the throat area. 
         [0024]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a gas turbine engine, the section includes an end wall having at least one undulation. 
         [0025]    In a further non-limiting embodiment of any of the foregoing methods of calculating a throat area of a section of a gas turbine engine, the method includes adjusting the throat area calculation based on an error value associated with a trailing edge location of the section. 
         [0026]    The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]      FIG. 1  illustrates a schematic, cross-sectional view of a gas turbine engine. 
           [0028]      FIGS. 2A and 2B  illustrate a segment of a section of a gas turbine engine. 
           [0029]      FIG. 3  schematically illustrates a throat area associated with a segment of a section of a gas turbine engine. 
           [0030]      FIG. 4  schematically illustrates a method of calculating a throat area of a segment of a section of a gas turbine engine. 
       
    
    
     DETAILED DESCRIPTION 
       [0031]      FIG. 1  schematically illustrates a gas turbine engine  20 . The exemplary gas turbine engine  20  is a two-spool turbofan engine that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems for features. The fan section  22  drives air along a bypass flow path B, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26 . The hot combustion gases generated in the combustor section  26  are expanded through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures. 
         [0032]    The gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine centerline longitudinal axis A. The low speed spool  30  and the high speed spool  32  may be mounted relative to an engine static structure  33  via several bearing systems  31 . It should be understood that other bearing systems  31  may alternatively or additionally be provided. 
         [0033]    The low speed spool  30  generally includes an inner shaft  34  that interconnects a fan  36 , a low pressure compressor  38  and a low pressure turbine  39 . The inner shaft  34  can be connected to the fan  36  through a geared architecture  45  to drive the fan  36  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  35  that interconnects a high pressure compressor  37  and a high pressure turbine  40 . In this embodiment, the inner shaft  34  and the outer shaft  35  are supported at various axial locations by bearing systems  31  positioned within the engine static structure  33 . 
         [0034]    A combustor  42  is arranged between the high pressure compressor  37  and the high pressure turbine  40 . A mid-turbine frame  44  may be arranged generally between the high pressure turbine  40  and the low pressure turbine  39 . The mid-turbine frame  44  can support one or more bearing systems  31  of the turbine section  28 . The mid-turbine frame  44  may include one or more airfoils  46  that extend within the core flow path C. 
         [0035]    The inner shaft  34  and the outer shaft  35  are concentric and rotate via the bearing systems  31  about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor  38  and the high pressure compressor  37 , is mixed with fuel and burned in the combustor  42 , and is then expanded over the high pressure turbine  40  and the low pressure turbine  39 . The high pressure turbine  40  and the low pressure turbine  39  rotationally drive the respective high speed spool  32  and the low speed spool  30  in response to the expansion. 
         [0036]    In a non-limiting embodiment, the gas turbine engine  20  is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  bypass ratio is greater than about six (6:1). The geared architecture  45  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low speed spool  30  at higher speeds, which can increase the operational efficiency of the low pressure compressor  38  and low pressure turbine  39  and render increased pressure in a fewer number of stages. 
         [0037]    The pressure ratio of the low pressure turbine  39  can be pressure measured prior to the inlet of the low pressure turbine  39  as related to the pressure at the outlet of the low pressure turbine  39  and prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  38 , and the low pressure turbine  39  has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines, including direct drive turbofans. 
         [0038]    In this embodiment of the exemplary gas turbine engine  20 , a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
         [0039]    Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram °R)/(518.7 °R)] 0.5 , where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
         [0040]    Each of the compressor section  24  and the turbine section  28  may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades  25 , while each vane assembly can carry a plurality of vanes  27  that extend into the core flow path C. The blades  25  of the rotor assemblies create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine  20  along the core flow path C. The vanes  27  of the vane assemblies direct the core airflow to the blades  25  to either add or extract energy. 
         [0041]      FIGS. 2A and 2B  illustrate a vane segment  50  that can be incorporated into a section  100  of a gas turbine engine, such as the gas turbine engine  20  of  FIG. 1 . In this embodiment, the section  100  is the turbine section  28  of the gas turbine engine  20  and the vane segment  50  is part of a vane assembly disposed within the turbine section  28 . However, other sections of a gas turbine engine  20  could also benefit from the teachings of this disclosure, including but not limited to, the compressor section and the mid-turbine frame. 
         [0042]    The vane segment  50  can include a single annular structure or a plurality of vane segments  50  could be mechanically attached to one another and annularly disposed about the engine centerline axis A to form a full-ring vane assembly. The vane segment  50  could include one or more airfoils (or vanes)  56 . In this particular embodiment, the section  100  includes a first vane segment  50 A and a second vane segment  50 B each having a single airfoil  56 . 
         [0043]    The first and second vane segments  50 A,  50 B include an outer diameter platform  52  and an inner diameter platform  54  spaced from the outer diameter platform  52 . The airfoils  56  extend between the outer diameter platform  52  and the inner diameter platform  54 . The airfoils  56  may extend radially across an annulus between the outer diameter platform  52  and the inner diameter platform  54 . Each airfoil  56  includes a leading edge  58 , a trailing edge  60 , a pressure side  62  and a suction side  64 . The airfoils  56  axially extend between the leading edge  58  and the trailing edge  60  and circumferentially extend between the pressure side  62  and the suction side  64 . 
         [0044]    A gas path  65  is communicated axially downstream through the gas turbine engine  20  along the core flow path C ( FIG. 1 ) in a direction that extends from the leading edges  58  toward the trailing edges  60  of the airfoils  56 . The gas path  65  represents the communication of core airflow across the airfoils  56  and through the spaces that extend between the adjacent airfoils  56  of the first and second vane segments  50 A,  50 B. The space that extends between adjacent airfoils  56  of the vane segments  50 A,  50 B is referred to as a throat area  66 . 
         [0045]    As best illustrated in  FIG. 2B , the inner diameter platform  54  and the outer diameter platform  52  of one or both of the first and second vane segments  50 A,  50 B include end walls  67  that establish the radially outer and radially inner flow path boundaries of the gas path  65 . One or more of the end walls  67  can include undulations  69  that provide a contoured end wall surface. The undulations  69  may reduce friction losses that can occur as the airflow of the gas path  65  is communication through the section  100 , thereby increasing engine efficiency. However, these undulations  69  can also add difficulty in calculating the throat area  66  using known throat area calculation methodologies. 
         [0046]      FIG. 3  schematically illustrates a throat area  66  of a vane segment  50 . The throat area  66  extends between a first airfoil  56 A and a second, adjacent airfoil  56 B of the vane segment  50 . The throat area  66  of the vane segment  50  is a planar opening with a periphery bounded in a radial direction by the outer diameter platform  52  (shown in  FIG. 2 ) and the inner diameter platform  54  and peripherally bounded in the circumferential direction by the suction side  64  of the first airfoil  56 A and the trailing edge  60  of the second airfoil  56 B (at the pressure side  62 ). Of course, a full ring vane assembly would include a plurality of throat areas that extend between adjacent airfoils. The throat area  66  of the vane segment  50  is an important parameter that may influence engine efficiency. Therefore, the throat area  66  must be periodically measured and calculated to verify clearances in the core flow path C. 
         [0047]      FIG. 4 , with continued reference to the previous Figures, schematically illustrates an exemplary method  101  of calculating a throat area  66  of a section of a gas turbine engine. In one embodiment, the method  101  can be used to calculate a throat area  66  associated with a turbine vane assembly of the gas turbine engine, although other sections could also benefit from the teachings of this disclosure. The method  101  may be particularly useful for calculating a throat area on segments of the gas turbine engine that include flow path end wall contouring (see, e.g., undulations  69  of end walls  67  illustrated by  FIG. 2B ). 
         [0048]    The method may begin at step block  102  by outlining a boundary  68  of the throat area  66 . In one embodiment, the boundary  68  is outlined using a CAD model of a section of the gas turbine engine. A person of ordinary skill in the art having the benefit of this disclosure would be able to select and outline the boundary  68  of the throat area  66  by referencing a CAD model. 
         [0049]    The boundary  68  is an outer periphery of the throat area  66 . In this embodiment, the boundary  68  extends radially between an outer diameter portion  70  and an inner diameter portion  72  and extends circumferentially between a suction side portion  74  and a pressure side portion  76 . The outer diameter portion  70  and the inner diameter portion  72  represent the locations where the throat area  66  is bounded by the outer diameter platform  52  and the inner diameter platform  54 , respectively (see  FIGS. 2A ,  2 B). The outer diameter portion  70  and the inner diameter portion  72  may be contoured to reflect any undulations  69  formed in the end walls  67  of the outer diameter platform  52  and/or the inner diameter platform  54 . In other words, the outer diameter portion  70  and the inner diameter portion  72  may extend non-linearly as shown. In addition, the suction side portion  74  of the boundary  68  represents the portion of the throat area  66  that extends adjacent a suction side  64  of a first airfoil  56 A of the section and the pressure side portion  76  represents the portion of the throat area  66  that extends adjacent the pressure side  62  of a second, adjacent airfoil  56 B of the section (see, for example, the vane segment  50  of  FIG. 3 ). The suction side portion  74  and the pressure side portion  76  are linear in this embodiment. 
         [0050]    Next, at step block  104 , a plurality of inspection points (marked in  FIG. 4  with a +sign) are selected along the boundary  68  of the throat area  66 . For example, a plurality of inspection points OD 1  through OD N  are selected along the outer diameter portion  70 , a plurality of inspection points ID 1  through ID N  are selected along the inner diameter portion  72 , a plurality of inspection points SS 1  through SS N  are selected along the suction side portion  74  and a plurality of inspection points PS 1  through PS N  are selected along the pressure side portion  76 . In one embodiment, the locations of these various inspection points may be specified in an engineering print or CAD model of the section of the gas turbine engine. However, the plurality of inspection points may also be selected using other known methodologies. 
         [0051]    The plurality of inspection points of the boundary  68  of the throat area  66  can be measured at step block  106 . For example, each of the plurality of inspection points can be measured on an actual casting of the section of the gas turbine engine using a coordinate measuring machine (CMM). Other measuring techniques are also contemplated as within the scope of this disclosure. The actual locations of the plurality of inspection points of the boundary  68  can be recorded at step block  108 , including, to the extent the location is incorrect, a recordation of an out-of-nominal distance associated with each of the plurality of inspection points. As discussed below, the CMM data is later used to calculate the area of a plurality of sections of the throat area  66 . 
         [0052]    At step block  110 , the throat area  66  is divided into a plurality of sections S 1  through S N . In one embodiment, the throat area  66  is radially divided into the plurality of sections S N  by drawing radial lines L 1  through L N  from at least a first portion of the plurality of inspection points OD 1  through OD N  to a corresponding second portion of the plurality of inspection points ID 1  through OD N . In this embodiment, the inspection points OD 2  and ID 2  are connected by radial line L 2  that defines a radial distance, inspection points OD 4  and ID 4  are connected by radial line L 4  that defines a radial distance, inspection points OD 6  and ID 6  are connected by radial line L 6  that defines a radial distance, inspection points OD 10  and ID 10  are connected by radial line L 10  that defines a radial distance, inspection points OD 12  and ID 12  are connected by radial line L 12  that defines a radial distance, and inspection points OD 13  and M B  are connected by a radial line L 13  that defines a radial distance. This is but one example of how the throat area  66  could be divided, and it should be understood that the throat area  66  could be divided into greater or fewer sections than are illustrated in this embodiment. The division of the throat area  66  into the plurality of sections S 1  through S N  may also depend on the design of the section for which the throat area is being calculated. That is, the throat areas  66  of different parts may be divided into different numbered and shaped sections. 
         [0053]    The plurality of sections S 1  through S N  that divide the throat area  66  may embody a variety of shapes. For example, the plurality of sections S 1  through S N  can be trapezoidal shaped. In another embodiment, the plurality of sections S 1  through S N  can be triangular shaped. In yet another embodiment, a first portion of the plurality of sections S 1  through S N  are trapezoidal shaped and a second portion of the plurality of sections S 1  through S N  are triangular shaped sections. Other shapes are also contemplated. For example, each shape may be one in which an area of the shape can be easily mathematically calculated. 
         [0054]    Next, at step block  112 , an individual area of each of the plurality of sections S 1  through S N  is calculated. A person of ordinary skill in the art having the benefit of this disclosure would understand how to calculate the area of each of the plurality of sections S 1  through S N . For example, as is known, the area of a triangle can be calculated using the formula A=½(b)(h), where b is the length of the base of the triangle and h is the height of the triangle. In another example, the area of a trapezoid can be calculated using the formula ½(b 1 +b 2 )h, where b 1  and b 2  are the lengths of the bases and h is the height. In this embodiment, the lengths of the radial lines L 1  through L N  can be used as the “h” values in the area formulas. These radial distances are measured and calculated from the CMM inspection data from step block  108 . At step block  114 , the individual areas of each of the plurality of sections S 1  through S N  can be summed to calculate the throat area  66 . 
         [0055]    An error value associated with the throat area  66  calculation can optionally be calculated at step block  116 . It may be necessary to calculate the error value associated with the throat area  66  calculation because the airfoil  56  of a section can be cast with a trailing edge  60  that is shorter or longer (due to shrinkage error) than defined by the engineering print or CAD model of the section. The calculated throat area  66  can be adjusted based on the error value calculated at step block  116 . The error value may also be added to account for slight inabilities of the formula to measure the exact flow area value that can result from a coarse number of divisions of the flow area surface. Additional methods may be incorporated to account for throat area  66  value effects from the trailing edge  60  location. 
         [0056]    Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
         [0057]    It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
         [0058]    The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.