Patent Publication Number: US-9404367-B2

Title: Gas turbine engine compressor rotor assembly and balancing system

Description:
TECHNICAL FIELD 
     The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a gas turbine engine compressor rotor assembly with a balancing system. 
     BACKGROUND 
     Gas turbine engines include compressor, combustor, and turbine sections. Rotating components of the gas turbine engine may need to be balanced due to limitations in component manufacturing. In particular the compressor rotor assembly may need to be balanced to reduce vibrations in the gas turbine engine. Larger compressor rotor assemblies may use a dynamic balancing system and method for balancing to reduce vibration and increase component reliability. 
     US Publication No. 2010135774, to Dezouche, discloses the balancing flyweights of a turbomachine rotor includes two pyramid shaped end parts each one having a base and an apex, and an intermediate part which connects the two bases of the end parts together. The two apexes are aligned on a longitudinal axis. The two end parts and the intermediate part exhibit, in cross section through a plane perpendicular to the longitudinal axis, cross-sections having polygonal shapes centered on said longitudinal axis. 
     The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors. 
     SUMMARY OF THE DISCLOSURE 
     A method for balancing a compressor rotor assembly including a forward weldment and an aft weldment is disclosed. The method includes pre-balancing the aft weldment of the compressor rotor assembly with compressor disks prior to populating the compressor disks with circumferentially installed compressor rotor blades. Pre-balancing the aft weldment includes measuring a rotational balance of the aft weldment. Pre-balancing the aft weldment also includes determining a number of underplatform weights needed and a location for each underplatform weight within a circumferential slot of one the compressor disks. Pre-balancing the aft weldment further includes mounting each underplatform weight in the determined location. 
     A gas turbine engine compressor rotor assembly with a balancing system includes a first stage compressor disk, a plurality of compressor disks, forward weights, and a plurality of underplatform weights. The first stage compressor disk has a cylindrical body. The first stage compressor disk includes a plurality of forward balancing holes circumferentially about the cylindrical body. The first stage compressor disk also includes a plurality of aft balancing holes circumferentially about the cylindrical body and located adjacent to the plurality of forward balancing holes. Each of the compressor disks includes a circumferential slot. Each circumferential slot includes a dovetail profile. Forward weights are configured to be installed in the plurality of forward balancing holes and the plurality of aft balancing holes. Each underplatform weight is configured to be installed within one or more of each of the circumferential slots. Each underplatform weight has a dovetail shape corresponding to the circumferential slot dovetail profile of one or more of the plurality of compressor disks. The plurality of underplatform weights includes two or more sizes. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic illustration of an exemplary gas turbine engine. 
         FIG. 2  is a perspective view of the compressor rotor assembly of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is a cross-sectional view of the forward weldment of the compressor rotor assembly of  FIG. 2 . 
         FIG. 4  is a cross-sectional view of the aft weldment of the compressor rotor assembly of  FIG. 2 . 
         FIG. 5  is a perspective view of a portion of the compressor rotor assembly of  FIG. 2  with circumferentially installed compressor rotor blades, an exemplary underplatform weight, and a compressor disk with a portion of the compressor disk cutaway to show the roots of the compressor rotor blades and the underplatform weight. 
         FIG. 6  is a perspective view of the underplatform weight of  FIG. 5 . 
         FIG. 7  is a side view of the underplatform weight of  FIG. 5 . 
         FIG. 8  is a flowchart of a method for balancing a gas turbine engine compressor rotor assembly, which includes pre-balancing the aft weldment and pre-balancing the forward weldment. 
         FIG. 9  is a flowchart of methods for balancing a gas turbine engine compressor rotor assembly, which includes balancing the assembled compressor rotor assembly. 
     
    
    
     DETAILED DESCRIPTION 
     The systems and methods disclosed herein include a gas turbine engine compressor rotor assembly with a balancing system. In embodiments, the compressor rotor assembly includes a forward weldment, an aft weldment, and a balancing system. The balancing system includes forward weights and underplatform weights. The forward weights may be installed into one of two rows of balancing holes, which may provide for a quicker and more accurate balancing of the forward weldment or the compressor rotor assembly. The underplatform weights may be installed between any circumferentially installed compressor rotor blades of the compressor rotor assembly, which may provide the ability to pre-balance the aft weldment and may provide for a quicker and more accurate balancing of the compressor rotor assembly. 
       FIG. 1  is a schematic illustration of an exemplary gas turbine engine. Some of the surfaces have been left out or exaggerated (here and in other figures) for clarity and ease of explanation. Also, the disclosure may reference a forward and an aft direction. Generally, all references to “forward” and “aft” are associated with the flow direction of primary air (air which is used in the Brayton cycle, the thermodynamic basis for gas turbine operation), unless specified otherwise. For example, forward is “upstream” relative to primary air flow, and aft is “downstream” relative to primary air flow. 
     In addition, the disclosure may generally reference a center axis  95  of rotation of the gas turbine engine, which may be generally defined by the longitudinal axis of its shaft  120  (supported by a plurality of bearing assemblies  150 ). The center axis  95  may be common to or shared with various other engine concentric components. All references to radial, axial, and circumferential directions and measures refer to center axis  95 , unless specified otherwise, and terms such as “inner” and “outer” generally indicate a lesser or greater radial distance from, wherein a radial  96  may be in any direction perpendicular and radiating outward from center axis  95 . 
     A gas turbine engine  100  includes an inlet  110 , a shaft  120 , a gas producer or “compressor”  200 , a combustor  300 , a turbine  400 , an exhaust  500 , and a power output coupling  600 . The gas turbine engine  100  may have a single shaft or a dual shaft configuration. 
     The compressor  200  includes a compressor rotor assembly  210 , compressor stationary vanes (“stators”)  250 , and inlet guide vanes  251 . The compressor rotor assembly  210  mechanically couples to shaft  120 . As illustrated, the compressor rotor assembly  210  is an axial flow rotor assembly. The compressor rotor assembly  210  may include a forward weldment  211  and an aft weldment  212 . The forward weldment  211  and the aft weldment  212  each include one or more compressor disk assemblies  219 . Each compressor disk assembly  219  includes a compressor rotor disk  220  (shown in  FIGS. 2, 3, and 4 ) that is circumferentially populated with compressor rotor blades. The forward weldment may also include the first stage compressor disk  221 , which may be coupled to the forward hub  213 . 
     Stators  250  axially follow each of the compressor disk assemblies  219 . Each compressor disk assembly  219  paired with the adjacent stators  250  that follow the compressor disk assembly  219  is considered a compressor stage. Compressor  200  includes multiple compressor stages. Inlet guide vanes  251  axially precede the first compressor stage. 
     The combustor  300  includes one or more injectors  350  and includes one or more combustion chambers  390 . 
     The turbine  400  includes a turbine rotor assembly  410  and turbine nozzles  450 . The turbine rotor assembly  410  mechanically couples to the shaft  120 . As illustrated, the turbine rotor assembly  410  is an axial flow rotor assembly. The turbine rotor assembly  410  includes one or more turbine disk assemblies  420 . Each turbine disk assembly  420  includes a turbine disk that is circumferentially populated with turbine blades. Turbine nozzles  450  axially precede each of the turbine disk assemblies  420 . Each turbine disk assembly  420  paired with the adjacent turbine nozzles  450  that precede the turbine disk assembly  420  is considered a turbine stage. Turbine  400  includes multiple turbine stages. 
     The exhaust  500  includes an exhaust diffuser  520  and an exhaust collector  550 . 
       FIG. 2  is a perspective view of the compressor rotor assembly  210  of  FIG. 1 . The compressor rotor assembly  210  may include a balancing system. The balancing system may include a forward balancing system  255 , compressor rotor blades, and underplatform weights  260  (shown in  FIGS. 5-7 ). 
     Forward balancing system  255  includes multiple forward balancing holes  242 , multiple ah balancing holes  243 , and forward weights  256 . A first group of balancing holes may be selected from the forward balancing holes  242  and the aft balancing holes  243 . The remaining forward balancing holes  242  and aft balancing holes  243  may comprise a second group of balancing holes. Alternatively, the forward balancing holes  242  may comprise the first group of balancing holes and the aft balancing holes  243  may comprise the second group of balancing holes. 
     Forward weights  256  may have various sizes, masses, and lengths. In one embodiment forward weights  256  have a ⅜ inch diameter and lengths of ¼ inch, ½ inch, or ¾ inch. Alternatively, other diameters may be used. 
     Compressor rotor blades may be axially installed compressor rotor blades (“axial blades”)  229  or circumferentially installed compressor rotor blades (“circumferential blades”)  230 . Compressor rotor blade sizes may be determined by the sizes of the compressor disks  220 . 
       FIG. 3  is a cross-sectional view of the forward weldment  211  of the compressor rotor assembly  210  of  FIG. 2 . The forward weldment  211  includes multiple compressor disks  220  including the first stage compressor disk  221  and the forward fastening compressor disk  223 . The first stage compressor disk  221  may be located at the forward end of forward weldment  211 . The first stage compressor disk  221  may have a cylindrical body  240  and may include a forward end  238 , an aft end  239 , an outer axial flange  237 , and an outer surface  241 . The outer axial flange  237  may extend axially forward from the cylindrical body  240 . The outer surface  241  may extend from the forward end  238  towards the aft end. A portion of the outer surface  241  may be on the outer axial flange  237 . 
     Radial flange  246  may extend radially outward from the cylindrical body  240 . Radial flange  246  may include axial slots  235  configured for mounting axial blades  229  (shown in  FIG. 2 ) to the first stage compressor disk  221 . The axial slots  235  may have a fir tree or dovetail cross-sectional shape. 
     The first stage compressor disk  221  may also include forward balancing holes  242  and aft balancing holes  243 . Each forward balancing hole  242  may extend radially inward through the outer surface  241 . Forward balancing holes  242  may be aligned circumferentially and evenly spaced about outer surface  241 . Each aft balancing hole  243  may extend radially inward through the outer surface  241 . Aft balancing holes  243  may be aligned circumferentially and evenly spaced about outer surface  241 . The aft balancing holes  243  may be adjacent to the forward balancing holes  242 , may be axially aft of the forward balancing holes  242 , and may be circumferentially offset or clocked relative to the forward balancing holes  242 . 
     The forward balancing holes  242  and the aft balancing holes  243  may be located near the center of gravity of the first stage compressor disk  221 . The aft balancing holes  243  may be closer to the center of gravity of the first stage compressor disk  221  than the forward balancing holes  242 . The forward balancing holes  242  and the aft balancing holes  243  may be threaded. In one embodiment the holes have a ⅜ inch diameter. Alternatively, other diameters may be used. 
     The forward balancing holes  242  may total more man twelve and less than thirty. The aft balancing holes  243  may total more than twelve and less than thirty. The number of forward balancing holes  242  and aft balancing holes  243  may correspond with the diameter of outer surface  241  or may correspond with the number of axial slots  235  in the first stage compressor disk  221 . The aft balancing holes  243  may be circumferentially offset or clocked by half of the angular distance between adjacent forward balancing holes  242 . The depth of the forward balancing holes  242  and the aft balancing holes  243  may correspond with the size of the forward weights  256  of the forward balancing system  255 . 
     In one embodiment the forward balancing holes  242  may total twenty-four, the aft balancing holes  243  may total twenty-four, and the aft balancing holes  243  may be circumferentially offset or clocked 7.5 degrees relative to the forward balancing holes  242 . The aft balancing holes  243  may be shifted 1.5 inches axially aft of the forward balancing holes  242 . In another embodiment the aft balancing holes  243  may be at least 0.75 inches deep. 
     The first stage compressor disk  221  may also include forward surface  244 , hub mounting holes  245 , and inner axial flange  248 . Forward surface  244  may be the axially forward facing surface adjacent to the outer surface  241 . Forward surface  244  may be on the outer axial flange  237 . Hub mounting holes  245  may extend aft through the forward surface  244 . In one embodiment, the hub mounting holes  245  are in the outer axial flange  237 . 
     The inner axial flange  248  may extend axially forward from the forward end  238 . The inner axial flange  248  may be located within the outer axial flange  237 . 
     The first stage compressor disk  221  may also include an aft welding member  226 . The aft welding member  226  may have an annular shape and may extend aft from the cylindrical body  240 . 
     The first stage compressor disk  221  may further include a bore  249 . The bore  249  may extend from the inner axial flange  248  at the forward end  238 , to the aft end  239 . The shaft  120  may pass through the bore  249  of the first stage compressor disk  221 . 
     The forward fastening compressor disk  223  may be located at the aft end of forward weldment  211 . The forward fastening compressor disk  223  may include a forward welding member  225  and forward weldment mounting holes  227 . The forward welding member  225  may have an annular shape and may extend forward from the forward fastening compressor disk  223 . The forward weldment mounting holes  227  may be located on an aft end of the forward fastening compressor disk  223  and may extend axially forward. In the embodiment shown in  FIG. 3 , forward fastening compressor disk  223  also includes a circumferential slot  236  for mounting circumferential blades  230  to forward fastening compressor disk  223 . Circumferential slot  236  extends completely around forward fastening compressor disk  223 . Circumferential slot  236  may have a fir tree or a dovetail shape. 
     The compressor disks  220  not located at the forward or aft end of the forward weldment may include a forward welding member  225  and an aft welding member  226 . The forward welding member  225  may have an annular shape and may extend forward from the compressor disk  220 . The aft welding member  226  may have an annular shape and may extend aft from the compressor disk  220 . The aft welding member  226  of the first stage compressor disk  221  may be welded to the forward welding member  225  of the subsequent compressor disk  220 . Each subsequent compressor disk  220  may be welded to the previous compressor disk  220  in a similar manner. The forward fastening compressor disk  223  may also be welded to the previous compressor disk  220  in a similar manner. In one embodiment the forward weldment  211  may include nine compressor disks  220 ; the forward fastening compressor disk  223  may be the ninth stage compressor disk. 
     Each compressor disk  220  of forward weldment  211  may include multiple axial slots  235  or a circumferential slot  236 . If the compressor disk  220  includes axial slots  235 , one axial blade  229  may be inserted into each axial slot  235 . If the compressor disk  220  includes a circumferential slot  236 , multiple circumferential blades may be inserted into the circumferential slot  236 . Underplatform weights  260  may be inserted into circumferential slot  236  between circumferential blades  230  (as shown in  FIG. 5 ). In the embodiment shown in  FIG. 3 , the first six compressor disks  220  include axial slots  235 , while the seventh, eighth, and ninth compressor disks  220  each include a circumferential slot  236 . 
       FIG. 4  is a cross-sectional view of the aft weldment  212  of the compressor rotor assembly of  FIG. 2 . The aft weldment  212  may include multiple compressor disks  220  including the last stage compressor disk  222  and the aft fastening compressor disk  224 . The aft fastening compressor disk  224  may include an aft welding member  226  and aft weldment mounting holes  228 . The aft welding member  226  may have an annular shape and may extend aft from the aft fastening compressor disk  224 . The aft weldment mounting holes  228  may be located on a forward end of the aft fastening compressor disk  224  and may extend axially aft. 
     The aft welding member  226  of the aft fastening compressor disk  224  may be welded to the forward welding member  225  of the subsequent compressor disk  220 . Each subsequent compressor disk  220  may be welded to the previous compressor disk  220  in a similar manner. The last stage compressor disk  222  may also be welded to the previous compressor disk  220  in a similar manner. In one embodiment the aft weldment  212  may include seven compressor disks  220 . In the embodiment shown in  FIG. 4 , the aft fastening compressor disk  224  is the tenth stage compressor disk and the last stage compressor disk  222  is the sixteenth stage compressor disk. 
     Each compressor disk  220  of aft weldment  212  may include multiple axial slots  235  or a circumferential slot  236 . If the compressor disk  220  includes axial slots  235 , one axial blade  229  may be inserted into each axial slot  235 . If the compressor disk  220  includes a circumferential slot  236 , multiple circumferential blades  230  may be inserted into the circumferential slot  236 . Underplatform weights  260  may be inserted into circumferential slot  236  between circumferential blades  230  (as shown in  FIG. 5 ). In the embodiment shown in  FIG. 4 , each aft weldment  212  compressor disk  220  includes a circumferential slot  236 . Some circumferential slots  236  in forward weldment  211  and aft weldment  212  may have different dovetail or fir tree cross sections. 
       FIG. 5  is a perspective view of a portion of the compressor rotor assembly  210  of  FIG. 2  with circumferential blades  230 , an exemplary underplatform weight  260 , and a compressor disk  220  with a portion of the compressor disk  220  cutaway to show the roots  234  of the circumferential blades  230  and the underplatform weight  260 . Each circumferential blade  230  may include an airfoil  231  and a base  232 . Each base  232  may include a platform  233  and a root  234 . Platform  233  attaches to an end of airfoil  231 . Root  234  extends from platform  233  in a direction opposite airfoil  231 . Root  234  may have a dovetail or fir tree shape that matches the dovetail or fir tree shape of circumferential slot  236 . 
     Each underplatform weight  260  is shaped to match the dovetail or fir tree shape of root  234 . The shape of each underplatform weight  260  may also match the profile of a circumferential slot  236 . The height of each underplatform weight  260  may be sized such that the top of the underplatform weight  260  does not contact platform  233 . The width of each underplatform weight  260  may be sized to fit between adjacent circumferential blades  230 . The width may be sized based on the tolerances of underplatform weight  260  and the root  234  of the circumferential blades  230  to ensure that underplatform weights  260  will fit between roots  234 . 
     The width of each underplatform weight  260  may also be sized to avoid too much space between each underplatform weight  260  and the adjacent circumferential blade roots. Too much space may allow underplatform weights to shift and alter the balance of the compressor rotor assembly  210 . Multiple underplatform weights  260  configurations and sizes may be used in the balancing system. For example, compressor disks  220  with circumferential slots may be divided into contiguous sections; each section includes one or more compressor disk  220 . A different set of underplatform weights  260  may be provided for each section. One embodiment includes four sections. The first section includes one compressor disk. The second section is adjacent to and downstream of the first section and includes two adjacent compressor disks. The third section is adjacent to and downstream of the second section and includes four adjacent compressor disks. A fourth section is adjacent to and downstream of the third section and includes three adjacent compressor disks. 
     In the embodiment shown in  FIGS. 2, 3, and 4 , first underplatform weights are used for the first section. The first section includes the seventh stage compressor disk. Second underplatform weights are used for the second section. The second section includes the eighth and ninth stage compressor disks. Third underplatform weights are used for the third section. The third section includes the tenth through the thirteenth stage compressor disks. Fourth underplatform weights are used for the fourth section. The fourth section includes the fourteenth through the sixteenth stage compressor disks. 
       FIG. 6  is a perspective view of the underplatform weight  260  of  FIG. 5 .  FIG. 7  is a side view of the underplatform weight  260  of  FIG. 5 . Referring to  FIGS. 6 and 7 , each underplatform weight  260  may include a top surface  261 , a bottom surface  262 , an upper face  263  at each end, a lower face  264  at each end, and two side surfaces  265 . The cross-section or profile of the dovetail shape may be a convex hexagon with two parallel sides. In the embodiment shown in  FIGS. 6 and 7 , the top surface  261  and the bottom surface  262  are parallel and define the two parallel sides of the hexagonal shape. The surfaces defining the hexagonal shape may have different lengths. For example, in the embodiment shown top surface  261  is longer than upper face  263  and upper face  263  is longer than lower face  264 . 
     Each upper face  263  may extend from an end of top surface  261  at an angle greater than 90 degrees and less than 180 degrees. Each lower face  264  may extend from an end of bottom surface  262  at an angle greater than 90 degrees and less than 180 degrees. The intersection of the upper face  263  and the lower face  264  at each end of each underplatform weight  260  may be at an angle between 90 degrees and 180 degrees. Side surfaces  265  extend from top surface  261  to bottom surface  262 . Side surfaces  265  may be perpendicular to top surface  261  and bottom surface  262 . Each end of underplatform weight  260  may be symmetrical. 
     The edges between surfaces and faces may be chamfered or rounded. In the embodiment shown in  FIGS. 5, 6, and 7 , the edges between top surface  261  and side surfaces  265  include a chamfer  266 ; the edges between upper face  263  and lower face  264 , bottom surface  262  and lower face  264 , and upper face  263  and side surfaces  265  are rounded. 
     INDUSTRIAL APPLICABILITY 
     Gas turbine engines may be suited for any number of industrial applications such as various aspects of the oil and gas industry (including transmission, gathering, storage, withdrawal, and lifting of oil and natural gas), the power generation industry, cogeneration, aerospace, and other transportation industries. 
     Referring to  FIG. 1 , a gas (typically air  10 ) enters the inlet  110  as a “working fluid”, and is compressed by the compressor  200 . In the compressor  200 , the working fluid is compressed in an annular flow path  115  by the series of compressor disk assemblies  219 . In particular, the air  10  is compressed in numbered “stages”, the stages being associated with each compressor disk assembly  219 . For example, “4th stage air” may be associated with the 4th compressor disk assembly  219  in the downstream or “aft” direction, going from the inlet  110  towards the exhaust  500 ). Likewise, each turbine disk assembly  420  may be associated with a numbered stage. 
     Once compressed air  10  leaves the compressor  200 , it enters the combustor  300 , where it is diffused and fuel  20  is added. Air  10  and fuel  20  are injected into the combustion chamber  390  via injector  350  and combusted. Energy is extracted from the combustion reaction via the turbine  400  by each stage of the series of turbine disk assemblies  420 . Exhaust gas  90  may then be diffused in exhaust diffuser  520 , collected and redirected. Exhaust gas  90  exits the system via an exhaust collector  550  and may be further processed (e.g., to reduce harmful emissions, and/or to recover heat from the exhaust gas  90 ). 
     Gas turbine engines and other rotary machines include a number of rotating elements. An imbalanced rotating element may cause vibration when rotating. Vibration in a rotating element may cause undesirable stresses in the rotating element. The stresses caused by the vibration may cause a fatigue failure in the rotating element or other related elements. Excessive vibration in a gas turbine engine may reduce reliability, may cause high bearing loads, and may lead to component failures. In a gas turbine engine excessive vibration may also cause the shaft to bend or suffer from fatigue failure. 
     Through research and testing it was determined that some larger gas turbine engines may need to include a more complex balancing system and method. A gas turbine compressor rotor assembly may be balanced with weights near the forward end, near the aft end and near the mid-plane of compressor assemblies. Due to the length of larger assemblies, more balancing locations may be needed to balance a larger assembly within a desired standard. 
     A suitable balancing method may be accomplished by increasing the number of balancing locations, while limiting the number of components used in the balancing system. The balancing system disclosed herein may increase the number of balancing locations by making underplatform weights  260  available for each compressor disk  220  with a circumferential slot  236 , and by providing forward balancing holes  242  and aft balancing holes  243  for forward weights  256 . Increasing the number of balancing locations may reduce the difficulty of balancing the forward weldment  211 , the aft weldment  212 , and the compressor rotor assembly  210  by providing more balancing options. The balancing system disclosed herein may limit the number of components used in the balancing system by using the same underplatform weights  260  in mom than one axial location or stage. Limiting the number of components in the balancing system may limit or reduce the complexity of the balancing system. Reduced complexity and reduced difficulty of a balancing system may reduce the balancing time and may increase the accuracy of the balancing system. 
     In the embodiment shown in  FIGS. 2, 3, and 4  the number of axial balancing locations totals twelve. This includes forward balancing holes  242 , aft balancing holes  243 , and compressor disks  220  in contiguous stages, stages seven through sixteen. However, the number of different components used in the embodiment shown in  FIGS. 2, 3, and 4  may be as low as five. This includes forward weights  256 , underplatform weights  260  for the stage seven compressor disk  220 , underplatform weights  260  for the compressor disks  220  in stages eight and nine, underplatform weights  260  for the compressor disks  220  in stages ten through thirteen, and underplatform weights  260  for the compressor disks  220  in stages fourteen through sixteen. This number may slightly increase if multiple sized forward weights  256  are used. 
     The balancing system disclosed herein may reduce the imbalance of the gas turbine engine leading to less vibration and more trouble-free operation. In particular, it was determined that the balancing system including the forward balancing system  255  and underplatform weights  260  may reduce vibration and may increase the reliability of the compressor rotor assembly  230 , the shaft  120 , and the associated bearings among other components. 
     Through research and development the location of the forward balancing holes  242  and the aft balancing holes  243  were determined. Misplacement of the forward balancing holes  242  and the aft balancing holes  243  may reduce the fatigue strength of the first stage compressor disk  221  and may reduce the overall reliability of the first stage compressor disk  221 . Variations in the cross-section throughout the first stage compressor disk  221 , such as variations resulting from the forward balancing holes  242  and aft balancing holes  243 , may lead to stress concentrations. These stress concentrations may cause cracking in the first stage compressor disk  221 . 
       FIG. 8  is a flowchart of a method for balancing the compressor rotor assembly  210 , which includes pre-balancing the aft weldment  212  at step  810  and may include pre-balancing the forward weldment  211  at step  820 . Balancing the compressor rotor assembly  210  may include the balancing system disclosed herein, which may include the embodiment shown in  FIGS. 2, 3, and 4 . 
     Pre-balancing the aft weldment  212  includes measuring the rotational balance of the aft weldment  212  at step  812 . Step  812  is followed by determining the number of underplatform weights  260  needed and the location for each underplatform weight  260  within a circumferential slot  236  of one the circumferentially loaded compressor disks  220  of the aft weldment  212  at step  814 . Step  814  is followed by mounting each underplatform weight  260  at the determined location at step  816 . Pre-balancing the aft weldment  212  occurs prior to joining the aft weldment  212  to the forward weldment  211 . Pre-balancing the aft weldment  212  also occurs prior to populating the aft weldment with compressor rotor blades. 
     Pre-balancing the forward weldment  211  includes measuring the rotational balance of the forward weldment at step  822 . Step  822  is followed by determining the number of forward weights  256  needed and the location for each forward weight  256  at step  824 . Step  824  is followed by mounting each forward weight  256  in the determined forward balancing hole  242  or aft balancing hole  243  at step  826 . 
     Pre-Balancing the forward weldment  211  may also include balancing the first stage compressor disk  221  prior to the first stage compressor disk  221  being welded to forward weldment  211 . Balancing the first stage compressor disk  221  may include measuring the rotational balance of the first stage compressor disk  221 . Measuring the rotational balance of the first stage compressor disk  221  may be followed by determining the number of forward weights  256  needed and the location for each forward weight  256 . The location for each forward weight  256  may be in a forward balancing hole  242  or in an aft balancing hole  243 . Balancing the first stage compressor disk  221  may also include mounting each forward weight  256  in the determined location. 
     The method for balancing the compressor rotor assembly  210  may also include balancing the assembled compressor rotor assembly at step  840 . The assembled compressor rotor assembly including the forward weldment and the aft weldment coupled together, and compressor rotor blades mounted to the forward weldment and the aft weldment.  FIG. 9  is a flowchart of a method for balancing the assembled compressor rotor assembly. Balancing the assembled compressor rotor assembly  210  includes measuring the compressor rotor balance at step  842 . Step  842  is followed by determining the number of forward weights  256  needed and the location for each forward weight  256  at step  844 . Step  842  is also followed by determining the number of underplatform weights  260  needed and the location for each underplatform weight  260  within a circumferential slot  236  of one the circumferentially loaded compressor disks  220  at step  846 . The underplatform weight  260  mounting locations may be at any circumferentially loaded compressor disk  220  rather than just at the circumferentially loaded compressor disks  220  located at the midplane and at the aft plane of the compressor rotor assembly  210 . The location of each underplatform weight  260  may determine which underplatform weight  260  is used, as the balancing system may use multiple underplatform weights  260 . 
     Steps  844  and  846  are followed by mounting each forward weight  256  in the determined forward balancing hole  242  or aft balancing hole  243 , and mounting each underplatform weight  260  at the determined location within the circumferential slots  236  at step  848 . 
     Balancing the assembled compressor assembly is performed after assembly of the compressor rotor assembly  210  including joining the forward weldment  211  to the aft weldment  212 , and mounting the compressor rotor blades to the forward weldment  211  and the aft weldment  212 . Compressor rotor blades may be weighed and sorted prior to mounting the compressor rotor blades at step  830 . Balancing the assembled compressor assembly may be performed before or after the gas turbine engine  100  is operated and tested. Balancing the assembled compressor assembly before testing the gas turbine engine  100  may be considered a shop balance, while balancing the assembled compressor assembly after testing may be considered a trim balance. 
     Steps  844  and  846  may also be followed by relocating compressor rotor blades based on the weight of the compressor rotor blades at step  850 . Step  850  may be limited to relocating the axial blades  229  of first stage compressor disk  221  and the second stage compressor disk. The axial blades  229  of the first two compressor stages may be the largest compressor rotor blades. These axial blades  229  may have a greater effect on the imbalance of the compressor rotor assembly  210 . 
     In some embodiments of the disclosed method, ¼ inch, ½ inch, or ¾ inch forward weights  256  are used in the aft balancing holes  243 , and ¼ inch or ½ inch forward weights  256  are used in the forward balancing holes  242 . In one embodiment, only the aft balancing holes  243  are used to pre-balance the first stage compressor disk  221 . 
     It is understood that the steps disclosed herein (or parts thereof) may be performed in the order presented or out of the order presented, unless specified otherwise. For example, pre-balancing the aft weldment  212  at step  810  may be performed prior to, after, or simultaneously to pre-balancing the forward weldment  211  at step  820 . 
     The preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. The described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. Hence, although the present disclosure, for convenience of explanation, depicts and describes a particular forward weldment, a particular aft weldment, particular forward weights, particular underplatform weights, and associated processes, it will be appreciated that other forward weldments, aft weldments, forward weights, underplatform weights and processes in accordance with this disclosure can be implemented in various other compressor rotor assemblies, configurations, and types of machines. Furthermore, there is no intention to be bound by any theory presented in the preceding background or detailed description. It is also understood that the illustrations may include exaggerated dimensions to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.