Patent Publication Number: US-8540482-B2

Title: Rotor assembly for gas turbine engine

Description:
BACKGROUND 
     This application relates generally to a gas turbine engine, and more particularly to a rotor assembly for a gas turbine engine. 
     Gas turbine engines include rotor assemblies having a plurality of rotating airfoils or blades. The rotor assemblies, especially in the high pressure compressor section, are subjected to a large strain range (e.g., creep-fatigue mechanism) during operation. The large strain range is induced during the engine flight cycle and is at least partially attributable to the extreme temperature differences between the relatively hot primary flowpath airflow that is communicated through the compressor section and the relatively cool compressor rotor assembly components. The large strain range acting on the rotor assembly can result in a relatively low fatigue life of such components. 
     Attempts to improve component fatigue life of the rotor assembly have included extracting primary flowpath air to cool the inner diameters of the compressor rotor assembly. However, this solution can compromise compressor efficiency. 
     SUMMARY 
     A rotor assembly for a gas turbine engine includes a rotor airfoil and a first rotor disk. The rotor airfoil extends along a radial axis. The first rotor disk includes an outer rim, a bore and a web extending between the outer rim and the bore. The first rotor disk is axially offset from the radial axis of the rotor airfoil. 
     In another exemplary embodiment, a gas turbine engine includes a section having alternating rows of rotating rotor airfoils and static stator vanes. A rotor assembly includes a first rotor disk and a second rotor disk. The first rotor disk and the second rotor disk each include a plurality of rotor airfoils. Each of the rotor airfoils are integrally formed with a bladed ring that is radially trapped between the first rotor disk and the second rotor disk. 
     In another exemplary embodiment, a method for providing a rotor assembly for a gas turbine engine includes positioning a rotor disk of the rotor assembly at a position that is axially offset relative to a radial axis of a rotor airfoil of the rotor assembly. 
     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 
         FIG. 1  illustrates a simplified cross-sectional view of a standard gas turbine engine; 
         FIG. 2  illustrates a cross-sectional view of a portion of the gas turbine engine; 
         FIGS. 3A-3C  illustrate additional cross-sectional views of a portion of the gas turbine engine; 
         FIG. 4  illustrates an example rotor assembly that includes a bladed ring; and 
         FIG. 5  illustrates another example rotor assembly including a bladed ring. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  shows a gas turbine engine  10 , such as a turbofan gas turbine engine, that is circumferentially disposed about an engine centerline (or axial centerline axis)  12 . The gas turbine engine  10  includes a fan section  14 , a compressor section  15  having a low pressure compressor  16  and a high pressure compressor  18 , a combustor  20 , and a turbine section  21  including a high pressure turbine  22  and a low pressure turbine  24 . This application can also extend to engines without a fan, and with more or fewer sections. 
     As is known, air is compressed in the low pressure compressor  16  and the high pressure compressor  18 , is mixed with fuel and burned in the combustor  20 , and is expanded in the high pressure turbine  22  and the low pressure turbine  24 . Rotor assemblies  26  rotate in response to the expansion, driving the low pressure and high pressure compressors  16 ,  18  and the fan section  14 . The low and high pressure compressors  16 ,  18  include alternating rows of rotating compressor rotor airfoils or blades  28  and static stator vanes  30 . The high and low pressure turbines  22 ,  24  include alternating rows of rotating turbine rotor airfoils or blades  32  and static stator vanes  34 . 
     It should be understood that this view is included simply to provide a basic understanding of the sections of a gas turbine engine  10  and not to limit the disclosure. This disclosure extends to all types of gas turbine engines  10  for all types of applications. 
       FIG. 2  shows a portion of the compressor section  15  of the gas turbine engine  10 . In this example, the portion shown is the high pressure compressor  18  of the gas turbine engine  10 . However, this disclosure is not limited to the high pressure compressor  18 , and could extend to other sections of the gas turbine engine  10 . 
     The illustrated compressor section  15  includes multiples stages of alternating rows of rotor assemblies  26 A- 26 H and stator vanes  30 A- 30 H. In this example, eight stages are shown, although the compressor section  15  could include more or less stages. The stator vanes  30 A- 30 H extend between each rotor assembly  26 . Each rotor assembly  26  includes a rotor airfoil  28  and a rotor disk  36 . The rotor disks  36  include an outer rim  38 , a bore  40 , and a web  42  that extends between the outer rim  38  and the bore  40 . 
     At least a portion of the rotor assemblies  26  include an axially offset rotor disk  36 . That is, the rotor disk  36  is axially offset (See rotor assembly  26 F) from a radial axis R of the rotor airfoil  28 . It should be understood that the axial offset of the illustrated rotor disks  36  is not shown to the scale it would be in practice. Instead, the axial offset is shown enlarged to better illustrate the positioning of the rotor disks  36  relative to the radial axis R of the rotor airfoils  28 . The actual distance of the axial offset will vary depending upon a number of factors including but not limited to airfoil positioning, the number of stages in compressor section  15 , bleed location requirements, the axial length of the compressor section  15  and the spacing requirements between adjacent rotor disks  36 . 
     In this example, the rear stages of the high pressure compressor  18  include rotor assemblies  26 E- 26 H having axially offset rotor disks  36 . However, each rotor assembly  26 A- 26 H could include an axially offset rotor disk  36 , or the axial displacement could be applied to only a portion of the stages (such as depicted in  FIG. 2 ). The stages that do not include an axially offset rotor disk  36  (in this example, rotor assemblies  26 A- 26 D) can include standard axial attachments in which the rotor disks  36  are substantially in-line with the radial axis R of the rotor airfoils  28 . 
     A tie shaft  51  is connected to the rotor assemblies  26 A- 26 H. The tie shaft  51  can be preloaded to maintain tension on the plurality of rotor assemblies  26 A- 26 H. The tie shaft  51  extends between a forward hub  53  and an aft hub  55 . In this example, the tie shaft  51  is threaded through the forward hub  53  and is snapped into the rotor disk  36  of the rotor assembly  26 H. Once connected between the forward hub  53  and the aft hub  55 , the preloaded tension on the tie shaft  51  is maintained with a nut  57 . 
       FIG. 3A  illustrates a portion of the compressor section  15  that includes the rotor assembly  26 F (and the rotor disk  36 E of adjacent rotor assembly  26 E). Each of the outer rim  38 , the bore  40  and web  42  of the rotor disk  36 F of rotor assembly  26 F are axially offset from the radial axis R of the rotor airfoil  28 . In this way, the outer rim  38 , the bore  40  and the web  42  of the axially offset rotor disk  36 F are each generally radially inward from the stator vane  30  and extend along a radial axis R 2  of the stator vane  30 . In one example, the outer rim  38 , the bore  40  and the web  42  are generally coaxial with the stator vane  30 . The outer rim  38  can also include a seal coating, such as Zirconium Oxide, to seal the interface between the stator vane  30  and the outer rim  38  to reduce the potential for damage to the stator vane  30 . The rotor disks  36  are axially displaced in a downstream direction (DD) relative to the rotor airfoils  28 , in this example. In another example embodiment, the rotor disks  36  are axially displaced in an upstream direction (UD) relative to the rotor airfoils  28  (see  FIG. 3B ). 
     Referring again to  FIG. 3A , in this example the radial axis R 2  that extends through the rotor disk  36  of rotor assembly  26 F is axially offset from the radial axis R of the rotor airfoil  28  by a distance X. An axially outermost portion  29  of the web  42  is axially offset from an axially outermost portion  31  of the rotor airfoil  28  by a distance X 2  such that no portion of the web  42  is positioned directly radially inwardly from the rotor airfoil  28 . In other words, the entire web  42  is fully offset from the radial axis R of the rotor airfoil  28  in a direction away from the rotor airfoil  28 . 
     The portion of the rotor assemblies  26  that include axially offset rotor disks  36  further include a bladed ring  44  (e.g., bling). In the example embodiment, the bladed rings  44  and the rotor airfoils  28  are integrally formed as a single, continuous piece with no mechanical attachments. That is, the rotor airfoils  28  are detached from a traditional integrally bladed rotor (IBR) and are instead formed as a single, continuous piece with the bladed rings  44 . The airfoils  28  extend radially outwardly from the bladed rings  44 . In this example, the axially outermost portion  29  of the web  42  is axially offset from an axially outermost portion  33  of the bladed ring  44 . 
     The bladed rings  44  can include a tangential style attachment which conforms to the profile of adjacent portions of the rotor disks  36  to radially trap the bladed rings  44 , and therefore, the rotor airfoils  28 , in the radial direction. In one example, the bladed rings  44  are sandwiched between the outer rims  38  of adjacent rotor disks  36 . Here, the bladed ring  44  is radially trapped between the rotors disk  36 E (e.g., a first rotor disk) and rotor disk  36 F (e.g., a second rotor disk) of rotor assemblies  26 E,  26 F. The bladed rings  44  can also be trapped between the webs  42  of adjacent rotor disks  36 . Friction forces between the bladed ring  44  and adjacent rotor disks  36  minimize any circumferential movement of the bladed ring  44  relative to the rotor disk  36 . The bladed rings  44  enable the airfoils  28  to be decoupled from the rotor disks  36 , thereby improving part life by relocating the notch feature (e.g., transition area of leading end and trailing end fillets of the airfoils  28  and the rotor disks  36 ) off of the rotor disks  36 . 
     The axially offset rotor disks  36  further include a spacer  46  that extends from the rotor disk  36 . In this example, a catenary spacer  46  extends from the web  42  of the rotor disk  36 . In another example, the spacer  46  is a cylindrical or conical spacer. The spacers  46  are positioned radially inwardly from the bladed rings  44  to provide radial load support for the rotor airfoils  28 . The spacers  46  are integrally formed with the rotor disk  36 . In one example embodiment, the spacers  46  extend in the upstream direction UD from the rotor disks  36 . In another example, the spacers  46  extend in the downstream direction DD from the rotor disks  36  (See  FIG. 3B ). 
     Referring to  FIG. 3C , the axial displacement of the outer rims  38 , bores  40  and webs  42  of the rotor disks  36  relative to the rotor airfoils  28  alters the fundamental load path of the airfoil radial pull (RP) and creates a non-direct path for the radial pull RP. For example, as best illustrated by rotor assembly  26 G, the modified load path runs in the radial direction D 1  along the span of the rotor airfoil  28 , then axially in a direction A 1  aft of the rotor airfoil  28 , and then radially along the rotor disk  36  in the direction D 2 . In other words, the radial pull of each rotor airfoil  28  runs axially along the airfoil  28  prior to moving down the web  42  and into the bore  40  of the rotor disk  36 . Accordingly, the modified load path minimizes the strain range that each rotor assembly  26  is subjected to during gas turbine engine  10  operation and otherwise enhances rotor response without the need to extract primary flowpath airflow to cool each rotor assembly  26  by effectively decoupling the rotor airfoils  28  from the rotor disks  36 . 
       FIG. 4  illustrates an example rotor assembly  26  including a bladed ring  44  that is represented as a full hoop ring. In this example embodiment, the bladed ring  44  extends circumferentially over 360° to form the full hoop ring. A plurality of rotor airfoils  28  are integrally formed with the full hoop bladed ring  44  as a single, continuous piece with no mechanical attachments. 
       FIG. 5  illustrates another example rotor assembly  126 . The rotor assembly  126  includes a segmented bladed ring  144 . Rather than extending in a full hoop, the segmented bladed ring  144  is apportioned into a plurality of separate components  144 A- 144 N that provide greater compliance to the rotor assembly  126 . The actual number of segmentations will vary depending upon design specific parameters. A plurality of rotor airfoils  28  are integrally formed with each segmented portion of the segmented bladed ring  144 . Any number of clusters of rotor airfoils  28  can be formed onto each component  144 A- 144 N of the segmented bladed ring  144 , including a single airfoil  28  per component  144 A- 144 N. 
     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 would 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.