Patent Publication Number: US-2023160395-A1

Title: Rotor Disk Having a Curved Rotor Arm for an Aircraft Gas Turbine

Description:
This claims the benefit of German Patent Application DE102021123173.6, filed on Sep. 7, 2021 and which is hereby incorporated by reference herein 
     The present invention relates to a rotor disk for a compressor of a gas turbine, in particular an aircraft gas turbine. 
     Directional words such as “axial,” “axially,” “radial,” “radially,” and “circumferential” are taken with respect to the machine axis of the gas turbine, unless explicitly or implicitly indicated otherwise by the context. 
     SUMMARY OF THE INVENTION 
     In a compressor, in particular a high-pressure compressor, of an (aircraft) gas turbine, axial clamping forces are transmitted by the rotor arm of one rotor disk to an axially adjacent rotor disk, so that the plurality of rotor disks, which are clamped against one another by axial tension forces, can be stabilized. In the case of axially adjacent rotor disks which differ significantly in diameter, the problem arises that a relatively large radial distance must be spanned by the rotor arm to support the acting clamping forces. Therefore, existing rotor arms have at least one sharp bend and possibly also variations in material thickness, which results in undesired stresses in the rotor arm. 
     It is an object of the invention to provide a rotor disk that enables axial force transmission with reduced stresses. 
     Accordingly, there is provided a rotor disk for a compressor of a gas turbine, in particular an aircraft gas turbine, the rotor disk having 
     a main body,   at least one rotor arm projecting from the main body in the axial direction,   the rotor arm having, in a sectional view taken in a sectional plane defined by the axial direction and the radial direction: a beginning portion merging into the main body;   an end portion remote from the main body and forming a free end in the axial direction, the beginning portion and the end portion being interconnected by an intermediate portion.   

     It is provided that the intermediate portion be curved with at least one radius of curvature. 
     The curved or bent configuration of the intermediate portion allows axially acting (clamping) forces to be transmitted with little stress. The curvature avoids sharp bends in the shape of the rotor arm, where local stress peaks may occur in the bend regions, which may undesirably result in material fatigue. It should be noted that the radius of curvature and a respective center of the curved intermediate portion are located in the specified sectional plane. 
     In the rotor disk, the beginning portion, the end portion and the intermediate portion may have substantially the same rotor arm thickness. In other words, the two substantially straight beginning and end portions and the curved intermediate portion together form a rotor arm of substantially the same thickness along its axial length. The rotor arm thickness may be about 0.3 cm to 1.3 cm. 
     In the rotor disk, at least one radially outwardly directed sealing fin may be disposed on the rotor arm. The at least one sealing fin is disposed radially opposite a sealing element which is attached to a stator or stator vane ring and usually has a honeycomb structure and into which the sealing fin may rub in certain operating conditions of the gas turbine in order to provide a seal. 
     In the rotor disk, the radius of curvature of the intermediate portion may be from about 2 cm to 6 cm, in particular about 2.5 cm to 5.1 cm. 
     A rotor blade disk for a compressor of a gas turbine, in particular an aircraft gas turbine, may have a rotor disk as described above, the rotor blade disk having a plurality of rotor blades arranged adjacent one another in the circumferential direction and connected to the rotor disk. 
     In the rotor blade disk, the rotor disk and the rotor blades may be formed integrally with each other, in particular as a blisk. 
     A compressor, in particular a high-pressure compressor, for a gas turbine, in particular an aircraft gas turbine, may have at least one rotor disk as described above or at least one rotor blade disk as described above. 
     An aircraft gas turbine may be equipped with such a compressor, in particular a high-pressure compressor. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention will now be described by way of example, and not by way of limitation, with reference to the accompanying drawings. 
         FIG.  1    is a simplified schematic representation of an aircraft gas turbine; 
         FIG.  2    is a simplified sectional view showing a portion of a compressor, specifically the region between two rotor blade disks; 
         FIG.  3    is an enlarged view of a rotor arm of  FIG.  2   . 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    shows, in simplified schematic form, an aircraft gas turbine  10 , illustrated, merely by way of example, as a turbofan engine. Gas turbine  10  includes a fan  12  surrounded by a schematically indicated casing  14 . Disposed downstream of fan  12  in the axial direction AR of gas turbine  10  is a compressor  16  that is accommodated in a schematically indicated inner casing  18  and may be single-stage or multi-stage. Disposed downstream of compressor  16  is combustor  20 . The flow of hot exhaust gas exiting the combustor then flows through the downstream turbine  22 , which may be single-stage or multi-stage. In the present example, turbine  22  includes a high-pressure turbine  24  and a low-pressure turbine  26 . A hollow shaft  28  connects high-pressure turbine  24  to compressor  16 , in particular a high-pressure compressor  29 , so that they are jointly driven or rotated. Another shaft  30  located further inward in the radial direction RR of the turbine connects low-pressure turbine  26  to fan  12  and to a low-pressure compressor  32  so that they are jointly driven or rotated. Disposed downstream of turbine  22  is an exhaust nozzle  33 , which is only schematically indicated here. 
     In the illustrated example of an aircraft gas turbine  10 , a turbine center frame  34  is disposed between high-pressure turbine  24  and low-pressure turbine  26  and extends around shafts  28 ,  30 . Hot exhaust gases from high-pressure turbine  24  flow through turbine center frame  34  in its radially outer region  36 . The hot exhaust gas then flows into an annular space  38  of low-pressure turbine  26 . Compressors  29 ,  32  and turbines  24 ,  26  are represented, by way of example, by rotor blade rings  27 . For the sake of clarity, the usually present stator vane rings  31  are shown, by way of example, only for compressor  32 . 
     The invention will now be described in more detail with simultaneous reference to  FIGS.  2  and  3   ,  FIG.  3    being an enlarged view of the portion designated III in  FIG.  2   . 
       FIG.  2    shows a rotor disk  40  having a main body  42  and a rotor arm  44 . Rotor arm  44  is connected to main body  42 . When viewed relative to the direction of air flow LR through an annular space  46  schematically indicated by short-dashed lines, another rotor disk  40   a  is disposed upstream of rotor disk  40 . The two rotor disks  40 ,  40   a  are clamped against one another. 
     Rotor arm  44  of rotor disk  40  bears against rotor disk  40   a  in axial direction AR and radial direction RR, which allows transmission of acting forces of the axial clamping. A rotor blade  48  is connected to rotor disk  40 . Rotor disk  40   a  also has a rotor blade  48   a  connected thereto. With regard to rotor blades  48  and  48   a , it should be noted that these blades may be formed integrally with the respective rotor disk  40  and  40   a , in particular as what is known as a blisk. Alternatively, however, it is also conceivable that rotor disks  40  and  40   a  may have openings formed therein in which rotor blade roots of rotor blades may be interlockingly received. 
     Rotor arm  44  can be divided into a beginning portion  44   a , an end portion  44   e , and an intermediate portion  44   z , as shown in  FIG.  3   . Beginning portion  44   a  is connected to main body  42  and extends obliquely to axial direction AR and to radial direction RR. Beginning portion  44   a  is substantially straight. 
     End portion  44   e  rests against the axially forward rotor disk  40   a . End portion  44   e  extends substantially parallel to axial direction AR and substantially orthogonal to radial direction RR. Due to the end portion  44   e  extending substantially parallel to axial direction AR, axially acting forces can be optimally transmitted and supported. In  FIGS.  2  and  3   , the flow of force along axial direction AR in rotor arm  44  and rotor disks  40 ,  42   a  is indicated in simplified form by a dash-dotted line KF. 
     The intermediate portion  44   z  extending between beginning portion  44   a  and end portion  44   e  is curved or bent and has an inner radius Ri and an outer radius Ra relative to a center MP. The two radii Ri and Ra are selected such that intermediate portion  44   z  has a substantially uniform rotor arm thickness RD. Beginning portion  44   a  and end portion  44   e  also have a rotor arm thickness RD that is substantially uniform. In other words, the entire rotor arm  44  has a continuous thickness RD that is maintained substantially constant. Radius of curvature Ri or Ra of intermediate portion  44   z  has a length of about 2 cm to 6 cm, in particular of about 2.5 cm to 5.1 cm. The substantially constant thickness RD of rotor arm  44  is about 0.3 to 1.3 cm. 
     The selected arrangement of the obliquely extending beginning portion  44   a  and the adjoining curved intermediate portion  44   z  allows forces acting due to the axial clamping to be optimally transmitted with little stress from the rotor disk  40  of larger diameter to the rotor disk  40   a  of smaller diameter, without local stress peaks occurring in rotor arm  44 , and specifically in intermediate portion  44   z . 
     Rotor arm  44  may have at least one sealing fin  50  provided thereon which, in an assembled state of a compressor, is disposed opposite an abradable sealing element of a stator or stator vane ring. 
     A rotor disk  40  having the curved rotor arm  44 , as described with reference to  FIGS.  2  and  3   , may be disposed, for example, in a high-pressure compressor  29  of an aircraft gas turbine  10 , as shown in  FIG.  1   . The rotor blades  48  and  48   a  may form part of a rotor blade ring  27  indicated in  FIG.  1   .  
     
       
         
           
               
               
             
               
                 LIST OF REFERENCE NUMERALS 
               
             
            
               
                 
                   10 
                 
                 aircraft gas turbine 
               
               
                 
                   12 
                 
                 fan 
               
               
                 
                   14 
                 
                 casing 
               
               
                 
                   16 
                 
                 compressor 
               
               
                 
                   18 
                 
                 inner casing 
               
               
                 
                   20 
                 
                 combustor 
               
               
                 
                   22 
                 
                 turbine 
               
               
                 
                   24 
                 
                 high-pressure turbine 
               
               
                 
                   26 
                 
                 low-pressure turbine 
               
               
                 
                   28 
                 
                 hollow shaft 
               
               
                 
                   29 
                 
                 high-pressure compressor 
               
               
                 
                   30 
                 
                 shaft 
               
               
                 
                   31 
                 
                 stator vane ring 
               
               
                 
                   32 
                 
                 low-pressure compressor 
               
               
                 
                   33 
                 
                 exhaust nozzle 
               
               
                 
                   34 
                 
                 turbine center frame 
               
               
                 
                   36 
                 
                 radially outer region 
               
               
                 
                   38 
                 
                 annular space 
               
               
                   40 ,40a 
                 rotor disk 
               
               
                 
                   42 
                 
                 main body 
               
               
                 
                   44 
                 
                 rotor arm 
               
               
                 
                   44 
                   a 
                 
                 beginning portion 
               
               
                 
                   44 
                   e 
                 
                 end portion 
               
               
                 
                   44 
                   z 
                 
                 intermediate portion 
               
               
                 
                   46 
                 
                 annular space 
               
               
                   48 . 48   a   
                 rotor blade 
               
               
                 
                   50 
                 
                 sealing fin 
               
               
                 AR 
                 axial direction 
               
               
                 LR 
                 direction of air flow 
               
               
                 MP 
                 center 
               
               
                 Ra 
                 outer radius 
               
               
                 RD 
                 rotor arm thickness 
               
               
                 Ri 
                 inner radius 
               
               
                 RR 
                 radial direction