High pressure rotor disk

Abstract

A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships.

What is claimed is: 
     
      
       1. A gas turbine engine comprising: 
       
         a compressor section including a high pressure compressor and a low pressure compressor; 
         a combustor in fluid communication with the compressor section; 
         a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65. 
       
      
     
     
      
       2. The gas turbine engine as recited in  claim 1 , wherein the ratio (DAY) is between 1.35 and 1.55. 
      
     
     
       3. The gas turbine engine as recited in  claim 1 , wherein the ratio (D/W) is 1.45. 
     
     
       4. The gas turbine engine as recited in  claim 1  A gas turbine engine comprising:
 a compressor section including a high pressure compressor and a low pressure compressor; 
 a combustor in fluid communication with the compressor section; and 
 a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65, and wherein the disk includes an outer diameter (OD) related to the bore diameter (D) according to a ratio (OD/D) that is between 2.95 and 3.25. 
 
     
     
       5. The gas turbine engine as recited in  claim 4 , wherein the ratio (OD/D) is between 3.04 and 3.20. 
     
     
       6. The gas turbine engine as recited in  claim 4 , wherein the ratio (OD/D) is 3.15. 
     
     
       7. The gas turbine engine as recited in  claim 1  A gas turbine engine comprising:
 a compressor section including a high pressure compressor and a low pressure compressor; 
 a combustor in fluid communication with the compressor section; and 
 a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65, and wherein the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between 2.25 and 3.00. 
 
     
     
       8. The gas turbine engine as recited in  claim 7 , wherein the ratio (d/D) is between 2.50 and 2.75. 
     
     
       9. The gas turbine engine as recited in  claim 7 , wherein the ratio (d/D) is 2.69. 
     
     
      
       10. A rotor disk for a gas turbine engine comprising: 
       
         an outer diameter (OD) related to a bore diameter (D) according to a ratio (OD/D) that is between 2.95 and 3.25. 
       
      
     
     
       11. The rotor disk as recited in  claim 10 , wherein the ratio (OD/D) is between 3.04 and 3.20. 
     
     
       12. The rotor disk as recited in  claim 10 , wherein the ratio (OD/D) is 3.15. 
     
     
       13. The rotor disk as recited in  claim 10 , wherein the bore diameter (D) is related to a bore width (W) according to a ratio (D/W) between 1.25 and 1.65. 
     
     
       14. The rotor disk as recited in  claim 13 , wherein the ratio (D/W) is between 1.53 and 1.55. 
     
     
       15. The rotor disk as recited in  claim 13 , wherein the ratio (DAY) is 1.45. 
     
     
      
       16. The rotor disk as recited in  claim 10 , wherein the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between 2.25 and 3.00. 
      
     
     
       17. The rotor disk as recited in  claim 16 , wherein the ratio (d/D) is between 2.50 and 2.75. 
     
     
       18. The rotor disk as recited in  claim 16 , wherein the ratio (d/D) is 2.69. 
     
     
      
       19. A method of fabricating a rotor disk for a gas turbine engine comprising; 
       
         forming a bore including a bore diameter (D) and a live rim diameter (d) with a ratio (d/D) of the live rim diameter (d) to the bore diameter (D) being between 2.25 and 3.00 
         forming at least one lug for mounting a blade at the live rim diameter (d); and 
         forming an outer diameter (OD). 
       
      
     
     
      
       20. The method as recited in  claim 19 , including forming the disk to include a ratio (OD/D) of the outer diameter (OD) to the bore diameter (D) between 2.95 and 3.25. 
      
     
     
       21. The method as recited in  claim 19 , including forming a bore including a bore diameter (D) and a bore width (W) in a direction parallel to an axis of intended rotation, wherein the bore diameter (D) is related to the bore width (W) according to a ratio (D/W) that is between 1.25 and 1.65. 
     
     
       22. The gas turbine engine as set forth in claim 4, wherein the ratio (D/W) is between 1.35 and 1.55. 
     
     
       23. The gas turbine engine as recited in claim 4, wherein the ratio (D/W) is 1.45. 
     
     
       24. The gas turbine engine as set forth in claim 7, wherein the ratio (D/W) is between 1.35 and 1.55. 
     
     
       25. The gas turbine engine as recited in claim 7, wherein the ratio (D/W) is 1.45. 
     
     
       26. A gas turbine engine, comprising:
 a high speed shaft mounted for rotation about a longitudinal axis;   a low speed shaft coupled to a fan of a fan section, the low speed shaft mounted for rotation concentrically with the high speed shaft, the low speed shaft configured to rotate at a lower speed than the high speed shaft;   a rotor disk of a turbine attached to the high speed shaft, the rotor disk of the turbine having a first bore diameter (D 1 ) related to a first bore width (W 1 ) according to a first ratio (D 1 /W 1 ) between 1.25 and 1.65; and   a rotor disk of a compressor attached to the high speed shaft, the rotor disk of the compressor having a second bore diameter (D 2 ) related to a compressor outer diameter (OD 2 ) according to a second ratio (OD 2 /D 2 ) between 2.95 and 3.25;
 wherein the fan has less than 26 fan blades; and 
 further comprising a speed change device, wherein a low pressure turbine attached to the low speed shaft is configured to drive the fan through the speed change device at a lower speed than the low speed shaft. 
   
     
     
       27. The gas turbine engine as recited in claim 26, wherein the fan section has a low fan pressure ratio measured across the fan blades alone of less than 1.45 at a flight condition of 0.8 Mach and 35,000 ft., wherein the fan section has a low corrected fan tip speed less than 1,150 ft./sec at the flight condition of 0.8 Mach and 35,000 ft., and wherein the low corrected fan tip speed is an actual fan tip speed divided by ((Tram ° R)/(518.7° R)) 0.5 . 
     
     
       28. The gas turbine engine as set forth in claim 26, wherein the turbine is a two stage turbine. 
     
     
       29. The gas turbine engine as set forth in claim 26, wherein the ratio (D 1 /W 1 ) is between 1.35 and 1.55. 
     
     
       30. The gas turbine engine as recited in claim 26, wherein the ratio (D 1 /W 1 ) is 1.45. 
     
     
       31. The gas turbine engine as recited in claim 26, wherein the ratio (OD 2 /D 2 ) is between 3.04 and 3.20. 
     
     
       32. The gas turbine engine as recited in claim 26, wherein the ratio (OD 2 /D 2 ) is 3.15. 
     
     
       33. A gas turbine engine, comprising:
 a high speed shaft mounted for rotation about a longitudinal axis;   a low speed shaft coupled to a fan of a fan section, the low speed shaft mounted for rotation concentrically with the high speed shaft, the low speed shaft configured to rotate at a lower speed than the high speed shaft;   a rotor disk of a turbine attached to the high speed shaft, the rotor disk of the turbine having a first bore diameter (D 1 ) related to a first bore width (W 1 ) according to a first ratio (D 1 /W 1 ) between 1.25 and 1.65;   a rotor disk of a compressor attached to the high speed shaft, the rotor disk of the compressor having a second bore diameter (D 2 ) related to a compressor outer diameter (OD 2 ) according to a second ratio (OD 2 /D 2 ) between 2.95 and 3.25;
 wherein the fan has less than 26 fan blades; and 
 wherein the fan section has a low fan pressure ratio measured across the fan blades alone of less than 1.45 at a flight condition of 0.8 Mach and 35,000 ft., wherein the fan section has a low corrected fan tip speed less than 1,150 ft./sec at the flight condition of 0.8 Mach and 35,000 ft., and wherein the low corrected fan tip speed is an actual fan tip speed divided by ((Tram ° R)/(518.7° R)) 0.5 . 
   
     
     
       34. The gas turbine engine as set forth in claim 33, wherein the turbine is a two stage turbine. 
     
     
       35. The gas turbine engine as set forth in claim 33, wherein the ratio (D 1 /W 1 ) is between 1.35 and 1.55. 
     
     
       36. The gas turbine engine as recited in claim 33, wherein the ratio (D 1 /W 1 ) is 1.45. 
     
     
       37. The gas turbine engine as recited in claim 33, wherein the ratio (OD 2 /D 2 ) is between 3.04 and 3.20. 
     
     
       38. The gas turbine engine as recited in claim 33, wherein the ratio (OD 2 /D 2 ) is 3.15.