Patent Document

This invention was made with Government support under F33657-99-2051-0008 awarded by the United States Air Force. The Government has certain rights in this invention. 

   FIELD OF THE INVENTION 
   This invention relates generally to gas turbine engines, and more particularly to gas turbine disk slots. 
   BACKGROUND OF THE INVENTION 
   Gas turbine engine disks commonly have slots for attaching blades which are generally axially oriented. These slots have a profile which mates with the roots of the blades, and have a configuration which will retain the blades in the slots under the applied centrifugal forces incurred in operation of the engine. The slot profiles are often of a “fir-tree” configuration to increase the load bearing area in the slot, although other configurations are also employed. 
   The turbine disk slots for mounting turbine blades typically have a a sharp edge entrance for airflow. The sharp edge entrance causes an unfavorable airflow separation at the slot inlet, and undesirably generates an increased heat transfer rate because of airflow reattachment. 
   SUMMARY OF THE INVENTION 
   In one aspect of the present invention, a gas turbine disk assembly comprises a turbine disk defining a plurality of turbine disk slots for accommodating turbine blades. The plurality of turbine disk slots each include an inlet having a rounded periphery at a bottom portion thereof. 
   In another aspect of the present invention, a gas turbine engine comprises a compressor section, a combustion section disposed downstream from the compressor section, and a turbine section disposed downstream from the combustion section. The turbine section includes a turbine disk defining a plurality of turbine disk slots for accommodating turbine blades. The plurality of turbine disk slots each include an inlet having a rounded periphery at a bottom portion thereof. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a side elevation schematic view of a gas turbine engine with the engine partially broken away to show a portion of the turbine section of the engine. 
       FIG. 2  is a related art partial cross-sectional, side elevation view of a gas turbine engine showing the location of turbine disk slots. 
       FIG. 3  is an enlarged front perspective view of the gas turbine engine of  FIG. 2  showing turbine disk slots. 
       FIG. 4  is an enlarged front perspective view of turbine disk slots embodying the present invention. 
       FIG. 5  is a cross-sectional, side view of a turbine disk slot embodying the present invention. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     FIG. 1  is a side elevation, simplified view of an example of a gas turbine engine  10 . The view is partially broken away to show elements of the interior of the engine. The engine  10  includes a compression section  12 , a combustion section  14  and a turbine section  16 . An airflow path  18  for working medium gases extends axially through the engine  10 . The engine  10  includes a first, low pressure rotor assembly  22  and a second, high pressure rotor assembly  24 . The high pressure rotor assembly  24  includes a high pressure compressor  26  connected by a shaft  28  to a high pressure turbine  32 . The low pressure rotor assembly  22  includes a fan and low pressure compressor  34  connected by a shaft  36  to a low pressure turbine  38 . During operation of the engine  10 , working medium gases are flowed along the airflow path  18  through the low pressure compressor  26  and the high pressure compressor  34 . The gases are mixed with fuel in the combustion section  14  and burned to add energy to the gases. The high pressure working medium gases are discharged from the combustion section  14  to the turbine section  16 . Energy from the low pressure turbine  38  and the high pressure turbine  32  is transferred through their respective shafts  36 ,  28  to the low pressure compressor  34  and the high pressure compressor  26 . 
   With reference to  FIG. 2 , a partial cross-sectional view of a turbine section is generally indicated by the reference number  40 . Within the area enclosed by circle  42 , the turbine section includes a plurality of turbine blades mounted on turbine disk slots. Turning to the enlarged view of  FIG. 3 , conventional turbine disk slots  44  for mounting turbine blades typically have a non-rounded or otherwise sharp-edged periphery  46  at a bottom portion  48  relative a front face  43   f  of a turbine disk  43  which produces a sharp edge entrance for airflow. The sharp edge entrance causes an unfavorable airflow separation at the slot inlet, and undesirably generates an increased heat transfer rate because of airflow reattachment. 
   Turning now to  FIG. 4 , a turbine disk  50  defines a plurality of turbine disk slots  52  embodying the present invention. Each turbine disk slot  52  defined by the turbine disk  50  includes an inlet  54  having a rounded periphery  56  relative a front face  50   f . The rounded periphery  56  is generally located at a bottom portion  58  of each turbine slot  52  disk. An extra machining process is employed to generate the rounded periphery  56  of the inlet  54 . A radius (r) of the rounded periphery  56  is based on a hydraulic diameter (D h ) of the slot  52 , which in turn is based on a cooling airflow area between the bottom portion  58  of the slot  52  and a bottom of a turbine blade. To maximize the effectiveness of the inlet  54  having the rounded periphery  56 , an r/D h  ratio of 0.16 is preferably used, but an r/D h  ratio that is either greater or lesser than 0.16 can be used without departing from the scope of the present invention. Because of the nature of the design, the entire edge of the inlet  54  of the slot  52  cannot be rounded. Instead, the full radius of the rounded periphery  56  extends approximately 180 degrees and then tapers down to points  60  as shown in  FIG. 4 . 
     FIG. 5  illustrates a cross-section of a turbine disk  70  in accordance with the present invention. The turbine disk  70  defines a slot  72  including a rounded periphery  74  at a turbine disk slot entrance adjacent to an aft face  76  of a forward cover plate  78 . The turbine disk  70  further defines a plurality of blade cooling passages  80  disposed on an opposite side of the turbine disk  70  relative to the slot  72 . 
   It has been discovered that a rounded periphery of an inlet of a turbine disk slot offers the following advantages: 
   1) Reduces inlet pressure loss because of the sharp edge entrance; 
   2) Minimizes and/or eliminates flow separation at the inlet; and 
   3) Reduces the increased heat transfer rate because of flow reattachment. 
   As will be recognized by those of ordinary skill in the pertinent art, numerous modifications and substitutions can be made to the above-described embodiment of the present invention without departing from the scope of the invention. Accordingly, the preceding portion of this specification is to be taken in an illustrative, as opposed to a limiting sense.

Technology Category: 2