Patent Publication Number: US-7210906-B2

Title: Internally cooled gas turbine airfoil and method

Description:
TECHNICAL FIELD 
   The invention relates to internally cooled airfoil structures within a gas turbine engine. 
   BACKGROUND 
   The design of gas turbine airfoils is the subject of continuous improvement, since design directly impacts cooling efficiency. In some gas turbine designs, the turbine airfoil chord is long relative to the airfoil length, resulting in a “short” &amp; “fat” airfoil. Traditional serpentine cooling passages need either to have increased number of turns to account for the additional area to cool, which results in increased pressure losses, or the individual passages must simply be wider, which leads to “dead” zones in which air tends to stagnate undesirably, thereby reducing cooling efficiency. Therefore, there continues to be a need for improved cooling for internally cooled gas turbine airfoils. 
   SUMMARY 
   In one aspect the invention provides an internally cooled airfoil for a gas turbine engine, the airfoil having a hollow section and a trailing edge, the airfoil comprising: 
   a plurality of partition walls located in the hollow section and defining internal cooling air passages, at least some of the passages extending from an inlet to at least one outlet adjacent to the trailing edge; and 
   at least one crossover located in the hollow section and being adjacent to the outlet, the crossover generally extending radially in the hollow section and having a distal end portion on an end of the airfoil distally opposite the inlets of the passages, the crossover being in fluid communication with at least two of said passages that are substantially parallel to each other, one of which said parallel passages being dedicated to supplying cooling air to the distal end portion of the crossover. 
   In another aspect the invention provides an internally cooled gas turbine airfoil comprising: 
   a hollow airfoil body having a first end, a second end and a trailing edge extending therebetween; and 
   a plurality internal passages defined in the hollow airfoil body, the passages including at least two passages extending from distinct inlets in the first end and in parallel communication with an exit plenum defined in the hollow airfoil body adjacent to the trailing edge, wherein the passages are disposed side-by-side and wherein a first one of said at least two passages communicates directly with a substantially larger portion of the exit plenum than a second. 
   In a further aspect the invention provides an airfoil for use in a gas turbine engine, the airfoil comprising a hollow section with passages adapted to direct an internally-circulating flow of cooling air, the airfoil including a trailing edge and at least one exit plenum adjacent to the trailing edge, the hollow section including partition walls dividing adjacent passages, the adjacent passages including at least two fluidly parallel cooling air paths upstream of and communicating in parallel with the exit plenum. 
   In a still further aspect the invention provides a method of cooling an airfoil of a gas turbine engine using an internally-circulating flow of cooling air, the airfoil including a trailing edge and at least one exit plenum adjacent to the trailing edge, the method comprising: 
   dividing the flow of cooling air in at least two fluidly parallel cooling air paths; and then 
   directing the cooling air paths parallelly through the exit plenum. 
   Still other aspects and inventions will be apparent in the appended description and figures. 

   
     DESCRIPTION OF THE DRAWINGS 
       FIG. 1  shows a generic gas turbine engine to illustrate an example of a general environment in which the invention can be used. 
       FIG. 2  is an isometric view of a turbine blade according to the invention, a portion of the blade being cut away to show some of the internal cooling passages in the airfoil thereof. 
       FIG. 3  is an enlarged side view of the internal passages shown in  FIG. 2 . 
       FIG. 4  is a view similar to  FIG. 3 , showing another embodiment. 
       FIG. 5  is a side view of a cooling passage which does not include the present invention. 
   

   DETAILED DESCRIPTION 
     FIG. 1  illustrates an example of a gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan  12  through which ambient air is propelled, a multistage compressor  14  for pressurizing the air, a combustor  16  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section  18  for extracting energy from the combustion gases. 
     FIG. 2  shows a turbine blade having an airfoil  20  according to one embodiment of the invention. Although a turbine blade is shown in  FIG. 2 , the present invention can be used in a compressor and turbine blades and vanes. The airfoil  20  extends from a root section  22  and comprises a hollow section  24  generally radially extending from the root section  22 . The root section  22  is mounted into a corresponding recess of a rotary support structure of the turbine disc (not shown). The shape of the hollow section  24  may depend on its location within the gas turbine engine  10 , the operating parameters of the gas turbine engine  10 , etc. 
   The root section  22  of the turbine blade includes one or more cooling air inlets receiving cooling air from a plenum located on the upstream side of the turbine disk. The cooling air inlet or inlets lead to the interior of the hollow section  24 . In use, relatively cool air, bled typically from the compressor  14 , is fed to the cooling air plenum through conventional means (not shown) and then enters through the root section  22 . The air enters internal passages (described below) to thereby cool the airfoil  20 . 
   Air exits through holes (not shown) provided for surface film cooling and through one or more preferably, a plurality of trailing edge exit holes  26  located adjacent to the trailing edge  28  of the airfoil  20 . 
     FIG. 3  illustrates an enlarged portion of  FIG. 2 . The hollow section  24  comprises a plurality of partition walls  30  configured and disposed to define internal air cooling passages  32 ,  34 ,  36  and  38  having respective inlets  32 A,  34 A,  36 A and  38 A. 
   Passages  36  and  38  are preferably independent from each other (i.e. in parallel) from inlet  36 A/ 38 A to intermediate plenum  41  and/or exit plenum  25 , but if desired may be in partial fluid communication using aperture(s) or other openings  60 , as shown in  FIG. 4 , depending on the design and operational requirements.  FIG. 4  schematically illustrates that one (or more) aperture(s)  60  can optionally be provided in one or more of the partition walls  30 . 
   In this application the term “crossover” is used to describe an internal wall which contains numerous openings permitting air to pass therethrough. The flow of cooling air is controlled by adjusting the size and number of these openings. At least one crossover is located at the rear of the hollow section  24 . The illustrated airfoil  20  is shown with a first crossover  40  and a second crossover  42 . The second crossover  42  is located between the first crossover  40  and the trailing edge  28 , and an intermediate plenum  41  is located therebetween. They are generally extending radially inside the hollow section  24 . An exit plenum  25  is interposed between second crossover  42  and exit holes  26 . 
   The first crossover  40  comprises what is generally referred to as a distal end portion  44 , which is located near the end of the first crossover  40  which is remote or distally opposite from inlets  36 A,  38 A of passages  36  and  38  (i.e. the upper end as depicted in  FIG. 4 ). The airfoil  20  is designed so that the first crossover  40  is preferably in fluid communication with at least two substantially spatially parallel passages  36 ,  38 , one of which preferably ends at the distal end portion  44 . As mentioned, the passages are preferably in “parallel” both spatially and fluidly, and are divided by a partition wall  30 . In particular, the passages  36  and  38  are divided by a bypass divider wall  31 . The flow of cooling air coming out of the trailing edge exhaust ports  26  is thus divided by one of the partition walls  30 , namely bypass divider wall  31 , which creates the “bypass” passage  36  and the “rear” passage  38 . The rear passage  38  can be further divided with additional partition walls  30  (not shown) to provide additional parallel passages. The bypass passage  36  is selected so as to minimize air stagnation therein, as described further below. In  FIG. 3 , the bypass passage  36  communicates with the distal end portion  44  of the first crossover  40 .  FIG. 4  illustrates that the partition wall  30  may include an extension  30 A between the bypass passage  36  and the rear passage  38  to second crossover  42 , so that air passing through the bypass passage  36  is directed to exit plenum  25  without flowing into the intermediary plenum  41 . 
   To assist an illustration of the operation of the present invention,  FIG. 5  shows a portion of a hollow section  24 ′ similar to  FIGS. 3 and 4 , but without the bypass passage  36  and bypass divider wall  31  shown in  FIGS. 3 and 4 . Due to the relatively wide chord of the airfoil, the passage  38 ′ feeding crossover  40 ′ and exit plenum  25 ′ are relatively wide. Passage  38 ′ is thus prone to the unintentional but unavoidable creation of an air “dead zone” of more or less stagnant air which undesirably decreases convective heat transfer to the cooling flow. By contrast, in  FIGS. 3 and 4 , the two narrower passages  36 ,  38  are substituted for the single passage  38 ′ of  FIG. 5 , and the bypass divider wall  31  between them is configured to direct air in passages  36  and  38  in a manner to substantially reduce the presence of an air “dead zone” therein. Benefit is thus is achieved without requiring a larger number of turns or a longer overall passage, and thus minimizes introduced aerodynamic losses. The presence of the bypass divider wall  31  between the bypass passage  36  and the rear passage  38  also strengthens the airfoil  20 , which is also particularly beneficial in a wide chord blade. 
   A new method of cooling an airfoil of a gas turbine engine comprises dividing the flow of cooling air directed to the exit plenum  25  in at least two parallel cooling air paths prior to directing the cooling air to the exit plenum  25 , preferably via a crossover  40 . One of the cooling air paths  36  is preferably directed to a distal end portion of the plenum  25 , while the other passage  38  is directed through the trailing edge inwardly therefrom relative to the inlets. This parallel geometry helps distribute the air to reduce stagnation and internal pressure losses. 
   The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, although application of the invention to a turbine blade is described and depicted herein, the invention may be applied to compressor and turbine blades and vanes. The invention can be used concurrently with other cooling techniques for increasing the heat transfer between the internal structures of the airfoil  20  and the cooling air. The various means for promoting internal heat transfer between the internal structures and the cooling air include dimples, trip strips, pedestals, fins, etc., all of which are intended to be indicated and schematically represented in  FIG. 3  as reference numeral  50 . Other techniques to introduce turbulence into the cooling air flow to promoting convective heat transfer may also be used, or none at all may be used. The crossovers may be omitted, if desired. Still other modifications will be apparent to those skilled in the art in light of a review of this disclosure and such modifications are intended to fall within the scope of the appended claims.