EP0475658A1

EP0475658A1 - Turbine blade airfoil with serial impingement cooling through internal cavity-forming ribs

Info

Publication number: EP0475658A1
Application number: EP91308047A
Authority: EP
Inventors: Ching-Pang Lee; Chung-Der Young
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1990-09-06
Filing date: 1991-09-03
Publication date: 1992-03-18

Abstract

A turbine blade airfoil (48) has pairs of air flow jet-producing orifices (44) formed through chordwise spaced transverse walls (46) extending between opposite side walls (50,52) of the airfoil and dividing a hollow chamber of the airfoil into a series of separate cavities (38). The pairs of orifices in first (46A,46B) and second (46C,46D,46E) pluralities of the transverse walls direct air flow in opposite directions from intermediate ones of the cavities toward leading (56) and trailing (62) edges of the airfoil. The pairs of orifices (44) in the first plurality of transverse walls (46A,46B) have pairs of axes (54A,54B)that diverge from one another in the direction of cooling air flow toward the leading edge (56) of the airfoil to cause successive impingement against portions of the opposite side walls (50,52) of successive cavities (38A,38B) of a first series of cavities by cooling air flow before exiting from the airfoil through exit holes (60) in the airfoil leading edge (56). The pairs of orifices (44) in the second plurality of transverse walls (46C,46D,46E) have pairs of axes (54C,54D,54E) that diverse from one another in the direction of cooling air flow toward the trailing edge (62) of the airfoil to cause successive impingement against portions of the opposite side walls (50,52) of successive cavities (38D,38E,38F) of a second series of cavities by cooling air flow before exiting from the airfoil through exit holes (64) in the airfoil trailing edge (62).

Description

BACKGROUND OF THE INVENTION

Field of the Invention

The present invention relates generally to gas turbine engine blades and, more particularly, to a turbine blade airfoil with serial impingement cooling through internal cavity-forming ribs.

Description of the Prior Art

Impingement cooling has been used in high pressure turbine stage nozzle vanes and rotor blades (hereinafter referred to as turbine blades) due to its high impingement heat transfer coefficient. Typically, cooling air flows into and upwardly through the blade shank, through internal serpentine cavities and passages formed in the airfoil, and out through exit holes in the wall of the blade airfoil.
In the past, impingement cooling was accomplished by inserting impingement inserts into the cavities of the airfoil. Impingement inserts typically have hollow configurations generally conforming to the interior contour of the respective cavity of the airfoil but in spaced relation to the airfoil wall. The inserts have jet air flow producing apertures in preselected locations. High pressure cooling air from the turbine compressor is directed into the inserts through the blade shank in a well known manner and is exhausted through such apertures to form jets of air striking the interior surfaces of the airfoil wall for impingement cooling. One example of this practice is disclosed in U. S. Patent No. 4,297,077 to Durgin et al.
However, two problems are associated with the use of impingement inserts. First, the inserts vibrate inside the blade and create metal fatigue. Second, to reduce a crossflow effect on the impingement heat transfer, the post impingement flows are usually permitted to bleed out of the airfoil right after the impingement and are used for film cooling. Thus, this practice of using impingement inserts fails to recapture any post impingement air flow and use it for more cooling before it is allowed to bleed out of the airfoil.
Consequently, a need still exists for improvement of impingement cooling techniques so that the problems associated with the use of impingement inserts can be avoided.

SUMMARY OF THE INVENTION

The present invention provides a serial impingement cooling arrangement designed to satisfy the aforementioned need. In accordance with the present invention, a turbine blade airfoil incorporates the serial impingement cooling arrangement in the internal cavity-forming ribs of the airfoil which both takes advantage of impingement air flow and recaptures the post impingement air flow and uses it for more cooling before it is allowed to bleed out of the airfoil.
Accordingly, the present invention is set forth in a turbine blade airfoil having leading and trailing edges and opposite side walls defining pressure and suction sides and merging together at said leading and trailing edges to define a hollow interior chamber for communication of cooling air flow to said side walls of said airfoil. The airfoil also has a plurality of holes through the leading and trailing edges to permit exit of cooling air from the hollow interior chamber of the airfoil. The present invention is directed to an impingement cooling arrangement which comprises: (a) a multiplicity of interior transverse walls spaced one to the next along the direction of a chord extending between the leading and trailing edges of the airfoil and disposed across the hollow interior chamber and rigidly connected with the opposite side walls so as to define a plurality of interior cavities in the chamber serially-arranged along the chord between the leading and trailing edges; and (b) means defining a pair of jet-producing orifices through each of the transverse walls for providing communication from one cavity to the next. Pairs of orifices in a first plurality of the transverse walls that define a first series of cavities have respective pairs of axes which diverge from one another in a first direction of cooling air flow toward the leading edge of the airfoil. The divergent relation of the orifice axes cause successive impingement against portions of the opposite side walls of successive cavities of the first series of cavities by portions of cooling air flow through the pairs of orifices in the first plurality of transverse walls before exiting from the airfoil through the exit holes in the leading edge of the airfoil. Pairs of orifices in a second plurality of the transverse walls that define a second series of cavities have respective pairs of axes which diverge from one another in a second direction of cooling air flow, opposite from the first direction, toward the trailing edge of said airfoil. The divergent relation of the orifice axes cause successive impingement against portions of the opposite side walls of successive cavities of the second series of cavities by portions of cooling air flow through the pairs of orifices in the second plurality of transverse walls before exiting from the airfoil through the exit holes in the trailing edge of the airfoil.
These and other features and advantages and attainments of the present invention will become apparent to those skilled in the art upon a reading of the following detailed description when taken in conjunction with the drawings wherein there is shown and described an illustrative embodiment of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

In the course of the following detailed description, reference will be made to the attached drawings in which:
Fig. 1 is a perspective view of a prior art turbine engine blade having holes in the blade airfoil for exit of cooling air therefrom.
Fig. 2 is an enlarged cross-sectional view of the prior art blade airfoil taken along line 2--2 of Fig. 1.
Fig. 3 is an enlarged longitudinal sectional view of the prior art blade taken along line 3--3 of Fig. 1.
Fig. 4 is a view similar to Fig. 3 but now illustrating the serial impingement cooling arrangement of the present invention.
Fig. 5 is a diagrammatic view of of an illustration of a cooling air flow circuit defined by the serial impingement cooling arrangement of the present invention in the blade airfoil of Fig. 4.
Fig. 6 is an enlarged fragmentary view of the blade taken along line 6--6 of Fig. 4.

DETAILED DESCRIPTION OF THE INVENTION

In the following description, like reference characters designate like or corresponding parts throughout the several views. Also in the following description, it is to be understood that such terms as "forward", "rearward", "left", "right", "upwardly", "downwardly", and the like, are words of convenience and are not to be construed as limiting terms.

Prior Art Turbine Engine Blade

Referring now to the drawings, and particularly to Figs. 1-3, there is illustrated a prior art gas turbine engine hollow blade, generally designated 10. The hollow blade 10 includes an airfoil 12 having pressure and suction sides 14, 16 and leading and trailing edges 18, 20, and a base 22 mounting the airfoil 12 to a rotor (not shown) of the engine (not shown). The base 22 has a platform 24 rigidly mounting the airfoil 12 and a dovetail root 26 for attaching the blade 10 to the rotor.
The airfoil 12 has opposite side walls 28, 30 defining the pressure and suction sides 14, 16 of the airfoil 12 which merge together at the leading and trailing edges 18, 20 of the airfoil 12 and are rigidly attached upright on the platform 24. The airfoil 12 also has an end cap 32 which closes the outer ends of the side walls 28, 30. The side walls 28, 30 and end cap 32 of the airfoil 12 contain small apertures or holes 34 which permit passage and exit of cooling air from the interior of the blade airfoil 12.
Cooling air flows into and upwardly through the base 22 of the blade 10 to the airfoil 12. The airfoil 12 includes a plurality of interior spaced ribs or transverse walls 36 which extend across the hollow interior of the airfoil 12 and rigidly interconnected with the opposite side walls 28, 30 so as to define a series of interior cavities 38 in the airfoil 12 in a hollow interior chamber 40 of the airfoil. As seen in Fig. 3, the ribs 36 extend vertically and alternately connect to and terminate short of the end cap 32 at their upper ends and of a solid portion 42 of the base 22. In such fashion, the ribs 36 define serpentine arrangements of cavities and passages within the interior of the airfoil 12 causing cooling air to flow along internal serpentine paths, as illustrated in Fig. 3, and exit through the holes 34 in the side walls 28, 30, leading and trailing edges 18, 20 and end cap 32 of the airfoil 12.

Serial Impingement Cooling Blade of Present Invention

In the prior art airfoil 12 described above, no impingement air flow occurs against the interior surfaces of the side walls 28, 30 of the airfoil 12. Consequently, inefficient and ineffective cooling of the turbine blade 10 is taking place.
Referring to Figs. 4 and 6, in accordance with the present invention, an arrangement of air flow jet-producing orifices 44 is provided in the transverse walls 46 of the turbine blade 48 for producing serial impingement cooling of the side walls 50, 52 of the blade 48. Otherwise the turbine blade 48 of Fig. 4 is the same as the turbine blade 10 of Figs. 1-3.
In Fig. 4, a pair of the orifices 44 are formed through each of the transverse walls 46 for providing communication from one interior cavity 38 to the next. A first plurality of the transverse walls 46A, 46B that define a first series of cavities 38A, 38B have pairs of orifices 44 with respective pairs of axes 54A, 54B which diverge from one another in a first direction of cooling air flow from an intermediate one of the cavities 38C toward the leading edge 56 of the airfoil 58. The divergent relation of the orifice axes 54A, 54B cause successive impingement against portions of the opposite side walls 50, 52 of successive cavities 38A, 38B of the first series of cavities by portions of cooling air flow jetting from the pairs of orifices 44 in the first plurality of transverse walls 46A, 46B before exiting from the airfoil 58 through the exit holes 60 in the leading edge 56 of the airfoil.
A second plurality of the transverse walls 46C, 46D, 46E that define a second series of cavities 38D, 38E, 38F have pairs of orifices 44 with respective pairs of axes 54C, 54D, 54E which diverge from one another in a second direction of cooling air flow, opposite from the first direction, from another intermediate one of the cavities 38G toward the trailing edge 62 of the airfoil 58. The divergent relation of the orifice axes 54C, 54D, 54E cause successive impingement against portions of the opposite side walls 50, 52 of successive cavities 38D, 38E, 38F of the second series of cavities by portions of cooling air flow through the pairs of orifices 44 in the second plurality of transverse walls 46C, 46D, 46E before exiting from the airfoil 58 through exit holes 64 in the trailing edge 62 of the airfoil.
It will be noticed that the larger the cross-sectional size of the given cavity 38 of the first and second series the greater is the divergent relation between the pair of axes 54 of the orifices 44 which communicate air flow into the cavity 38.
The serial impingement airfoil 58 of Fig. 4 thus has two circuits, a forwardly directed one 66 and a rearwardly directed one 68, as seen in Fig. 5. The number of circuit branches can be varied depending on the design. The impingement orifices 44 are drilled on the cavity ribs or transverse walls 46 and oriented to directly impinge on either the pressure or the suction side wall surfaces. The post impingement air will flow through the succeeding impingement holes to impinge on the surfaces of the next cavity 38 without creating the cross flow penalty in the same cavity. This design will allow the further usage of post impingement air before it bleeds out of the airfoil 58 through the exit holes 60, 64. At the end of the impingement circuit, the air can be bled out of the airfoil for either film cooling or recirculated for regenerative purposes. The impingement orifices 44 can be either cast or drilled during the fabrication process.
Because the impingement process is in series, the cooling capacity of air will be fully utilized and the cooling efficiency will be higher when compared to the present impingement insert designs. The impingement transverse walls between cavities not only provide the impingement purpose but also continue to serve as the airfoil structure to carry the mechanical and thermal loads.
It is thought that the present invention and many of its attendant advantages will be understood from the foregoing description and it will be apparent that various changes may be made in the form, construction and arrangement of the parts thereof without departing from the spirit and scope of the invention or sacrificing all of its material advantages, the forms hereinbefore described being merely preferred or exemplary embodiments thereof.

Claims

In a turbine blade airfoil having leading and trailing edges and opposite side walls defining pressure and suction sides and merging together at said leading and trailing edges to define a hollow interior chamber for communication of cooling air flow to said side walls of said airfoil, said airfoil also having a plurality of holes through said leading and trailing edges to permit exit of cooling air from said hollow interior chamber of said airfoil, an impingement cooling arrangement, comprising:
(a) a multiplicity of interior transverse walls spaced one to the next chordwise between said leading and trailing edges of said airfoil and disposed across said hollow interior chamber and rigidly connected with said opposite side walls so as to define a plurality of interior cavities in said chamber serially-arranged chordwise between said leading and trailing edges; and

(b) means defining a pair of jet-producing orifices through each of said transverse walls for providing communication from one cavity to the next;

(c) said pairs of orifices in a first plurality of said transverse walls that define a first series of said cavities having respective pairs of axes which diverge from one another and intersect said chord in a first direction of cooling air flow toward said leading edge of said airfoil to cause successive impingement against portions of said opposite side walls of successive cavities of said first series of cavities by portions of cooling air flow through said pairs of orifices in said first plurality of transverse walls before exiting from said airfoil through said exit holes in said leading edge of said airfoil;

(d) said pairs of orifices in a second plurality of said transverse walls that define a second series of said cavities having respective pairs of axes which diverge from one another and intersect said chord in a second direction of cooling air flow, opposite from said first direction, toward said trailing edge of said airfoil to cause successive impingement against portions of said opposite side walls of successive cavities of said second series of cavities by portions of cooling air flow through said pairs of orifices in said second plurality of transverse walls before exiting from said airfoil through said exit holes in said trailing edge of said airfoil.
The impingement cooling arrangement as recited in Claim 1 wherein the air flow in said first direction originates from a first one of said cavities of said first series located intermediately between said leading and trailing edges.
The impingement cooling arrangement as recited in Claim 2 wherein the air flow in said second direction originates from a first one of said cavities of said second series located intermediately between said leading and trailing edges.
The impingement cooling arrangement as recited in Claim 3 wherein said first one of said cavities of said first series is located adjacent said first one of said cavities of said second series.
The impingement cooling arrangement as recited in claim 1 wherein the larger the cross-sectional size of a given one cavity of said first and second series the greater is the divergent relation between said pair of axes of said orifices which communicate air flow into said cavity.
An engine turbine blade including a base and an airfoil rigidly supported on said base, said airfoil having opposite side walls defining pressure and suction sides of said airfoil, said walls merging together at leading and trailing edges of said airfoil to define a hollow interior chamber for communication of cooling air flow from said base of said blade to said airfoil, said airfoil also having an end cap attached to outer ends of said walls opposite said base to close said hollow interior chamber of said airfoil and a plurality of holes at least through said leading and trailing edges of said airfoil walls to permit passage and exit of cooling air from said hollow interior chamber of said airfoil, a serial impingement cooling arrangement, comprising:
(a) a multiplicity of interior transverse walls spaced one to the next chordwise between said leading and trailing edges of said airfoil and disposed across said hollow interior chamber and rigidly connected with said opposite side walls so as to define a plurality of interior cavities in said chamber serially-arranged chordwise between said leading and trailing edges; and

(b) means defining a pair of jet-producing orifices through each of said transverse walls for providing communication from one cavity to the next;

(c) said pairs of orifices in a first plurality of said transverse walls that define a first series of said cavities having respective pairs of axes which diverge from one another in a first direction of cooling air flow from a first one of said cavities of said first series located intermediately between said leading and trailing edges of said airfoil toward said leading edge of said airfoil to cause successive impingement against portions of said opposite side walls of successive cavities of said first series of cavities by portions of cooling air flow through said pairs of orifices in said first plurality of transverse walls before exiting from said airfoil through said exit holes in said leading edge of said airfoil;

(d) said pairs of orifices in a second plurality of said transverse walls that define a second series of said cavities having respective pairs of axes which diverge from one another in a second direction of cooling air flow, opposite from said first direction, from a first one of said cavities of said second series located intermediately between said leading and trailing edges of said airfoil toward said trailing edge of said airfoil to cause successive impingement against portions of said opposite side walls of successive cavities of said second series of cavities by portions of cooling air flow through said pairs of orifices in said second plurality of transverse walls before exiting from said airfoil through said exit holes in said trailing edge of said airfoil.
The impingement cooling arrangement as recited in Claim 6 wherein said first one of said cavities of said first series is located adjacent said first one of said cavities of said second series.
The impingement cooling arrangement as recited in Claim 6 wherein the larger the cross-sectional size of a given one cavity of said first and second series the greater is the divergent relation between said pair of axes of said orifices which communicate air flow into said cavity.