Internally grooved turbine wall

A turbine wall includes an outer surface for facing combustion gases, and an opposite inner surface for being impingement air cooled. A plurality of adjoining ridges and grooves are disposed in the inner surface for enhancing heat transfer by the impingement cooling air.

BACKGROUND OF THE INVENTION 
The present invention relates generally to gas turbine engines, and, more 
specifically, to turbine cooling therein. 
In a gas turbine engine, air is pressurized in a compressor, mixed with 
fuel in a combustor and ignited for generating hot combustion gases, which 
flow downstream through one or more turbine stages for extracting energy 
therefrom. A high pressure turbine (HPT) firstly extracts energy from the 
gases for powering the compressor. And, additional energy is typically 
extracted from the gases by a low pressure turbine (LPT) which typically 
powers a fan disposed upstream from the compressor. 
The HPT includes a stationary turbine nozzle which directly receives the 
combustion gases from the combustor for redirecting the gases into a row 
of rotary turbine blades extending radially outwardly from a rotor disk. 
The nozzle includes a plurality of circumferentially spaced apart stator 
vanes which complement the performance of the rotor blades. 
Both the vanes and blades are suitably configured as a airfoils which 
cooperate for maximizing efficiency of extraction of energy from the 
combustion gases which flow thereover. The vane and blade airfoils have 
generally concave pressure sides and opposite, generally convex suction 
sides which extend axially between corresponding leading and trailing 
edges thereof and radially over their radial span. 
The nozzle vanes extend radially between annular outer and inner bands 
which confine the combustion gases therebetween. The blade airfoils extend 
from their radially inner roots to their radially outer tips which are 
spaced closely radially inwardly from a surrounding annular turbine 
shroud. The shroud is stationary and defines the outer boundary for the 
combustion gases which flow past the rotating blade airfoils. 
Since the stator vanes, rotor blades, and turbine shrouds are directly 
exposed to the combustion gases, they require suitable cooling for 
maintaining their strength and ensuring suitable useful lives thereof. 
These components are typically cooled by channeling thereto corresponding 
portions of air bled from the compressor which is substantially cooler 
than the hot combustion gases. Various cooling techniques are used in 
cooling gas turbine engine components. Film cooling is one technique 
wherein air is channeled through inclined film cooling holes to form a 
film of cooling air between the outer or exposed surfaces of the 
components and the hot combustion gases which flow thereover. 
Impingement cooling is another technique wherein the cooling air is 
initially directed substantially normal to the inner surfaces of these 
components in impingement thereagainst for removing heat therefrom by 
convection heat transfer. The inner surfaces may be smooth for impingement 
cooling, or may include three dimensional turbulators in the form of 
cylindrical pins, bumps, or dimple depressions. These turbulators increase 
the effective surface area of the inner surfaces from which heat may be 
extracted. The turbulators are typically small in size for reducing any 
adverse pressure drop caused thereby for ensuring cooling efficiency. 
Since turbine vanes, blades, and shrouds are formed of high strength 
metals, they are typically manufactured by casting for achieving maximum 
material strength and precision of the small features thereof, including 
any turbulators which may be used therein. 
The vanes and blades are hollow for channeling therethrough the cooling air 
in several radially extending passages. The passages may be individually 
fed with cooling air or may be arranged in serpentine legs through which 
the cooling air flows. Impingement cooling for the vanes is typically 
provided by placing perforated impingement baffles inside corresponding 
internal passages therein. The cooling air is first channeled inside the 
baffle and then laterally through its perforations for impingement against 
the inner surface of the vane. 
Since turbine blades rotate during operation, an integral rib or bridge may 
be provided between its pressure and suction sides for defining an 
integral baffle having holes or perforations through which the cooling air 
is directed in impingement against the inner surface of the blade airfoil, 
typically along the leading edge. 
Both the vane and blade airfoils may be similarly cast in view of their 
common airfoil configurations with internal radial passages. The internal 
passages are defined by corresponding ceramic cores surrounded by wax 
which defines the configuration of the final airfoil. The wax is then 
surrounded by a ceramic shell, and subsequently removed in the lost wax 
method. Molten metal is then poured between the shell and core and 
solidifies in the form of the desired airfoil. The ceramic shell and cores 
are then removed to expose the cast airfoil. 
The ceramic cores themselves are produced in a separate casting process 
using a metallic core die precisely formed with the mirror features to be 
produced in the outer surface of the core. A typical core die may be 
formed in two or more halves with an internal passage being defined 
therebetween and extending along the span axis thereof. A ceramic slurry 
or paste is injected under significant pressure in the open end of the die 
to fill the die, after which the resulting ceramic core is removed and 
cured. 
The same core die is used repeatedly for casting multiple copies of the 
airfoils. However, the injection of the ceramic slurry into the die 
eventually leads to wear therein. Wear is most pronounced for three 
dimensional features such as the turbulators for enhancing impingement 
cooling, which turbulators of the core die are abraded over extended use. 
Once the die is worn, a new die must be manufactured at considerable 
expense. 
Accordingly, it is desired to provide improved impingement cooling features 
in a turbine component, which can reduce core die wear in a preferred 
embodiment. 
BRIEF SUMMARY OF THE INVENTION 
A turbine wall includes an outer surface for facing combustion gases, and 
an opposite inner surface for being impingement air cooled. A plurality of 
adjoining ridges and grooves are disposed in the inner surface for 
enhancing heat transfer by the impingement cooling air.

DETAILED DESCRIPTION OF THE INVENTION 
Illustrated in FIG. 1 is a portion of a gas turbine engine 10 which is 
axisymmetrical about a longitudinal or axial centerline axis 12. The 
engine includes a multistage axial compressor 14 configured for 
pressurizing air 16, portions of which are bled for later use in cooling 
the engine. 
The major portion of the air from the compressor is channeled to an annular 
combustor 18, shown in aft part, wherein the air is mixed with fuel and 
ignited for generating hot combustion gases 20 which flow downstream into 
a high pressure turbine (HPT). The turbine includes an annular turbine 
nozzle having a plurality of circumferentially spaced apart stator vanes 
22 extending radially between annular outer and inner bands. 
The high pressure turbine also includes a row of rotor blades 24 which 
extend outwardly from a supporting rotor disk, and are secured thereto by 
integral axial dovetails. Surrounding the rotor blades 24 is an annular 
turbine shroud 26 typically formed of a plurality of circumferentially 
adjoining arcuate shroud segments. 
During operation, the combustion gases 20 are discharged from the combustor 
between the nozzle vanes 22 for flow in turn between the downstream rotor 
blades 24 which extract energy therefrom for in turn rotating the 
supporting disk, which in turn powers the compressor 14. The combustion 
gases then flow downstream through a low pressure turbine, with the first 
nozzle stage thereof being illustrated, which also includes one or more 
rows of turbine blades (not shown) which extract additional energy from 
the gases for typically powering a fan (not shown) upstream of the 
compressor. 
The engine 10 as above described is conventional in configuration and 
operation. The engine is also conventional in bleeding corresponding 
portions of the pressurized air 16 for use in cooling various turbine 
components such as the nozzle vanes 22, HPT rotor blades 24, and the HPT 
shroud 26. These components are typically cooled by convection, film 
cooling, and impingement cooling in conventional manners for maximizing 
cooling efficiency of the air while minimizing pressure losses therein. 
In accordance with the present invention, impingement cooling features for 
the vanes 22, blades 24, and shroud 26 may be varied for obtaining various 
performance and casting advantages. 
More specifically, FIG. 2 illustrates one of the turbine nozzle vanes 22 in 
accordance with an exemplary embodiment of the present invention. The vane 
22 is in the form of an enclosing wall 28 which defines an airfoil. The 
vane has an outer surface 30 defining a generally concave pressure side 
and an opposite, generally convex suction side which face the combustion 
gases 20 which flow thereover during operation. The vane outer surface 30 
extends radially or longitudinally along a span axis 32, and axially or 
laterally along a chord axis 34 between an upstream leading edge 36 and 
downstream trailing edge 38 of the vane. 
The vane wall 28 also includes an opposite internal or inner surface 40 
which defines a radially extending inner passage or cavity 42 extending 
along the span axis for channeling the cooling air 16 therethrough. 
In accordance with the present invention, the vane inner surface 40 
includes a plurality of adjoining ridges 44 and grooves 46 for improving 
heat transfer and impingement cooling from the available air, as well as 
providing improvements in vane casting in a suitable embodiment. 
The ridges 44 and grooves 46 are parallel to each other and preferably 
directly adjoin each other side-by-side for increasing surface area 
available for cooling by the cooling air 16 without introducing 
appreciable pressure losses therein. The vane is heated from the outside 
by the combustion gases 20 which flow thereover, with the cooling air 16 
being provided inside the vane for internal cooling thereof. Without the 
ridges and grooves, a smooth inner surface of the vane has limited heat 
transfer surface area for being cooled. By introducing the relatively 
small ridges and grooves, a significant increase in surface area inside 
the vane is obtained from which the cooling air 16 may extract additional 
heat from the underlying vane wall 28 for improving the cooling thereof 
during operation. 
FIG. 3 illustrates an enlarged view of a typical cross section of a portion 
of the vane wall 28. In one embodiment, each of the ridges 44 has a width 
A, and each of the grooves 46 has a width B, with the ridges and grooves 
being generally equal in width. 
Each of the ridges 44 has a height C, which is the same as the 
corresponding depth of the adjoining groove 46, which is sufficiently tall 
for both increasing effective surface area and interrupting the boundary 
layer of cooling air formed along the vane inner surface during operation. 
As shown schematically in FIG. 3, a boundary layer 16b of the air 16 will 
form during operation over the inner surface of the vane. The boundary 
layer is typically turbulent and has a thickness D during operation. The 
ridges 44 are preferably sized in height C to slightly exceed the boundary 
layer thickness D for increasing heat transfer cooling during operation, 
without introducing excessive pressure losses due to excess height. For 
example, the height C of the ridges 44 may be in the exemplary range of 
about 15-25 mils. Correspondingly, the ridge width A and the groove width 
B may each also be in this exemplary range of about 15-25 mils. These 
small values are sufficient for exceeding the height of the cooling air 
boundary layer formed inside the vanes during operation and providing a 
substantial increase in surface area available for cooling without 
significant pressures losses associated therewith. 
The ridges 44 and grooves 46 illustrated in the exemplary embodiment of 
FIG. 3 are sized and configured to increase the surface area of the vane 
inner surface 40 by about 100%. Since the ridges and grooves have 
substantially equal width and height, the two sides bounding each ridge 
and groove effectively double the available surface area subject to 
cooling by the air 16. 
In the exemplary embodiment illustrated in FIG. 3, the ridges 46 are 
semicircular or convex in cross section at their tops and meet the grooves 
46 which are also semicircular, but concave at their bottoms. The ridges 
and grooves are thusly complementary with each other having compound side 
surfaces transitioning from concave to convex at their mid-heights having 
inflection points. This configuration reduces stress concentrations while 
providing smooth contours along which the cooling air 16 may flow parallel 
along the lengths of the ridges and grooves, and in cross-flow laterally 
thereacross from ridge to ridge. 
FIG. 4 illustrates an alternative embodiment of the ridges and grooves of 
FIG. 3 designated 44b, and 46b, respectively. In this embodiment, the 
ridges 44b are triangular in cross section, and correspondingly the 
adjoining grooves 46b are triangular in cross section in a sawtooth 
pattern, with small radii at the tips of the ridges and the bases of the 
grooves. 
FIG. 5 illustrates yet another embodiment of the ridges and grooves of FIG. 
3 designated 44c and 46c, respectively. In this embodiment, the ridges 44c 
are flat along their tops between adjacent grooves 46c, with both the 
ridges 44c and grooves 46c being rectangular in cross section in a 
square-wave form. 
In this embodiment, the grooves 46c are flat at their bases between 
adjacent ridges 44c, with the sidewalls extending perpendicularly between 
the tops of the ridges and the bottoms of the grooves also being flat. 
With equal widths and heights of the ridges and grooves illustrated in 
FIG. 5, the available surface area subject to cooling is double that of 
the surface without the ridges and grooves therein. 
FIG. 6 illustrates yet another embodiment of the ridges and grooves of FIG. 
3 designated 44d, and 46d, respectively. In this embodiment, the ridges 
44d are semicircular or convex in cross section, and the adjoining grooves 
46d are flat therebetween and aligned along the maximum diameters thereof. 
FIG. 7 illustrates yet another embodiment of the ridges and grooves of FIG. 
3 designated 44e and 46e, respectively. The ridges 44e are flat in cross 
section at their tops and adjoin semicircular or concave grooves 46e. 
In the five exemplary embodiments illustrated in FIGS. 3-7, the ridges and 
grooves are parallel to each other and preferably continuous along their 
lengths for basically defining two dimensional components which vary in 
configuration solely along their cross sections, while being identical 
along their lengths. These various configurations may be readily formed in 
the vane 22 illustrated in FIG. 2 for improving internal cooling thereof 
without introducing significant pressure losses. 
In FIG. 2, the inner surface 40 of the airfoil wall defines the inner 
cavity 42 which extends radially along the span axis 32 at the upstream or 
forward end of the vane at the leading edge 36. And, an additional one of 
the inner cavities 42 may also be formed in the aft end of the vane near 
the trailing edge 38, with the two internal cavities beings separated by 
an integral rib extending between the pressure and suction sides. 
In the forward cavity 42, the ridges 44 and grooves 46 preferably extend 
radially or along the span axis 32 over those portions of the vane inner 
surface for which additional cooling is desired. In FIG. 2, the ridges are 
disposed continuously over the inner surface behind the leading edge 36 
and downstream behind the forward portions of the pressure and suction 
sides. 
A particular advantage of the span ridges 44 and span grooves 46 is their 
ability to not only improve cooling heat transfer inside the vane during 
operation, but also reduce wear in the corresponding core die used for 
casting thereof. 
FIG. 8 illustrates schematically a core die 48 used for making a ceramic 
core 50 which in turn is used for casting the forward cavity of the vane 
illustrated in FIG. 2. The core die 48 is typically in the form of a two 
piece metal shell having an inner cavity 48a matching the vane inner 
surface 40 in the forward cavity 42 illustrated in FIG. 2. The same ridges 
44 and grooves 46 found in the vane 22 of FIG. 2 are initially provided in 
the core die 48 illustrated in FIG. 8. This is typically accomplished by 
precision milling of these features therein. 
The core die 48 illustrated in FIG. 8 has a longitudinal axis 52 and is 
open at its top end for defining an inlet for receiving a ceramic slurry 
or paste 54 conventionally injected therein by a suitable ceramic injector 
56. The ceramic 54 is injected into the cavity 48a along the span axis 52 
for completely filling the cavity therewith. The ridges 44 and grooves 46 
in this preferred embodiment extend parallel to the longitudinal axis 52 
along which the ceramic is injected. 
Since the ceramic is injected along the lengths of the ridges and grooves, 
they are subject to relatively less wear than if the ceramic were injected 
transversely across the ridges from side to side. By injecting the ceramic 
along the lengths of the ridges and grooves, the core die 48 may be used 
repetitively with reduced friction wear for enhanced life. 
The resulting ceramic 54 is suitably cured to form the core 50 on which are 
formed grooves 50a which are mirror images to the span ridges 44, and 
ridges 50b which are mirror images of the span grooves 46. The ceramic 
core 50 is then used in conjunction with a second such core to define the 
forward and aft vane cavities, with a cooperating outer ceramic shell for 
casting the vane 22 illustrated in FIG. 2 in a conventional manner using 
the lost wax process. 
A particular advantage of the ridges and grooves illustrated in FIG. 2 is 
their ability to improve impingement cooling inside the vane 22. The vane 
22 preferably also includes an impingement baffle 58 which is disposed 
inside the inner cavity 42. The impingement baffle 58 may have any 
conventional configuration and is typically in the form of a thin metal 
shell perforated with impingement holes. The baffle 58 is spaced generally 
perpendicularly from the ridges 44 for impinging a portion of the cooling 
air 16 thereagainst. 
An enlarged section of the impingement baffle 58 spaced from the vane wall 
28 is illustrated in FIG. 3. The baffle is suitably mounted inside the 
vane for providing a baffle spacing E across which the cooling air 16 is 
directed in jets from the baffle apertures for impingement against the 
ridges and grooves. 
The ridges 44 are relatively small for improving impingement cooling 
without introducing undesirable pressure losses therefrom. The height C of 
the ridges is preferably smaller than the baffle space in E. Preferably, 
the ridge height C is about an order of magnitude less than the baffle 
spacing E. As indicated above, the ridge height C is within the exemplary 
range of about 15-25 mils, with the baffle spacing E being in an exemplary 
range of about 100-150 mils. The ridges 44 and grooves 46 increase surface 
area effective for impingement cooling, and thereby increase the heat 
transfer cooling of the vane inner surface 40. The post-impingement air 16 
may flow longitudinally along the lengths of the grooves 46 as well as in 
cross-flow over the ridges 44. 
Referring again to FIG. 2, two impingement baffles 58 may be used in the 
forward and aft vane cavities for correspondingly providing impingement 
cooling therein. The aft vane cavity may also include the ridges and 
grooves for enhancing impingement cooling. As indicated above, the ridges, 
such as those in the forward cavity of the vane 22 of FIG. 2, preferably 
extend along the span axis 32 for reducing core die wear. 
However, the ridges and grooves may have other orientations as desired. For 
example, the ridges and grooves illustrated in the aft cavity of the vane 
22 in FIG. 2 are inclined between the span axis 32 and the chord axis 34. 
They are still effective for improving impingement cooling although they 
are prone to more wear in the corresponding core die than ridges formed 
solely along the span axis. Since the ridges and grooves are relatively 
small in height and are symmetrical along their lengths, core die wear is 
nevertheless relatively little for this configuration. 
As indicated above, the nozzle vanes 22 and impingement baffles 58 therein 
may have any conventional configuration which may obtain improved cooling 
performance by the introduction of the cooperating ridges 44 and grooves 
46 in various embodiments. The vanes 22 may have other conventional forms 
of cooling in addition thereto such as various rows of film cooling holes 
60 extending through the vane walls along the pressure and suction sides 
thereof as desired. The spent impingement cooling air from the forward and 
aft vane cavities is conveniently discharged through the film cooling 
holes 60 for effecting cooling air films on the external surface of the 
vane for providing a barrier against the heating effects of the combustion 
gases 20 which flow over the vanes. 
The ridges and grooves may be used in other components of the turbine for 
improving impingement cooling thereof. For example, FIG. 9 illustrates a 
portion of the first stage turbine blade 24 which may be modified to 
incorporate the ridges and grooves. Like the vane 22 illustrated in FIG. 
2, the blade 24 illustrated in FIG. 9 is also in the form of an airfoil 
suitably configured for its specific function. Accordingly, similar 
components of the vane 22 and blade 24 are labeled with the same reference 
numerals. 
For example, the blade 24 illustrated in FIG. 9 includes a wall 28 defining 
a corresponding airfoil having an outer surface 30 exposed to the 
combustion gases 20 during operation. The outer surface 30 includes a 
generally concave pressure side, and an opposite generally convex suction 
side which extend longitudinally or radially along a span axis 32, and 
laterally along a chord axis 34. 
The blade airfoil includes an inner surface 40 defining an inner cavity 42 
extending longitudinally along the span axis 32 from the root to the tip 
of the blade for channeling the cooling air 16 against the backside of the 
leading edge in impingement thereagainst. 
The blade airfoil typically includes several of the inner cavities between 
the leading and trailing edges 36,38 of the airfoil which may be 
configured in various conventional manners for internally cooling the 
blade. For example, some of the inner cavities may be linked together to 
provide serpentine cooling with or without corresponding wall turbulators 
therein. 
Since the leading edge 36 of the rotor blade first encounters the 
combustion gases 20, it typically includes a dedicated cooling circuit 
therefor. By introducing the ridges 44 and grooves 46 in the leading edge 
cavity 42 of the blade 24, improved cooling may be obtained in an 
otherwise conventional rotor blade, also including rows of the film 
cooling holes 60. 
Since the blade 24 rotates during operation, whereas the vane 22 is 
stationary during operation, an impingement baffle is introduced in the 
blade illustrated in FIG. 9 by an integral, perforated rib or bridge 58b 
which extends between the pressure and suction sides to define the leading 
edge forward cavity 42. By positioning the bridge baffle 58b adjacent the 
forward cavity 42, the impingement holes in the baffle direct a portion of 
the cooling air 16 in the axial direction toward the inner surface 40 
around the blade leading edge 36. The impingement air thusly engages the 
ridges 44 and grooves 46 inside the blade leading edge for improving 
impingement cooling thereat in the same manner as provided in the vane 
illustrated in FIG. 2. 
The ridges and grooves illustrated in FIG. 9 may have any of the 
configurations disclosed for the vane 22 described above for also enjoying 
the benefits therefrom. For example, referring to FIG. 3 in addition to 
FIG. 9, the height C of the ridges 44 for the turbine blade is also 
preferably smaller than the corresponding baffle spacing E between the 
inside of the blade leading edge 36 and the bridge baffle 58b over most of 
the leading edge. The ridges and grooves may be introduced wherever 
desirable in the leading edge cavity 42, and may additionally cooperate 
with the conventional film cooling holes 60 extending through the airfoil 
wall which receive spent impingement air from the cavity. 
In the exemplary embodiment illustrated in FIG. 9, the ridges 44 extend 
along the direction of the chord axis 34 instead of along the span axis 
32. Since the blade rotates during operation, the cooling air 16 channeled 
therethrough is subject to centrifugal force including Coriolis forces 
which produce secondary flow fields that may additionally enhance cooling 
by cooperating with the chord ridges 44. However, the ridges 44 may 
alternatively be oriented solely along the span axis 32 similar to those 
illustrated in the forward cavity of the FIG. 2 vane, or may be inclined 
as in the aft cavity of the FIG. 2 vane. 
FIG. 10 illustrates yet another application of the ridges 44 and grooves 46 
applied to the segments of the turbine shroud 26. The shroud and its 
segments may have any conventional configuration but for the introduction 
of the ridges 44 and grooves 46 therein. Each segment of the shroud 26 
typically includes forward and aft rails which engage complementary 
forward and aft hooks for mounting the shroud in the turbine case as 
illustrated in FIG. 1. The central portion of the shroud hangar, 
designated 58c, channels air radially inwardly through a corresponding 
impingement baffle for impingement cooling the shroud in a conventional 
manner. 
As shown in FIG. 10, the shroud segment is in the form of an arcuate panel 
or wall 28 having an outer surface 30 which is arcuate and faces radially 
inwardly above the row of turbine blades 24 as shown in FIG. 1. The shroud 
wall 28 has an inner surface 40 which faces radially outwardly and is open 
and exposed to the cooling air 16 directed thereagainst. The cooling air 
16 is isolated behind or inside the shroud 26 radially above the blade row 
for providing impingement cooling of the shroud. The ridges 44 and grooves 
46 are disposed in the shroud inner surface 40 for enhancing impingement 
cooling thereof in basically the same manner as indicated above for the 
vanes 22 and blades 24. Like those other embodiments, the ridges 44 and 
grooves 46 may have any of the configurations disclosed above and suitable 
orientations as desired. 
For example, the ridges 44 and grooves 46 preferably extend 
circumferentially along the shroud inner surface 40 in the direction of 
blade rotation. In this way, additional cross-flow advantages of the spent 
impingement air are obtained as the air is channeled through film cooling 
holes (not shown) in the shroud panel or around the forward and aft rails 
thereof. The spent impingement cooling air is also readily distributed 
circumferentially around the circumference of the shroud without 
significant pressure loss along the lengths of the ridges and grooves. 
By the simple introduction of the two-dimensional ridges 44 and 
corresponding grooves 46 in otherwise conventional turbine components, 
improved impingement cooling may be obtained without significant pressure 
losses. And, advantages in casting may also be obtained. For spanwise 
directed ridges and grooves in the vanes and blades, the corresponding 
core dies therefor enjoy less wear and may be used for producing more 
vanes and blades over their useful life. The turbine shrouds 26 are also 
typically cast in the lost wax process, without the need for core dies in 
view of their different configuration, and die wear is not a concern. 
While there have been described herein what are considered to be preferred 
and exemplary embodiments of the present invention, other modifications of 
the invention shall be apparent to those skilled in the art from the 
teachings herein, and it is, therefore, desired to be secured in the 
appended claims all such modifications as fall within the true spirit and 
scope of the invention.