A gas turbine engine turbine blade includes an airfoil having pressure and suction sides joined together at leading and trailing edges and extending from a root to a tip. The airfoil also includes a longitudinally extending flow channel for receiving compressed air, with the flow channel including a chordally extending, imperforate septum dividing the flow channel into first and second laterally spaced apart flow chambers. The blade further includes a supply manifold for receiving the compressed air at a single pressure which is disposed in flow communication with both the first and second chambers for channeling thereto respective first and second portions of the compressed air. The septum is used to reduce or eliminate radial velocity distortion, or blowoff of film cooling air from the suction side.

The present invention relates generally to gas turbine engine turbine 
blades, and, more specifically, to a turbine blade having improved 
cooling. 
BACKGROUND OF THE INVENTION 
A conventional gas turbine engine includes a compressor for providing 
compressed air to a combustor wherein it is mixed with fuel and ignited 
for generating combustion gases. The combustion gases are channeled 
firstly through a high pressure turbine and typically through a low 
pressure turbine disposed downstream therefrom for extracting energy to 
drive the compressor and provide output power in the form of a combustion 
gas exhaust jet or shaft power for rotating a fan for generating thrust 
for powering an aircraft in flight. The efficiency of the engine is 
directly related to the high pressure turbine inlet temperature of the 
combustion gases channeled thereto. Since the combustion gases are 
considerably hot, the turbine rotor blades are typically hollow and 
provided with conventional film cooling holes for providing effective 
cooling thereof for ensuring a useful operating life. A portion of the 
compressed air from the compressor is suitably channeled to the turbine 
blades for cooling, and since such compressed air does not undergo 
combustion in the combustor, the net efficiency of the engine is 
decreased. 
Accordingly, turbine blades are continually being improved for reducing the 
required amount of compressor cooling air channeled thereto for improving 
overall efficiency of the engine while still providing acceptable life of 
the blade. For example, in advanced gas turbine engines being presently 
considered, substantial reductions in cooling airflow through the turbine 
blades is being considered along with a substantial increase in the 
turbine rotor inlet temperature for substantially improving the operating 
efficiency of the engine. This will require a substantial increase in heat 
transfer effectiveness in cooling the turbine blades using the reduced 
amount of cooling airflow. 
A conventional turbine blade includes a concave pressure side and a convex 
suction side over which the static pressure of the combustion gases 
channeled thereover varies significantly. Compressed air is conventionally 
channeled upwardly through the blade dovetail and into conventional 
serpentine passages through the blade airfoil for the convective coding 
thereof. The several passages within the blade airfoil may have various 
heat transfer enhancement turbulator ribs or pins for increasing the heat 
transfer coefficient over that for a smooth wall. Furthermore, 
conventional film cooling holes are selectively provided around the 
surface of the airfoil as required for forming suitable film cooling air 
layers to protect the airfoil against the hot combustion gases. Since the 
leading edge of the airfoil is typically subjected to the highest heat 
flux from the combustion gases, it typically requires the greatest 
protection from the heat and the highest heat transfer enhancement from 
the compressed air being channeled through the airfoil. 
In one conventional airfoil having a mid-chord radial flow channel, the 
compressed air is normally channeled therethrough with a fairly uniform 
parabolic radial velocity distribution with a maximum velocity near the 
center of the channel and lower velocities adjacent the pressure and 
suction side interior walls of the airfoil. However, it has been observed 
that by providing a plurality of film cooling holes along the radial 
channel to eject uniformly part of the inlet channel air flow out the 
airfoil external surface on only one side of the channel, such as the 
pressure side, with the other, suction side of the channel being 
imperforate, the radial velocity distribution of the compressed air 
through the channel is distorted with the peak velocity being moved 
transversely toward the film cooling holes. Correspondingly, the velocity 
of the compressed air adjacent the opposite, suction side decreases for a 
given average flow rate through the channel. Accordingly, the decreased 
compressed air velocity adjacent the imperforate wall results in a 
decrease of the convective heat transfer enhancement of that wall which 
typically requires an increase in the average velocity of the compressed 
air to offset the decrease at the wall to ensure effective cooling 
thereof. 
Furthermore, the turbine blade must also be designed to provide an adequate 
backflow margin to ensure that the combustion gases are not allowed to 
backflow through the film cooling holes into the blade airfoil during 
operation. The pressure of the compressed air inside the airfoil is, 
therefore, predeterminedly selected to be suitably larger than the 
pressure of the combustion gases flowing over the airfoil to ensure the 
forward flow of the compressed air from the interior of the airfoil 
through the film cooling holes to the exterior of the airfoil. Since the 
leading edge region of a typical high pressure turbine blade includes film 
cooling holes along both the pressure and suction sides of the airfoil, 
the pressure of the compressed air inside the airfoil must be suitably 
large to provide an effective backflow margin through the pressure side 
film cooling holes adjacent the leading edge which are subject to the 
highest pressure from the combustion gases flowable over the airfoil. 
However, since the pressure of the combustion gases channeled over the 
suction side of the airfoil is necessarily lower than that over the 
pressure side, the pressure ratio of the compressed air inside the airfoil 
relative to the suction side adjacent the leading edge is relatively high, 
and higher than the pressure ratio across the pressure side at the leading 
edge, which increases the ejection velocity of the film cooling air 
through the suction side leading edge film cooling holes. This may lead to 
a condition known as blowoff wherein the film cooling air initially breaks 
free from the airfoil suction side as it is ejected from the film cooling 
holes before reattaching to the suction side downstream therefrom. This 
leads to a decrease in the air film effectiveness and cooling capability 
of the film cooling air in this region. 
In both of the above situations, the internal heat transfer coefficient and 
the film cooling effectiveness of the compressed air channeled through the 
airfoil are decreased, which requires even more air, for example, to 
ensure acceptable cooling of the turbine blade, which decreases overall 
efficiency. Furthermore, it is desirable to channel the compressed air 
through the airfoil with as little pressure losses therein as possible to 
further increase the overall efficiency of the engine. 
Since a high pressure turbine blade typically includes several serpentine 
flow passages therein and film cooling holes along the pressure and 
suction sides thereof, as well as discharge holes through the trailing 
edge thereof, the blade must be suitably tested to ensure that suitable 
flow is obtained through all of the holes. In a turbine blade having two 
or more independent cooling air channels or circuits therethrough, for 
example one circuit channeling cooling air from the mid-chord and out 
through the leading edge holes, and a second circuit channeling air 
through the mid-chord and out the trailing edge holes, the several holes 
of each circuit must be independently plugged, for example, to test the 
operation of the other circuit. This is a relatively complex procedure 
required for ensuring effective operation of all the apertures through the 
blade airfoil. 
SUMMARY OF THE INVENTION 
A gas turbine engine turbine blade includes an airfoil having pressure and 
suction sides joined together at leading and trailing edges and extending 
from a root to a tip. The airfoil also includes a longitudinally extending 
flow channel for receiving compressed air, with the flow channel including 
a chordally extending, imperforate septum dividing the flow channel into 
first and second laterally spaced apart flow chambers. The blade further 
includes a supply manifold for receiving the compressed air at a single 
pressure which is disposed in flow communication with both the first and 
second chambers for channeling thereto respective first and second 
portions of the compressed air.

DESCRIPTION OF THE PREFERRED EMBODIMENT(S) 
Illustrated in FIG. 1 is an exemplary embodiment of a gas turbine engine, 
high pressure turbine rotor blade 10 having a conventional dovetail 12 for 
joining the blade 10 to a conventional rotor disk (not shown) having a 
complementary dovetail slot through which is axially inserted the dovetail 
12. In this exemplary embodiment, the dovetail 12 includes an inlet supply 
channel or manifold 14 for receiving relatively cool, compressed air 16 
from a conventional gas turbine engine compressor (not shown). 
The blade 10 further includes an airfoil 18 having a first, outwardly 
concave, pressure side 20 and a second, outwardly convex, suction side 22 
joined together at a leading edge 24 and a trailing edge 26 spaced axially 
therefrom, with the sides 20, 22 extending longitudinally or radially from 
a root 28 to a radially outer tip 30. The root 28 is integral with the top 
of the dovetail 12, with the blade 10 further including a conventional 
platform 32 at the root 28, which platform 32 provides a portion of the 
radially inner boundary for combustion gases 34 which are conventionally 
channeled over the airfoil 18 from the leading edge 24 to the trailing 
edge 26. As the combustion gases 34 are channeled over the airfoil 18 they 
effect a relatively high pressure on the airfoil pressure side 20 and a 
relatively low pressure on the airfoil suction side 22 which is used for 
rotating the disk containing the blades 10 in conventional fashion. 
Referring also to FIG. 2, the airfoil 18 includes an arcuate camber or 
chordal line 36 extending generally equally between the pressure and 
suction sides 20 and 22 from the leading edge 24 to the trailing edge 26. 
In accordance with one embodiment of the present invention, the airfoil 18 
includes a longitudinally extending mid-chord flow channel 38 extending 
upwardly from the inlet channel 14 and disposed in flow communication 
therewith for receiving the compressed air 16 therefrom. The mid-chord 
channel 38 is spaced along the chordal line 36 generally equally between 
the leading edge 24 and the trailing edge 26 in this exemplary embodiment 
and includes a chord-wise or chordally extending, imperforate first rib or 
septum 40 dividing the flow channel 38 into a first portion or chamber 38a 
disposed adjacent to the pressure side 20, and a second portion or flow 
chamber 38b spaced laterally from the first chamber 38a generally 
perpendicularly to the chordal line 36 and disposed adjacent to the 
suction side 22. 
As shown in FIGS. 3 and 4, the first septum 40 extends radially inwardly 
from the tip 30 to the base of the dovetail 12 to bifurcate or divide both 
the mid-chord channel 38 into the two chambers 38a, 38b, and the inlet 
channel 14 into two chambers 14a and 14b being disposed in flow 
communication with the first and second chambers 38a, 38b, respectively 
for channeling thereto first and second portions 16a, 16b of the 
compressed air 16, respectively. In this embodiment of the invention, the 
inlet channel 14 is a supply manifold which receives the compressed air 16 
at a single static pressure P.sub.s with this common pressure air then 
being channeled radially upwardly through the inlet chambers 14a, 14b into 
the two mid-chord chambers 38a, 38b. 
Referring again to FIGS. 2-4, the airfoil pressure side 20 adjacent the 
first chamber 38a includes a plurality of longitudinally spaced apart, 
pressure side, mid-chord film cooling holes 42 disposed in two vertically 
extending rows, for example. The mid-chord film cooling holes 42 extend 
laterally through the pressure side 20 in flow communication with the 
first chamber 38a for receiving therefrom the compressed air first portion 
16a for conventionally film cooling the airfoil pressure side 20 adjacent 
the first chamber 38a. The cooling holes 42 are preferably inclined 
upwardly toward the blade tip 30 and rearwardly toward the trailing edge 
26 as they extend from the first chamber 38a to the pressure side 20. 
The airfoil suction side 22 adjacent the second chamber 38b is imperforate 
in this embodiment of the invention and is characterized by the absence of 
any film cooling holes extending therethrough to the second chamber 38b 
for preventing flow of the compressed air second portion 16b from being 
discharged from the second chamber 38b directly through the airfoil 
suction side 22 at the second chamber 38b. 
In a conventional turbine blade, the mid-chord flow channel 38 would not 
include the septum 40 and the compressed air 16 would simply be channeled 
radially upwardly therethrough for cooling the interior surfaces of both 
the pressure side 20 and the suction side 22 by convective heat transfer. 
The mid-chord channel 38 may include conventional turbulators 44 for 
example along the inner surfaces of the pressure and suction sides 20 and 
22 for enhancing the convective heat transfer coefficient for improved 
cooling of the airfoil 18. However, as above described in the Background 
section, the mid-chord film cooling holes 42 cause the radial velocity 
distribution of the compressed air 16 to distort toward the pressure side 
20 within the mid-chord channel 38, with the peak velocity being disposed 
nearer the pressure side 20 than the suction side 22, with a resulting 
reduction in velocity of the compressed air 16 adjacent the suction side 
22 which decreases the convective heat transfer rate at that location. By 
providing the first septum 40 in accordance with the present invention, 
and separately channeling the compressed air first and second portions 
16a, and 16b to the respective first and second chambers 38a and 38b, the 
radial velocity distortion through the mid-chord channel 38 may be reduced 
for improving the convective heat transfer coefficient adjacent the inner 
surface of the suction side 22 in the second chamber 38b. The distortion 
effect of the mid-chord film cooling holes 42 will then be localized only 
in the first chamber 38a, with the septum 40 blocking their adverse effect 
in radial velocity distribution in the second chamber 38b. 
Accordingly, the radial velocity distribution through the second chamber 
38b will be a substantially undistorted parabolic distribution with a wall 
velocity adjacent the suction side 22 being greater than it otherwise 
would be without the septum 40. In this way, an improved heat transfer 
coefficient is obtained along the inside surface of the suction side 22 in 
the second chamber 38b for improving the overall efficiency of cooling the 
airfoil 18. In this exemplary embodiment, the septum 40 is preferably 
disposed in the mid-chord channel 38 substantially equidistantly between 
the airfoil pressure and suction sides 20 and 22 for maximizing the flow 
area of the individual first and second chambers 38a and 38b for 
minimizing pressure losses therefrom. 
Referring again to FIGS. 1 and 2, an alternate embodiment of the present 
invention includes a leading edge flow channel 46 extending longitudinally 
adjacent to the leading edge 24 between the leading edge 24 and the 
mid-chord channel 38. The leading edge channel 46 similarly includes a 
chordally extending, imperforate second rib or septum 48 dividing the 
leading edge channel 46 into a first portion or flow chamber 46a disposed 
adjacent to the pressure side 20, and an independent second portion or 
flow chamber 46b spaced laterally from the first chamber 46a generally 
perpendicularly to the chordal line 36 and disposed adjacent to the 
suction side 22. A plurality of leading edge, pressure side film cooling 
holes 50 are longitudinally spaced apart from each other and disposed in 
three exemplary vertical rows extending through the pressure side 20 in 
flow communication with the first chamber 46a. 
The airfoil 18 further includes a plurality of longitudinally spaced apart, 
leading edge, suction side film cooling holes 52, also known as gill 
holes, with three exemplary vertical rows being illustrated, extending 
through the suction side 22 in flow communication with the second chamber 
46b of the leading edge channel 46. A plurality of auxiliary, leading 
edge, suction side, film cooling holes 54 may be disposed between the 
leading edge 24 and the film cooling holes 52 in flow communication with 
the first chamber 46a for providing high pressure film cooling air 
immediately adjacent to the leading edge 24 for the effective film cooling 
thereof. 
More specifically, the region of the airfoil 18 adjacent the leading edge 
24 is subject to the highest heat flux from the combustion gases 34 and 
requires the most effective film cooling capability with a suitable amount 
of backflow margin. Since the leading edge first chamber 46ais disposed 
adjacent to the leading edge 24 and the pressure side 20, the film cooling 
holes 50 discharge to the highest pressure area of the combustion gases 34 
channeled over the airfoil 18. They, therefore, require the highest 
pressure of cooling air in the first chamber 46a for maintaining an 
adequate backflow margin and for obtaining acceptable film cooling 
therefrom. 
Accordingly, a mid-chord supply manifold 56 is disposed aft of the leading 
edge channel 46 and forward of the mid-chord channel 38 and extends 
longitudinally between the airfoil root 28 and tip 30. A first 
longitudinally extending rib 58 separates the leading edge channel 46 from 
the mid-chord supply manifold 56, a second longitudinally extending rib 60 
separates the mid-chord supply manifold 56 from the mid-chord channel 38, 
and a third longitudinally extending rib 62 separates the mid-chord 
channel 38 from a longitudinally extending trailing edge flow channel 64. 
The trailing edge flow channel 64 may have any conventional design 
including the two channel embodiment illustrated. 
The first rib 58, the second septum 48, and the leading edge portions of 
the airfoil 18 define the leading edge channel 46 and its two chambers 46a 
and 46b. The first and second ribs 58 and 60 along with respective 
portions of the airfoil 18 define the mid-chord supply manifold 56. The 
second and third ribs 60 and 62, the first septum 40, and the respective 
portions of the airfoil 18 define the mid-chord channel 38 including its 
two chambers 38a and 38b. The third rib 62 along with the trailing edge 
portions of the airfoil 18 define the trailing edge channel 64 which is 
disposed in flow communication with a plurality of longitudinally spaced 
apart trailing edge holes 66. 
As illustrated in FIGS. 2 and 3, the mid-chord first chamber 38a preferably 
includes a single outlet 68 disposed adjacent to the airfoil tip 30 in 
flow communication with the trailing edge channel 64 for channeling the 
compressed air first portion 16a therethrough solely in an aft direction 
into the trailing edge channel 64 from the first chamber 38a. The 
compressed air first portion 16a then flows through the trailing edge 
channel 64 and is discharged from the airfoil 18 through the trailing edge 
holes 66 in one cooling air flow circuit from the mid-chord first chamber 
38a. 
The mid-chord second chamber 38b in this exemplary embodiment includes a 
single outlet 70 as shown in FIGS. 2 and 4, disposed adjacent to the 
airfoil tip 30 in flow communication with the mid-chord supply manifold 56 
and in turn the leading edge channel 46 for channeling the compressed air 
second portion 16b therefrom solely in a forward direction into the 
mid-chord supply manifold 56 and in turn into both leading edge channel 
first and second chambers 46a and 46b. 
The compressed air second portion 16b, therefore, fills the mid-chord 
supply manifold 56 with compressed air at a substantially single static 
pressure which is then, in turn, further divided for flow through the 
leading edge first and second chambers 46a and 46b. More specifically, a 
plurality of longitudinally spaced apart first inlet holes 72 extend 
through the first rib 58 and are disposed in flow communication between 
the mid-chord supply manifold 56 and the leading edge first chamber 46a 
for channeling thereto a portion of the compressed air second portion 16b 
designated compressed air third portion 16c. 
A plurality of longitudinally spaced apart second Inlet holes 74 also 
extend through the first rib 58 and are disposed in flow communication 
between the mid-chord supply manifold 56 and the leading edge second 
chamber 46b for channeling another portion of the compressed air second 
portion 16b designated compressed air fourth portion 16d into the second 
chamber 46b. The first inlet holes 72 are predeterminedly sized to provide 
the compressed air third portion 16c into the leading edge first chamber 
46a with reduced pressure but providing a minimum pressure ratio across 
the pressure side film holes 50 for maintaining a predetermined amount of 
backflow margin. The second inlet holes 74 are predeterminedly sized for 
metering or reducing pressure of the compressed air fourth portion 16d 
inside the leading edge second chamber 46b to decrease the pressure ratio 
across the suction side film holes 52 at the second chamber 46b to reduce 
or eliminate blowoff of the compressed air fourth portion 16d off the 
suction side 22. In this way, compressed air may be selectively provided 
at different pressures with a greater pressure in the leading edge first 
chamber 46a than in the leading edge second chamber 46b to correspond with 
the difference in pressure of the combustion gases 34 over the pressure 
side 20 and the suction side 22 for providing effective pressure ratios 
for obtaining acceptable backflow margin through the several film cooling 
holes 50, 52, and 54 as well as reducing or eliminating blowoff of the 
compressed air discharged from the leading edge, suction side film cooling 
holes 52. 
Accordingly, the chordally extending first and second septums 40 and 48 may 
be provided in accordance with the present invention to bifurcate both the 
mid-chord channel 38 and the leading edge channel 46 for separately 
tailoring the performance thereof. In the mid-chord channel 38, the first 
septum 40 reduces the radial velocity distortion due to ejection of the 
compressed air through the mid-chord film cooling holes 42 on the pressure 
side 20 for enhancing the heat transfer coefficient along the inside 
surface of the airfoil suction side 22 in the second chamber 38b for 
example. The second septum 48 bifurcates the leading edge channel 46 to 
allow different pressures to be effected in the two chambers 46a and 46b 
to better match the pressures in the combustion gases 34 flowing 
separately over the airfoil pressure side 20 and suction side 22 for 
maintaining acceptable backflow margin while reducing or eliminating 
blowoff from the leading edge, suction side film cooling holes 52. 
Accordingly, the blade 10 in accordance with the present invention provides 
an improvement in heat transfer for more effectively and more efficiently 
cooling the airfoil 18. The blade 10 may therefore be operated at higher 
turbine rotor inlet temperatures for a given flowrate of cooling air, or a 
reduction in the amount of compressed air 16 may be obtained for a given 
blade temperature for improving overall efficiency while obtaining 
acceptable cooling, or both. 
Furthermore, the bifurcated mid-chord channel 38 forming the two separate 
flow circuits one extending rearwardly through the trailing edge channel 
64 and the trailing edge holes 26 and the other extending forwardly 
through the leading edge channel 46 and the several film cooling holes 50, 
52, and 54 allows for relatively easier air quantitative quality-control 
testing during the manufacturing process. By selectively plugging 
separately the mid-chord first chamber 38a or the mid-chord second chamber 
38b at the respective inlet chambers 14a 14b the respective flow circuit 
may be obstructed so that the other flow circuit may be evaluated for 
providing a given amount of airflow therethrough. By being able to block 
either the mid-chord first chamber 38a or the second chamber 38b , 
blocking of all the individual film cooling holes or trailing edge holes 
of the different circuits is not required for measuring the airflow 
through the unobstructed circuit. 
The airfoil 18 may include in any of its passages any additional heat 
transfer enhancement structures such as the turbulators 44 shown in the 
mid-chord channel 38 for improving the overall heat transfer ability and 
efficiency of the blade 18. All of the several film cooling holes 42, 50, 
52, and 54 are preferably tapered or conical with an increased hole flow 
area from inlet to exit for controlling the rate of the compressed air 
channeled therethrough and for diffusing the compressed air to reduce its 
film hole exit velocity as it is ejected along the outer surfaces of the 
airfoil 18 for further improving the cooling air film effectiveness. And, 
although the invention has been described with respect to a turbine rotor 
blade, it may also be used with respect to turbine stator vanes. 
Furthermore, although the airfoil 18 has been described with two separate 
cooling air circuits to the respective leading and trailing edges 24, 26 
with the mid-chord channel 36 being an inlet to the airfoil 18 for 
receiving the compressed air 16 from the dovetail inlet channel 14, 
various alternate cooling air circuits may also be used and bifurcated 
with analogous septums 40, 48. 
While there have been described herein what are considered to be preferred 
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. 
Accordingly, what is desired to be secured by Letters Patent of the United 
States is the invention as defined and differentiated in the following 
claims: