Turbomachinery airfoil with optimized heat transfer

A blade or vane for a gas turbine engine includes a primary cooling system (42) with a series of medial passages (44, 46a, 46b, 46c, 48) and an auxiliary cooling system (92) with a series of cooling conduits (94). The conduits of the auxiliary cooling system are parallel to and radially coextensive with the medial passages and are disposed in the peripheral wall (16) of the airfoil between the medial passages and the airfoil external surface (28). The conduits are chordwisely situated in a zone of high heat load (104, 106) so that their effectiveness is optimized. The conduits may also be chordwisely coextensive with some of the medial passages so that coolant in the medial passages is protected from excessive temperature rise. The chordwise dimension C of the conduits is limited so that potentially damaging temperature gradients do not develop in the airfoil wall (16).

TECHNICAL FIELD 
This invention pertains to coolable turbomachinery components and 
particularly to a coolable airfoil for a gas turbine engine. 
BACKGROUND OF THE INVENTION 
The blades and vanes used in the turbine section of a gas turbine engine 
each have an airfoil section that extends radially across an engine 
flowpath. During engine operation the turbine blades and vanes are exposed 
to elevated temperatures that can lead to mechanical failure and 
corrosion. Therefore, it is common practice to make the blades and vanes 
from a temperature tolerant alloy and to apply corrosion resistant and 
thermally insulating coatings to the airfoil and other flowpath exposed 
surfaces. It is also widespread practice to cool the airfoils by flowing a 
coolant through the interior of the airfoils. 
One well known type of airfoil internal cooling arrangement employs three 
cooling circuits. A leading edge circuit includes a radially extending 
impingement cavity connected to a feed channel by a series of radially 
distributed impingement holes. An array of "showerhead" holes extends from 
the impingement cavity to the airfoil surface in the vicinity of the 
airfoil leading edge. Coolant flows radially outwardly through the feed 
channel to convectively cool the airfoil, and a portion of the coolant 
flows through the impingement holes and impinges against the forwardmost 
surface of the impingement cavity. The coolant then flows through the 
showerhead holes and discharges over the leading edge of the airfoil to 
form a thermally protective film. A midchord cooling circuit typically 
comprises a serpentine passage having two or more chordwisely adjacent 
legs interconnected by an elbow at the radially innermost or radially 
outermost extremities of the legs. A series of judiciously oriented 
cooling holes is distributed along the length of the serpentine, each hole 
extending from the serpentine to the airfoil external surface. Coolant 
flows through the serpentine to convectively cool the airfoil and 
discharges through the cooling holes to provide transpiration cooling. 
Because of the hole orientation, the discharged coolant also forms a 
thermally protective film over the airfoil surface. Coolant may also be 
discharged from the serpentine through an aperture at the blade tip and 
through a chordwisely extending tip passage that guides the coolant out 
the airfoil trailing edge. A trailing edge cooling circuit includes a 
radially extending feed passage, a pair of radially extending ribs and a 
series of radially distributed pedestals. Coolant flows radially into the 
feed passage and then chordwisely through apertures in the ribs and 
through slots between the pedestals to convectively cool the trailing edge 
region of the airfoil. 
Each of the above described internal passages--the leading edge feed 
channel, midchord serpentine passage, tip passage and trailing edge feed 
passage--usually includes a series of turbulence generators referred to as 
trip strips. The trip strips extend laterally into each passage, are 
distributed along the length of the passage, and typically have a height 
of no more than about 10% of the lateral dimension of the passage. 
Turbulence induced by the trip strips enhances convective heat transfer 
into the coolant. 
The above described cooling arrangement, and adaptations of it, have been 
used successfully to protect turbine airfoils from temperature related 
distress. However as engine designers demand the capability to operate at 
increasingly higher temperatures to maximize engine performance, 
traditional cooling arrangements are proving to be inadequate. 
One shortcoming of a conventionally cooled airfoil is its possible 
unsuitability for applications in which the operational temperatures are 
excessive over only a portion of the airfoil's surface, despite being 
tolerable on average. Locally excessive temperatures can degrade the 
mechanical properties of the airfoil and increase its susceptibility to 
oxidation and corrosion. Moreover, extreme temperature gradients around 
the periphery of an airfoil can lead to cracking and subsequent mechanical 
failure. 
Another shortcoming is related to the serpentine passage. A serpentine 
passage makes multiple passes through the airfoil interior. Accordingly, 
it takes more time for coolant to travel through a serpentine than to 
travel through a simple radial passage. This increased coolant residence 
time is usually considered to be beneficial since it provides an extended 
opportunity for heat to be transferred from the airfoil to the coolant. 
However the increased residence time and accompanying heat transfer also 
significantly raise the coolant's temperature as the coolant proceeds 
through the serpentine, thereby progressively diminishing the coolant's 
effectiveness as a heat sink. If the engine operational temperatures are 
high enough, the diminished coolant effectiveness can offset the benefits 
of lengthy coolant residence time. 
A third shortcoming is related to the desirability of maintaining a high 
coolant flow velocity, and therefore a high Reynolds Number, in internal 
cooling passages perforated by a series of coolant discharge holes. The 
accumulative discharge of coolant through the holes is accompanied by a 
reduction in the velocity and Reynolds Number of the coolant stream and a 
corresponding reduction in convective heat transfer into the stream. The 
reduction in Reynolds Number and heat transfer effectiveness can be 
mitigated if the cross sectional flow area of the passage is made 
progressively smaller in the direction of coolant flow. However a 
reduction in the passage flow area also increases the distance between the 
perimeter of the passage and the airfoil surface, thereby inhibiting heat 
transfer and possibly neutralizing any benefit attributable to the area 
reduction. 
A fourth shortcoming affects the airfoils of blades, but not those of 
vanes. Blades extend radially outwardly from a rotatable turbine hub and, 
unlike vanes, rotate about the engine's longitudinal centerline during 
engine operation. The rotary motion of the blade urges the coolant flowing 
through any of the radially extending passages to accumulate against one 
of the surfaces (the advancing surface) that bounds the passage. This 
results in a thin boundary layer that promotes good heat transfer. However 
this rotational effect also causes the coolant to become partially 
disassociated from the laterally opposite passage surface (the receding 
surface) resulting in a correspondingly thick boundary layer that impairs 
effective heat transfer. Unfortunately the receding passage surface may be 
proximate to a portion of the airfoil that is subjected to the highest 
temperatures and therefore requires the most potent heat transfer. 
It may be possible to enhance the heat transfer effectiveness in a 
conventional airfoil by providing a greater quantity of coolant or by 
using coolant having a lower temperature. In a gas turbine engine, the 
only reasonably available coolant is compressed air extracted from the 
engine compressors. Since the diversion of compressed air from the 
compressors degrades engine efficiency and fuel economy, extraction of 
additional compressed air to compensate for ineffective airfoil heat 
transfer is undesirable. The use of lower temperature air is usually 
unfeasible since the pressure of the lower temperature air is insufficient 
to ensure positive coolant flow through the turbine airfoil passages. 
Improved heat transfer can also be realized by employing trip strips whose 
height is greater than 10% of the passage lateral dimension. However this 
approach is unattractive for rotating blades since the trip strips are 
numerous and the aggregate weight arising from the use of enlarged trip 
strips unacceptably amplifies the rotational stresses imposed on the 
turbine hub. 
SUMMARY OF THE INVENTION 
It is, therefore, a primary object of the invention to provide a coolable 
airfoil for a turbine blade or vane that requires a minimum of coolant but 
is nevertheless capable of long duration service at high temperatures. 
It is a further object of the invention to provide a coolable airfoil whose 
heat transfer features are customized to the temperature distribution over 
the airfoil surface. 
It is another object of the invention to provide a coolable airfoil that 
enjoys the heat absorption benefits of a serpentine cooling passage 
without experiencing excessive coolant temperature rise. 
It is an additional object of the invention to provide a coolable airfoil 
whose coolant passages diminish in cross sectional area to maintain a high 
Reynolds Number in the coolant stream, but without inhibiting heat 
transfer due to increased distance between the perimeter of the passage 
and the airfoil surface. 
It is still another object of the invention to provide a coolable airfoil 
having features that compensate for locally impaired heat transfer arising 
from rotational effects. 
According to the invention, a coolable airfoil has an auxiliary cooling 
system that supplements a primary cooling system by absorbing excess heat 
in a predetermined zone of high heat load. 
According to one aspect of the invention, a coolable airfoil includes a 
primary cooling system comprising one or more medial passages bounded in 
part by a peripheral wall of the airfoil, and an auxiliary cooling system 
comprising one or more cooling conduits disposed in the peripheral wall 
and chordwisely situated in a zone of high heat load. 
According to another aspect of the invention, the primary cooling system 
includes an array of medial passages, at least two of which are 
interconnected to form a serpentine passage, and the auxiliary conduits 
are chordwisely coextensive with at least one of the medial passages to 
thermally insulate coolant flowing through the medial passage. 
According to still another aspect of the invention, the chordwise dimension 
of the auxiliary conduits is no more than a predetermined multiple of the 
distance from the conduits to the external surface of the airfoil so that 
thermal stresses arising from the presence of the conduits are minimized. 
In one embodiment of the invention, the auxiliary cooling system comprises 
at least two auxiliary conduits with a radially extending interrupted rib 
separating chordwisely adjacent conduits. 
In another embodiment of the invention, an array of trip strips extends 
laterally from a portion of the perimeter surface of the conduits to a 
height that exceeds about 20% of the conduit lateral dimension and is 
preferably about 50% of the conduit lateral dimension. 
The airfoil of the present invention is advantageous in that it can 
withstand sustained operation at elevated temperatures without suffering 
thermally induced damage or consuming inordinate quantities of coolant. 
More specifically, the airfoil is suitable for use in an environment where 
the temperature distribution over the airfoil's external surface is 
spatially nonuniform. Additional specific advantages include the airfoil's 
decreased susceptibility to the loss of coolant effectiveness that 
customarily arises from factors such as lengthy coolant residence time, 
progressively diminishing coolant stream Reynolds Number, and adverse 
rotational effects. 
The foregoing features and advantages and the operation of the invention 
will become more apparent in light of the following description of the 
best mode for carrying out the invention and the accompanying drawings.

BEST MODE FOR CARRYING OUT THE INVENTION 
Referring to FIGS. 1-4 a coolable turbine blade 10 for a gas turbine engine 
has an airfoil section 12 that extends radially across an engine flowpath 
14. A peripheral wall 16 extends radially from the root 18 to the tip 22 
of the airfoil 12 and chordwisely from a leading edge 24 to a trailing 
edge 26. The peripheral wall 16 has an external surface 28 that includes a 
concave or pressure surface 32 and a convex or suction surface 34 
laterally spaced from the pressure surface. A mean camber line MCL extends 
chordwisely from the leading edge to the trailing edge midway between the 
pressure and suction surfaces. 
The illustrated blade is one of numerous blades that project radially 
outwardly from a rotatable turbine hub (not shown). During engine 
operation, hot combustion gases 36 originating in the engine's combustion 
chamber (also not shown) flow through the flowpath causing the blades and 
hub to rotate in direction R about an engine longitudinal axis 38. The 
temperature of these gases is spatially nonuniform, therefore the airfoil 
12 is subjected to a nonuniform temperature distribution over its external 
surface 28. In addition, the depth of the aerodynamic boundary layer that 
envelops the external surface varies in the chordwise direction. Since 
both the temperature distribution and the boundary layer depth influence 
the rate of heat transfer from the hot gases into the blade, the 
peripheral wall is exposed to a chordwisely varying heat load along both 
the pressure and suction surfaces. In particular, a zone of high heat load 
is present from about 0% to 20% of the chordwise distance from the leading 
edge to the trailing edge along the suction surface, and from about 10% to 
75% of the chordwise distance from the leading edge to the trailing edge 
along the pressure surface. Although the average temperature of the 
combustion gases may be well within the operational capability of the 
airfoil, the heat transfer into the blade in the high heat load zone can 
cause localized mechanical distress and accelerated oxidation and 
corrosion. 
The blade has a primary cooling system 42 comprising one or more radially 
extending medial passages 44, 46a, 46b, 46c and 48 bounded at least in 
part by the peripheral wall 16. Near the leading edge of the airfoil, feed 
passage 44 is in communication with impingement cavity 52 through a series 
of radially distributed impingement holes 54. An array of "showerhead" 
holes 56 extends from the impingement cavity to the airfoil surface 28 in 
the vicinity of the airfoil leading edge. Coolant C.sub.LE flows radially 
outwardly through the feed passage and through the impingement cavity to 
convectively cool the airfoil, and a portion of the coolant flows through 
the impingement holes 54 and impinges against the forwardmost surface 58 
of the impingement cavity to impingement cool the surface 58. The coolant 
then flows through the showerhead holes and discharges as a thermally 
protective film over the leading edge of the airfoil. The cross sectional 
area A of the feed passage diminishes with increasing radius (i.e. from 
the root to the tip) so that the Reynolds Number of the coolant stream 
remains high enough to promote good heat transfer despite the discharge of 
coolant through the showerhead holes. 
Midchord medial passages 46a, 46b and 46c cool the midchord region of the 
airfoil. Passage 46a, which is bifurcated by a radially extending rib 62, 
and chordwisely adjacent passage 46b are interconnected by an elbow 64 at 
their radially outermost extremities. Chordwisely adjacent passages 46b 
and 46c are similarly interconnected at their radially innermost 
extremities by elbow 66. Thus, each of the medial passages 46a, 46b and 
46c is a leg of a serpentine passage 68. Judiciously oriented cooling 
holes 72 are distributed along the length of the serpentine, each hole 
extending from the serpentine to the airfoil external surface. Coolant 
C.sub.MC flows through the serpentine to convectively cool the airfoil and 
discharges through the cooling holes to transpiration cool the airfoil. 
The discharged coolant also forms a thermally protective film over the 
pressure and suction surfaces 32, 34. A portion of the coolant that 
reaches the outermost extremity of passage 46a is discharged through a 
chordwisely extending tip passage 74 that guides the coolant out the 
airfoil trailing edge. 
Trailing edge feed passage 48 is chordwisely bounded by trailing edge 
cooling features including ribs 76, 78, each perforated by a series of 
apertures 82, a matrix of posts 83 separated by spaces 84, and an array of 
pedestals 85 defining a series of slots 86. Coolant C.sub.TE flows 
radially into the feed passage and chordwisely through the apertures, 
spaces and slots to convectively cool the trailing edge region. 
An auxiliary cooling system 92 includes one or more radially continuous 
conduits, 94a-94h (collectively designated 94), substantially parallel to 
and radially coextensive with the medial passages. Each conduit includes a 
series of radially spaced film cooling holes 96 and a series of exhaust 
vents 98. The conduits are disposed in the peripheral wall 16 laterally 
between the medial passages and the airfoil external surface 28, and are 
chordwisely situated within the zone of high heat load, i.e. within the 
sub-zones 104, 106 extending respectively from about 0% to 20% of the 
chordwise distance from the leading edge to the trailing edge along the 
suction surface 34 and from about 10% to 75% of the chordwise distance 
from the leading edge to the trailing edge along the pressure surface 32. 
Coolant C.sub.PS, C.sub.SS flows through the conduits thereby promoting 
more heat transfer from the peripheral wall than would be possible with 
the medial passages alone. A portion of the coolant discharges into the 
flowpath by way of the film cooling holes 96 to transpiration cool the 
airfoil and establish a thermally protective film along the external 
surface 28. Coolant that reaches the end of a conduit exhausts into the 
flowpath through exhaust vents 98. 
The conduits 94 are substantially chordwisely coextensive with at least one 
of the medial passages so that coolant C.sub.PS and C.sub.SS absorbs heat 
from the peripheral wall 16 thereby thermally shielding or insulating the 
coolant in the chordwisely coextensive medial passages. In the illustrated 
embodiment, conduits 94d-94h along the pressure surface 32 are chordwisely 
coextensive with both the trailing edge feed passage 48 and with legs 46a 
and 46b of the serpentine passage 68. The chordwise coextensivity between 
the conduits and the trailing edge feed passage helps to reduce heat 
transfer into coolant C.sub.TE in the feed passage 48. This, in turn, 
preserves the heat absorption capacity of coolant C.sub.TE thereby 
enhancing its ability to convectively cool the trailing edge region as it 
flows through the apertures 82, spaces 84 and slots 86. Similarly, the 
chordwise coextensivity between the conduits and legs 46a, 46b of the 
serpentine passage 68 helps to minimize the temperature rise of coolant 
C.sub.MC during the coolant's lengthy residence time in the serpentine 
passage. As a result, coolant C.sub.MC retains its effectiveness as a heat 
transfer medium and is better able to cool the airfoil as it flows through 
serpentine leg 46c and tip passage 74. Consequently, the benefits of 
lengthy coolant residence time are not offset by excessive coolant 
temperature rise as the coolant progresses through the serpentine. 
The auxiliary conduits are chordwisely distributed over substantially the 
entire length, L.sub.S +L.sub.P, of the high heat load zone, except for 
the small portion of sub-zone 104 occupied by the impingement cavity 52 
and showerhead holes 56 and a small portion of sub-zone 106 in the 
vicinity of serpentine leg 46c. However the conduits may be distributed 
over less than the entire length of the high heat load zone. For example, 
auxiliary conduits may be distributed over substantially the entire length 
L.sub.S of the suction surface sub-zone 104, but may be absent in the 
pressure surface sub-zone 106. Conversely, conduits may be distributed 
over substantially the entire length L.sub.P of the pressure surface 
sub-zone 106 but may be absent in the suction surface sub-zone 104. 
Moreover, conduits may be distributed over only a portion of either or 
both of the subzones. The extent to which the conduits of the auxiliary 
cooling system are present or absent is governed by a number of factors 
including the local intensity of the heat load and the desirability of 
mitigating the rise of coolant temperature in one or more of the medial 
passages. In addition, it is advisable to weigh the desirability of the 
conduits against any additional manufacturing expense arising from their 
presence. 
Referring primarily to FIG. 1A, Each auxiliary conduit 94 has a lateral 
dimension H and a chordwise dimension C and is bounded by a perimeter 
surface 108, a portion 112 of which is proximate to the external surface 
28. The chordwise dimension exceeds the lateral dimension so that the 
cooling benefits of each individual conduit extend chordwisely as far as 
possible. The chordwise dimension is constrained, however, because each 
conduit divides the peripheral wall into a relatively cool inner portion 
16a and a relatively hot outer portion 16b. If a conduit's chordwise 
dimension is too long, the temperature difference between the two wall 
portions 16a, 16b may cause thermally induced cracking of the airfoil. 
Therefore the chordwise dimension of each conduit is limited to no more 
than about two and one half to three times the lateral distance D from the 
proximate perimeter surface 112 to the external surface 28. Adjacent 
conduits, such as those in the illustrated embodiment, are separated by 
radially extending ribs 114 so that the inter-conduit distance I is at 
least about equal to lateral distance D. The inter-conduit ribs ensure 
sufficient heat transfer from wall portion 16a to wall portion 16b to 
attenuate the temperature difference and minimize the potential for 
cracking. 
Each inter-conduit rib 114 is interrupted along its radial length so that 
coolant can flow through interstices 124 to bypass any obstruction or 
constriction that may be present in a conduit. Obstructions and 
constrictions may arise from manufacturing impression or may be in the 
form of particulates that are carried by the coolant and become lodged in 
a conduit. 
An array of trip strips 116 (only a few of which are shown in FIGS. 3 and 4 
to preserve the clarity of the illustrations) extends laterally from the 
proximate surface 112 of each conduit. Because the conduit lateral 
dimension H is small relative to the lateral dimension of the medial 
passages, the conduit trip strips can be proportionately larger than the 
trip strips 116' employed in the medial passages without contributing 
inordinately to the weight of the airfoil. The lateral dimension or height 
H.sub.TS of the conduit trip strips exceeds 20% of the conduit lateral 
dimension H, and preferably is about 50% of the conduit lateral dimension. 
The trip strips are distributed so that the radial separation s.sub.ts 
(FIG. 4) between adjacent trip strips is between five and ten times the 
lateral dimension (e.g. H.sub.TS) of the trip strips and preferably 
between five and seven times the lateral dimension. This trip strip 
density maximizes the heat transfer effectiveness of the trip strip array 
without imposing undue pressure loss on the stream of coolant. 
The airfoil may also include a set of radially distributed coolant 
replenishment passageways 122, each extending from a medial passage (e.g. 
passage 44, 46a and 48) to the auxiliary cooling system. Coolant from the 
medial passage flows through the passageways 122 to replenish coolant that 
is discharged from the conduits through the film cooling holes 96. The 
replenishment passageways are situated between about 15% and 40% of the 
airfoil span S (i.e. the radial distance from the root to the tip) but may 
be distributed along substantially the entire span if necessary. The 
quantity and distribution of replenishment passageways depends in part on 
the severity of the pressure loss experienced by coolant flowing radially 
through the conduit or conduits being replenished. If the conduit imposes 
a high pressure loss, a disproportionately large fraction of the coolant 
will discharge through the film cooling holes rather than proceed radially 
outwardly through the conduit. As a result, a large quantity of 
passageways will be necessary to replenish the discharged coolant. 
However, it is undesirable to have too many passageways since coolant 
introduced into a conduit by way of a replenishment passageway diverts 
coolant already flowing through the conduit and encourages that coolant to 
discharge through film cooling holes upstream (i.e. radially inwardly) of 
the passageway. If the diverted coolant still has a significant amount of 
unexploited heat absorption capability, then the coolant is being used 
ineffectively, and engine efficiency will be unnecessarily degraded. 
The replenishment passageways 122 are aligned with the interstices 124 
distributed along the inter-conduit ribs 114 rather than with the conduits 
themselves. This alignment is advantageous since the replenishment coolant 
is expelled from the passageway as a high velocity jet of fluid. The fluid 
jet, if expelled directly into a conduit, could impede the radial flow of 
coolant through the conduit thereby interfering with effective heat 
transfer into the coolant. 
During engine operation, coolant flows into and through the medial passages 
and auxiliary conduits as described above to cool the blade peripheral 
wall 16. Because the conduits are situated exclusively within the high 
heat load zone, rather than being distributed indiscriminately around the 
entire periphery of the airfoil, the benefit of the conduits can be 
concentrated wherever the demand for aggressive heat transfer is the 
greatest. Discriminate distribution of the conduits also facilitates 
selective shielding of coolant in the medial passages, thereby preserving 
the coolant's heat absorption capacity for use in other parts of the 
cooling circuit. Such sparing use of the conduits also helps minimize 
manufacturing costs since an airfoil having the small auxiliary conduits 
is more costly to manufacture than an airfoil having only the much larger 
medial passages. The small size of the conduits also permits the use of 
trip strips whose height, in proportion to the conduit lateral dimension, 
is sufficient to promote excellent heat transfer. 
The cooling conduits also ameliorate the problem of diminished coolant 
stream Reynold's Number due to the discharge of coolant along the length 
of a medial passage. For example, the presence of suction surface conduits 
94a, 94b, 94c allow the peripheral wall thickness t (FIG. 1) between 
leading edge feed passage 44 and airfoil suction surface 34 to be greater 
than the corresponding thickness in a prior art airfoil. As a result, the 
radial reduction in flow area A of the leading edge feed passage 44 is 
proportionally greater in the present airfoil than in a similar leading 
edge feed channel in a prior art airfoil. Consequently, high coolant 
stream Reynold's Number and corresponding high heat transfer rates can be 
realized along the entire length of passage 44 despite the discharge of 
coolant through showerhead holes 56 and film cooling holes 96. Moreover, 
the suction surface conduits 94a, 94b, 94c compensate for any loss of heat 
transfer from the peripheral wall attributable to the increased thickness 
t. 
The invention also helps to counteract the impaired heat transfer arising 
from rotational effects in turbine blades. During engine operation, a 
blade having an airfoil as shown in FIG. 1 rotates in direction R about 
the engine centerline 38. Coolant flowing radially outwardly, for example 
through leading edge feed passage 44, therefore tends to be urged against 
advancing surface 126 while also becoming partially disassociated from 
receding surface 128. The disassociative influence promotes the 
development of a thick aerodynamic boundary layer and concomitantly poor 
heat transfer along the receding surface. The presence of conduits 94a, 
94b, 94c compensates for this adverse rotational effect. A similar 
compensatory effect could, if desired, be obtained adjacent to the 
midchord and trailing edge passages 46a, 46b, 46c and 48. However the 
coolant in these passages is subjected to a lower heat load than the 
coolant in passage 44 and is adequately protected by the cooling film 
dispersed by film cooling holes 72. 
Various changes and modifications can be made without departing from the 
invention as set forth in the accompanying claims. For example, although 
the midchord medial passages are shown as being interconnected to form a 
serpentine, the invention also embraces an airfoil having independent or 
substantially independent midchord medial passages. In addition, 
individual designations have been assigned to the coolant supplied to the 
passages and conduits since each passage and conduit may each be supplied 
from its own dedicated source of coolant. In practice, however, a common 
coolant source may be used to supply more than one, or even all of the 
passages and conduits. A common coolant source for all the passages and 
conduits is, in fact, envisioned as the preferred embodiment.