A coolable airfoil structure having internal heat transfer features through which cooling air is flowed under operative conditions is disclosed. Various construction details and features are developed which affect the castability and core effectiveness of the airfoil after manufacture. In one particular embodiment, the airfoil has a plurality of heat transfer members disposed in the rearmost section of the trailing edge region which comprises a single impingement rib, a single row of pedestals and a single row of chordwisely extending flow dividers.

TECHNICAL FIELD
 This invention relates to coolable airfoil structures of the type used in
 high temperature rotary machines, and more specifically, to heat transfer
 elements in the trailing edge region of such airfoils. The concepts
 disclosed have application to airfoils for turbine vanes and for turbine
 blades.
 BACKGROUND OF THE INVENTION
 An axial flow rotary machine, such as a gas turbine engine for an aircraft,
 has a compression section, a combustion section, and a turbine section. An
 annular flow path for working medium gases extends axially through the
 sections of the engine. The gases are compressed in the compression
 section to raise their temperature and pressure. Fuel is burned with the
 working medium gases in the combustion section to further increase the
 temperature of the hot, pressurized gases. The hot, working medium gases
 are expanded through the turbine section to produce thrust and to extract
 energy as rotational work from the gases. The rotational work is
 transferred to the compression section to raise the pressure of the
 incoming gases.
 The compression section and turbine section have a rotor which extends
 axially through the engine. The rotor is disposed about an axis of
 rotation Ar. The rotor includes arrays of rotor blades which transfer
 rotational work between the rotor and the hot working medium gases. Each
 rotor blade has an airfoil for this purpose which extends outwardly across
 the working medium flow path. The working medium gases are directed
 through the airfoils. The airfoils in the turbine section receive energy
 from the working medium gases and drive the rotor at high speeds about an
 axis of rotation. The airfoils in the compression section transfer this
 energy to the working medium gases to compress the gases as the airfoils
 are driven about the axis of rotation by the rotor.
 The engine includes a stator disposed about the rotor. The stator has an
 outer case and arrays of stator vanes which extend inwardly across the
 working medium flowpath. The outer case extends circumferentially about
 the working medium flow path to bound the flow path. The stator has seal
 elements for this purpose, such as a circumferentially extending seal
 member which is disposed radially about the rotor blades. The seal member
 is in close proximity to the tips of the rotor blades to form a seal that
 blocks the leakage of working medium gases from the flowpath.
 The arrays of stator vanes are disposed upstream of the arrays of rotor
 blades in both the compression section and turbine section. The stator
 vanes each have an airfoil for guiding the working medium gases to the
 rotor blades as the gases are flowed along the flow path. The airfoils of
 the stator vanes and the rotor blades are designed to receive, interact
 with and discharge the working medium gases as the gases are flowed
 through the engine.
 As a result, the airfoils are bathed in hot working medium gases during
 operation of the rotor blades causing thermal stresses in the airfoils.
 These thermal stresses affect the structural integrity and fatigue life of
 the airfoil. In addition, rotational forces acting on the rotor blade
 cause additional stresses in the rotor blade, such as mechanical stresses,
 as the rotor blade is driven about the axis of rotation.
 One way to reduce stresses in rotor blades downstream of the combustion
 section in the high pressure turbine is to cool the rotor blades by
 flowing cooling fluid through the airfoil. The cooling fluid removes heat
 from the airfoil decreasing thermal stresses and avoiding unacceptably
 high temperatures for the material used for the walls of the airfoil. Each
 rotor blade has one or more openings at its inner end for receiving the
 cooling air.
 One source of cooling fluid is working medium gases from the compression
 section (pressurized air). The fluid bypasses the combustion process and
 is at a much lower temperature than the working fluid in the turbine
 section. The cooling fluid is flowed through and around various structures
 within the turbine section.
 The use of such a cooling fluid has a negative impact on the aerodynamic
 efficiency of the gas turbine engine. This occurs because the compressed
 cooling fluid bypass the combustion section where energy is added to the
 fluid and enters the flowpath with little transfer of useful energy from
 the compressed fluid to the turbine stages. The loss of efficiency is
 balanced by increased durability of the parts and higher combustion
 temperatures that increase cycle efficiency. This balancing emphasizes the
 need to use efficiently the cooling fluid drawn from the compressor
 section.
 The turbine airfoils have complex internal passages for receiving and
 discharging the cooling fluid. As the cooling fluid passes through these
 passages, heat is transferred from internal surfaces to the cooling fluid.
 The surfaces include surfaces on heat transfer members, such as trip
 strips and pedestals, to increase heat transfer between the cooling fluid
 and the turbine airfoil. The cooling fluid then exits into the flow path
 through cooling holes distributed about the airfoil.
 An example of such a rotor blade is shown in U.S. Pat. No. 4,474,532
 entitled "Coolable Airfoil For a Rotary Machine", issued to Pazder and
 assigned to the assignee of this application. Another example of a
 coolable rotor blade is shown in U.S. Pat. No. 4,278,400 issued to Yamarik
 and Levengood entitled "Coolable Rotor Blade" and assigned to the assignee
 of this application. Each of these rotor blades is provided with a
 plurality of cooling air passages on the interior of the blade. Cooling
 air is flowed through the passages to the rearmost portion of the rotor
 blade, commonly referred to as the trailing edge, from whence the cooling
 air is exhausted into the working medium flowpath.
 U.S. Pat. No. 5,368,441 issued to Sylvestro entitled "Turbine Airfoil
 Including Diffusing Trailing Edge Pedestals" has a plurality of flow
 dividers (teardrop shaped pedestals) in the trailing edge region of the
 blade. The pedestals place the interior of the airfoil of the rotor blade
 in flow communication with the exterior of the blade. The flow of cooling
 fluid directly cools the interior of the blade by impingement on the
 internal surfaces and convection as the flow proceeds through cooling
 channels. The trailing edge is cut back on the pressure side to uncover a
 diffusing region at the end of the flow dividers. A channel is formed
 between the flow dividers to diffuse cooling air over the suction surface
 of the airfoil and provide a film of cooling air over the suction wall.
 The film cooling is effective but its effectiveness decreases if the flow
 separates from the walls of the diffusing section and turbulently mixes
 with the hot working medium gases.
 Other constructions have a plurality of pedestals in the trailing edge
 region to increase heat transfer. Increasingly intricate constructions
 have been formed by casting the airfoil with a lost wax process. These
 constructions may be formed upstream of flow dividers at the trailing edge
 of the airfoil but in close proximity to increase heat transfer to the
 cooling fluid. These include constructions in which pairs of spanwisely
 extending ribs in the trailing edge region have a plurality of spanwisely
 spaced orifices for directing cooling air on adjacent internal structure.
 Impingement cooling results, improving heat transfer in the trailing edge
 of the airfoil. The ribs in such constructions are relatively small in the
 chordwise direction but relatively long in the spanwise direction. The
 small orifices when coupled with the need during the casting process to
 have wax where metal will form and core material (slurry) only where the
 small orifices are formed in the casting results in only small amounts of
 core material being disposed in the rib location of the casting core.
 The ribs may cause difficulties during the casting process. During the
 casting process of such an airfoil, the ceramic core (which defines the
 openings and structure in the trailing edge region) is disposed on the
 interior of an airfoil shaped mold. Molten metal is poured around the
 core, rushing into the mold during the pouring process. The molten metal
 fills openings in the core to form solid structure and flows around the
 solid ceramic core material to form holes, such as the orifices in the
 ribs. As the molten metal enters the structure, portions of the core in
 the trailing edge region where the heat transfer elements are intricate
 and delicately formed, may collapse resulting in an unusable casting.
 Accordingly, the complexity of the heat transfer members is balanced by
 the need to cast these features.
 The above notwithstanding, scientists and engineers working under the
 direction of Applicants Assignee have sought to develop cooling
 constructions for the trailing edge region of a rotor blade which have
 acceptable castability during manufacture and have acceptable levels of
 blade stress and metal temperature under operative conditions.
 SUMMARY OF INVENTION
 The invention is in part predicated on the recognition that overall
 aerodynamic considerations establish the external contour of an airfoil
 while the internal contour depends on casting and internal cooling needs.
 As a result, the arrangement of heat transfer features on the interior may
 be combined and located within the airfoil to improve core strength,
 castability of the airfoil, and efficiency of the engine.
 According to the present invention, an airfoil has a plurality of heat
 transfer members disposed in the rearmost section of the trailing edge
 region which comprises a single impingement rib, a single row of pedestals
 and a single row of chordwisely extending flow dividers, each member
 having an overall hydraulic diameter that is less than the hydraulic
 diameter of the preceding member and located to increase the velocity of
 the flow rearwardly and for reasons of castability and core strength
 during manufacture, and strength of the airfoil and cooling effectiveness
 for the airfoil after manufacture.
 In accordance with one detailed embodiment of the present invention, the
 chordwise length of the trailing edge region is less than five times the
 length Lf in the cutback trailing edge (that is, the uncovered diffusing
 section of the flow dividers).
 In accordance with one embodiment of the present invention, a plurality of
 film cooling holes extend from the interior to the exterior of the airfoil
 on either chordwise side of the pedestals and the entrance of the film
 cooling hole is in flow communication with a supply passage between the
 pedestals and the rib and the exit is outwardly of the supply passage
 between the pedestal and the flow dividers for delivering a film of
 cooling air along the pressure wall upstream of the cutback trailing edge.
 In accordance with one detailed embodiment of the invention, the supply
 passage and film cooling holes are located within a length measured from
 the pressure wall lip which is less than four times the length Lf in the
 cut back trailing edge (uncovered diffusing section) and the exit of the
 film cooling holes is within two and half times the length Lf.
 A primary feature of the present invention is the chordwise length of the
 trailing edge region. Another feature is the length from the pressure wall
 lip to the supply passage for cooling air for the flow dividers. Still
 another feature is the lateral distance between the pressure wall and
 suction wall which contracts rapidly from the rib to the circular
 pedestals and is relatively constant from the circular pedestals to the
 flow dividers over the rearmost fluid supply passage. Another feature is
 the location of the three arrays of heat transfer members in the trailing
 edge region such that the trailing edge region is provided with two arrays
 of heat transfer members that provide strength during operative conditions
 of the airfoil (rib, flow dividers) and two arrays of heat transfer
 members (pedestals, spanwisely spaced dividers) that provide strength to
 the core during fabrication of the airfoil.
 A primary advantage of the present invention is the durability and cooling
 efficiency of the trailing edge region of the airfoil which results from
 increasing the velocity of the cooling fluid as it moves rearwardly in the
 airfoil due to the cooperation and arrangement of the heat transfer
 members. Still another advantage is the castability of the airfoil which
 results from having at least two heat transfer members having a level of
 core material during the casting process which supports the core and
 strengthens the core during handling and casting of the airfoil. Still
 another feature is the durability of the airfoil which results from
 positioning the heat transfer members in the trailing edge region such
 that the heat transfer members provide material that extends between the
 pressure wall and the suction wall to support the trailing edge region
 during fabrication, handling, and operative use of the airfoil. An
 advantage of one particular embodiment having a larger angle of diffusion
 on the inwardly facing wall of the flow dividers than the outwardly facing
 wall is the rate of heat transfer which results from the thinner wall of
 the downstream portion of the flow divider acting as a more efficient fin
 to remove heat from the pressure wall and the suction wall and transfer
 the heat to cooling fluid in channels bounded by the sidewall. Still
 another advantage of this one embodiment is the durability and level of
 thermal gradients during transient conditions between the pressure and
 suction walls and the flow divider as compared to constructions which have
 thicker flow dividers.
 The foregoing features and advantages of the present invention will become
 more apparent in light of the following detailed description of the best
 mode for carrying out the invention and accompanying drawings.

BEST MODE
 FIG. 1 is a schematic, side elevation view of a rotary machine 10, such as
 a turbofan gas turbine engine. The engine is disposed about an axis of
 symmetry A and has an axis of rotation Ar. The engine includes a
 compression section 12, a combustion section 14, and a turbine section 16.
 An annular, primary flowpath 18 for working medium gases extends axially
 through the sections of the engine. A by-pass flowpath 20 is outward of
 the primary flow path.
 The engine is partially broken away to show a stator 22 and a rotor 24 in
 the compression section 12 and the turbine section 16. The stator 22
 includes having an outer case 26 which extends circumferentially about the
 primary flowpath. The stator includes arrays of stator vanes, as
 represented by the stator vane 28 in the compression section and the
 stator vane 32 in the turbine section. Each stator vane has an airfoil 34,
 36 which extends inwardly from the outer case to direct the flow of
 working medium gases as the gases pass through the compression section and
 the turbine section.
 The rotor has arrays of rotor blades, as represented by the rotor blade 38
 in the compression section 12 and the rotor blade 42 in the turbine
 section 16. Each rotor blade has an airfoil 44, 46 which extends radially
 outwardly across the working medium flow path into close proximity with
 the stator 22. In the turbine section, the rotor has a rotor assembly 48
 which includes a rotor disk 52. The rotor blades engage the rotor disk and
 are carried by the disk about the axis of rotation Ar.
 FIG. 2 is a perspective view of the rotor blade 42 from the high pressure
 portion of the turbine section 16. The rotor blade is disposed about a
 spanwisely extending axis As. The turbine blade has a root 56 having an
 opening 58. The opening adapts the airfoil 46 to be in flow communication
 with a source of cooling air under operative conditions. Each rotor blade
 has a platform 62 outwardly of the root. The platform inwardly bounds the
 working medium flowpath 18. The airfoil has an opening (not shown) which
 adapts the airfoil to receive cooling air from the root of the rotor
 blade. Cooling passages (not shown) extend internally of the airfoil and
 are in flow communication with the exterior of the airfoil through
 openings, as represented by openings such as the film cooling holes 64 and
 the channels 66 in the exterior of the airfoil.
 The airfoil has a leading edge 68, a leading edge region 72, a trailing
 edge 74, and a trailing edge region 76. A plurality of airfoil sections,
 as represented by the airfoil section 78, are disposed chordwisely about
 the spanwise axis As to define the contours of the airfoil (as used
 herein, plurality means an indefinite number of two or more). The airfoil
 has a chordwise direction C and a spanwise direction S that provide
 reference directions. The spanwise direction is generally perpendicular to
 the axis of rotation Ar. The spanwise extent of the airfoil is divided
 into an inner portion 82, an outer portion 84, and a midspan portion 86
 that extends spanwisely from the inner portion to the outer portion. The
 inner portion typically ranges from the platform to about the thirty
 percent (30%) span location of the airfoil, the midspan portion ranges
 from about the thirty percent (30%) span location to the seventy percent
 (70%) span location, and the outer or tip portion ranges from the seventy
 percent (70%) span location to the tip of the airfoil.
 The airfoil 46 has a suction wall 88 extending chordwisely and spanwisely
 about the back side of the airfoil. The suction wall extends from the
 leading edge region 72 to the trailing edge region 76. The suction wall
 has a suction wall lip 92 disposed in the trailing edge region. The
 airfoil has a pressure wall 94 extending chordwisely and spanwisely about
 the opposite side (front side) of the airfoil. The pressure wall extends
 from the leading edge region to the trailing edge region.
 FIG. 3 is a sectional view of the rotor blade 42 shown in FIG. 2 with
 portions of the rotor blade broken away except for the trailing edge
 region 76 and for structure adjacent the trailing edge region. As shown in
 FIG. 2 and in FIG. 3, the pressure wall 94 has a pressure wall lip 96
 which is chordwisely (axially) upstream of the suction wall lip 92. The
 pressure wall lip and the suction wall lip define a cutback trailing edge
 .The pressure wall is spaced laterally from the suction wall 88 leaving an
 internal cavity 98 for cooling fluid therebetween. The cavity is in flow
 communication with the opening for cooling fluid in the root through the
 internal opening in the airfoil. A spanwisely extending passage 102 ducts
 the flowpath 104 for cooling fluid in the spanwise direction. The
 spanwisely extending flowpath supplies cooling fluid to the trailing edge
 74 after the flowpath turns chordwisely into the trailing edge region 76.
 FIG. 3, FIG.4. and FIG. 5 show the relationship of the pressure and suction
 walls 88, 94 to structural heat transfer members in the trailing edge
 region. The heat transfer members are represented by a rib 106, an array
 of pedestals 108, and an array of flow dividers 112. The airfoil section
 78 is taken in the midspan portion 86 of the airfoil. A plurality of
 airfoil sections disposed about the spanwise axis As define the contour of
 the airfoil.
 FIG. 4 is a perspective view partially broken away of the trailing edge
 region. FIG. 5 is a schematic perspective view corresponding to FIG. 4.
 The pressure wall is broken away in FIG. 5 to show the relationship of the
 heat transfer members to the suction wall in the midspan portion of the
 airfoil.
 FIG. 3 and FIG. 4 show the internal passage 102 which extends spanwisely in
 the airfoil and which has the spanwisely extending flowpath 104 for
 cooling fluid. The trailing edge region 76 is the portion of the airfoil
 that receives the spanwisely flowing cooling fluid and directs the fluid
 rearwardly in the chordwise direction to the trailing edge 74. In the
 installed condition, the chordwise direction with respect to the engine is
 generally axial. Thus, a trailing edge region extends rearwardly from the
 location at which the flow becomes chordwise to the trailing edge.
 The heat transfer members 106, 108, 112 extend laterally to join the
 suction wall 88 to the pressure wall 94. As shown, the structural heat
 transfer members disposed in the trailing edge region also bound the
 trailing edge region. These three rearmost members 106,108,112 of the
 airfoil include the single rib 106 which extends spanwisely in the
 trailing edge region. The rib extends laterally from the suction wall to
 the pressure wall. The rib has a leading edge 114 and a trailing edge 116.
 The rib has an upstream height Hru, a downstream height Hrd and an average
 height Hr. The rib has a plurality of orifices 118 having an inner
 diameter Rd extending therethrough. The orifices are bounded by a curved
 surface 122. The curved surfaces are somewhat similar to the inner curved
 surface of a torus. The orifices are spaced spanwisely by a distance Pr
 leaving rib material extending therebetween. The orifices are spaced
 laterally from the walls leaving rib material extending between the walls
 and the openings.
 FIG. 6 shows an enlarged view of the portion of the trailing edge region
 shown in FIG. 3 and a plan view taken along the line 6--6 of FIG. 3. As
 shown in FIG. 5 and FIG. 6, the single array of pedestals 108 is spaced
 chordwisely from the rib 106 leaving a first chordwisely extending supply
 passage 124 for cooling fluid therebetween. The array of pedestals extends
 spanwisely in the airfoil. Each pedestal extends from the pressure wall 94
 to the suction wall 88 to join the walls. Each pedestal has a leading edge
 126 and a trailing edge 128. Each pedestal is spaced spanwisely by a
 distance Pp from the adjacent pedestal leaving an opening 132
 therebetween. The opening between pedestals extends chordwisely between
 the pedestals and from the suction wall to the pressure wall. Each
 pedestal has an upstream height Hpu, a downstream height Hpd and an
 average height Hp.
 The plurality of film cooling holes 64 extend from the exterior of the
 airfoil through the pressure wall 94 to the first supply passage 124. This
 places the interior of the airfoil in flow communication with the exterior
 to provide a film of cooling air outwardly of the pedestals in a region of
 increased thickness for the pressure wall.
 The array of flow dividers 112 is disposed in the cutback trailing edge and
 is spaced chordwisely from the pedestals 108. This leaves a second
 chordwisely extending supply passage 134 between the pedestals 108 and the
 flow dividers. Cooling channels 66 are disposed between adjacent flow
 dividers and provide fluid communication between the trailing edge region
 76 of the rotor blade and the flowpath 18. Each of the flow dividers is
 teardrop shaped in cross-section. Each of the flow dividers has a rounded
 leading edge 136 and an axially converging section 138. The rounded
 leading edge faces chordwisely inward toward the second supply passage
 134. As shown in FIG. 6. the rounded leading edge is a semicircle in
 cross-section having a diameter Df. In the embodiment shown, the leading
 edge forms an upstream portion of the flow divider and terminates at a
 downstream edge 139, along the flat part of the flow divider and tangent
 to the semicircle. Other noncircular blunt shapes may be equally
 applicable to the leading edge of the flow dividers.
 The axially converging section 138 of the flow dividers includes a pair of
 converging sidewalls 142, 144, each of which extends from the downstream
 end 139 of the upstream portion. The pair of converging sidewalls is
 represented by a first inwardly facing sidewall 144 and a second outwardly
 facing sidewall 142.
 The converging walls nearly converge at their downstream end 146 with the
 opposing angled side. As shown in FIG. 4, FIG. 5, and FIG. 6, the
 converging sidewalls do not completely converge to a point but approach
 each other. The angled sides or sidewalls of adjacent flow dividers
 cooperate with the pressure wall 94 and suction wall 92 to define a
 covered diffusing section 148 between adjacent flow dividers. The angle
 .alpha. at which each of the walls converges is also the angle of
 diffusion for the flow and for the channel 66. The angle .alpha. and the
 suggested magnitude which lies in a range of two degrees to twelve
 degrees, was discussed at length in U.S. Pat. No. 5,368,441 issued to
 Sylvestro entitled "Turbine Airfoil Including Diffusing Trailing Edge
 Pedestals" the disclosure of which is incorporated herein by reference.
 The covered diffusing section 148 begins upstream of the pressure wall lip
 96 and extends downstream between the pressure wall and the suction wall.
 An uncovered diffusing section 152 is defined by the angled sidewalls and
 the suction wall. The uncovered diffusing section extends downstream of
 the pressure wall lip 96 and to the suction wall lip 92. The uncovered
 diffusing section of the flow dividers has a length Lf extending
 rearwardly from the pressure wall lip.
 The length Lf of the uncovered diffusing section 152 and the minimum
 hydraulic diameter Dhm of the channel 66 are closely related to the heat
 transfer requirements of the trailing edge region 76. These two dimensions
 may be used to non-dimensionalize the physical relationships of the heat
 transfer members in the trailing edge region. The hydraulic diameter of a
 section is defined as Dhm=(4*A)/P, that is, four times the flow area of
 the section divided by the perimeter, where Dhm is the minimum diameter at
 the section, A is the cross-sectional area of the channel, and P is the
 perimeter of the channel. The location of the minimum hydraulic diameter
 is at a section generally perpendicular to the flow. In this embodiment
 with a circular leading edge, the section passes through the center of
 each semicircle defining the leading edge 136 of two adjacent flow
 dividers.
 The length Lf of the flow divider in the uncovered diffusing section 152 is
 less than or equal to five times the minimum hydraulic diameter Dhm.
 The length Lf of the flow divider in the uncovered diffusing section may
 possibly be shorter depending on the angle of diffusion. The length is
 suggested to lie in a range of three to four times the minimum hydraulic
 diameter Dhm (3 Dhm.ltoreq.Lf&lt;4 Dhm). Alternate embodiments are discussed
 in FIG. 8 and FIG. 9 which have the potential for decreasing the length
 Lf.
 The spanwisely extending part of the flowpath for cooling fluid 104 in the
 spanwise supply passage 102 is bounded by the leading edge 114 of the rib
 106. The trailing edge region extends from the leading edge of the rib to
 the suction wall lip and such that the heat transfer members are
 coextensive with the trailing edge region. The total length of the
 trailing edge region Lpte as measured from the pressure wall lip (pressure
 wall trailing edge) is less than five times the length Lf of the uncovered
 diffusing section 152 or less than six times the length Lf as measured
 from the suction wall lip.
 FIG. 7 is a side elevation view of the pedestal 108 shown in FIG. 6. The
 leading edge 126 has a cylindrically shaped portion 154 which extends
 approximately eleven mils (0.011 inches). The cylindrically shaped portion
 transitions into a blend radius of about five mils (0.005 inches). At the
 trailing edge 128, the cylindrical portion has decreased in height to
 approximately four mils (0.004 inches) with the blend radius of about five
 mils (0.005 inches). Accordingly, the lateral height of the pedestal is
 less than the adjacent blend radius. The height of the pedestal at the
 leading edge is about twenty mils (0.020 inches) and at the trailing edge
 about fifteen mils (0.015 inches). The pedestal has a diameter of
 twenty-five mils (0.025 inches) as measured perpendicular to the
 cylindrical portion of the pedestal.
 In the embodiment shown in FIG. 3, the rib 106 has a height of Hru 51 mils
 (0.051 inches) and a height of 44 mils (0.044 inches) Hrd at the trailing
 edge. The axial or chordwise thickness of the rib is 30 mils (0.030
 inches). The diameter Rd of the orifice 118 is 32 mils (0.032 inches). The
 spanwise spacing between the openings is 73 mils (0.073 inches). The first
 supply passage 124 has a chordwise length La of 87 mils (0.087 inches).
 The pedestals are spaced apart by a distance Pp of 73 mils (0.073 inches).
 The length Lb is the diameter of the pedestals and is 25 mils (0.025
 mils). The second supply passage 134 for supplying cooling fluid to the
 flow dividers 112 has a length Lc of 72 mils (0.072 inches). The flow
 dividers have a leading edge that has a diameter Df of 31 mils (0.031
 inches). The covered diffusing section 148 has a length Lc of 77 mils
 (0.077 inches) over which the converging sidewalls 142, 144 are covered.
 The length Lf of the uncovered diffusing section 152 is equal to 66 mils
 (0.066 mils). The minimum hydraulic diameter Dhm of channel 66 is about 19
 mils (0.019 inches) in the midspan portion 86 of the airfoil 46. The
 suction wall has a thickness Ts which ranges from about forty three mils
 (0.043 inches) at the rib 106 to about forty mils (0.040 inches) at the
 pedestals 108 and thirty five mils at the flow dividers 112. The pressure
 wall 94 has a thickness of about thirty mils (0.038 inches) at the rib
 106, forty mils (0.040 inches) at the pedestals 108, and a much thinner
 twenty-nine mils (0.029 inches) at the flow dividers 112.
 FIG. 8 shows an alternate embodiment 156 of the flow dividers 112 and is
 shown in a converging diffusing section 158. The pair of converging
 sidewalls 162,164 are converging in the downstream direction such that
 spanwisely facing sidewalls of adjacent dividers diverge to define the
 diffusing section 158 between adjacent dividers. Each flow divider has a
 first converging sidewall 162 which faces spanwisely inwardly. The first
 converging sidewall is angled away from the direction of flow between flow
 dividers with an angle .alpha. as measured with respect to a line parallel
 to the axis of rotation Ar. In the embodiment shown, the angle .alpha.
 might also be measured with respect to a reference line perpendicular to a
 spanwise line connecting the centers of the leading edge circle. A second
 converging sidewall faces spanwisely outwardly. The second converging
 sidewall is angled away from the direction of flow between flow dividers
 with an angle .beta. as measured with respect to a line parallel to the
 axis Ar. The angle .alpha. of the inwardly facing sidewall is greater than
 the angle .beta. of the outwardly facing sidewall. The suggested magnitude
 for the diffusing angle .alpha. of the flow divider 156 lies in a range of
 about two (2) degrees to about fifteen (15) degrees
 (2.degree..ltoreq..alpha..ltoreq.15.degree.). The angle .beta. of the
 outwardly facing second converging sidewall lies in a range of one degree
 to twelve degrees (1.degree..ltoreq..beta..ltoreq.12.degree.). The range
 of the diffusion angle is due to the relationship of the angle .alpha. and
 the angle .beta. and is discussed below.
 The pressure wall lip 166 is aligned with the suction wall lip 168. Thus,
 the diffusing section extends from upstream to at least one of the lips.
 In an alternate embodiment, the pressure wall lip might be upstream of the
 suction wall lip as shown by the broken lines 166a to form a cutback
 trailing edge.
 FIG. 9 is an alternate embodiment 170 of the flow dividers 156 shown in
 FIG. 8. The flow divider 170 has a constant thickness section 172. The
 constant thickness section includes a pair of parallel sidewalls 174 which
 are radially spaced from each other. Each of the parallel sidewalls is
 also radially spaced from and parallel to a parallel sidewall of an
 adjacent flow divider. The parallel sidewalls of adjacent flow dividers,
 in conjunction with the pressure wall and suction wall, define constant
 area channels 176 of constant width between the flow dividers.
 The converging section of the flow dividers includes a pair of converging
 sidewalls 176, 178. In this embodiment, the angle .alpha.2 of the radially
 outer flow divider 178o is slightly greater than the angle .alpha.1 if the
 radially inner flow divider 170i.
 During operation of the gas turbine engine shown in FIG. 1, hot working
 medium gases are flowed along the working medium flowpath 18 over the
 airfoil 46. Cooling fluid is flowed through internal passages to cool the
 airfoil. The cooling fluid follows flowpaths for cooling fluid, such as
 the flowpath 104 in passage 102. A portion of this cooling fluid is
 ejected rearwardly through the cutback trailing edge 74, again along
 flowpath 104, and into the flowpath 18. The passage 102 acts as a manifold
 during this process for the trailing edge region 76, ducting cooling fluid
 outwardly and distributing the cooling air to the rib 106.
 The cooling fluid enters the orifices 118 of the rib 106 which increases
 the velocity and the turbulence to the flow. The hydraulic diameter of the
 flow path decreases and the heat transfer coefficient increases as the
 fluid passes through the rib. Heat is transferred from the rib to the
 fluid.
 As the flow enters the first supply passage 124, the flowpath 104 for
 cooling fluid rapidly contracts across the first supply passage. The
 average height of the pedestal Hp is less than forty percent of the
 average height of the rib Hr. This causes the flow area and hydraulic
 diameter of the flowpath to decrease rapidly. Again, the velocity of the
 flow and increases and the convective heat transfer coefficient increases.
 The effect continues with the downstream height Hpd of the pedestal 108
 being less than eighty percent (80%) of the upstream height of the
 pedestal Hpu to contract the area of the flowpath 104 through the
 pedestals. Again, the hydraulic diameter of the flowpath 104 decreases as
 the flow path passes between the pedestals. The relatively small
 downstream height of the pedestal also decrease the average height of the
 second supply passage 134.
 The decreasing size of the flowpath 104 continues. The height Hfu of the
 flow divider 112 at the leading edge of the flow divider is less than
 ninety percent (90%) of the downstream height Hpd of the pedestal. This
 further decreases the average height of the second supply passage. In
 fact, the lateral height of the second supply passage Hsp2 (as measured
 along a line extending perpendicular to the spanwise distribution of the
 array of pedestals and the array of flow dividers), is equal to the height
 Hfu of the leading edge of the flow divider 112 over at least three
 fourths of the length of the second supply passage. This is made possible
 by locating the bulk of the second supply passage so far rearwardly that
 the interior surface of the suction wall and pressure wall are parallel
 and in close proximity to each other. As a result, the second supply
 passage has a low aspect ratio of width divided by height. The low aspect
 ratio causes a lowering of the hydraulic diameter and further increases
 the heat transfer coefficient to effectively use cooling fluid in this
 critical location of the airfoil.
 One advantage using a design having a rib-spaced pedestal-spaced flow
 divider configuration is the ability to dispose the design in the
 laterally close (narrow) portion of the trailing edge region 76. The
 design makes this possible because the rib is made small to the extent
 permitted by casting consideration. The same considerations permit use of
 a pedestal smaller than the rib. Thus, the same casting consideration
 allow cylindrical pedestals and their spacing in the even tighter lateral
 space between the pressure wall and suction wall in close proximity to the
 flow dividers. Accordingly, a feature is the chordwise location of the
 heat transfer members one to the other, and their spacing from the
 trailing edge region between the rapidly converging pressure wall and
 suction wall of the rearmost portion of the airfoil.
 A particular advantage of the rib-pedestal-flow divider configuration
 occurs during casting and during handling and use of the airfoil. The
 configuration provides for increased castability because the array of
 pedestals and the flow divider provide more core material in the trailing
 edge region during the casting process than does the rib. The increased
 core material is provided by core material disposed spanwisely between the
 array of adjacent pedestals to form the pedestal openings and core
 material disposed chordwisely and spanwisely between the array of flow
 dividers to form the channels. This increased material at the very end of
 the trailing dengue region provides support during the casting process and
 handling of the core for the airfoil before casting.
 The rib and flow divider have a relatively larger mass of material than the
 pedestal in the finished turbine blade. During handling and processing of
 the airfoil after casting and during operative conditions, the impingement
 rib and the flow dividers cooperate with the pressure wall and suction
 wall to form a box-like structure for strengthening the trailing edge
 region of the airfoil. The pedestals are disposed chordwisely between the
 rib and dividers and reinforce the box-like structure.
 One variable feature of the present embodiment that effects rearwardly
 locating the heat transfer members 106,108,112 is the distance that the
 flow divider 112 extends upstream of the pressure wall lip 96. The leading
 edge 136 of the flow divider is spaced upstream from the pressure wall lip
 96 by a length Ld,e. The length Ld,e is equal to the length Ld of the
 covered channel of the divider. The length Le is the length (radius) of
 the curved edge 136 of the flow divider.
 The length Ld,e is less than one and one half times the length Lf of the
 uncovered diffusing section 152downstream of the pressure wall lip 96.
 This places the second supply passage 134, the array of pedestals 108 and
 the first supply passage 124 in close proximity to the pressure wall lip
 96 of the trailing edge region and between the rapidly converging pressure
 wall and suction wall. The length Lc is the length of the second supply
 passage. The length Ld, Le, and Lc extends from the pressure wall lip to
 the downstream side of the array of pedestals Pd and is less than three
 times the length Lf of the uncovered diffusing section 152.
 Shortening the length Ld places the first supply passage 124 in proximity
 to the pressure wall lip 96 (trailing edge) and makes possible the
 rearward location of the film cooling passage 64. The film cooling passage
 extends from the first supply passage upstream of the pedestals to a
 location on the exterior of the pressure wall which is just downstream of
 the array of pedestals. On either side of the pedestal, the film cooling
 passage is less than a length of the diameter of the pedestal Dp away from
 the pedestal at the surfaces of the pressure wall.
 The length to the film cooling passge at the exterior of the pressure wall
 is Lpfc. Lpfc is less than two and a half times the length Lf from the
 pressure wall lip. This provides a film of cooling fluid through film
 cooling holes at a location with respect to the pressure wall which is
 chordwisely aligned with the second supply passage.
 As the cooling fluid exits the cutback trailing edge, the cooling fluid
 flows through the channels 66 defined by the spanwisely disposed flow
 dividers 112. Heat is transferred directly from the pressure wall 94 and
 suction wall 88 to the cooling fluid and indirectly through the flow
 dividers. The exiting cooling fluid first impinges upon the flow divider
 leading edge 136 to provide impingement type cooling of the flow divider
 and to indirectly transfer heat from the pressure wall and suction wall.
 The cooling fluid then passes through the channel 66 in which additional
 heat is transferred. The cooling fluid then enters the covered diffusing
 section 148 where the diffusing of the cooling fluid begins.
 As the cooling fluid is diffused, the static pressure within the cooling
 fluid increases and the velocity of the cooling fluid decreases. The
 covered diffusing region permits the cooling fluid to begin diffusing
 before engaging with the working fluid flowing externally to the airfoil
 portion and over the suction surface lip. Upon exiting the covered
 diffusing region the cooling fluid continues diffusing over the suction
 surface lip and provides a film of cooling fluid over the suction surface
 lip. This film of cooling fluid provides a buffer between the hot working
 fluid and the suction surface lip and cools the suction surface wall.
 Beginning the diffusion upstream of the pressure surface lip and at
 velocities which are increased by reducing flow area provides means to
 initiate controlled diffusion before the diffusing cooling fluid exits the
 airfoil portion and is engaged by the hot working fluid flowing over the
 airfoil portion. Controlled diffusion upstream of the pressure lip results
 in an orderly and efficient diffusing film of cooling fluid over the
 suction surface lip.
 The angled sides of the converging region form an angle .alpha. with a line
 parallel to the straight sides of the constant thickness section, as shown
 in FIG. 6. It is suggested that the diffusion angle .alpha. in a range of
 two degrees to ten degrees (2.degree.&lt;.alpha.&lt;10.degree.. Diffusion angles
 .alpha. less than two degrees (2.degree.) may not provide sufficient
 diffusion and diffusion angles .alpha. greater than ten (10.degree.) may
 result in flow separation from the angled sides depending upon other flow
 characteristics through the channels.
 A particular advantage of the embodiment shown in FIG. 8 and FIG. 9 is the
 ability to have a greater angle of diffusion along the inwardly facing
 first converging sidewall 162, 176 than along the outwardly facing
 converging sidewalls 164,178 of the adjacent flow divider. An angle
 .alpha. of two (2) degrees to about fifteen (15) degrees
 (2.degree..ltoreq..alpha..ltoreq.15.degree.) is suggested with a smaller
 magnitude for the angle .beta.. An angle .beta. for the outwardly facing
 second converging sidewall is suggested to lie in a range of one degree to
 twelve degrees (1.degree..ltoreq..beta..ltoreq.12.degree.).
 The difference in angles is made posssible by rotation of the rotor blade
 about the axis of rotation Ar. Rotation at the high angular velocities of
 a typical rotary machine causes the inwardly facing surface to exert a
 centripetal force on the cooling fluid as it passes through the channels
 66. This centripetal force causes the fluid to exert an equal and opposite
 centrifugal force on the sidewall. These forces delay and even prevent
 separation of the flow as compared to equivalent flow at the same angle of
 diffusion on the outwardly facing sidewall. Alternatively, this enables
 greater diffusion on the outer, inwardly facing wall for a given rearward
 axial length. Thus , the same amount of diffusion may occur but in a
 shorter axial length Lf, and creates a shorter diffusing section 148,152.
 A shorter, or more angled flow divider, has reduced mass which increases
 the ability of the flow divider to act as a heat transfer rib or fin. It
 also decreases the mass of material subjected to thermal gradients in the
 material and resulting thermal stresses. Accordingly angling the outer
 wall outwardly with an angle greater than the angle of the outwardly
 facing inner wall provides for more effective utilization of the cooling
 fluid and increased durability of the trailing edge region of the rotor
 blade.
 Although the invention as shown in FIGS. 2-7 is applied to a turbine blade,
 it should be obvious to those skilled in the art that this embodiment of
 the invention is equally applicable to turbine vanes or any turbomachine
 airfoil having internal cooling and a cooling exit disposed along the
 trailing edge of the turbine airfoil.
 Although the invention has been shown and described with respect with
 exemplary embodiments thereof, it should be understood by those skilled in
 the art that various changes, omissions, and additions may be made
 thereto, without departing from the spirit and scope of the invention.