Patent Publication Number: US-2023160318-A1

Title: Vane ring assembly for a gas turbine engine with dedicated through-flow vanes

Description:
FIELD OF THE DISCLOSURE 
     The present disclosure relates generally to gas turbine engines, and more specifically to directing airflow through the turbine section of a gas turbine engine. Further, it relates to directing cooling air through vane rings in the turbine section of a gas turbine engine. 
     BACKGROUND 
     Gas turbine engines are used to power aircraft, watercraft, generators, and the like. Gas turbine engines typically include a compressor, a combustor, and a turbine. The compressor compresses air drawn into the engine and delivers high pressure air to the combustor. In the combustor, fuel is mixed with the high pressure air and is ignited. Products of the combustion reaction in the combustor are directed into the turbine where work is performed to drive the compressor and, sometimes, an output shaft. Left-over products of the combustion are exhausted out of the turbine at high temperatures and may provide thrust in some applications. 
     Compressors and turbines typically include alternating stages of static vane ring assemblies and multiple rotating wheel assemblies. Each static vane ring assembly provides structural mounting, transmits reaction loads, and contains a radially disposed set of airfoils designed to channel combustion gases through the primary flow path annulus into the turbine blades which extract energy from the flow. 
     SUMMARY 
     The present disclosure may comprise one or more of the following features or combinations thereof. 
     According to one aspect of the present disclosure, a vane ring assembly adapted for use in a gas turbine engine includes an airfoil set, an outer band, and an inner band. The airfoil set extends radially across a primary flow path annulus arranged around a central reference axis. The outer band extends around the airfoil set to define a radially-outer boundary of the primary flow path annulus. The inner band extends around the central reference axis at the airfoil set to define a radially-inner boundary of the primary flow path annulus. The inner band is shaped to block the in-flow cooling air in the in-flow internal cavities from moving radially-inward of the primary flow path annulus. 
     In some embodiments, the airfoil set may include a plurality of in-flow vanes. Each of the in-flow vanes may have an in-flow transfer passage configured to receive the in-flow cooling air from a radially-outwardly opening in-flow aperture. The in-flow transfer passage may discharge cooling air into the primary flow path annulus via cooling holes formed on an exterior surface of the in-flow vanes to cool the in-flow vanes. 
     In some embodiments, the airfoil set may further include a plurality of through-flow vanes. Each of the through-flow vanes may have a through-flow transfer passage configured to receive through-flow cooling air from a radially-outwardly opening through-flow aperture. The through-flow vanes may discharge at least some of the through-flow cooling air radially inwardly of the primary flow path annulus into an air feed passage, cooling the through-flow vane. The air feed passage may cool other components of the gas turbine engine axially forward or aft of the vane ring assembly or may prevent primary hot gas ingestion. 
     In some embodiments, the outer band may be formed to define the in-flow apertures from which the in-flow internal cavities of the in-flow vanes receive the in-flow cooling air. The outer band may be formed to include the through-flow aperture through which the through-flow internal cavities of the through-flow vanes receive the through-flow cooling air. In some embodiments, the in-flow vane and through-flow vane may be integrated into a one-piece monolithic component. 
     In some embodiments, the inner band may cause all cooling air received by the in-flow internal cavities of the in-flow vanes to be discharged into the primary flow path annulus. The inner band may be formed to define a plurality of purge air metering orifices. Each of the purge air metering orifices may be located in fluid communication with a corresponding through-flow transfer passage of one of the through-flow vanes. The purge air metering orifices may be shaped so that at least some of the cooling air received by the through-flow internal cavities of the through-flow vanes is discharged radially-inwardly of the primary flow path annulus. The purge air may be delivered to other components upstream or downstream of the vane ring assembly. This may allow the cooling air to be re-used to cool the components or prevent primary hot gas ingestion after passage through the plurality of through-flow vanes. 
     In some embodiments, the in-flow transfer passage of the in-flow vane may be formed to include a plurality of turbulators. Each of the turbulators are arranged on a surface of the in-flow transfer passage in contact with the in-flow cooling air. The turbulators may be configured to cause heat transfer by inducement of turbulent air flow over the surface contacting the in-flow cooling air. 
     In some embodiments, the through-flow transfer passage of the through-flow vanes may have, on a surface contacting the through-flow cooling, air a plurality of turbulators. The numerical number of total cooling features may be fewer than the numerical total of cooling features disposed on any surface in contact with the in-flow air within the in-flow transfer passage. In some embodiments, the through-flow transfer passage of the through-flow vanes may be free of turbulators on all surfaces contacting the through-flow cooling air. 
     In some embodiments, the trailing edge of the through-flow vanes may be formed to include cooling holes. In some embodiments, the trailing edge of the through-flow vanes may be formed without the cooling holes. 
     In some embodiments, at least part of the trailing edges may be composed of a solid or hollow first material distinct. The first material may be distinct from a second material forming the remainder of the through-flow vane. The second material may be able to withstand, without degradation in structure or shape, temperatures in excess of those withstood by the second material from which the remainder of the through-flow vanes are composed. In some embodiments, the first material of the trailing edges of the through-flow vanes may be formed of a ceramic matrix composite material. 
     In some embodiments, the vane ring seal assembly may further include a seal. The seal may be located radially inward of the inner band. 
     In some embodiments, the seal may be configured to partition an upstream air feed passage in fluid communication with components upstream of the vane ring assembly. The seal may partition a second downstream air feed passage in fluid communication with components downstream of the vane ring assembly. 
     In some embodiments, the vane ring assembly may further include a plurality of purge air metering orifices on the inner band. Each of the purge air metering orifices may be shaped to distribute through-flow cooling air at preselected levels between the upstream air feed passage and the downstream air feed passage. 
     In some embodiments, the vane ring assembly may include at least one purge air metering orifice on the inner band. At least one purge air metering orifice may be shaped to regulate volume of through-flow cooling air conducted through the through-flow vanes. 
     According to another aspect of the present disclosure, a vane ring assembly adapted for use in a gas turbine engine includes a plurality of in-flow vanes and a plurality of through-flow vanes. The in-flow vanes extend radially across a primary flow path annulus and are arranged around a central reference axis. The plurality of through-flow vanes may extend radially across the primary flow path annulus. The plurality of through-flow vanes may be interspersed among the plurality of in-flow vanes around the central reference axis. 
     In some embodiments, the plurality of in-flow vanes may have an in-flow transfer passage. The in-flow transfer passage may be configured to receive in-flow cooling air from a radially-outwardly opening in-flow aperture. The in-flow transfer passage may discharge the cooling air into the primary flow path annulus via cooling holes. The cooling holes may be formed at a trailing edge of the in-flow vanes to cool the in-flow vanes. 
     In some embodiments, the plurality of through-flow vanes may each have a through-flow transfer passage. The through-flow transfer passage may be configured to receive through-flow cooling air from a radially-outwardly opening through-flow aperture. The through-flow transfer passage may discharge at least some of the through-flow cooling air radially inwardly of the primary flow path annulus. This discharge may cool the through-flow vanes and other components of the gas turbine engine axially forward or aft of the vane ring assembly. The discharge may also prevent primary hot gas ingestion. 
     In some embodiments, the in-flow transfer passage of the in-flow vanes may have, on a surface contacting the in-flow cooling air, a plurality of turbulators. The turbulators may be configured to cause heat transfer by inducement of turbulent air flow over the surface contacting the in-flow cooling air. 
     In some embodiments, the through-flow transfer passage of the through-flow vanes may have, on a surface contacting the through-flow cooling air, a plurality of turbulators. The numerical number of total turbulators may be fewer than the numerical total of turbulators disposed on any surface in contact with the in-flow air within the in-flow transfer passage. The through-flow transfer passage of the through-flow vanes may be free of turbulators on all surfaces contacting the through-flow cooling air. 
     In some embodiments, the trailing edge of the through-flow vanes may be formed to include cooling holes. In some embodiments, the trailing edge of the through-flow vanes may be formed without the cooling holes. 
     In some embodiments, at least part of the trailing edges may be composed of a solid or hollow first material. A second material may form the remainder of the through-flow vane. The first material may be able to withstand, without degradation in structure or shape, temperatures in excess of those withstood by the second material. The first material of the trailing edges of the through-flow vanes may be formed of a ceramic matrix composite. 
     In some embodiments, the through-flow vanes may further include one or more purge air metering orifices. Each of the purge air metering orifices may be sized to control air that moves radially inwardly of the primary flow path annulus. 
     In some embodiments, the purge air may move in an upstream air feed passage, in fluid communication with components upstream of the vane ring assembly. In some embodiments, the purge air may move in a downstream air feed passage, in fluid communication with components downstream of the vane ring assembly. 
     In some embodiments, the in-flow vane and through-flow vane may be integrated into a one-piece monolithic component. 
     These and other features of the present disclosure will become more apparent from the following description of the illustrative embodiments. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a cut-away perspective view of a gas turbine engine showing that a typical engine includes a fan driven by an engine core having a compressor, a combustor, and a turbine and suggesting the turbine includes a plurality of rotationally fixed vane ring assemblies which are partially illustrated in  FIG.  2   ; 
         FIG.  2    is a cutaway perspective view of a vane ring segment adapted for inclusion in the turbine section indicated in  FIG.  1    showing that the vane ring segment includes two distinct vane types: a through-flow vane, which allows through-flow cooling air to flow radially through it into an air feed passage, and an in-flow vane, which transmits no flow to the air feed passage, but which ejects in-flow cooling air into a primary flow path annulus; 
         FIG.  3    is a cross-sectional view of the in-flow vane type depicted in  FIG.  2    having heat transfer augmenting features (e.g., impingement jets and pins, or vortex generators, dimples, turbulators, etc.) disposed on an inner surface of the in-flow transfer passage showing in-flow cooling air ejected only into the primary flow path annulus; 
         FIG.  4    is a cross-sectional view of the through-flow vane type depicted in  FIG.  2    showing both through-flow cooling air optionally ejected into the primary flow path annulus at the trailing edge of a through-flow vane, and through-flow cooling air discharged into the air feed passage radially-inward of the through-flow vane; 
         FIG.  4   a    is a cross-sectional view of the through-flow vane type with the through-flow transfer passage divided into a forward through-flow transfer passage and an aft through-flow transfer passage which allow tailoring the cooling augmentation in each passage to better match the external heat flux on the through-flow vane and/or manage the temperature increase of the through-flow air. 
         FIG.  5    is a diagrammatic view showing the through-flow air entering the through-flow aperture of the outer band and being transmitted via the through-flow transfer passage through a purge air metering orifice so that the through flow air cools the through-flow vane and is ejected into the air feed passage while in-flow air enters the in-flow aperture of the outer band and is directed for cooling of the in-flow vane before being ejected into the primary flow path annulus; 
         FIG.  6    is a cutaway perspective view of a second vane ring segment adapted for inclusion in the turbine section indicated in  FIG.  1    showing that the vane ring segment again includes two distinct vane types: a second embodiment of the through-flow vane, which allows through-flow cooling air to flow through it into the air feed passage without allowing through-flow air to enter the primary flow path annulus, and an in-flow vane, which transmits no flow to the air feed passage, but which ejects in-flow cooling air into the primary flow path annulus; 
         FIG.  7    is a cross-sectional view of a second embodiment of the in-flow vane type depicted in  FIG.  3    showing turbulators augmenting heat transfer of in-flow cooling air before ejecting in-flow cooling air only into the primary flow path annulus; and 
         FIG.  8    is a cross-sectional view of a second embodiment of the through-flow vane type shown in  FIG.  6   , having its trailing edge comprised of a ceramic matrix composite material, showing through-flow cooling air ducted into the through-flow transfer passage without entrance into the primary flow path annulus and final delivery to the feed air passage. 
     
    
    
     DETAILED DESCRIPTION 
     For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to a number of the embodiments illustrated in the drawings and specific language will be used to describe the same. 
     An illustrative gas turbine engine  100  includes a fan  102 , a compressor  104 , a combustor  106  and a turbine  108  as shown in  FIG.  1   . The turbine  108  receives hot expanding gases from the combustor  106  and extracts energy therefrom to drive the fan  102  in order to provide thrust to propel an aircraft. In other embodiments, the compressor  104 , the combustor  106 , and the turbine  108  provide an engine core  110  that can power a turbo shaft configured for aerospace power generation or other applications. 
     A vane ring assembly  10  providing part of the turbine  108  of the gas turbine engine  100  is shown, in part, in  FIG.  2   . The vane ring assembly  10  includes an in-flow vane  12  and a through-flow vane  14  as shown in  FIG.  2   . The in-flow vane  12  includes an in-flow airfoil  41 , an in-flow outer platform  56 , and an in-flow inner platform  64  as shown in  FIG.  2   . The through-flow vane  14  includes a through-flow airfoil  42 , a through-flow outer platform  57 , and a through-flow inner platform  65  as shown in  FIG.  2   . 
     The vane ring assembly  10  is illustratively made up of multiple segments  11  arranged adjacent to one another around a central axis  101  to form a ring as suggested in  FIG.  1    and  FIG.  2   . In the exemplary construction, three in-flow vanes  12  and one through-flow vane  14  are integrated into a single cast component or “quad,” as shown in  FIG.  2   . When the segments  11  are assembled, the full hoop of the vane ring assembly  10  includes an inner band  60 , an outer band  50 , and an airfoil set  40 . 
     In one embodiment, the inner band  60  includes a continuous structural material surface interspersed with purge air metering orifices  61  formed to align with, or be integral to, the through-flow inner platform  65  of the through-flow vane  14  as shown in  FIG.  4   . The inner band  60  may be formed to include the purge air metering orifices  61  at regular or symmetric intervals or the patterns may be diverse as long as each corresponds in axial and circumferential placement to the through-flow vanes  14  as shown in  FIG.  2   . The inner band  60  is formed to define at least part of a radially-inner boundary of a flow path annulus  46  by combining the continuous structural material surface with a series of in-flow inner platforms  64 , through-flow inner platforms  65 . The continuous structural material surface forming the radially-inner boundary of the flow path annulus may be formed of the same monolithic material via a casting or subtractive machining process. The inner band  60  also forms at least part of a radially-outer boundary to an upstream air feed passage  62  and a downstream air feed passage  63 , as shown in  FIG.  4    and  FIG.  5   . The upstream air feed passage  62  and the downstream inner feed passage  63  each deliver cooling air to respective components in their flow paths, as shown in  FIG.  5   . 
     The outer band  50  is formed to include in-flow apertures  51  and through-flow apertures  52  at regular or symmetric intervals corresponding to placement of in-flow vanes  12  and through-flow vanes  14 , respectively, as shown in  FIG.  2   . The outer band  50  is formed to define at least part of a radially-outer boundary of the primary flow path annulus by circumferentially combining a series of in-flow outer platforms  56  and through-flow outer platforms  57  each adjacent to either platform type, substantially continuous and optionally formed of the same monolithic material via casting or subtractive manufacturing. The outer band  50  may form at least part of a radially-outer boundary to compressor cooling air  55  in two optionally distinct paths, through-flow air  53  and in-flow air  54 , as shown in  FIG.  2    and  FIG.  5   . 
     The airfoil set  40  includes in-flow airfoils  41  and through-flow airfoils  42  fixedly attached to both the inner band  60  and the outer band  50  and extending radially there between, as shown in  FIG.  3    and  FIG.  4   , respectively. For purposes of this disclosure, the structure extending between the inner band  60  and outer band  50  is regarded as the airfoil belonging to each vane type. One embodiment of the in-flow airfoil type includes turbulators  43  on a surface of an in-flow transfer passage  45  contacting the in-flow air  54  as shown in  FIG.  3   . This embodiment of the in-flow vane  12  also utilizes an impingement tube  44  to diffuse in-flow air  54  throughout an in-flow transfer passage  45 , as shown in  FIG.  3   . 
     The in-flow transfer passage  45  transmits air from the in-flow aperture  51  to the primary flow path annulus  46 , as shown in  FIG.  3    and  FIG.  4   . In this embodiment, the in-flow airfoil  41  has a series of cooling holes  49  to inject in-flow air  54  into the primary flow path annulus  46  from any surface pores forming a film row on an exterior of the in-flow airfoil  41  or at the trailing edge  47  of the in-flow airfoil  41 , as shown in  FIG.  3   . This injection of in-flow air  54  cools the thin structure at the trailing edge  47  of the in-flow airfoil  41  and produces a film of cooling air to insulate the trailing edge  47  from combustion gases otherwise capable of melting the material of the in-flow airfoil  41 , as shown in  FIG.  3   . 
     The turbulators  43  constitute any surface deformation or structures inserted into flow paths, such as pin fins, fins, bumps, or the like and may be referred to interchangeably as cooling features. These features act to interrupt the boundary layer and promote mixing of the cooling flow. This flow structure enhances heat transfer and consequent cooling of the surfaces contacting the in-flow air. The turbulators  43  also function to increase the surface area available for heat transfer. 
     The through-flow airfoils  42  include a through-flow transfer passage  48 , as shown in  FIG.  4   . The through-flow transfer passage  48  transmits through-flow air  53  from the through-flow apertures  52  to both the upstream air feed passage  62  and the downstream air feed passage  63  through the purge air metering orifices  61 , as shown in  FIG.  4   . 
     The purge air metering orifices  61  are designed to ensure that the thermal expansion characteristics in the through-flow vanes  14  are suitable for the material properties thereof during every operational state of the gas turbine engine. In this embodiment, the through-flow airfoil  42  has a series of cooling holes  49  to inject only a portion of the through-flow air  53  into the primary flow path annulus  46  at the trailing edge  47  of the through-flow airfoil  42 , as shown in  FIG.  4   . This injection of in-flow air  54  cools the thin structure at the trailing edge  47  of the through-flow airfoil  42  and produces a film of cooling air to insulate the through-flow airfoil  42 , as shown in  FIG.  4   . 
     The in-flow air  54  and the through-flow air  53  may be similar in temperature, possibly bled from the same location on the compressor  104 . The volumetric majority of the through-flow air  53  continues through the through-flow transfer passage  48  into the upstream air feed passage  62  and the downstream air feed passage  63 , as shown in  FIG.  4   . In contrast, the entirety of the in-flow air  54  is injected into the primary flow path annulus  46 , as shown in  FIG.  3   . In this embodiment, turbulent flow is induced by virtue of the increased velocity of the through-flow air  53 . Accordingly, the surface of the through-flow transfer passage  48  which is in contact with the through-flow air  53  may be free of turbulators  43 , as illustrated in  FIG.  4   . 
     The through-flow transfer passage  48  may be divided into two chambers, a forward through-flow transfer passage  48 ′ and an aft through-flow transfer passage  48 ″ as illustrated in  FIG.  4   a   . The forward through-flow transfer passage  48 ′ may have a corresponding forward through-flow aperture  52 ′ and a forward purge air metering orifice  61 ′ while the aft through-flow transfer passage  48 ″ may have a corresponding aft through-flow aperture  52 ″ and a forward purge air metering orifice  61 ″, each passage with a certain type or density of cooling features and one with a different type or density of cooling features or without cooling features depending on the external heat flux into the through-flow vane  14 ′, as illustrated in  FIG.  4   a   . Optionally, heat transfer augmentation features  43 ′ may be included in either through-flow transfer passage  48 ′,  48 ″ to achieve the desired metal temperature of the through-flow airfoil  42 ′ allowing tailoring of the cooling geometry to achieve the desired vane metal temperature or manage the temperature increase of the through-flow air  53 . 
     The vane ring assembly  10  may be manufactured as a single monolithic structure by the process of investment casting, additive layer manufacturing, subtractive machining, welding, or any combination thereof. Additionally, it may be advantageous to manufacture sections, using the above methods, in either quint, quad, triple, double, or single vane units. Resulting units may be assembled to form a complete vane ring assembly  10  by arranging each manufactured unit of vane type with a frequency achieving the required cooling through-flow air  53  displacement. Sometimes, manufacturing pairs of in-flow vanes  12  and through-flow vanes  14  in separate casting or machining processes is cost effective and should be considered by one of skill—having updated information concerning specific manufacturing capability. Sometimes, manufacturing through-flow vanes  14  integrated with or between in-flow vanes  12  can help manage sealing and/or differences in thermal expansion within the vane ring assembly  10 . 
     The vane ring assembly  10  channels two gas paths. First, the vane ring assembly  10  channels the primary flow path annulus  46  toward turbine blades which are designed to extract energy from the combustion gases. This first gas path ultimately produces the thrust and power gas turbine engines are known for—driving the fan  102 , feeding the compressor  104 , and thus enabling simultaneous intake, compression, combustion and exhaust. Second, the vane ring assembly  10  internally conducts a radial through-flow air  53  within the through-flow transfer passage  48 , into an upstream air feed passage  62  and a downstream air feed passage  63 , as illustrated in  FIG.  4   . This through-flow air  53  is used for cooling turbine blades, other vane ring assemblies  10 , internal structures of the rotating components, and to prevent hot gas ingestion from the primary flow path. The through-flow air  53  enables the gas turbine to combust at higher temperatures and operate more efficiently. 
     Depending on the engine architecture, the through-flow air  53  may become the principal air used to cool rotor wheels and vane ring assemblies  10 , as in the present disclosure. A flow path according to this architecture is shown in  FIG.  1    and diagrammatically illustrated in  FIG.  5   . This approach typically results in the through-flow air  53  being divided between two distinct flow passages within a single vane. A standard vane ring assembly is generally constructed of a single vane type, each ducting through-flow air  53  and in-flow air  54  and sealing between the two flow paths. One means of achieving flow separation is by use of an impingement tube  44 , as shown in  FIG.  3   . An impingement tube  44  is often sealed by a weld to the outer band  50  and adjoining to the inner band  60  with a slip fit. As a consequence of dissimilar thermal growth between the impingement tube  44  and adjoining components, leaks occur, which compromise efficiency of air transfer and cooling power of the delivered air. 
     The current disclosure teaches a vane ring assembly  10  comprising vanes of two distinct types. The first type, and most numerous, are the in-flow vanes  12  which transfer in-flow air  54  through their trailing edges  47  into the primary flow path annulus  46  and cool the associated in-flow airfoil  42 , as shown in  FIG.  3   . A second type, numbering less than the in-flow vanes  12  per vane ring assembly  10 , are the through-flow vanes  14 . These are dedicated as tubes to deliver cooling through-flow air  53  to the upstream air feed passage  62  and downstream air feed passage  63 , a cavity located radially inside the inner band  60 , and shown in  FIG.  2   . 
     A dedicated through-flow vane  14  can allow the vane to have a high enough flow rate to sufficiently cool itself without the addition of turbulators  43  or at least to cool itself with a simpler cooling geometry than the in-flow vane  12 , simplifying manufacturing of each through-flow vane  14 . Since the entire through-flow vane  14  is dedicated to ducting through-flow air  53 , problematic sealing around impingement tubes is avoided, thus reducing the adverse effects of leakage compromising heat transfer. Without the need for an impingement tube, metering of the through-flow air  53  volume may be achieved by configuration of the through-flow aperture  51 . Further, without the complexities of an impingement tube, high-temperature materials used for the trailing edge  47 , though unable to be cast into the through-flow vane  14 , may be considered for secondary assembly and may eliminate the need for cooling holes  49 . 
     Another illustrative vane ring assembly  10  adapted for use in a gas turbine engine  100  is shown in  FIG.  6   . A vane ring assembly  210  is substantially similar to the vane ring assembly  10  shown in  FIGS.  1 - 5    and described herein. Accordingly, similar reference numbers in the  200  series indicate features that are common between the vane ring assembly  10  and the vane ring assembly  210 . The description of the vane ring assembly  10  is hereby incorporated by reference to apply to the vane ring assembly  210 , except in instances when it conflicts with the specific description and drawings of the vane ring assembly  210 . 
     Unlike the vane ring assembly  10 , the vane ring assembly  210  includes a trailing edge  247  which is composed of a ceramic matrix composite material. The trailing edge  247  is attached to mounting structures on a through-flow vane  214 , as shown in  FIG.  8   . In this illustrative embodiment the trailing edge  247  has no cooling holes  49 , such as those illustrated in the first embodiment of the through-flow vane  14 . Increased heat transfer from the additional flow provided by the through-flow air  53  coupled with a high-temperature capability material eliminated the need for cooling of the trailing edge  247 . 
     Unlike the through-flow vane  14 , the through-flow vane  214  as illustrated in  FIG.  8    suggests optional turbulators  243  disposed on the surface contacting through-flow air  53  to enhance heat transfer. It is suggested in  FIG.  8    that a lower density of turbulators (or no turbulators) are required in a through-flow transfer passage  248  than in the in-flow transfer passage  45  illustrated in  FIG.  3   . 
     Unlike the purge air metering orifice  61 , the purge air metering orifice  261  is comprised of a plurality of orifices, as shown in  FIG.  8   . It is also suggested that the air exiting the through-flow transfer passage  248  is divided by a flow path barrier  270  which causes one flow to move toward the upstream air feed passage  262  and the other confined to the downstream air feed passage  263 , as shown in  FIG.  8   . The purge air metering orifices  261  are designed to ensure the through-flow air  53  within the through-flow transfer passage  248  circulates to remove heat from surfaces with sufficient uniformity. 
     While the disclosure has been illustrated and described in detail in the foregoing drawings and description, the same is to be considered as exemplary and not restrictive in character, it being understood that only illustrative embodiments thereof have been shown and described and that all changes and modifications that come within the spirit of the disclosure are desired to be protected.