Abstract: A coolable airfoil is provided which includes an internal cavity, an external wall, a plurality of first cooling apertures, and a plurality of second cooling apertures. The external wall includes a suction side portion and a pressure side portion. The wall portions extend chordwise between a leading edge and a trailing edge and spanwise between an inner radial surface and an outer radial surface. The first cooling apertures are disposed in the external wall adjacent the trailing edge, exiting the airfoil through the pressure side portion. The second cooling apertures are disposed in the external wall adjacent the trailing edge, exiting the airfoil through the suction side portion.

Abstract: A hollow airfoil is provided having a leading edge, a trailing edge, and a wall including a suction side portion and a pressure side portion. The wall, which includes an interior surface and an exterior surface, surrounds a first cavity and a second cavity, separated from one another by a rib extending between the suction side and pressure side wall portions. The first cavity is contiguous with the leading edge. The airfoil further includes a coolant flow splitter attached to the wall interior surface within the first cavity, and at least one metering orifice disposed in the rib. The metering orifice(s) are substantially aligned with the coolant flow splitter, such that cooling air passing through the metering orifice(s) encounters the flow splitter. The flow splitter splits the cooling air flow and directs it along the wall interior surface.

Abstract: A turbine blade includes a hollow airfoil extending from an integral dovetail. The airfoil includes sidewalls extending between leading and trailing edges and longitudinally between a root and a tip. The sidewalls are spaced apart to define a flow channel for channeling cooling air through the airfoil. The tip is tapered longitudinally above at least one of the sidewalls and decreases in thickness.

Abstract: A mechanism for cooling the platform for the drive blades of a gas turbine uses a simple configuration which reliably cools the platform. The mechanism includes cooling channels in the interior of the platform which open out from one of the cooling air channels for cooling the turbine blades and which exit the platform through the edge nearest the tail. Cooling channels in the platform open out from the entrance to blade cooling channels, travel from the head of the blade along the blade sides, and exit through the edge near the tail of the blade. This structure diverts a portion of the cooling air entering the blade from the cooling channel in the base in order to cool the platform. Cooling air channels may extend from an enclosed air space below the platform to the upper surface of the platform at the front or rear side of the blade. Air channels may also extend on the rear of the turbine blade obliquely from the underside of the platform to the trailing edge of the platform.

Abstract: In a cooling system for the leading-edge region of a hollow gas-turbine blade, a duct (3), through which flow occurs longitudinally, extends from the blade root up to the blade tip and is defined in the region of the blade body (4) by the inner walls of the leading edge (5), the suction side (6) and the pressure side (7) and by a web (8). The inner walls of the suction side and the pressure side are provided with a plurality of ribs (9), which run slantwise and at least approximately in parallel. The suction-side ribs and the pressure-side ribs are offset from one another over the blade height. The ribs (9) run radially inward from the web (8) in the direction of the leading edge (5), merge into the radial in the region of the leading edge and are led around the leading edge. The deviation of the ribs (9) from the slant into the radial is effected with the smallest possible radius.

Abstract: A gas turbine engine airfoil includes first and second sidewalls joined together at opposite leading and trailing edges, and extending longitudinally from a root to a tip. The sidewalls are spaced apart from each other to define in part first and second adjoining flow chambers extending longitudinally therein, and defined in additional part by corresponding first and second partitions disposed between the sidewalls. The second partition is common to both chambers, and both partitions include respective pluralities of first and second inlet holes sized to meter cooling air therethrough in series between the chambers.

Abstract: In connection with a turbine rotor disk in whose disk grooves air-cooled turbine blades have been inserted, at least two cooling air channels, respectively extending from the same disk front face, terminate in each disk groove. The outlet openings of two cooling air channels in each disk groove preferably lie essentially next to each other in a common sectional plane, which is normal to the disk axis. Because of this it is possible to supply a larger cooling air flow without drastically increasing the weakening, or respectively the stress of the disk in the groove bottom.

Abstract: The invention relates to a cooled shroud in a gas turbine stationary blade which is able to flow a cooling air in the entire area of an inner shroud for cooling thereof. Three stationary blades are fixed to the inner shroud 2, a cover 13, 14 is provided to form a space 21 and space 22a, 22b and 22c, respectively. The cooling air is introduced through an independent air passage 3A of a leading edge of each stationary blade into the spaces 22a, 22b and 22c and is flown therefrom through a tunnel 18 and air reservoirs 19-2, 19-3 and 19-4 to be blown out of a trailing edge while cooling surfaces of the shrouds and the trailing edges. Also, a portion of the cooling air from the space 22b is flown into the space 21 through a tunnel 11, a leading edge side passage 12 and an endmost tunnel 11 and is then blown out of the trailing edge through a tunnel 18 and an air reservoir 19-1, so that the leading edge portion, the endmost portion and the endmost trailing edge portion are cooled.

Abstract: A turbine vane-system cooling system uses three internal cooling cavities 1, 12, 13) separated by two radial walls (9, 10). The upstream cavity (11) uses a helical ramp (30) and is fed through an intake (22) at the vane root (3). The middle cavity (12) also is fed at the vane root (3) and includes a compartmented, multi-perforated lining (40). The air is exhausted from each compartment through impact orifices and enters the succeeding compartment through slots (42) and then is finally exhausted through a vane-head orifice (21). The vane side walls opposite the downstream cavity (13) have double skins with bridging elements. The air passes through these double skins but circulates centrifugally in the upstream portion (15) of the downstream cavity (13) and enters this cavity's downstream portion (16) to be exhausted through slots (19) in the trailing edge (6). A third wall (14) divides the downstream cavity (13) into two parts (15, 16).

Abstract: A gas turbine stationary blade having a cooling medium passage and outer and inner shrouds in a blade section, comprising a hollow portion in the blade section, the hollow portion including a plurality of chambers divided in the chord direction, the cooling medium passage being formed by connecting the divided chambers in series, one of the chambers located in the center with respect to the chord direction being formed as a sealing air passage in order to allow a cooling medium to flow from the outside of the outer shroud to the bore of the inner shroud.

Abstract: 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).

Abstract: Disclosed is an integral single-cast multi wall structure including a very thin wall and a second thin wall. There is a passageway interposed between the pair of walls of the structure, and a high thermal conductivity member extends into said passageways and thermally couples the walls. The high thermal conductivity member increases the heat transfer between the walls of the structure. The present invention further includes a method for casting an integral structure having very thin walls that utilizes the high thermally conductive member in the casting process to hold the pattern and cores in alignment.

Abstract: An internally air cooled turbine blade for a gas turbine engine of the type including a trailing edge section, leading edge section and mid chord section wherein each section includes a straight through radial passage communicating cooling air from the root to the tip of the blade, and the radial passages (feed channels) on the pressure side and suction side supply the cooling air to the film cooling holes in the airfoil surface and the radial passage (feed chamber) in the mid chord section replenishes the feed channels with cooling air through replenishment cooling holes interconnecting the feed channels and feed chamber. The rotation of the blade imparts a centrifugal pumping action to the air within the feed chamber for maximizing the cooling effectiveness of the cooling air. The air discharging at the tip cools the tip of the blade and provides tip aerodynamic sealing and the leading edge and trailing edge are similarly fed cooling air.

Abstract: A gas turbine engine turbine blade and method of manufacture includes an airfoil having pressure and suction sides joined together at leading and trailing edges. The airfoil further includes a supply channel for receiving compressed air, and has an elongate recess disposed in at least one of the pressure and suction sides which is separated from the supply channel by a partition. The partition includes a plurality of spaced apart flow metering holes disposed in flow communication with the supply channel. A transpiration strip is fixedly joined to the airfoil in the recess and spaced from the partition to define a plenum for receiving the compressed air from the metering holes. The transpiration strip is pervious for channeling the compressed air from the plenum in a blanket of film cooling boundary layer air therefrom.

Abstract: A bulkhead in the form of a rib extending spanwise adjacent the leading edge of the airfoil of a turbine blade of a gas turbine engine serves to stiffen the airfoil to attain protection from impact by foreign objects, but doesn't materially affect cooling effectiveness in the event of a puncture of the skin. Holes extending spanwise in the rib reduce the mass, allow radial flow, and restrict flow into the leading edge. This construct allows a reduction in the thickness of the skin of the airfoil, and hence a weight reduction.

Abstract: A gas turbine component that includes a very thin wall. The gas turbine component being of a single-piece, single-cast structure. In one form of the gas turbine component it is formed of a superalloy composition and is capable of withstanding gases impinging upon the walls at temperatures greater than 4000.degree. F.

Abstract: A stator vane for a gas turbine engine wherein the airfoil section includes a plurality of pockets with an impingement hole at the upstream end of a passageway formed in the pocket and a slot formed on the opposite end of the pocket for discharging a film of cooling air along the outer surface of the air foil. The impingement hole and slot being arranged so that the flow of air in the passageway is in counterflow indirect heat exchange relationship with the gas path, thus placing the hottest part of the metal forming the pocket with the coolest air in the pocket.

Abstract: A turbine airfoil has a mesh cooling hole arrangement which includes first and second pluralities of cooling holes extending between internal and external surfaces of an airfoil side wall at least at a pressure side and extending from an internal chamber to the airfoil exterior. The cooling holes of each plurality extend generally parallel to one another. The cooling holes of the first and second pluralities intersect so as to define a plurality of spaced apart internal solid nodes in the side wall having pairs of opposite sides interconnected by pairs of opposite corners. The spaced nodes define a multiplicity of hole portions of the cooling holes which extend between and along opposite sides of adjacent nodes and a plurality of flow intersections which interconnect the hole portions of the cooling holes and are disposed between the corners of adjacent nodes.

Abstract: A turbine airfoil having a cut-back trailing edge and a plurality of diffusing flow dividers upstream of the cut-back trailing edge is disclosed. Various construction details are developed which provide ejection of a diffusing film of cooling fluid over a cut-back trailing edge. In one particular embodiment, a turbine airfoil includes a plurality of radially spaced flow dividers extending between a pressure wall and a suction wall. Each flow divider includes a rounded leading edge, a pair of parallel sidewalls downstream of the leading edge, and a pair of converging sidewalls downstream of the parallel sidewalls. Adjacent sidewalls of adjacent flow dividers form flow channels having a constant area channel and a diffusing section. The diffusing section includes a covered portion upstream of the cut-back trailing edge and an uncovered portion extending over the cut-back trailing edge.

Abstract: A turbo-machine vane having a plurality of internal cavities for the flow of a cooling fluid, and a plurality of openings through the outer wall of the vane for communicating the internal cavities with the outside of the vane. Two rows of openings are provided in the leading edge of the vane, one row on each side of the central line of the leading edge end extending parallel with the central line, and each opening of each row is oriented to direct cooling fluid which flows through it from the interior of the vane away from the central line relative to the rows.

Abstract: A gas turbine engine blade includes an airfoil having first and second sides and a dovetail extending from the airfoil root. The airfoil includes a tip having a tip floor with first and second tip walls extending from the floor and spaced apart to define therebetween a tip pienum. The first tip wall is recessed at least in part from the airfoil first side to define an outwardly facing tip shelf, with the tip shelf and the first tip wall defining therebetween a trough. A plurality of cooling holes extend through the tip floor at the tip shelf for channeling cooling air from a flow channel inside the airfoil into the trough for cooling the blade tip.

Abstract: A perforated sheet for the promotion of film cooling a gas turbine engine comprises two separate layers of laminate material which are placed in superposed abutting relationship. The first layer has a plurality of apertures of given cross-sectional area therein. The second layer also has a plurality of apertures therein which have larger cross-sectional areas than the corresponding apertures in the first layer. The apertures in the first layer are in fluid communication with the apertures in the second layer to form passageways through the perforated sheet. These passageways permit a cooling flow of fluid to pass through the perforated sheet so it can be discharged as a cooling fluid film along the inner surface of the perforated sheet. The apertures in each layer are independently formed to produce the passageways. The apertures can easily be machined to the desired size, shape, and angle to optimize the cooling film effectiveness for any particular application.

Abstract: A blade device for turbo-engines, comprising a blade shell, a blade core and a blade base, has a blade core comprising a bundle of small tubes which provides the blade device with high stability and stiffness. The arrangement has the advantage that the blade device can be used universally in the case of almost all blade devices, and the blade core permits a heat exchanger function.

Abstract: An improved design for a structure such as a turbine vane or blade subjected to extreme temperature levels and/or gradients resulting in extreme deformation stresses within the structure. The present structures comprise an interior skeletal support having a plurality of spaced, flexible support legs which extend out and are attached to spaced areas of a flexible outer surface skin which is subjected to extreme temperature levels or temperature gradients during use. The extreme stresses normally developed by the expansion of the surface skin are avoided by the ability of the skin to move minutely and flex the spaced support legs to variable extents and in predetermined directions.

Abstract: A nozzle guide vane assembly for a gas turbine engine comprises inner and outer porous sheet metal platform rings and a plurality of aerofoils extending between the rings, each of the aerofoil portions having mounting means at its inner and outer extremities by which the aerofoils, and thus the rings, are supported from fixed engine structure.

Abstract: Known porous laminates of the kind wherein ambient atmosphere each side of the laminate is connected via holes and passageways, the latter lying within the laminate thickness in planes parallel with the faying faces, suffer from airflow energy loss which arises from the changes in direction undergone by the airflow while in transit from one side of the laminate to the other side thereof. The invention provides extra holes on the high pressure side of the laminate, over those points in the passages wherein the greatest change in direction of flow occurs, so as to re-energize the airflow which has reached those points via the relevant holes of the known arrangement.

Abstract: A turbine blade spar wall surrounds a coolant plenum of the blade and is cast with an intermediate thickness exceeding its final or design thickness. The outer surface of the spar wall is machined to form a plurality of incomplete holes extending to bottoms located at the final thickness dimension of the spar wall. A pre-formed sheath of porous metal is fit closely around the spar wall outer surface with coolant pores in the sheath communicating with the incomplete holes. A high pressure inert gas is introduced into the coolant plenum concurrently with application of compressive forces to the sheath as both the sheath and the spar wall are raised to a high temperature whereby the sheath is diffusion bonded to the spar wall. The inert gas pressure in the coolant plenum reinforces the spar wall against the compressive forces.

Abstract: A cooled vane or blade for a gas turbine engine has a cooling arrangement for its trailing region which can be accommodated in the relatively thin section available. In this arrangement the trailing region of the hollow interior of the blade is divided off from the remainder by a partition which may be apertured to allow cooling air to enter the compartment thus formed. The concave, pressure flank of the compartment is cooled by arrays of film cooling holes while the convex, suction flank has a perforated plate spaced therefrom to provide impingement cooling. The suction flank is therefore unapertured and there is no disturbance of the high speed airflow in this region. The spent impingement air leaves the aerofoil via a slot and may pass over pedestals en route to cool the entire trailing edge.

Abstract: A spacer extends between and engages the rear face of a first disk and front face of a second disk of co-rotating rotors of a two stage turbine. The spacer engages the disks radially inwardly of the blade root slots of each disk and defines an intermediate cooling air compartment between the disks, radially inward of the spacer. Each disk includes blade root slots, each having a blade root disposed therein and defining a cooling air passageway across each slot from the front to rear face of each disk. Cooling air from a compartment upstream of the first disk is directed downstream through the passageways across the slots and thence radially inwardly into the intermediate compartment between the disks. From the intermediate compartment the cooling air travels radially outwardly and then axially through the cooling air passageways across the second disk blade root slots.

Abstract: In a hollow airfoil, thin metal baffle sheets are bonded to opposing suction and pressure sides of a longitudinally extending cooling air cavity. The longitudinally extending downstream edges of these opposing baffle sheets are closely spaced apart defining a cooling air outlet slot for the cavity which may, for example, feed cooling air to the trailing edge portion of the airfoil. If the rearward edges of either baffle sheet become unbonded from the inside cavity wall the edge of the baffle sheet might bend inwardly and close the outlet slot. To prevent this possibility a plurality of pedestals extend outwardly from each side of the cavity wall opposite the rearward edge of each baffle sheet and extending substantially to the inwardly facing surface of that baffle sheet trapping it against the wall.

Abstract: It is an object of the invention to provide a film cooling apparatus of increased effectiveness and efficiency. In accordance with the invention, a cooling fluid is injected into a hot flowing gas through a passageway in a wall which contains and is subject to the hot gas. The passageway is slanted in a downstream direction at an acute angle to the wall. A cusp shape is provided in the passageway to generate vortices in the injected cooling fluid thereby reducing the energy extracted from the hot gas for that purpose. The cusp shape increases both film cooling effectiveness and wall area coverage. The cusp may be at either the downstream or upstream side of the passageway, the former substantially eliminating flow separation of the cooling fluid from the wall immediately downstream of the passageway.

Abstract: An improved high pressure turbine rotor blade and tip cap structure therefor is provided, which comprises, a plurality of metallic layers bonded to the tip end of the blade, each layer having a peipheral shape conforming to the camber of the blade, the layers defining a plurality of radially outwardly opening, serpentine-shaped passageways for passage of coolant fluid through the periphery of the tip end of the blade.

Abstract: A high-temperature, extremely high-speed ceramet composite blade comprising a ceramic blade jacket seated from above on a metallic core having a crossweb near its upper end for securing the jacket on the core by one or more metallic retaining pins extending through the crossweb and secured to the core. A head on each bolt is located within a chamber formed above the cross-web and surrounded by the radially projecting end of the blade jacket. A porous abradable coating is secured to the head of the bolt and substantially fills the chamber. The coating projects slightly beyond the end of the blade jacket and is supplied and permeated with cooling air via a cooling duct extending through the retaining pin and metallic core.

Abstract: A transpiration cooled ceramic blade for a gas turbine is shown wherein a spar or strut member defining a root portion and an airfoil portion provides the main structural component of the blade. The air foil portion contains longitudinal grooves in the surface in flow communication with an air flow passage in the root portion and a flexible perforated ceramic tape is wrapped around the air foil portion with the perforations therein in registry with the grooves in the core. The flexible ceramic tape and the strut assembly are heated initially to a low temperature to drive off the binder forming the tape and then heated to a relatively high temperature to fuse the ceramic component of the tape together and to the strut to form a unitary blade structure with internal air flow paths and transpiration cooling orifices through the skin.

Abstract: Improved structure and method for an internally cooled, laminated stator or turbine blade for turbomachinery includes internal cooling passage configurations within each lamina which promote different forms of cooling of the internal or external surfaces of the blade.

Abstract: A transpiration cooled blade for a gas turbine engine is assembled from a plurality of individual airfoil-shaped hollow ceramic washers stacked upon a ceramic platform which in turn is seated on a metal root portion. The airfoil portion so formed is enclosed by a metal cap covering the outermost washer. A metal tie tube is welded to the cap and extends radially inwardly through the hollow airfoil portion and through aligned apertures in the platform and root portion to terminate in a threaded end disposed in a cavity within the root portion housing a tension nut for engagement thereby. The tie tube is hollow and provides flow communication for a coolant fluid directed through the root portion and into the hollow airfoil through apertures in the tube. The ceramic washers are made porous to the coolant fluid to cool the blade via transpiration cooling.

Abstract: A transpiration cooled ceramic blade for a gas turbine is shown wherein a spar or strut member defining a root portion and an airfoil portion provides the main structural component of the blade. The air foil portion contains longitudinal grooves in the surface in flow communication with an air flow passage in the root portion and a flexible perforated ceramic tape is wrapped around the air foil portion with the perforations therein in registry with the grooves in the core. The flexible ceramic tape and the strut assembly are heated initially to a low temperature to drive off the binder forming the tape and then heated to a relatively high temperature to fuse the ceramic component of the tape together and to the strut to form a unitary blade structure with internal air flow paths and transpiratin cooling orifices through the skin.

Abstract: A gas turbine blade constituted of a supportive metallic blade core and a thin-walled ceramic blade airfoil, in which the airfoil is supported against a tip plate of the blade core. The blade core consists of rod or wire-shaped pins which have widened bases at their radially inner ends. Through these widened bases, the pins are retained in a metallic adapter slidable into a turbine disc.

Abstract: Improved structure and method for an internally cooled, laminated stator or turbine blade for turbomachinery includes internal cooling passage configurations within each lamina which promote different forms of cooling of the internal or external surfaces of the blade.

Abstract: A transpiration air cooled combustor assembly for a gas turbine engine includes an annular liner of laminated metal with an inner sheet and an outer sheet having a plurality of mechanically formed holes therein on either side of a mechanically pressed waffle patterned core sheet with offset depressings and dimples on either face thereof; the dimples have raised lands bonded to the inner and outer sheets; small cross passages are drilled in the core sheet so that the margins of the cross passages are located in spaced relationship to the land surfaces thereby to prevent burr formation disruption of the bond joints; the core sheet has a total metal mass equivalent to the orginal metal mass prior to press displacement of metal to form the depressions and dimples therein except for the metal removed by formation of the cross passages which communicate offset depressions on opposite sides of the core sheet to form a tortuous intercommunicating flow path through said annular liner between holes in the inner and out

Abstract: Improved structure and method for an internally cooled, laminated stator or turbine blade for turbomachinery includes internal cooling passage configurations within each lamina which promote different forms of cooling of the internal or external surfaces of the blade.

Abstract: A blade is disclosed having an upper airfoil configuration and a lower platform with a root constructed with conventional fir-tree design. The forward part of the blade is formed of radial wafers wherein desired configurations of cooling passageways can be formed on mating surfaces of the wafers so that bonding of these wafers together forms intricate internal passageways. Film cooling is shown on the pressure side of the blade, while convection cooling and some film cooling is shown on the suction side. The radial wafers of the forward part are open at the center thereof to provide a cavity in a finished blade for a cooling fluid to flow from an opening in the root thereof to the internal passageways. The cavity extends to the top of the blade and is covered by a tip cap. The cap extends over the side wafers of the rearward part. The rearward part of the blade is formed of side wafers with one forming the suction side of the blade, while the other forms the pressure side of the blade.

Abstract: Methods are disclosed for more practical construction of liquid-cooled buckets able to efficiently transport heat energy from the inside of the airfoil skin surface in contact with hot gas to the outer surface of preformed tubes recessed into the bucket core, through which tubes liquid coolant is passed during operation. The bucket is made of a series of preformed solid components, which are assembled, consolidated and then converted into a unified structure. In each arrangement illustrated one of the preformed solid components is a flat bimetallic sheet comprising an erosion, corrosion resistant layer and a layer of high thermal conductivity, these layers being joined by an optimized metallurgical bond.

Abstract: A material suitable for making combustion chambers for gas turbine engines comprises at least two abutting sheets of perforated material, the perforation being out of alignment and interconnected by a series of channels formed on one or both of the abutting surfaces of abutting sheets. The total cross-sectional area of the perforations in at least one sheet is at least double the total cross-sectional area of the perforations in the remaining sheets or sheets per unit area.

Abstract: A segment of a turbomachine nozzle is tangentially held in position by a pair of lugs extending radially from the outer and inner bands, respectively, at opposite ends of the segment, and resting against the adjacent support structure. In this way, the stress in the vanes and bands is maintained at a low level to thereby allow the use of high temperature, relatively low strength materials and reduced cooling measures. The lugs form integral end caps on the vane's one end and the cooling air enters the opposite end thereof.

Abstract: A coolable wall element which is adapted for combined impingement and transpiration cooling in environments where the pressure differential across the wall element differs substantially with physical position along the wall is disclosed. Techniques varying the proportion of impingement to transpiration cooling along the wall are developed. The wall element is shown in one embodiment as forming a portion of the wall of an air-foil adapted for use in the turbine section of a gas turbine engine.

Abstract: An improved hollow vane for a gas turbine is cooled by a flow of cooling air through the interior of the vane. Flow guiding plates inserted in the interior of the vane direct coolant flow first to the leading edge of the vane for maximum cooling at that location and then to discharge openings in the trailing edge of the vane.

Abstract: A thermally highly stressed, cooled component, more particularly, a blade for turbine engines, and a method for manufacturing the blade. A component of the above-mentioned type in which a central supporting core formed of a solid material has connected thereto several short, radially outwardly projecting ridges which, in turn, carry an outer shroud concentrically encompassing the supporting core, and which is made of a through-porous material, whereby the supporting core, ribs, and outer shroud are cast in a single piece.

Abstract: An airfoil cooling system for use in a gas turbine engine having high turbine inlet temperatures is disclosed. Various construction details designed to prevent thermal deterioration are developed. The system is built around impingement, film, and convective cooling techniques which are combined to limit the temperature of the airfoil material and to reduce thermal gradients within the component.

Abstract: A turbine blade includes a longitudinally extending cavity into which air for cooling the interior wall of the blade is admitted. The air passes from the cavity through each of a group of inserts each of which is constituted by a plurality of stepped overlapping walls with spaced air passages therethrough, these walls establishing therebetween corresponding cooling spaces interconnected by the passages and which are bounded at one side by the blade wall. The cooling air passes in succession through the cooling spaces of each insert and is discharged from the insert and thence from the blade through an outlet port formed in the blade wall and which is connected to the last cooling space of each insert through which the air is passed. The passages through the overlapping walls also provide impingement cooling of the blade wall. The outlet ports for the cooling air are formed in the suction side of the blade as well as in its trailing edge.