Abstract: In an air-cooled turbine blade (10) which has a shroud-band element (11) at the blade tip, the shroud-band element (11) extending transversely to the blade longitudinal axis, hollow spaces (16, 16′, 17, 17′) for cooling being provided in the interior of the shroud-band element (11), which hollow spaces (16, 16′, 17, 17′) are connected on the inlet side to at least one cooling-air passage (18) passing through the turbine blade (10) to the blade tip and open on the outlet side into the exterior space surrounding the turbine blade (10), the hollow spaces (16, 16′, 17, 17′) and the shroud-band element (11) are matched to one another in shape and dimensions in order to reduce the weight of the shroud-band element (11).

Abstract: A turbine airfoil includes pressure and suction sidewalls joined together at leading and trailing edges, and extending from root to tip. The tip includes a tip rib extending from a tip cap enclosing an internal flow channel between the sidewalls. An aligned pair of first and second holes extend through the tip cap and rib, respectively, for discharging coolant from the flow channel.

Abstract: A turbine airfoil includes pressure and suction sidewalls extending between leading and trailing edges and from root to tip. The tip includes a floor bounding an internal cooling channel within the airfoil which channels cooling air. The airfoil tip includes a first rib adjacent the pressure sidewall, a second rib spaced therefrom to define a first slot, and a third rib adjacent the suction sidewall to define a second slot with the second rib. The tip floor includes feed holes extending between the cooling channel and the first slot for supplying cooling air therein for discharge over the second rib towards the third rib.

Abstract: A gas turbine engine rotor blade 18 includes a dovetail 22 and integral airfoil 24. The airfoil includes a pair of sidewalls 28,30 extending between leading and trailing edges 32,34, and longitudinally between a root 36 and tip 38. The sidewalls are spaced laterally apart to define a flow channel 40 for channeling cooling air through the airfoil. The tip includes a floor 48 atop the flow channel, and a pair of ribs 50,52 laterally offset from respective sidewalls. The ribs are longitudinally tapered for increasing cooling conduction thereof.

Abstract: An airfoil having a trailing edge cavity formed by a leading wall and a trailing edge connected by a pair of side walls which converge at the trailing edge to define a cooling passage of substantially triangular cross section; a plurality of guide vanes arranged within the passage, spaced from the leading wall and trailing edge, and configured so that cooling gas flow introduced a generally radial direction is forced to flow in a direction toward the trailing edge.

Abstract: A turbine blade includes a hollow airfoil having a squealer tip rib extending outboard from a tip cap enclosing the airfoil. An impingement baffle is spaced inboard from the tip cap to define a tip plenum therebetween. Impingement holes extend through the baffle and are directed at the junction of the rib and cap for impinging a coolant jet thereagainst.

Abstract: A tip shroud (11) for a moving blade to be used at a downstream stage of a gas turbine. The tip shroud is extended in its creep lifetime by feeding it effectively with cooling air. In the tip shroud (11), there are cut-off portions (12, 13) which are passed by a hot combustion gas so that they are influenced by thermal stress. On the shroud faces, there are formed grooves (31, 32; 33, 34), and cooling guide covers (21, 22) are mounted in the grooves. The cooling guide covers (21, 22) are closed at one side with members (21a, 22a) and open at an end thereof. The covers function to cover the air holes (17) so that the cooling air is fed to the cut-off portions (12, 13) individually. The metal temperature rise at the cut-off portions (12, 13), as will be influenced by the hot combustion gas, can be suppressed so as to extend the creep lifetime of the tip shroud.

Abstract: A turbine airfoil includes a squealer rib extending outwardly from a tip cap to define a tip cavity thereatop. A thermal insulator is disposed in the tip cavity atop the tip cap.

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 method for preparing a Ni base superalloy inner wall surface of a body open end, such as an end of a turbomachinery blade, and an end plate, such as a blade tip cap, for brazing together at a rim of the end plate includes electrochemically removing oxides from the inner wall surface. The end plate is prepared, at least at its rim, by first removing surface and subsurface oxides, for example by mechanical abrading or a combination of such abrading and acid cleaning. Then at least the rim is electroplated with Ni which is heated to diffuse the Ni into the rim substrate. This provides an improved combination of surfaces for brazing, for example with a Ni base brazing alloy. After brazing the rim to the inner wall, there is provided an article with an improved relatively low oxide brazed joint, including less than about 20 volume % oxides, and a plate rim of substantially Ni along with elements diffused from the brazing alloy and the rim.

Abstract: A turbine assembly comprises a plurality of rotating blade portions in a spaced relation with a stationery shroud. The rotating blade portions comprise a root section, a tip portion and an airfoil. The tip portion has a pressure side wall and a suction side wall. A number of flow discouragers are disposed on the blade tip portion. In one embodiment, the flow discouragers extend circumferentially from the pressure side wall to the suction side wall so as to be aligned generally parallel to the direction of rotation. In an alternative embodiment, the flow discouragers extend circumferentially from the pressure side wall to the suction side wall so as to be aligned at an angle in the range between about 0.degree. to about 60.degree. with respect to a reference axis aligned generally parallel to the direction of rotation.

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 first-stage turbine blade includes an airfoil having a profile according to Table I. The airfoil has a plurality of cooling air passages extending linearly from the root portion to the tip portion of the airfoil. The blade includes a shank having a pair of cavities in communication through the blade dovetail with a plenum in the wheel space for supplying cooling air to the passages in the airfoil. The cooling passages in the airfoil terminate in a recess at the tip portion which has an opening adjacent the trailing edge of the airfoil and along the suction side to enable egress of cooling air into the hot gas stream on the low pressure side of the airfoil. The majority of the cooling passages are turbulated. Certain of those passages are arranged in rows lying adjacent to the pressure and suction sides of the airfoil.

Abstract: A coolable blade (10) essentially comprises a blade root (11) and a blade body (1), which is composed of a pressure-side wall (6) and a suction-side wall (5). They are connected to one another essentially via a trailing-edge region (4) and a leading-edge region (3) in such a way that at least one hollow space (2) used as a cooling-fluid passage is formed. An essentially radially running, diverging cooling passage (7) is arranged in the trailing-edge region (4).

Abstract: A gas turbine engine rotor blade includes an airfoil having a concave side wall and a convex side wall joined together at leading and trailing edges. Concave and convex tip walls extend from adjacent the leading edge along the respective concave and convex side walls to adjacent the trailing edge and are spaced apart to define a tip cavity therebetween. A hole or channel is disposed in a trailing edge tip region connecting the tip cavity to the trailing edge for channeling cooling fluid through the trailing edge tip region.

Abstract: A rotor blade includes a dovetail and an airfoil joined thereto. The airfoil includes first and second spaced apart sides joined together laterally at opposite leading and trailing edges, and spanwise at a root and opposite tip. A serpentine cooling circuit extends inside the airfoil for channeling air therethrough for cooling the blade. The serpentine circuit includes first and second passes and a first bend therebetween for firstly receiving the cooling air in turn from the dovetail. A tip circuit is disposed between the tip and the serpentine circuit at the first bend for separating the tip from the first bend and providing cooling thereof near the trailing edge.

Abstract: A turbine blade including an airfoil section having a double-wall construction for side-wall impingement cooling on the pressure side and a multi-pass serpentine along the suction side of the blade, is described. More particularly, and in one embodiment, the airfoil section includes a pressure side wall and a suction side wall which are joined together at a leading edge and a trailing edge. The blade also includes a leading edge, or tip, and a trailing edge, or tail. The airfoil section also includes a leading edge cavity having a plurality of radial film air holes, and an inner cavity which is a three pass serpentine. As cooling air flows along the passageways, it convectively cools the portions of the turbine blade adjacent these passageways. The airfoil section further includes a trailing edge cavity to cool the trailing edge flow region of the airfoil section.

Abstract: A pressurized porous surface near the leading edge of a rotorcraft blade, is designed to be used as an active control device which alleviates the aerodynamics of blade vortex interactions (BVI) and thus the impulsive BVI noise levels and signature. The pressurized porous surface can be actuated on an azimuth-dependent deployment schedule or actuated continuously. The pressurized porous surface is supplied with either positively pressurized air, negatively pressurized air, or a combination of both, when actuated. The pressurized porous surface targets the local blade aerodynamics, rather than the vortex strength or blade/vortex separation distance.

Abstract: A turbine blade includes an airfoil having an internal cooling circuit therein. The airfoil extends from a root to a tip cap, and includes laterally opposite pressure and suction sides extending between a leading edge and an opposite trailing edge over which is flowable a combustion gas. A pair of squealer tips extend radially upwardly from the tip cap along the pressure and suction sides, and are spaced apart between the leading and trailing edges to define an upwardly open tip cavity. At least one of the squealer tips includes a slot extending radially inwardly to the tip cap, with the slot also extending along the squealer tip between the leading and trailing edges. A plurality of spaced apart supply holes extend radially through the tip cap in the slot in flow communication with the cooling circuit for channeling the coolant into the slot for cooling the squealer tip.

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 method and device for collecting energy from wind includes an airfoil having an outer surface defining a leading edge, a central section of greatest breadth adjacent to which a Bernoulli effect reduced pressure region results as an air flow passes over the airfoil from the leading edge to a trailing edge, an air passage contained within the outer surface, and a plurality of edge nozzles penetrating the outer surface within the section of greatest breadth. The edge nozzles each have an outlet orifice oriented toward the trailing edge and an inlet orifice feeding into the air passage and being positioned to communicate between the air passage and the Bernoulli effect reduced pressure region. The airfoil is positioned in the wind with the leading edge facing substantially into the wind and with the wind passing over the edge nozzles. An airflow-driven turbine capable of converting an airflow into rotational mechanical energy is also provided, the turbine being in airflow communication with the air passage.

Abstract: An internally cooled turbine blade for a gas turbine engine is modified at the leading and trailing edges to include a dynamic cool air flowing radial passageway with an inlet at the root and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length.

Abstract: In a gas turbine hollow cooled rotor blade, a plurality of cooling holes 3 and 5 are provided which communicate between a cooling air passage in the blade and the tip squealer portion 1 on the pressure side and between the cooling air passage in the blade and a position in the vicinity of the suction side of a tip cap 4.

Abstract: A self bleeding rotor blade and method of operation are disclosed for reducing boundary layer thickness for improved performance. The rotor blade includes an outer surface configured for pressurizing air flowable thereover with bleed apertures being disposed therein for bleeding a portion of boundary layer air from the outer surface during operation and thereby decreasing its thickness for improving aerodynamic performance of the blade.

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 jet engine component includes a body having a wall portion with an external surface exposed to hot gas flow and an internal surface exposed to a cooling air flow. The engine component incorporates an arrangement of cooling holes defined through the wall portion between the external and internal surfaces thereof to permit flow of cooling air from the hollow interior through the wall portion to the exterior of the component. Each cooling hole includes at least one flow inlet at the internal surface of the wall for receiving the cooling air flow, at least a pair of flow outlets at the external surface of the wall for discharging the cooling air flow, and at least a pair of flow branches extending through the wall portion and between the flow inlet and the flow outlets for permitting passage of the cooling air flow from the flow inlet to the flow outlets. In one V-shaped configuration, the flow branches merge and intersect with one another at the flow inlet.

Abstract: A device for preparing liquid to thin pulpy media having an inner chamber 1 is substantially symmetrical about longitudinal axis, and the inner chamber has a neck end and a passage end. An outer circumference of the inner chamber tapers down from the passage and to the neck end. An outer chamber surrounds the inner chamber, and the outer chamber is also substantially symmetrical about the longitudinal axis. The outer chamber communicates with the inner chamber through an opening in the neck of the inner chamber, and an opening in the passage end of the inner chamber. A centrifuge wheel is positioned inside the outer chamber and outside the inner chamber. The centrifuge wheel has an intake pipe extending from a center of the centrifuge wheel, and this intake pipe extends into the neck opening of the inner chamber. The centrifuge wheel and the intake pipe are substantially symmetrical with, and rotate about the longitudinal axis.

Abstract: A blade array for turbomachines comprises blades disposed between an inner wall and an outer wall. The inner wall and/or the outer wall is provided with a suction port near at least some blades. This port has a first end situated along the upper surface in a region of the blade extending from the point of maximum curvature to the neck of the passage between said blade and the adjacent blade. The port enabling the efficiency to be increased is oriented along an isobar line. Its length is such that the second end is spaced from the upper surface of the blade by a distance between one quarter and one half the width of the neck of the inter-blade passage.

Abstract: A mixer is disclosed for absorbing and mixing gases in liquids, or for mixing a liquid with another liquid and/or atomizing a liquid in a gas. The mixer includes an impeller with blades which are bent backward, the blades being surrounded on both sides by a front wall comprising an exhaust opening and a back wall sealing off at least the entire conveying area. A sectional chamber, which is built together with the coaxial hollow main shaft stub, is constructed in the shaft of the impeller and at least two blades are assembled with a pipe for propellant. The inner end of the pipe adjoins a sectional chamber and a nozzle is provided on the outer end. The impeller is closed off with a casing, injectors coaxial with primary nozzles are led through a casing, a mixing nozzle of the injectors is attached on three sides to the casing of the impeller, the front wall and the back wall while connected with a blade on the fourth side.

Abstract: An apparatus and method for varying the cooling air flow within a gas turbine bucket tip. Each turbomachinery blade contains at least one radial cooling passage located between a hollow interior and an outside of the blade. A transversely-oriented elongate channel connects at least one of the cooling passages with the trailing edge of the blade. Secured within the transversely-oriented channel is a pin which extends into at least one of the radial cooling passages to partially obstruct the cooling fluid flow depending on the needs of the design.

Abstract: A turbine aerofoil blade is provided with a shroud member at its radially outer extent for engagement with the corresponding shroud members of adjacent aerofoil blades. A cooling air duct within the shroud member directs cooling air to the abutment region of the shroud member so as to provide impingement cooling thereof. Ports provided in the cooling air duct direct a flow of cooling air from the duct over the outer surface of the abutment region so as to provide additional cooling thereof.

Abstract: An improved turbine blade tip with diffusion cooling holes in the tip is disclosed. One particular embodiment of the invention provides an improved squealer blade tip with diffusion cooling holes in the tip. Yet another embodiment of the invention provides blade tip diffusion cooling holes comprising a first cylindrical portion coupled to a second conical portion.

Abstract: A rotor blade particularly adapted for turbine engines includes various injection holes on the blade tip and near the blade tip and on the root plane and near the root plane so directed as to reduce the tip leakage flow crossing the tip and to control the boundary layer by means of fluid curtain and entrainment effects.

Abstract: A dirt removal means for air cooled blades for removing dirt at the blade of a gas turbine engine. The means is a stamped out sheet metal that defines a deflector or plurality of deflectors adapted to overlie the inlets of the blade's internal cooling passages for diverting the dirt entrained air through a high velocity air stream and discharging the dirt entrained air while allowing the cleaner portion of air to turn the corner of the deflector and flow through the blades cooling passages. These means can be applied to existing blades.

Abstract: Dirt removal means for an internally air cooled blade for a gas turbine engine of the type that includes multi passes some of which are serpentine as made in situ by casting the ribs so as to be angularly disposed at the root of the blade and a straight through passageway with an opening at the tip. The velocity of the airstream in the straight through passageway being sufficiently high to carry the dirt through the opening and the angularly disposed ribs overlying the entrance to the other passageway so as to divert the dirt entrained airstream into the straight through passageway while allowing relatively dirt-free air to turn the angle of the rib to enter the serpentine passes.

Abstract: An airfoil-shaped blade assembly having a metallic core, thin coolant liner and ceramic blade jacket includes variable-size cooling passages and a circumferential stagnant air gap to provide a substantially cooler core temperature during high temperature operations.

Abstract: In a turbomachine having stationary and rotating bladed stages, the blades of the rotating bladed stage are cooled by ducting cooling fluid into an interior cavity formed in each blade. Each rotating blade has a tip portion which is closely adjacent an annular shroud to avoid hot gas products bypassing the blades. The required close clearance between the blade tips and the annular shroud has caused manufacturers to provide a recessed blade tip cap and an open plenum at the blade tip to improve cooling flow. This invention minimizes the height requirements of sidewalls to the open plenum by providing an opening in the blade sidewalls whereby cooling air can be discharged from the interior blade cavity and the blade plenum without regard to the clearance between the blade tip and the annular shroud.

Abstract: The blade has an aerofoil body 10 having a leading edge surface 11 which is cooled by air passing through a helical first passage 12 having first portions passing close to said leading edge surface and alternating with second portions passing through a more nearly central part of the blade section remote from said leading edge. A spanwise but straight second passage 15 extends through the blade in a position within the helical passage and closer to the second than the first portions thereof. Heat abstracted from the leading edge by air flow in said first portions is transferred by the flow to the second portions and from there through the blade material to the flow in the straight passage.

Abstract: A hybrid blade for a fluid flow engine has a U-shaped core of metal or a metal alloy and a ceramic outer jacket which forms together with a mounting plate a unitary, single piece structure. The legs of the core straddle the mounting plate and a heat insulating member is inserted between a crosspiece of the core and the mounting plate. This structure permits relative movement between the core and the jacket to compensate for different heat expansion coefficients.

Abstract: The invention comprises a cooled turbine rotor blade having an improved blade tip structure. A groove is provided in the trailing edge end of the blade tip on those turbine blades whose trailing edge is too thin to support an extension of the blade walls to form a blade tip cavity which extends to the tip of the trailing edge of the blade. The groove protects adjoining exhaust apertures from closure by a blade tip smear.

Abstract: Hollow fluid cooled turbine blades of the multi-pass type used in gas turbine engines have suffered from separation of cooling fluid from longitudinally extending walls defining longitudinally extending passages as the cooling fluid is turned through a turning passage from one longitudinally extending passage to an adjacent one. The invention provides an aerodynamically shaped end portion arranged at an angle with respect to the remainder of the wall to increase the turning radius, and a turning vane is positioned in turning passage to form a convergent passage with the end portion of wall. The turning vane has a large turning radius and a large surface area for retention of a boundary layer of cooling fluid thereon. The turning vane directs the cooling fluid flowing through the convergent passage and over the turning vane substantially parallel to the wall to reduce separation in order to reduce pressure losses.

Abstract: An improved blade tip with an abrasive coating is disclosed. The blade tip is included in a rotor blade which is rotatable with respect to a stationary surface. The tip has a contour which is effective for producing a normal loading component on the coating if the tip contacts the surface while rotating. In a specific form of the present invention, the tip comprises an end wall extending radially outwardly from the perimeter of the outer end of the rotor blade and a concave surface bounded by the end walls and extending radially inwardly therefrom.

Abstract: A muilti-stage turbine having shrouded turbine blades which has a stationary guide grid arranged between the rotor shroud and the turbine casing. By virtue of the construction of the guide grid, leakage and cooling air (escape air) of the rotor blade can be adapted to the low whirl of the main stream passing through the rotor to effectively prevent misaligned entry into the following turbine stage.

Abstract: A hot gas impinged turbine blade suitable for use under super-heated gas operating conditions has a hollow ceramic blade member and an inner metal support core extending substantially radially through the hollow blade member and having a radially outer widened support head. The support head has radially inner surfaces against which the ceramic blade member supports itself in a radial direction on both sides of the head. The radially inner surfaces of the head are inclined at an angle to the turbine axis so as to form a wedge or key forming a dovetail type connection with respectively inclined surfaces of the ceramic blade member. This dovetail type connection causes a compressive stress on the ceramic blade member during operation, whereby an optimal stress distribution is achieved in the ceramic blade member.

Abstract: An improved high pressure turbine rotor blade and tip cap structure therefor is provided which comprises a tip end closure for the blade bonded to the end wall of the blade casting and including a base plate member and a pair of upstanding side walls defining a peripheral contour coincidental with and faired to the cambered side walls of the casting, the base plate member having a plurality of radially outwardly opening passageways therethrough and disposed along the chord of the blade and communicating with a channel included in the radially outwardly facing surface of the casting along a chord of the blade, the passageways disposed along axes at angles to the base plate whereby coolant fluid flowing therethrough is directed against the tip side wall surfaces; and opening is provided through the end wall of the casting in the form of a slit or a plurality of holes along the chord of the blade and intersecting the channel, to define an outlet through the end wall for passage of coolant fluid through the blade

Abstract: A hollow turbine blade 10 of aerofoil shaped cross section is provided with walls 18 of varying thicknesses to provide thermal slugging at those regions around the aerofoil profile where the blade experiences high thermal fluxes. The thicker regions of the blade 10 are provided with cooling channels 22 which are elongated in a direction extending from the external surface of the wall towards the internal surface of the wall. The thermal slugging slows down the rate of change of blade wall temperature to a value similar to the rates elsewhere in the blade section. The cooling channels 22 ensure that the temperature gradient across the wall 18 is low around the profile of the aerofoil.

Abstract: A ceramic blade assembly including a corrugated-metal partition situated in the space between the ceramic blade element and the post member, which corrugated-metal partition forms a compliant layer for the relief of mechanical stresses in the ceramic blade element during aerodynamic and thermal loading of the blade and which partition also serves as a means for defining contiguous sets of juxtaposed passages situated between the ceramic blade element and the post member, one set being open-ended and adjacent to exterior surfaces of the post member for directing cooling fluid thereover and the second set being adjacent to the interior surfaces of the ceramic blade element and being closed-off for creating stagnant columns of fluid to thereby insulate the ceramic blade element from the cooling air.

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: The turbine blades of a high pressure turbine in a gas turbine jet engine are cooled by cooling air injected into cooling channels in the blade using angled nozzles supplied with high pressure air from the engine compressor. The nozzles project the high pressure air across a lower-pressure gap into slots which lead to internal channels in the turbine disk without requiring the use of seals to produce a pressure differential to encourage the flow of cooling air. Maximum air transfer into the internal channels in the turbine blades occurs when the air impacting slots in the turbine wheel is travelling at an angle with respect to the slots which is approximately parallel to the axes of the internal channels. A relatively narrow range of angular misalignment for aiming the nozzles is permitted. The pressure of the air from the plenum may be reduced by employing two or more flow channels in series before projecting the air across the gap to the turbine wheel.

Abstract: The invention comprises a cooled turbine rotor blade having an improved blade tip structure. A recessed tip is provided at the leading edge end of the blade tip on downstream turbine blades which are too narrow to support a blade tip cavity over the entire exterior surface of the blade tip without interfering with cooling airflow from apertures in the exterior surface. The recessed tip structure protects apertures therein from blockage by a blade tip smear and does not substantially reduce the performance efficiency of the blade.