Abstract: A gas turbine engine rotor disc (20) is provided with a plurality of cooling air supply passages (32) which are generally radially extending but are inclined in the downstream direction. Each passage (32) has a cross-sectional configuration which renders the ends (34) of the passages (32) less likely than normal to act as the site of hoop-stress induced cracks.

Abstract: A cavity judiciously dimensioned and located adjacent the tip's surface discharge port of internally cooling passage of the airfoil of the turbine blade of a gas turbine engine and extending from the pressure surface to the back wall of the discharge port guards against the contamination and plugging of the discharge port.

Abstract: A rotor disc (15) has an axially enlarged hub portion (18) with a bore (20) through which a turbine shaft (19) extends in a coaxial, radially spaced apart relationship, the annular gap between the hub portion (18) and the shaft (19) receives a flow of temperature regulating fluid. The shaft (19) is provided with axially extending elongate projections (22) to increase the turbulence and relative velocity of the temperature regulating fluid adjacent the projections (22).

Abstract: In a compressor rotor assembly including a plurality of rotor disks affixed to a rotor body, and wherein at least one radially directed cooling tube is supported from a radially outer end within a rotor shaft, with an opposite, radially inner free end extending toward a longitudinal axis of the rotor, an improvement comprising a ring member loosely secured to the tube at the free end.

Abstract: An arrangement for the thermal protection of a rotor, fitted with moving blades, of a high-pressure compressor. The roots of the moving blades are inserted and locked in peripheral grooves which are at an axial distance from one another. The rotor has, between two adjacent peripheral grooves for the moving blades, at least one further peripheral groove having a hook extending over the entire periphery of the rotor. At least two plate-shaped heat-accumulation segments, each having at least one root which has a contour adapted to the hook of the rotor so that both form a contact area, are pushed radially one after the other into the further peripheral groove and locked therein. A cavity forming an insulating layer extends between the heat-accumulation segments and the rotor and between the moving-blade roots and the rotor, in which insulating layer the velocity of the compressor air is greatly reduced.

Abstract: A lightweight high temperature rotor blade attachment structure for use in a gas turbine engine. The lightweight high temperature rotor blade attachment lug being cast of a single crystal alloy and the lug is then bonded to a conventional nickel based wheel. A circular arc firtree is utilized to connect the insertable turbine blade between a pair of circumferentially spaced lugs that have been bonded to the turbine disk. In an alternate form of the present invention the attachment lug includes an internal cooling passage for receiving cooling fluid from a compressor. More particularly, the present invention discloses a single crystal attachment lug that is bonded to a powdered metal nickel alloy rotor disk and includes internal cooling passages, and in one form is designed for use with turbine blades that do not have a platform.

Abstract: In a sealing arrangement between two discs of a gas turbine engine a split sealing arrangement comprising a pair of sealing formations each made up of a plurality of sealing segments is provided. Each of the segments has a root accommodated in serrations of the discs. The roots of the sealing segments are positively connected to roots of the rotor blades by dovetail or T-shaped tongues and grooves. The sealing segments are also provided with internal cooling passages.

Abstract: Because a gas turbine stationary blade is exposed to a high speed, high temperature gas flow, its service life is shortened due to thermal stress and high temperatures. The present invention provides a gas turbine stationary blade unit with reduced thermal stress, a high cooling effect and an elongated service life. A blade (103) is formed by a thin plate panel (110) formed with a reinforcing element (105) disposed on an inner side or cooling side of the blade (103). Thereby, thermal stress occurring at the blade (103) is reduced as well as the fluid force acting on the blade (103) by an operating gas (G) is transmitted to a stationary blade holding portion (104) via the reinforcing element (105). Thus, the thickness or rigidity of a connection portion of the blade (103) and an outer shroud (101) need not be increased, and a thermal deformation and an unhomogenous distribution of thermal stress can be avoided.

Abstract: A labyrinth disk includes a main stiffener placed in the middle of the rim immediately below labyrinth elements. Attachment elements are preferably in the form of a bayonet attachment system using teeth fixed on the labyrinth disk crown and teeth fixed on the rotor. Attachment by bolting may optionally be used. The disk may be utilized with turbojets, on the cooling circuit, on the upstream side of the high pressure turbine.

Abstract: A cooled rotor assembly for a gas turbine engine includes rotor blades (48) and a plurality of cover plates (86a) retained in the blade attachment slots (34) of a rotor disk (30). Each blade has a platform (46) with a trailing edge (58), and each cover plate includes at least one conduit (120, 126) for drawing leakage air from a plenum (102) in front of the disk and impinging that air against the platform trailing edges (58) to cool the tailing edges. In the preferred embodiment, each conduit (120) is defined by the rear face of a blade root (40) and a groove (122) cast into the cover plate. The invention makes productive use of leakage air by using that air to cool the platform trailing edges rather than allowing it to flow unbeneficially into the engine gaspath (24) in the vicinity of the platform leading edges (56).

Abstract: Disclosed is a gas turbine rotor (5) having multi-stage moving blades, each fitted to a disc (12), characterized in comprising: an inner cavity (9) and an outer cavity (8) provided between each disc (12); a blade cooling passage (6) of each moving blade except a rearmost stage moving blade (4); a cooling steam supply passage (13) extending from a rear portion of the rearmost stage moving blade (4) to a leading edge portion of a foremost stage moving blade (1) in the rotor (5); a bifurcation passage (16) provided in the disc (12) portion of the foremost stage so as to connect at its proximal end to the cooling steam supply passage (13) and to bifurcate at its distal end so that one bifurcation thereof (16a) connects to one end of the blade cooling passage (6) and the other bifurcation (16b) connects to the outer cavity (8); a blade return passage (17) connecting at its proximal end to the other end of the blade cooling passage (6) and at its distal end to the inner cavity (9); a cavity connecting passage (20)

Abstract: A turbine rotor includes a rotor disk having a plurality of dovetail slots defining therebetween respective disk posts. Each post has a top land and a plurality of tangs spaced radially therebelow inside adjacent ones of the slots. A plurality of turbine blades each having a dovetail are mounted in respective ones of the dovetail slots for radially retaining the blades. A platform is joined integrally with the dovetail and extends circumferentially to adjoin adjacent ones of the platforms to define a radially inner flowpath. An airfoil extends integrally from the platform for extracting energy from combustion gases flowable thereover. A thermal barrier coating is fixedly bonded to each of the disk post top lands for providing thermal insulation between the platforms and the disk posts for reducing disk post temperature.

Abstract: The inter-disk cavity between turbine rotor disks is used to pressurize cooling air. A plurality of ridges extend radially outwardly over the face of the rotor disks. When the rotor disks are rotated, the ridges cause the inter-disk cavity to compress air coolant flowing through the inter-disk cavity en route to the rotor blades. The ridges eliminate the need for an external compressor to pressurize the air coolant.

Abstract: An internally cooled airfoil for a gas turbine engine includes a leading edge region, a mid-chord region, a transition region, and a trailing edge region. The mid-chord region has a double wall configuration with the two outer walls and two inner walls. The trailing edge region has a conventional single wall configuration. The transition region has a three wall configuration designed to provide a gradual transition from the four wall configuration to the two wall configuration and to minimize high stresses therein.

Abstract: This device embodies a centrifugal compressor secured to a joint-carrier disk secured to an upstream flange of the rotor disk to be cooled. The recompressed air is ejected through channels near the roots of the blades of the disk.

Abstract: A vapor cooled gas turbine has a cooling system including a vapor supply port and a vapor recovery port, and the cooling system is formed so that vapor from the supply port is supplied to blades through a central supply passage in a rotor and the vapor having cooled the blades is recovered from the recovery port through a recovery passage spaced outwardly from the supply passage.

Abstract: A turbine airfoil has a mesh cooling hole arrangement which includes first and second pluralities of cooling holes formed within the interior structure of the side walls of the airfoil so as to extend between and along but not intersect spaced internal and external surfaces of the side walls extending between leading and trailing edge portions of the airfoil. 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 walls having pairs of opposite sides interconnected by pairs of opposite corners. The spaced nodes define a multiplicity of hole portions of the cooling holes extending between and along opposite sides of adjacent nodes and a plurality of flow intersections interconnecting the hole portions of the cooling holes and being disposed between the corners of adjacent nodes.

Abstract: In a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of the compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, the improvement comprising means for supplying cooling air at least to a peripheral rim of the last stage of the compressor.

Abstract: A jet engine component, such as an aircraft gas turbine engine rotor blade or a scramjet engine fuel injector. The component has a wall portion including a first surface exposable to a cooler, higher static pressure fluid and a second surface exposable to a hotter, lower static pressure gas flow flowing across the second surface. The component further includes a generally straight film coolant passageway having an inlet on the first surface and an outlet on the second surface. The second surface has a seamless groove which is open substantially entirely along its longitudinal dimension extending from the outlet along the lower static pressure gas flow for improved film cooling of the second surface.

Abstract: A wall adapted for use in a gas turbine engine between a first and a hotter second fluid includes a first side over which is flowable the first fluid, and an opposite second side over which is flowable the second fluid. An elongate slot extends inwardly from the second side and is disposed in flow communication with a plurality of longitudinally spaced apart holes extending inwardly from the first side. The holes are disposed at a compound angle relative to the second side for discharging the first fluid obliquely into the slot and at a shallow discharge angle from the slot along the second side. In a preferred embodiment, the slot has an aft surface including a plurality of longitudinally spaced apart grooves extending from the holes to the wall second side.

Abstract: A turbine blade has a cooling air flow path specifically directed toward cooling the platform portion of the blade root. Two cooling air passages are formed in the blade root platform just below its upper surface. Each passage extends radially outward from an inlet that receives a flow of cooling air and then extend axially along almost the entire length of the platform. Each passage also has an outlet formed in the downstream face of the platform that allows the cooling air to exit the platform and enter the hot gas flow path. The passages are formed in portions of the platform that overhang the shank portion of root.

Abstract: A cooling system for use in a rotor assembly of a gas turbine engine, the rotor assembly including an annular disk rotatable about a centerline axis of the engine and a plurality of blades mounted on the disk. The disk includes alternating posts and slots, with each slot receiving a blade dovetail. The cooling system comprises an axially extending plenum disposed inward of each blade dovetail and a thermal isolation chamber, formed by a seal body, disposed over the outer surface of each disk post. An annular forward blade retainer is attached to the disk and sealed with adjacent structures via inner and outer seals. Alternative structures are disclosed for diverting the cooling air from the plenums, past the inner and outer seals, and into the isolation chambers during operation of the engine so as to cool each disk point, without compromising the structural integrity of the forward blade retainer.

Abstract: An apparatus for modulating a flow of cooling air from a compressor to a turbine in a gas turbine engine is disclosed. The apparatus includes a support member which forms a channel for receiving the cooling flow and has a throat at a downstream end with a radial orientation. A radial flow inducer is connected to the support member adjacent to the channel downstream end for accelerating the cooling flow through the throat of the support member. Apparatus for modulating the quantity of the cooling air flowing through the throat and into the turbine is provided in the form of an axially translating endwall, wherein variation of the throat area is accomplished through modulation of the endwall with respect to the support member. The support member may also house either a stationary, unmodulated axial flow inducer or a second modulated radial flow inducer in addition to the modulated radial flow inducer.

Abstract: A hollow turbine blade has cooling air feed passages in the wall defining the central cavity. The wall surfaces of the feed passages slope away from the central cavity in a diverging manner, so as to reduce the step difference in metal thickness between the passages relative to the thickness of metal dividing each passage and the central cavity. The temperature gradient from the blade external surface and the wall surface of the central cavity is thus also reduced, avoiding thermal stresses.

Abstract: Apparatus is provided for creating heat by electromagnetic induction between relatively rotating components. In RAT applications, part of the nose cone (34) is formed from electrically conductive material (106) which is in thermally conductive relation to the nose cone exterior. A magnetic field is created by components (60) that are spaced apart from the electrically conductive material and relatively rotatable with respect to that material. The electrically conductive material intersects the magnetic field, at least in part, such that eddy currents are created in the electrically conductive material by relative rotation with respect to the magnetic field. These eddy currents create heat which is conducted to the exterior surface of the RAT nose cone. Preferably, the magnetic field is created by a pair of spaced-apart permanent magnets (60) mounted on a support plate (62) inside the RAT hub (28) and immediately behind the nose cone.

Abstract: Particulates are removed from the coolant used to cool gas turbines by subjecting the coolant to the centrifugal acceleration of the gas turbine rotor. Axially oriented channels are formed in the rotor at a radial distance from its longitudinal axis. The channels are in fluid communication with a coolant supply passage so that coolant coming onboard the rotor will first pass through the channels. The high centrifugal force produced by rotor rotation causes particulates in the coolant passing through the channel to move radially outward and deposit on a fixed surface in the channel. Thus, the particulates are collected in the channel and removed from the coolant discharged from the channel. The channels may be provided with removable liners.

Abstract: A hot air/cold air dual-mode electric fan including a vane assembly having an electric heating coil in the mica sheets of each vane, a bakelite mount fixedly secured to the vane assembly and having a circuit board with backward female contacts, a barrel mounted inside a holder in front of a motor and having separated carbon brushes, a shaft mounted within the barrel and coupled to the motor and having male contacts respectively connected to the female contacts and separated copper rings respectively disposed in contact with each carbon brush, and a control circuit for controlling power supply to the electric heating coils and regulating its heating temperature, so that the vane assembly produces currents of cold air when the motor is started and power supply is cut from the electric heating coils by the control circuit, or currents of hot air when the motor is started and power supply is connected to the electric heating coil by the control circuit.

Abstract: An apparatus for sealing a gap between adjacent blades in a rotor assembly for a gas turbine engine is provided. The rotor assembly includes a plurality of blades circumferentially disposed around a disc. Each of the blades includes an airfoil, a root, and a platform extending outward in a lateral direction in a transition area between the root and the airfoil. The disc includes a plurality of complementary recesses circumferentially distributed around the disc for receiving the blade roots. The gaps are formed between edges of adjacent platforms. The platforms collectively form a flow path for primary fluid flow passing by the airfoil side of the platforms and secondary fluid flow passing by the root side of the platforms. The apparatus comprises a thin plate body and apparatus for conducting secondary flow between the thin plate body and root side surfaces of adjacent blade platforms, and thereafter into the gap.

Abstract: A cavity judiciously dimensioned and located adjacent the tip's surface discharge port of internally cooling passage of the airfoil of the turbine blade of a gas turbine engine and extending from the pressure surface to the back wall of the discharge port guards against the contamination and plugging of the discharge port.

Abstract: A heat pipe, which may for instance be in the form of an aerofoil blade for a gas turbine engine, is provided with capillary means which is in two different forms. One form is constituted by a porous material and the other constituted by grooves formed in the surface of the internal wall of the heat pipe.

Abstract: Apparatus for heating and circulating air within a room, the apparatus including a ceiling fan having fan blades, a heater housing disposed vertically above the ceiling fan, a heating element disposed within the heater housing, intake ports provided in the heater housing for transferring air from the room into the interior of the heater housing, the air intake ports being disposed vertically above the ceiling fan, and air output ports provided in the heater housing for transferring air from the interior of the heater housing into the room, the air output ports being disposed vertically below the ceiling fan and outside of the radial sweep of the fan blades.

Abstract: A turbine blade includes a plurality of cooling passages each having a turbulated section of the passage preferentially located along the portion of the turbine blade subjected to the highest temperature. Thus, turbulent air flow is provided in intermediate sections of the blade to enhance the heat exchange relation with the metal of the blade. The bores of the cooling passages adjacent the tip and root portions are smooth and provide adequate cooling in those sections at a lower heat exchange relationship. The cooling passage bores are formed by an electrochemical machining process using an elongated electrode with a chemical electrolyte for forming enlarged cavities within the blade.

Abstract: The present invention discloses a system for cooling turbine rotor disk posts in a gas turbine engine with a preferred embodiment illustrating an application for a high pressure turbine stage 1 rotor disk. The disk includes a plurality of circumferentially alternating posts and slots disposed about the periphery of the disk with each slot receiving a dovetail of a radially extending blade. The cooling system comprises a plurality of seal bodies with each seal body including a relatively low volume thermal isolation chamber positioned over the top of a corresponding disk post; a plurality of axially extending blade cooling plenums supplied with cooling air with one of the plenums positioned radially inward of each of the blades; a plurality of shallow, radially extending slots formed in a relatively low stressed axially facing surface of each of the blade dovetails for diverting cooling air from the blade cooling plenums to the thermal isolation chambers.

Abstract: A hand-held mixing device comprising an electric motor, an elongated shaft extending from the motor, a housing enclosing the motor and a major portion of the shaft, a blade at the free end of the shaft, a heating element extending from the housing, the heating element formed to have at least one coil surrounding the blade, the coil being dimensioned to be spaced from the outer edge of the blade and to form a generally annular gap thereabout. The motor and heating element are electrically connectable to an electrical power source wherein the electrical power to the motor and to the heating element are independently controllable.

Abstract: The bucket 10 of a gas turbine includes first and second steam supply passageways 32 and 34 for directing cooling steam radially outwardly along the bucket. Steam return passageways 40, 42 and 44 lie adjacent the trailing edge for returning cooling steam to its source. A shroud is disposed at the bucket tip and has cross-over passages communicating the supply steam from the first and second passageways to the fifth, third and fourth passageways, respectively.

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: An active cooling mechanism for a turbine disk is disclosed. Various construction details are developed which disclose a turbine section having a rotor assembly, a disk with an attachment means, a radially movable heat shield disposed therebetween, and a cooling passage which directs cooling air over the forward surface and radially outer surface of the attachment means. During operation, rotational forces urge the heat shield radially outward to create a gap between the heat shield and the attachment means, the gap defining a portion of the cooling passage.

Abstract: The steam transfer arrangement includes dovetail connections between a turbine bucket 14 and rotor wheel 12 wherein a steam supply passage 20 in the rotor wheel registers with a steam return passageway 28 in the bucket and a steam return passageway 30 in the bucket lies in registry with a steam return passage in the wheel. At the interface of the dovetail connection, each registering passage and passageway includes a sleeve 50, 52 disposed in an enlarged recess 54, 56, the sleeves having annular spherical sealing surfaces for engagement with annular spherical seats on the turbine bucket. The sleeves are compressible radially relative to the bucket to enable the bucket to be secured to the wheel and expand radially outwardly to ensure sealing fit between the sealing surfaces and seats.

Abstract: An anti-rotation system for an interstage seal located in a gas turbine engine of a type having a turbine section including a first stage disk and a second stage disk, each having a plurality of dovetail slots. The interstage seal is located between the disks and includes a web portion having a forward arm and an aft arm, the aft arm having a bayonet connection for preventing relative axial movement between the second stage disk and the seal The interstage seal anti-rotation system includes a key positioned in at least one dovetail slot and having a tab for interconnecting with a slot formed in the aft arm bayonet connection. The key prevents relative circumferential movement between the interstage seal and the second stage disk An aft seal plate is provided for retaining the key in the interconnected position between the seal and the second stage disk.

Abstract: A gas turbine engine having a turbine rotor assembly with a free standing sideplate assembly is disclosed. Various construction details are developed which provide a sideplate assembly which is not radially or axially supported by the web or rim of the adjacent disk. In one particular embodiment, a rotor assembly includes a rotor disk, having a rim, a web, and a bore, and a sideplate assembly, having a web and a bore. The web of the sideplate is radially supported by the bore of the sideplate and includes a disk seal and an aperture. The disk seal is engaged with the rotor disk and has an axially directed seal force provided by an axially interfering fit between the sideplate and rotor disk. The aperture provides fluid communication between a source of cooling fluid and the rotor disk.

Abstract: An electrical an (10) for the icing-susceptible stator vane stage (16) located in an air inlet section (12) of a turbomachine (not shown). A plurality of stator vanes (20) which are electrically isolated from a plurality of structural shrouds (30, 32) are grouped in one continuous, or a plurality of contiguous arcs about a fanshaft which is generally coincidental with the turbomachine's longitudinal axis (X--X). Low voltage alternating current is provided by an alternator (not shown). This current is directed through a plurality of electrical conductors (26) and is further distributed through each stator vane (20) which is either electrically conductive or employs an electrically conductive medium disposed proximate to an icing susceptible surface of an airfoil contour (21). The electrical current flow, which is generally directed in a radial orientation through each stator vane (20) causes a direct electrically-resistive heating effect to the airfoil contour surfaces (21) of the stator vanes (20).

Abstract: Sheet Metal thermal control dampers are disposed in the cavities between turbine blades to bear against the undersides of opposed blade platforms portions of angularly adjacent blades under centrifugal loading to dampen blade vibrations and to seal the gaps between platforms against the radial flow of working fluid into the interblade cavities. The dampers are configured to cooperate with forward and aft seals to also block the axial flow of working fluid through the interblade cavities. Holes are provided in the dampers through which cooling air introduced into the interblade cavities can flow into contact with the platform undersides and cool the platforms, thereby controlling temperature rise in the blade attachment structure.

Abstract: The present invention is a device that can obtain fresh water from fog. The fog water collecting device comprises a vertical shaft, windmill elements for producing rotational force from wind, a cylindrical rotating structure supported so as to rotate about the vertical shaft, a water collecting vessel secured at the bottom of this rotating structure, a plurality of flexible rods one end of each is fixed at the upper portion of the rotating structure and the other end of each is directed toward the water collecting vessel and, a receiving conduit for obtaining water collected in the water collecting vessel. When the rotating structure turns from fog containing wind, water droplets sequentially adhere and accumulate on the flexible rods, then collect in the water collecting vessel, and are subsequently directed toward an external destination via the conduit.

Abstract: A forward seal assembly located in a turbine section of a turbine engine, the turbine section having a disk including a web, a bore, and a forward shaft integral with the web, wherein the forward seal assembly includes a face plate extending from the forward shaft to an outer periphery of the disk and includes orifices for conveying cooling air to the web. The face plate includes a plurality of radially inwardly-extending tabs shaped to engage radially outwardly-extending tabs located on the web to form a bayonet connection. The forward seal assembly further includes radially extending vanes to convey cooling air along the disk, and is secured by inward and outward bayonet connections, the inward connection including locking pins which are arranged to perform a balanced function.

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 rotor having a bladed disc and for use in a fluid flow machine has plates extending between and supported by the blades to protect the disc rim from erosion and foreign object damage and also, when the rotor is a gas turbine, to protect the disc rim from the heating effects of high temperature gasses. The plates are held between ridges on the blade faces.

Abstract: An air fan includes a plurality of fan blades each fan blade drilled with a plurality of perforations in the fan blade for reducing weight of each fan blade for reducing the electric energy required for driving the fan motor of the air fan. Each fan blade may be reinforced by at least a rib on a surface of each fan blade for enhancing its strength and preventing its deformation.

Abstract: An interstage thermal shield assembly for a gas turbine engine. The assembly includes an axially-extending thermal shield positioned between first and second stage disks to form a seal therebetween. The thermal shield includes complementary hook members shaped to engage slotted hook members formed on the first stage disk to form a bayonet connection, and complementary lip members for engaging lobe members formed on the second stage disk by virtue of split rings, thereby eliminating the need for a bolted connection between the thermal shield and disks and facilitating attachment and removal of the thermal shield. The thermal shield includes an annular impeller which is positioned rearwardly of the first stage disk, the impeller including bayonet connection with the disk which restrains the impeller from axial deflection, but permits radial deflection in response to thermal changes.

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: A method and apparatus for reducing thermal distress and creep of a disk post in a gas turbine engine, the disk post being defined between an adjacent pair of blade roots of a respective pair of turbine blades in a turbine disk. The blade roots extend radially outward of the turbine disk and each terminate in a blade platform to form a cavity above the disk post. A seal generally covers the cavity for preventing a flow of combustion gases over the disk post. The seal includes axial segments defining a channel over a radially outer surface of the disk post. A flow of insulative air is directed into the channel defined over the disk post and diffused to reduce its velocity. The insulative air effectively reduces heat transferred from the blade platforms and combustion gases to the disk post.