Abstract: A runner for a hydraulic turbine or pump includes a hub and a plurality of blades extending from the hub at spaced intervals therearound. Each blade includes an inner edge secured to the hub and a distal outer edge, a leading edge and an opposed trailing edge, and a curved suction surface and an opposed curved pressure surface. At least one of the blades is fabricated from a curved pressure-side member and a curved suction-side member secured together. The pressure-side member includes at least a substantial portion of the curved pressure surface and an opposed first inner surface. The suction-side member includes at least a substantial portion of the curved suction surface and an opposed second inner surface. The first and second inner surfaces face each other, and at least a portion of the first inner surface is spaced from at least a portion of the second inner surface to form a cavity within the blade. A method of making the hollow blade is also disclosed.

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 turbine, adaptable to enhance the levels of dissolved gas such as oxygen in water flowing therethrough, includes a rotating shaft and a runner secured to the shaft. The runner includes a crown, a band and a plurality of runner blades. Each blade is defined by an inner edge and a distal outer edge, a leading edge and a trailing edge separated by a water directing surface. At least one of the blades includes a leading blade portion with a first slot along a rear edge thereof and a trailing blade portion with a second slot along a front edge thereof. The leading and trailing blade portions are fixedly secured together along their front and rear edges so that the first and second slots cooperate to form an integral gas passage in the at least one blade. The at least one blade further includes a means for discharging gas from the channel to a location adjacent the trailing edge.

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 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: A turbine blade squealer tip has a cooling passage arrangement which includes a peripheral end wall surrounding a central portion of an end cap of the squealer tip, an outer peripheral groove defined in an outer surface of the peripheral end wall being spaced outwardly from the central portion of the end cap, and a multiplicity of holes extending through the end wall from the outer groove to an internal source of cooling air flow within the turbine blade, bypassing the central portion of the end cap of the squealer tip. The outer groove which receives the flow of cooling air from outlets of the holes tends to trap the air therein so as to form an annular air seal about the outer surface of the squealer tip end wall which impedes leakage of hot gas flow past the squealer tip.

Abstract: A circulation control aerofoil having an internal chamber for receiving during operation air for exhaust through a circulation control air exhaust slot defined by spaced-apart slot lips is provided with a plurality of actuators of electrically deformable material adapted during operation to cause local bending of the aerofoil structure to which they are attached to selectively adjust the width of the air exhaust slot. The actuators may be piezo-electric actuators.

Abstract: A gas turbine engine turbine airfoil includes first and second opposite sidewalls joined together at leading and trailing edges and extending from a root to a tip thereof. The sidewalls have respective outer surfaces over which is flowable combustion gases during operation. The sidewalls are spaced apart to define an internal flow passage for receiving a cooling fluid. A plurality of ejection slots are arranged in a column through the first sidewall and have flow inlets for receiving cooling fluid from the internal flow passage and ejecting the fluid from respective flow outlets along the outer surface. At least some of the ejection slots are inclined at an acute ejection angle which varies in magnitude for preferably matching the respective inclination of combustion gas streamlines flowable over the airfoil surface.

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 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 hollowed, internally cooled airfoil per gas turbine engines with an improved internal configuration for pronounced impingement cooling. Internal ribs extend across the hollowed interior and cooling passages through these ribs cause impingement cooling of the next-adjacent rib as well as the internal surfaces of both the pressure side and suction side of the airfoil.

Abstract: A flexible spoiler is arranged on the suction side of a wing and spoiler is made of a narrow flexible fold tape covering a tunnel which is sealed against the outer airstream and extends inside the wing from the root to the tip thereof. When the fold tape is deployed from a conform position it forms a trip producing stall. Deploying and conforming are achieved by varying the tunnel pressure. The pressure is essentially generated and varied by the tunnel aircolumn under centrifugal acceleration and throttled by an inlet throttle near the wing root and an outlet throttle near the wing tip. Opening the inner throttle increases the tunnel pressure to deploy the fold tape in the case of maximum speed and/or shaft torque of the wind-turbine.

Abstract: A circulation control rotor system for aircraft in which compressed air is directly supplied at generally constant pressure to plenums in the rotor blades and discharged through area controlled slots, the area controlling means involving camming surfaces and being actuated by cyclic, collective, aircraft vibration and flight stability inputs.

Abstract: An impeller forming a unit with the spherical rotor of a sphero-pump consists of a disc with vanes and an inlet member, both made by injection molding from resin. A rotor plate made from a thick gauge sheet metal covers the vanes and is held in position by rivet type protrusions on the inlet member and compensates the weight of the rotor.

Abstract: A turbine plant compressor disc is formed from two halves which are welded together near the rim and define an internal cavity in which a centripetal accelerator is mounted for inducing cooling air to flow from the compressor through inclined passages in the rim on the downstream side of the blades carried by the rim, through the cavity of the disc, and towards the turbine to be cooled in the vicinity of the turbine shaft. The centripetal accelerator is composed of sectors which are removably mounted in the cavity of the disc, the sectors forming a perforated ring disposed in the central bore of the disc and each sector having a flat web extending radially outwards in the inner cavity of the disc to form radial centripetal channels communicating with the air passages in the rim.

Abstract: The penetrating members (21, 35, 37, 38, 39) are disposed inside the cooling passage (8) at positions so as to penetrate through the cooling passage (8), and have both of ends thereof fixed to the rear side wall (17) and the body side wall (18). The projecting members (33) (34A, 34B) are disposed inside the cooling passage (8) so as to project from the rear side wall (17) or the body side wall (18). The rectification ribs (36) are disposed at intermediate positions between the penetrating members (21, 35, 37, 38, 39) or the projecting members (33, 34A, 34B). The penetrating members (21, 35, 37, 38, 39) or the projecting members (33, 34A, 34B) disturbs the main stream (3a) of the cooling fluid (3) near the internal partition (19) in both directions of the leading edge wall (16) and the first internal partition (19) and make the main stream (3a ) non-uniformly or causes its channelling.

Abstract: An apparatus for aerating water discharging from a hydraulic turbine comprises an air distribution apparatus secured to the trailing edges of a turbine runner. The air distribution apparatus comprises a manifold which contains a bounded air channel extending the length of the runner blade trailing edge. A plurality of holes are formed along the mainfold extending between the channel and the turbine interior. The channel is in air flow communication with a source of oxygen.

Abstract: A sheet metal electrode for forming a shaped, diffusing film coolant passage has a straight, longitudinally extending forward section forming an acute angle with a flat middle section, which, in turn, extends from a rear section which is the base of the electrode. The middle section includes a pair of edges rearwardly diverging from the longitudinal direction and from which extend side walls substantially perpendicular to the middle section. Longitudinally extending edges of the side walls are in the same plane as a lower, flat surface of the forward section such that the surface of the passage formed by such lower surface of the forward section is in the same plane as a surface of the diffusing portion of the passage which is formed by the longitudinally extending edges of the side walls.

Abstract: A compressor disc assembly with an integral centripetal accelerator for the exhaustion of air for use in a cooling device of a gas turbine, the bleeding of air at the surface of the compressor drum being effected by orifices situated in the platforms of the blades and in the median plane of the compressor disc. The assembly is formed from two half discs. The centripetal accelerator includes several sectors detachably mounted in the space defined between the two half discs of the compressor, each sector being formed from two half shells having on their internal face ribs in the form of centripetal blades by which the half shells are assembled together. The invention is utilized for cooling a gas turbine by bleeding air from the compressor.

Abstract: A high wing, twin jet gyrodyne having air jet reaction, contrarotating rotors powered with bypass air from twin turbofan engines. Tail pipes from the turbofan engines are disposed parallel to and at opposite sides of a longitudinal duct for discharging bypass air rearwardly. Control valve means can control supply of bypass air selectively to the hollow rotor mast for driving the rotors and to the longitudinal air discharge duct. The aft portions of the tail pipes and of the air discharge duct are connected by a horizontal stabilizer and rudders are located in the slipstream discharged from the tail pipes and the air discharge duct. The wings have drop tip sections providing flotation outriggers in their downwardly projecting positions.

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: A system for cooling a gas turbine by bleeding air from the compressor and channeling the air at the center of the engine towards the turbine in order to cool it. The air is bled from the plane of the disc bearing the blades of the compressor, said disc having radial ducts for the flow of the air. The invention is utilized in gas turbine engines to cool the turbine.

Abstract: Apparatus and propeller for distributing a gaseous, liquid or powdered material in a liquid by rotation in the liquid of a propeller having two or more blades. The blades have a perforated hollow portion connected to a source for the gas. The propeller is rotated at a cavitation producing velocity, and for the distribution of the material in the liquid use is made of the pressures occurring upon the implosion of the cavitations.

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 hollow cooled blade or vane for a gas turbine engine is provided with cooling on certain areas of its interior surface by an impingement plate which is supported from and sealed to the blade interior by a pair of ribs.

Abstract: An axial supersonic compressor comprises a casing and a hub rotating in the asing and carrying blades. The suction surface of each blade is formed with a zone in which the curvative changes and which corresponds to a supersonic-subsonic shock wave. A channel formed in each blade and opening in said zone is connected to boundary layer aspiration means.

Abstract: A cooling arrangement for high temperature turbines, particularly the first stage nozzles of gas turbine engines. Cooling is effected by compressor discharge air taken from the combustion chamber jacket of the engine. The shrouds of the nozzle have a base ring and a porous facing spaced from the base ring and exposed to the motive fluid. Cooling air fed through openings in the base ring flows through the porous facing into the motive fluid path to cool the facing by transpiration. The vanes are cooled by an arrangement providing for a mixture of impingement and transpiration cooling. Compressor discharge air is admitted between a baffle in the interior of the hollow vane and the high pressure face of the vane. Some of the air flows through the porous high pressure face for transpiration cooling. The rest of the air is discharged through a row of openings in the baffle to impinge on and cool the interior of the leading edge of the vane, which is imperforate.

Abstract: This invention cures the problem attendant the ingestion of oil in the engine of a fan-turbine engine combustion propulsor should oil leak from the fan blade seal by capturing and judiciously venting the leakage.