Abstract: Disclosed is a runner for a Francis-type hydraulic turbine, comprising a ring, a hub, and a number of blades all having a curved shape and being attached to the ring and hub. Each blade has an inlet edge and an outlet edge. The blades are characterized in that, in the direction of turbine rotation, the junction point of each blade inlet edge at the ring is located forwardly of the inlet edge attachment point for that blade at the hub, and the junction point of each blade outlet edge at the hub is located forwardly of the outlet edge attachment point for that blade at the ring.

Abstract: A plain blank is sized to produce an airfoil. The blank is plastically formed to include a pocket blending smoothly in one side thereof. Leading and trailing edges of the airfoil are benched. And, filler is placed into the pocket to effect an aerodynamic profile for the airfoil.

Abstract: A blade for an axial fluid machine comprises an effective blade portion having a root portion and a tip portion and projecting blade portions. The projecting blade portions are defined by axis reference lines extending to an upstream side with respect to a fluid flow from at least one of the root portion and the tip portion of the effective blade portion and by axes obliquely extending from ends of the axis reference lines toward a leading edge of the effective blade portion. The projecting blade portions are formed continuously to and integrally with the leading edge of the effective blade portion so that the effective blade portion and the projecting blade portions have substantially the same maximum blade thickness.

Abstract: A blank is milled using a rotating mill. A groove having a bottom is first cut across the blank to partially expose a sidewall in the blank. The sidewall is then cut by the mill. By alternately repeating cutting of the groove and sidewall, the groove is milled deeper in the blank to form a pocket along which the sidewall extends. Since the sidewall is formed in steps, improved efficiency and accuracy is obtained therefor.

Abstract: A swept fan blade design for the low pressure compressor rotor or fan rotor stage of a ducted fan gas turbine engine has a leading edge swept forward near the hub up to a first radial height, then rearward up to second radial height and finally is swept forward again up to the blade tip. The aerodynamic effect is to produce a mid-height bias to the airflow enabling the tip region to be given increased twist and to possess increased resistance to foreign object damage. The design also provides a rear radial blade stacking axis to help reduce internal stresses due to centrifugal forces.

Abstract: A high aspect ratio, low solidity fan blade of an axial flow air cycle machine has a local tip center of gravity axis that has been shifted ahead of an elastic axis. The center of gravity axis can be shifted by clipping the tip of the blade at the trailing edge, by undercutting a rear portion of the pressure surface, or by sweeping forward leading and trailing edges of the blade by about five degrees (without clipping or undercutting the blade).

Abstract: A hollow airfoil (34) for a gas turbine engine having a leading edge (50), a trailing edge (52), a pressure side (54), and a suction side (56) includes a solid region of enhanced chordal thickness (78), and buttresses (82, 84, 86, 88). The present invention airfoil is preferentially strengthened so as to reduce the risk of airfoil fracture due to impact of airfoil with successive airfoils during an airfoil loss condition.

Abstract: An optimum axial distance is secured by varying the distance between a nozzle blade and a moving blade along the length of a nozzle blade (11). The nozzle blades (11) are curved so that a middle portion of each nozzle blade has a section (b2) dislocated in the flowing direction of a fluid which flows through the fluid passage relative to the root section (b3) and the tip section (b1) of the blade with respect to a circumferential direction and an axial direction. Outlet flow angle of the nozzle is varied with distance along the blade length to vary optimum axial distance properly.

Abstract: A method of modifying the vibration resonance characteristics of a workpiece includes the steps of determining a vibratory resonance condition frequency of the workpiece, determining the mode shape of said vibratory resonance condition, locating an area on the workpiece that includes a maximum curvature for the determined mode shape and vibratory resonance condition frequency; then laser shock peening the located area to create residual compressive stresses within the workpiece to shift the determined vibratory resonance condition frequency. A workpiece such as a gas turbine engine blade is also disclosed.

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 turbine blade (18) for a gas turbine engine includes a machining datum (60) that extends radially from a pocket (54) in the outer shroud (36) of the blade. The datum is spaced from a sidewall (64) of the pocket so that the datum is peripherally continuous irrespective of whether the blade is in a prefinished state or in a completely finished state. Because the datum's peripheral continuity survives the original manufacturing process, the datum is available for use in post-manufacturing inspection and repair operations.

Abstract: An airfoil is provided having a cross-sectional geometry which includes a first wall, a second wall disposed opposite the first wall, a leading edge, a trailing edge disposed opposite the leading edge, and a first cavity. The first cavity is disposed between the first and second walls, and the leading and trailing edges. The cross-sectional geometry extends between a first and a second end, and the airfoil is formed from discontinuously reinforced aluminum.

Abstract: A vane for a fan including a platform, a root and a blade. The platform has opposite outer and inner surfaces, and opposite upper and lower surfaces. The upper surface has a camber shape, and the lower surface has a concave shape corresponding to the camber shape of the upper surface. The root is fixed to the inner surface of the platform and adapted to be fixed to a hub of the fan. The blade is fixed to the outer surface of the platform at a position shifted toward the lower surface from a center line between the upper and lower surfaces.

Abstract: A ceramic blade for a hybrid type gas turbine blade has a dovetail portion, a platform portion formed on the dovetail portion and blade portions formed on the platform portion. The number of the blade portions formed on one platform portion is two or more. The upper surface of the platform portion is shaped into an arc-like form, and the dovetail portion is linearly formed in a tangential direction to a turbine rotation direction. By the utilization of this blade, the hybrid type gas turbine blade can inhibit the leakage of a gas and can be easily manufactured and is excellent in durability.

Abstract: A turbocharger compressor impeller structured as a combination elevated auxiliary blades fan surrounded by a curved inducer which provides a high gas volume low speed turbocharger compressor to deliver pressured gas into a combustion engine.

Abstract: A rotor blade pair for rotation around an axial centerline is provided which includes a platform, a first and a second airfoil, and a root having a first and second wall. The platform has an inner and an outer radial surface. The first and second airfoils extend out from the outer radial surface of the platform. The root walls extend out from said inner radial surface of the platform, and are integrally connected to one another, forming a hollow between the walls and the inner radial surface. The first wall is substantially aligned with the first airfoil and the second wall is substantially aligned with the second airfoil.

Abstract: Aerofoil members such as the stator vanes and rotor blades of an axial flow compressor are defined by a number of nominal sections. The leading edge of each nominal section of the blade is set so that an airflow passing therethrough is incident at the optimum angle. The optimum angle is chosen so that the maximum velocities of the airflow at the leading edge are equal. Once the optimum angle has been found for each section of the blade the sections are stacked to define a blade having a leading edge the aerodynamic profile of which ensures that the airflow is incident at the optimum angle along the length of the blade from the root to the tip. The efficiency of the axial flow compressor is improved.

Abstract: A heavy-duty gas turbine includes a compressor; a combustion liner; a turbine blade in a single stage or multi-stages; and a turbine nozzle provided in correspondence to the turbine blade. The turbine blade has a dovetail secured to a turbine disk and has an overall length of not less than 180 mm, and it is made of a single-crystal Ni-base alloy whose .gamma. phase is a single crystal. Operating gas temperature is not less than 1400.degree. C., and metal temperature of a first blade is not less than 1000.degree. C. under working stress.

Abstract: A fixed blade assembly for an axial flow turbine, in which each blade (23) is secured between a radially inner ring (21) and a radially outer ring (22). The trailing edge (25) of each blade is straight and leans at an angle .theta. away from a radial line (28) from the turbine axis through the inner end of that edge and in the circumferential direction that the suction side (24) of the blade faces. The blades (23) have this (negative) angle of lean in the range 5 degrees to 12 degrees.

Abstract: An axial flow compressor for use in a gas turbine or other industrial application having an improved efficiency is provided through modification of the blade shape and profile of the blades subject to a subsonic high velocity inlet flow so as to avoid shock losses in the blade rows, and improve the total efficiency. The curvature distribution on the suction surface of the stator blade rows or rotor blade rows of the axial flow compressor is adapted to have a local minimum in a region toward the leading edge of the blade, and then a local maximum. Thereby, an excessive increment in velocity on the suction surface from the leading edge to the position of the maximum velocity can be suppressed, and the occurrence of a shock wave can be avoided so as to minimize pressure loss in the blade rows, thus increasing the total efficiency of the axial compressor.

Abstract: A turbine device and a method of driving the turbine device are disclosed. The turbine device includes an admission channel, a turbine, and an injection channel. The turbine device may also include a regulator. The turbine is driven by injecting a primary fluid into the admission channel at a given velocity and simultaneously causing a secondary fluid to flow into the admission channel at a lower velocity. The primary fluid and the secondary fluid form a mixture in the admission channel, which flows toward the turbine. The velocity of the mixture is less than that of the primary fluid, while the mass flow of the mixture is approximately equal to the sum of the mass flows of the primary and secondary fluids. The regulator compares the rotational speed of the turbine to a target speed and regulates parameters associated with the turbine device if the rotational speed of the turbine and the target speed differ by more than a predetermined amount.

Abstract: In a steam turbine rotor wheel and bucket dovetail joint construction wherein the wheel dovetail includes a plurality of radially adjacent hooks interconnected by respective necks, each having an upper and lower fillet pair of substantially identical fillet, an improvement wherein each upper fillet has a compound radius.

Abstract: A rotor blade for a gas turbine engine includes a bowed surface on a tip region of the suction side thereof. The curvature of the bowed surface progressively increases toward the tip of the blade. The bowed surface results in a reduction of tip leakage through a tip clearance from the pressure side to the suction side of the blade and reduces mixing loss due to tip leakage.

Abstract: A heavy-duty gas turbine includes a compressor; a combustion liner; a turbine blade in a single stage or multi-stages; and a turbine nozzle provided in correspondence to the turbine blade. The turbine blade has a dovetail secured to a turbine disk and has an overall length of not less than 180 mm, and it is made of a single-crystal Ni-base alloy whose .gamma. phase is a single crystal. Operating gas temperature is not less than 1400.degree. C., and metal temperature of a first blade is not less than 1000.degree. C. under working stress.

Abstract: A steam turbine blade having an airfoil portion and root portion by which the blade is affixed to a rotor. The geometry of the blade airfoil is configured to minimize energy loss through the row of blades and reduce the weight of the airfoil. The airfoil has a leading edge and a trailing edge defining a chord therebetween. The chord is reduced linearly from the base of the airfoil to 50% of the airfoil height. However, the chord remains essentially constant from 50% of the airfoil height to the airfoil tip. The root is fir tree shaped and has four sets of tangs and grooves that are configured to minimize the stresses in the root.

Abstract: A method and apparatus for increasing blade fatigue strength in a turbine engine. The blades in the turbine engine are configured so as to reduce stress at the tip of the blades during operation of the turbine engine, thus increasing blade fatigue strength. This stress reduction helps to counteract the detrimental effect of abrasive tip coatings on blade fatigue strength. In one embodiment, the tip of the blade is chamfered in order to reduce the stress on the tip of the blade. An abrasive coating is then applied to the tip of the blade to assist the blade in seating into an abradable outer air seal. In another embodiment, an abrasive coating is applied to a center portion of the tip of the blade, with the periphery of the abrasive coating being set back from the opposing surfaces of the blade.

Abstract: A next-to-the-last stage steam turbine bucket has a profile according to the table set forth in the specification. The buckets also include continuous couplings at the bucket mid-point including nubs on each bucket projecting from opposite sides thereof in generally circumferentially extending directions. A sleeve is disposed between each pair of adjacent buckets and has open opposite ends for receiving the nubs of the adjacent buckets. The cross-sections of the nubs and open ends of the sleeve are generally complementary and non-circular to prevent sleeve rotation during turbine operation. The buckets have a continuous cover including cover elements having tenons projecting from opposite sides for reception in corresponding openings in the tips of adjacent buckets.

Abstract: A rotatable turbine frame includes annular outer and inner bands having a plurality of circumferentially spaced apart struts extending therebetween. An annular drive shaft is fixedly joined to the inner band for transmitting output torque therefrom, and the struts are backwardly radially inclined relative to the direction of rotation of the frame so that gas flow between the struts tends to straighten the inclined struts to effect a compressive load component therein.

Abstract: A gas turbine engine fan blade having an airfoil portion and a dovetail root portion is enhanced by having spaced apart supporting ribs extend into the dovetail root portion. The supporting ribs define a plurality of hollow cavities extending into the dovetail root portion of the fan blade. The extended cavities reduce the weight of each blade and also improve the integrity of each blade.

Abstract: A last-stage steam turbine bucket having a profile according to the charts set forth in the specification.

Abstract: A blade for a gas turbine has a leading edge portion which has an arc-shaped cross-section, and a maximum thickness portion of the blade is located within this arc. With this blade configuration, an abrupt acceleration of hot gas at the leading edge portion of the blade is restrained, so that the velocity of the hot gas at the blade surface can be made low. Therefore, the heat transfer coefficient from the hot gas to the blade wall is lowered, so that the amount of cooling air required to be passed through the interior of the blade can be reduced.

Abstract: A steam turbine blade is provided in which the geometry of the blade airfoil is configured to reduce the weight of the blade while maintaining adequate thermodynamic performance. The blade has a high camber angle to increase its stiffness and an airfoil portion having pressure and suction surfaces. The curvature of the portion of the pressure surface between the base and 75% blade height has a point of inflection in an area adjacent the leading edge so that the curvature switches sign. The curvature upstream of the point of inflection is concave, while the curvature downstream of the point of inflection is convex.

Abstract: A steam turbine blade is provided in which the geometry of the blade airfoil is configured to reduce the weight of the blade while maintaining adequate thermodynamic performance. The blade has a high camber angle to increase its stiffness and an airfoil portion having pressure and suction surfaces. The curvature of the portion of the pressure surface between 25% and 75% blade height has a point of inflection in an area adjacent the leading edge so that the curvature switches sign. The curvature upstream of the point of inflection is concave, while the curvature downstream of the point of inflection is convex.

Abstract: A fluid impeller blade with lowered stress and increased useful life has an edge extending from the hub of the blade and forming, in part, a boundary for axial fluid flow. The edge, at least at its extremity, is spaced axially into the blade, at an angle of about 0.5.degree. to about 20.degree. from the radial line through the edge at the hub of the blade, whereby the mass of blade material exerting centrifugal force on the edge at the blade hub during rotation of the impeller is reduced.

Abstract: An axial-flow turbine has at least one row of bowed guide vanes (7) and at least one row of rotor blades. The bowing of the guide vanes (7) over the vane height is selected at right angles to the chord and is directed towards the pressure side of the respectively adjacent guide vane in the peripheral direction. The guide vanes are tapered in their radial extent. Secondary losses, which occur due to the deflection of the boundary layers in the guide vanes, are reduced by this measure.

Abstract: Steam turbine nozzle blades and buckets are profiled in accordance with Tables I and II or multiples of the X,Y and R coordinates of the charts. The profiles of the nozzle blades impose a parabolic flow distribution at the throat and the buckets accommodate the incoming flow. The nozzle blades and buckets are over-cambered at their roots and tips to constrict flow from the roots and tips into the mid-regions of the blades and buckets. Improved aerodynamic efficiencies are obtained by directing the flow away from the end walls of the blades and buckets toward the efficient mid-region of the blades and buckets whereby flow velocity and angle leaving the nozzle has a non-linear distribution and the bucket velocity leaving angles are similarly non-linear.

Abstract: A turbine blade comprises an airfoil/shank having a shank region and an airfoil region extending upwardly therefrom and two dovetail/platforms bonded thereto. The first dovetail/platform has a first dovetail region bonded to a first side of the shank region of the airfoil/shank, with a first platform region extending outwardly from the first dovetail region. The second dovetail/platform has a second dovetail region bonded to a second side of the shank region of the airfoil/shank and a second platform region extending outwardly from the second dovetail region in a direction generally opposite to that of the first platform region.

Abstract: A regenerative toric pump in which undesirable noise generation and leakage through clearance gaps between the impeller and housing is minimized includes an impeller having vanes lying in general planes radiating from the impeller axis disposed at variable spacings from each other in a geometrically balanced pattern. Recesses in one of opposed side surfaces on the impeller and housing are arranged in a pattern such as to minimize leakage through the clearance gap between those surfaces from points in the pump chamber which are at different pressures.

Abstract: A last-stage steam turbine bucket having a profile according to Charts I-XXXXI of Table I.The bucket is divided into a number of profile sections each of which is a predetermined radial distance from the root section. Each point on a profile section at the predetermined radial distance from the root is defined by an X-Y coordinate and a radius R along the arc of a circle connecting adjacent points.

Abstract: A turbofan gas turbine engine fan stage having alternating relatively rugged, preferably hollow titanium, blades and staggered relatively less rugged, preferably composite, blades. The blades are staggered such that the leading edges of the rugged titanium blades are positioned forward of the leading edges of the less rugged composite blades. One embodiment provides that all fan blades have trailing edges at the same axial position while another embodiment provides that all fan blades are of the same size and shape.

Abstract: A high performance steam turbine vane having a novel airfoil shape. The pressure surface of the airfoil has a substantially straight region proximate the trailing edge that provides the airfoil with adequate thickness to prevent distortion during forging. The novel shape on the suction surface is such that a first region of the suction surface downstream of the gauging point is essentially straight. Traveling in the upstream direction, the first region is followed by a second region that has a large amount of curvature. The second region is followed by a third region that has a lower curvature than the second region. The third region is followed by a fourth region of constantly increasing curvature that increases to a curvature greater than the curvature of the second region. The fourth region is followed by a fifth region, ending at the leading edge, of essentially constant large curvature that includes the maximum curvature throughout the suction surface.

Abstract: A last-stage steam turbine bucket having an actual and theoretical profile according to Charts II-XVI. The last-stage turbine bucket has a theoretical profile according to Charts I-XVII.

Abstract: Freestanding rotor blades are mixed tuned by varying the dimensions of the blades at two locations. First, a profiled tip includes a machined-out strip, having a height which can be varied to achieve the desired tuning effect. The longer or deeper the profile, the higher the frequency will become. This tuning technique is used in conjunction with a broadened base section. Blades having a shorter profile tip are machined beginning at the base section, so as to remove enough mass from between the base section and a next upper section to provide the desired tuning effect. The lower frequency blade has a thinner base section and a shorter profile tip.

Abstract: A first hollow titanium MMC driveshaft portion is diffusion bonded to a third portion which in turn is fastened to a second superalloy portion via circumferentially positioned pins which are tightly fitted within aligned holed formed in the second and third driveshaft portions. The first, second and third portions are concentric with respect to the driveshaft centerline, and the pins are retained by means of a concentric retaining cylinder.

Abstract: A high performance steam turbine blade is provided in which the geometry of the blade airfoil is configured to minimize energy loss through the row of blades and control the radial distribution of the degree of reaction. The velocity of the steam over the blade surface is minimized to reduce friction losses, rapid deceleration of the steam velocity as it expands toward the trailing edge is avoided to prevent boundary layer separation and the reaction distribution is adjusted so that a relatively high reaction occurs at the base of the airfoil, thereby reducing secondary flow, and a relatively low reaction occurs at tip of the airfoil, thereby minimizing tip leakage. Specifically, the radial reaction distribution is such that in the hub region at least 20% of the stage pressure drop occurs in the blade row and in the tip region less than 50% of the stage pressure drop occurs in the blade row.

Abstract: A last-stage steam turbine bucket having a profile according to Charts II-XIII, XV-XXVII and XXIX of Table I.

Abstract: An apparatus of stationary blade for an axial flow turbine includes a diverging flow channel for flowing an elastic fluid and stationary blades which are fixed at the diverging flow channel and are curved in perpendicular direction to the flow direction of the elastic fluid. The stationary blades form the same tangential lean angles at the leading edge and the trailing edge of the stationary blade corresponding to the flow direction of the elastic fluid, and an axial flow turbine has the apparatus of stationary blades thereof. The stationary blade distributes the fluid in the flow channel uniformly and, consequently, improves the efficiency of the flow.

Abstract: A hollow core fan blade for a gas turbine engine, having a continuous leading edge, is fabricated using a four-sheet superplastic forming/diffusion bonding process which results in a cost-efficient and lightweight, yet strong, structure. The rotor blade is comprised of a face sheet which has a 180 degree bend therein so that the two face sheet ends are aligned. Bonded to opposing sides of the face sheet are first and second core sheets, between which is the hollow core. To fabricate the blade, a core sheet assembly is inserted inside the prepared face sheet, thereby forming a Ti-Pack (titanium pack) assembly having a plurality of pressure-tight cells. The Ti-Pack is inserted into a cavity within a die, after which the rotor blade, having predetermined design characteristics, is superplastically formed by heating the die and selectively pressurizing the plurality of cells.

Abstract: There is provided a rotor arrangement comprising primary blades 16 mounted on a support disc 15 and secondary blades 18 mounted on a support disc 17. The secondary blades are disposed between respective primary blades with some axial overlap. The set of secondary blades is movable relative to the set of primary blades so that each secondary blade can be moved between end positions abutting the associated pair of primary blades.

Abstract: A stationary blade of a steam turbine having a rotor and an inner cylinder for mounting the stationary blade in a row with plural identical stationary blades, comprising an airfoil having a leading edge, a trailing edge, a pressure-side concave surface and suction-side convex surface extending between the leading and trailing edges. A stagger angle being defined by as an angle of a chord between the leading and trailing edges to a longitudinal axis of the rotor; an outer ring for connecting a proximal end of the airfoil to the inner cylinder; an inner ring connected to a distal end of the airfoil; and a seal assembly carried by the inner ring and sealingly engaging the rotor; wherein the stagger angle ranges from about 42.degree. at the distal end of the airfoil to about 52.degree. at the proximal end.