Showing papers on "Turbine blade published in 1971"

The Computation of Transonic Flow Through Two-Dimensional Gas Turbine Cascades

Abstract: Steady transonic flow through two-dimensional gas turbine cascades is efficiently predicted using a time-dependent formulation of the equations of motion An integral representation of the equations has been used in which subsonic and supersonic regions of the flow field receive identical treatment Mild shock structures are permitted to develop naturally without prior knowledge of their exact strength or position Although the solutions yield a complete definition of the flow field, the primary aim is to produce airfoil surface pressure distributions for the design of aerodynamically efficient turbine blade contours In order to demonstrate the accuracy of this method, computed airfoil pressure distributions have been compared to experimental resultsCopyright © 1971 by ASME

Local Heat Transfer Measurements on a Gas Turbine Blade

Abstract: This paper presents an experimental method for determining the variation of the local heat transfer coefficient around gas turbine blades. The method involves the accurate determination of the dist...

Turbine blade seal arrangement

Abstract: A seal arrangement for a turbine assembly including a turbine wheel, having a plurality of circumferentially arranged blades with adjacent edges of the blades being axially arranged above passages in the wheel assembly and retaining members arranged on each axial side of the wheel, a seal assembly between each adjacent pair of blades including an elongated seal member with means for maintaining the seal member in alignment with an axial recess formed by adjacent edges of the blade so that the seal member is urged radially outwardly into sealing engagement with the blades during rotation of the wheel and annular seal rings for maintaining sealing engagement between the retaining members and the plurality of blades.

Dual flow cooled turbine arrangement for gas turbine engines

Abstract: Turbine blading for a gas turbine engine includes a dual source of turbine cooling airflow wherein a portion of the high pressure compressor discharge airflow is directed to cool the most critical temperature areas of the turbine blade and a portion of the low pressure compressor interstage airflow is directed to cool the less critical temperature areas of the same turbine blade. The invention herein described was made in the course of or under a contract or subcontract thereunder (or grant) with the Department of the Navy.

Gas turbine temperature profiling structure

Abstract: An axial flow gas turbine having a structure including a fuel nozzle for providing any desired turbine inlet temperature profile according to the mechanical stress on the rotating turbine blades. The fuel nozzle may be a combined multiple fuel gas and liquid type or a single fuel type, an important feature of both being the provision of fuel injection asymmetrically into the combustion chamber to establish a desired ignited fuel temperature pattern therein which continues down to the turbine inlet. This asymmetrical fuel supply into the combustion chamber is obtained by providing a number or size of fuel jets according to the temperature desired; for example, to provide a temperature gradient decreasing in an inward turbine radial direction through the combustor to correspond to the desired turbine blade inlet temperature profile. The angular jet direction also may be varied for different parts of the fuel injection pattern to obtain a further spatial control of the fuel injection distribution into the combustor. These fuel distribution control variables may be used singly or in any combination, and the same or different ones may be used for the gas and liquid jets in multiple fuel nozzles. In addition, a further modification of the intake temperature profile may be obtained by providing relatively cool jets of gaseous fluid through spaced orifices in the radially inner sides of the transition passages which direct the high temperature working gas to the turbine blades.

Sideplate for turbine blade

Abstract: Sideplate structure for sealing a chamber between adjacent turbine blades in the periphery of a rotor disc is disclosed. The periphery of the disc is provided with side entry channels for receiving the root portions of rotor blades having radial cooling holes extending from the root portions through the tips of the blades. The channels provide an axially extending passageway communicating with the cooling holes whereby a cooling fluid flows through the passageway and into the holes to cool the blades. Sealing plates are provided for closing the downstream end of the passageways.

Seal ring for gas turbine

Abstract: A gas turbine blade tip ring seal, which surrounds the blade tips to minimize turbine gas leakage around the blades, that incorporates a thermal protection system that allows modulation of the seal leading edge temperature level and thermal response to minimize seal deterioration during turbine transient operating conditions.

Turbine blade damper

Abstract: A VIBRATION DAMPER FOR USE IN A TURBINE ROTOR OF A HIGH TEMPERATURE GAS TURBINE IS DISCLOSED. THE DAMPER WHICH IS X SHAPED IN OVERALL CROSS SECTION WITH A PAIR OF DAMPING LEGS, A PROTRUDING ASSEMBLY LUG, AND A PAIR OF POSITIONING LEGS IS LOCATED IN THE TURBINE ROTOR IN A CAVITY FORMED BETWEEN THE EXTENDED NECKS OF A PAIR OF ADJACENT TURBINE BLADES IN ONE DIRECTION AND END COVER PLATES IN THE OTHER DIRECTION.

Adjustable-blade turbine

Abstract: A gas turbine engine is shown in which the turbine nozzle vanes and turbine outlet guide vanes are simultaneously pivotal to obtain greater efficiencies over a broad range of gas stream energy levels.

Fiber blade attachment

Abstract: A fiber composite compressor blade having spanwise fiber elements cast within an appropriate composite material wherein at the root end of the blade the fiber elements are looped over transverse pins extending across a hollow root fitting. The root fitting is generally formed to dovetail into peripheral slots extending across a compressor disc. The exterior of each blade is enclosed by layers of crossply fiber cloth and the cavity within the root fitting is filled with an appropriate potting compound to respectively provide strength and rigidity to the blade.

Transpiration cooled turbine blade with metered coolant flow

Abstract: A turbine stator blade having a porous skin covering a strut member with channels for cooling air fed from the root, and a removable metering plate attached to the root and having orifices of various sizes for transmitting air to the blade channels, the metering plate being selected for having appropriate orifices for the cooling requirements of blades in differing temperature and pressure environments.

Guide vanes for supersonic turbine blades

Abstract: Guide vanes are installed between blades of the last stage of blading of a large steam turbine to form converging-diverging openings to minimize losses in this portion of the turbine. The guide vanes are fastened to adjacent blades through a shroud ring and through a lashing ring to connect the blades in such a manner as to reduce vibration and allow thermal expansion of the rotating blade assembly.

Low pressure end diffuser for axial flow elastic fluid turbines

Abstract: An improved low pressure end diffuser for axial flow elastic fluid turbines, such as steam turbines, is disclosed. A housing is provided on the outer fairing member of the diffuser, which housing defines a vented chamber. The outer fairing member is provided with openings communicating between the chamber and the annular area within the turbine casing adjacent the leading edges of the last stage turbine blades. Injection slots are formed in the outer fairing member to communicate between the chamber and the exhaust outlet downstream of the trailing edges of the last stage turbine blades. The injection slots are formed in such a manner that the fluid will flow from the chamber along a major portion of the inner surface of the outer fairing member whereby fluid boundary layer on the fairing surface will be accelerated to prevent separation of the boundary layer from the surface thereby resulting in improved diffuser performance.

Hollow turbine blade having diffusion bonded therein

Abstract: This invention relates to turbine blades having diffusion bonded pins or inserts in the holes which are formed in the turbine blades during casting.

Axial positioner and seal for cooled rotor blade

Abstract: An axial positioning and sealing means for axially positioning turbine blades with respect to a rotor disc and for sealing the downstream side of cooling fluid flow passageways below the root portions of the blades is disclosed. The periphery of the rotor disc is provided with axial channels for receiving the root portions of turbine blades having radial cooling bores extending from the root portions through the tips of the blades. The channels are formed to provide an axially extending passageway communicating with the cooling bores whereby a cooling fluid flows through the passageway and into the bores to cool the plates. Sideplates means are provided for retaining the blades to the rotor disc. Pads or protrusions are provided on each steeple of the rotor disc, with the pads being positioned and shaped such that the side faces of two adjacent pads define a radially extending, radially outwardly tapering channel over each of the passageways. Each positioning and sealing plate has tapered sides complementary in shape to the tapered channel whereby when said rotor disc is rotated at high speeds, the centrifugal force acting on said plate will urge the plate radially outwardly of the rotor disc to thereby wedge the plate against the pad and the root portion of the turbine blades. Locking means are provided for securing each plate against unlimited radial inward movement with respect to the rotor disc.

Turbine cooling control valve

Abstract: The stator of a gas turbine includes an arrangement for controlling flow of cooling air to the vanes of the second stage turbine nozzle from the engine compressor. Control is by a valve defined by two rings extending around the exterior of the turbine nozzle which have different coefficients of thermal expansion so that the gap between them varies in accordance primarily with the temperature of the cooling air which tends to increase with the power output of the engine and thus with the need for cooling air.

Vibration dampening for long twisted turbine blades

Abstract: Lashing long twisted turbine blades in groups with a small Vshaped gap between the groups, which tends to close due to the untwisting of the blades, when they are subjected to high centrifugal forces, provides a substantially continuous lashing ring with good vibration dampening characteristics, which is sufficiently flexible to allow for thermal expansion and contraction, and cooperates with the vibration reducing characteristic of joining the blades into groups to minimize stress vibrations in such blades.

Blade tip closure

Abstract: A hollow laminated sheet metal transpiration-cooled porous turbine blade has its free end closed to prevent escape of cooling fluid from the end of the blade. This closure is made by separating the inner layer of the blade wall from the other layers and bending it inward from each face of the blade so that the inner layers from each face of the blade meet approximately along the mean camber line of the blade, where they are welded together to provide the closure. The closure is spaced from the blade tip, located within the cooled blade walls.

Articulated sleeve for turbine bucket lashing

Abstract: A BLADE LASHING DEVICE FOR A TURBINE WHEEL HAVING CLOSELY ADJACENT BLADES FORMED WITH A LUG ON EACH BLADE FACE. A SLEEVE IS INTERPOSED BETWEEN LUGS ON OPPOSING BLADE FACES TO CONNECT TURBINES BLADES. THE SLEEVE IS FORMED WITH CLEARANCES SO THAT IT WILL FORM AN ARTICULATED JOINT BETWEEN INTERCONNECTED TURBINE BLADES.

Gas turbine blade arrangement

Abstract: A gas turbine is formed with a first and a second stage, the first stage contains wheels on which turbine blades are concentrically arranged around compressor blades and the second stage contains wheels on which turbine blades are concentrically arranged around webs. A combustion chamber is arranged to receive air after it has been compressed by the compressor blades in the first stage and has been guided but not further compressed by the webs in the second stage. After the compressed air is heated in the combustion chamber it passes over the turbine blades in the second stage and then over the turbine blades in the first stage. The turbine blades in the first stage are hollow and air is withdrawn from the compressor side in the first stage and passed into the interior of the turbine blades to cool them.

Design and Experimental Evaluation of a High-Temperature Radial Turbine

Abstract: : The report describes the design, fabrication and test of a radial turbine designed to produce 219.6 Btu/lb stage work at 87.5% efficiency, with a 5:1 stage pressure ratio. Turbine inlet gas conditions at design point were 257. 5 psia and 2300F. The resulting turbine configuration consisted of an air- cooled, 12-bladed rotor designed for 67,000 rpm, and a 20-vaned air-cooled nozzle section of a reflex-type (supersonic) design. Both parts were designed as IN100 (PWA 658) investment castings. As part of the preliminary design effort, a fabrication study was conducted to evaluate feasible methods of casting the turbine nozzle and rotor. Results showed that the nozzle section could be cast as an integral assembly, but fabrication of the rotor as an integral casting was much more difficult. Bicasting was evaluated as an alternate method of fabricating the rotor, and results showed substantial advantages for the bicasting technique. However, neither method could produce designed rotor properties, and testing was conducted with structurally limited rotors. A test rig was designed and fabricated by the contractor. The test rig consisted of a supercharged gas generator, which had the capability of controlling the turbine load by varying the compressor flow rate. Burner testing preceded turbine testing.

Blade or vane for a gas turbine engine

Abstract: The invention relates to a blade or vane for a gas turbine engine which comprises a carbon fibre reinforced matrix material which is further strengthened by strong ductile metal wires whose modulus of elasticity is similar to that of the carbon fibres.

Shroud section joined to turbine blade - by electron beam welding using stamped out shim or clip

Abstract: A butt joint is employed and the blade material at least adjacent to the joint and at least on the side facing the electron beam is reinforced by a piece of identical material metal, which takes up excess beam energy, and through which the beam passes welding it to the urea next to the joint surface. The piece of metal used consists of a shim with a cut-out having the bucket cross section. The method ensures a good joint between the shrouding and the bucket section of the blade units for turbines.

Device for attaching turbine blades to a rotor

Abstract: An annular array of circumferentially spaced rotatable turbine blades is securely fastened to a rotor by a plurality of pins having an hourglass shaped cross section; the pins fit into a plurality of axially disposed, registering grooves in the blades and rotor and provide multiple blade roots and reduce the probability of a blade flying off the rotor as fatigue cracks must initiate and propagate through a plurality of the roots to have such a failure.