George Pask has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).
Abstract: A nozzle guide vane for a gas turbine engine comprises a hollow ceramic vane member which includes an aerofoil portion and ceramic platforms. The ceramic vane member is supported by engagement of its interior surface with external supporting surfaces on a metallic load-bearing spine adapted to be supported from fixed structure of the engine.
Abstract: A nozzle guide vane assembly for a gas turbine engine comprises inner and outer porous sheet metal platform rings and a plurality of aerofoils extending between the rings, each of the aerofoil portions having mounting means at its inner and outer extremities by which the aerofoils, and thus the rings, are supported from fixed engine structure.
Abstract: The casings of gas turbine engines are subject to loadings which distort the casing and adversely affect the running clearance between rotating and static components contained within the casings. The invention provides an engine having inner and outer housings the inner housing containing one of the engine compressors. The inner casing is mounted on a load transmitting and isolating structure which allows thrust and torque to be transmitted to the outer casing, but isolates the inner casing from distortions of the outer casing. The load transmitting and isolating structure includes a torque reaction and mounting means at the upstream end of the inner casing, a mounting means and a thrust reacting means mounted between the ends of the inner casing. The mounting means are adjacent the bearings of the rotor shaft of the compressor and are not affected to any appreciable extent by outer casing distortion.
Abstract: To enable shroud segments in a gas turbine rotor blade stage to operate at high temperatures with an adequate margin of safety, the shroud segments are mounted and cooled such that they are thrust outwards against seatings on surrounding high strength support structure by the gas pressure in the turbine passage, the shroud segments being provided with strengthening ribs or the like so that the outward thrust is transferred to the support structure through a plurality of load paths distributed over the outer sides of the shroud segments as necessary to avoid overstressing.
Abstract: A load distribution member for the turbine of a gas turbine engine comprises an annular diaphragm between regions P.sub.1 and P.sub.2 of high and low gas pressure respectively which is in the form of a part of a complete thin shelled toroid having a circular cross-section area of revolution. The radially outer edge of the annular diaphragm has a ring member attached thereto which is attached in turn to the undersides of an annular array of stator aerofoil vanes. The radially inner edge of the annular diaphragm has a ring member attached thereto which carries one element of an annular gas seal. The other element of the seal is attached to a rotary part of the turbine. The configuration of the part-toroid shape of the diaphragm is selected so that the axial forces exerted by the diaphragm upon the radially inner ring member are zero, thereby minimizing the degree of axial deflection of the seal element.
Abstract: A casing structure for a gas turbine engine comprises an outer load-bearing casing within which is supported an inner casing which forms the static structure of a compressor. Bearing panels carried from the outer casing carry the rotor of the compressor and its associated turbine. In order to mitigate the undesirable consequences of the bending under load of the outer casing, the inner casing is mounted at a forward section by forward mounting means which maintain it concentric with the rotor axis and at a rearward section by rearward mounting means which maintain it parallel with a section of the outer casing which retains its parallelism with the rotor axis even when the casing bends. The forward mounting illustrated is a dogged engagement while the rearward mounting utilizes a parallel series of links as a parallel motion linkage.
Abstract: An aerofoil blade/disc assembly suitable for the compressor or turbine of a gas turbine engine comprising a rotor disc which has a circumferential, part-circular cross-section channel provided in its periphery for the reception of the correspondingly part-circular cross-section roots of an annular array of aerofoil blades. The arrangement is such that upon disc rotation, the aerofoil blades pivot with respect thereto until their platforms sealingly engage a circumferential radially extending flange provided on the disc periphery. A seal is provided between the aerofoil blades and the disc which reduces the leakage of gases between the aerofoil blade roots and their retaining channel.
Abstract: In order to damp possible vibration of a rotor blade for a gas turbine engine, the blade has a hollow tip portion with an internal surface extending across the direction of the centrifugal field on the blade. A weight comprising e.g. Silicon Nitride or Carbide is caused by centrifugal force to bear on this surface. Relative motion and hence friction between the weight and the surface serves to damp vibration of the blade.
December 13, 1982
Date of Patent:
November 27, 1984
George Pask, John H. R. Sadler, John F. Coplin
Abstract: A shroud assembly is pneumatically operated to adjust the clearance between the shroud segments and the tips of the rotor blades of an associated rotor. The assembly comprises a casing to which is secured a wall member which defines a chamber whose pressure may be varied by a valve. The wall member carries shroud segments some of which define the static wall with which the rotor blades co-operate. In order to allow proper movement of the shroud segments the support means which carry the segments from the wall are arranged to be small in axial extent compared with that of the wall.
Abstract: The invention relates to fuel burners and the combination of fuel burners and annular combustion equipment particularly for use in gas turbine engines.The present invention seeks to provide a fuel burner that requires a relatively small access hole in the engine casing in order to reduce stress concentrations in the casing.The fuel burner comprises a fuel feed arm and a fuel injector. The fuel feed arm and the fuel injector are cylindrical, and are arranged coaxially in end-to-end abutting relationship. The fuel feed arm has internal passages for the supply of fuel to the fuel injector, and the fuel injector has an air duct passing normally through the fuel injector for the flow of air into a cooperating combustion chamber. The fuel injector supplies fuel into the air duct to mix with the air flowing therethrough.The invention also provides combustion chamber equipment for use in conjunction with the fuel burners.
Abstract: A blade assembly for use in an axial flow compressor or turbine. The blade assembly includes a disc provided with a circumferential blade root retaining channel in its periphery and an annular array of aerofoil blades having roots located in the channel so that gaps are defined between adjacent roots. Each root is provided with two circumferentially extending sealing skirts which are arranged so that they cooperate with the blade roots to define an annular sealing surface which is in operation in sealing engagement with an annular portion of the internal surface of the root retaining channel.
Abstract: A cooled wall structure for a gas turbine engine comprises a perforated and an inner wall in which the walls are capable of relative movement to cope with the thermal strains experienced by the combustion chamber during operation of the engine. The inner wall comprises a number of wall elements attached to the outer wall in the manner of overlapping tiles. Each wall element is immovably secured to the outer wall at the mid-point of its downstream end and the sides of each wall element are movably attached to the outer wall adjacent the sides of the downstream end of the wall element.The upstream end of each wall element is located between the outer wall and an adjacent flow in either an upstream or a downstream direction between the walls.The wall elements can have a plurality of raised lands to increase the surface area of the elements and to protect the incoming cooling air against the cross-flow of cooling air already flowing in the wall structure.
Abstract: An annular seal for a gas turbine engine comprises a stacked plurality of Belleville washers with at least some of their abutting peripheries held together by U-section rings. If only those peripheries pointing toward a lower pressure area are retained together, gas pressure will hold the other peripheries together. Because the U-section rings do not prevent the washers from changing the angle of their section, relatively large axial motion between the sealed members may be tolerated.
Abstract: An aerofoil member such as a blade or vane for a gas turbine engine comprises a hollow aerofoil with a plug of porous material filling a region of the hollow interior. The porous material forms at least part of the external surface of the aerofoil so that cooling fluid can flow from the interior of the aerofoil through the plug to the external surface of the aerofoil. The plug is preferably at the leading or trailing edge of the aerofoil. Methods of making the member include investment casting using a porous ceramic core to form the porous plug, or using sintered particulate material to form the porous plug.
December 16, 1981
Date of Patent:
December 27, 1983
John H. R. Sadler, George Pask, Wilfred H. Wilkinson
Abstract: A seal for use between the turbine rotor of a gas turbine engine and associated static structure comprises a ring of low-friction material which cooperates with the surface of the rotor to form an air bearing. The ring carries a sealing member which cooperates with a surface of the rotor to form a seal. The air bearing ensures that the sealing member is maintained at a constant spacing from the rotor, and this spacing may thus be maintained at a low value.
Abstract: A cooled shroud for a gas turbine engine comprises an annular metallic supporting member having holes therethrough for the flow of cooling air and a layered coating on its inner face. The layered coating comprises a first layer of porous material through which the cooling air may permeate, and a second layer of impermeable ceramic covering all but selected areas of the surface of the first layer. In this way the cooling air is largely constrained to flow along the porous material to provide good cooling with relatively low air flow.
Abstract: A device for controlling the clearance between the blade tips of a gas turbine engine turbine rotor and associated segmented shrouds comprises a control ring which is secured to the engine casing by a plurality of dowels. The control ring is supplied with cooling fluid to its hollow interior and is covered with a thermally insulating layer, whereby the rate of movement of the shroud segments can be maintained substantially equal to that of the adjacent turbine rotor.
February 22, 1980
Date of Patent:
March 2, 1982
Thomas Steel, Noel H. Hooke, George Pask
Abstract: A stator vane assembly for a gas turbine engine comprises a plurality of aerofoil sectioned vanes supported from a circumferentially extending support member or casing via a circumferentially extending platform or platforms. Each vane is attached at one end to one said platform which is carried from the supporting member, each platform being relatively flexible so that deflections of its respective aerofoil sectioned vane cause it to deform. The mounting between the platform and the supporting member is such as to damp motion of the platform and thus of the aerofoil sectioned vane; in some instances resilient material may be interposed between the platform and supporting member to provide the damping.
Abstract: A bladed rotor for a gas turbine engine comprises a rotor disc having a plurality of blade retaining slots in its periphery and a rotor blade mounted in each slot, and sealing means between the rotor and the adjacent static structure comprising an annular projection from adjacent the disc periphery adapted to coact with an annular feature on the static structure. The annular projection is made up of a plurality of abutting segments, alternate ones of which project from the disc and one of the rotor blades respectively.
Abstract: Combustion equipment for a gas turbine engine comprises an annular flame tube having a circumferential array of upstream projecting pots or vortex generators at its upstream end, an annular diffusing passageway located upstream of the pots and terminating adjacent the upstream ends of the pots, and a plurality of tubular ducts, one extending from the annular diffusing passageway to the upstream end of each pot.