Abstract: The spectral content of a transient stress wave event occuring on a plate-like structure is determined by locating a transducer on the surface of the structure at a position remote from the location of the transient stress wave event. Lamb waves resulting from and characteristic of the transient stress wave event propogate through the plate-like structure in a dispersive manner to be intercepted by the transducer. Analysis of the transducer output enables the spectral content of the transient stress wave event to be realized.

Abstract: A hollow aerofoil cross-section aircraft propeller blade is provided with a tow of KEVLAR fibres internally thereof. The fibres which are anchored to the blade root and tip portions are arranged to be passive under all normal operating conditions of the blade. However in the event of a structural failure of the blade between its root and tip portions the fibres serve to contain the blade tip together with any portion of the blade attached thereto. The invention is additionally applicable to prop-fan and fan blades for gas turbine engines.

Abstract: The blade has an aerofoil body 10 having a leading edge surface 11 which is cooled by air passing through a helical first passage 12 having first portions passing close to said leading edge surface and alternating with second portions passing through a more nearly central part of the blade section remote from said leading edge. A spanwise but straight second passage 15 extends through the blade in a position within the helical passage and closer to the second than the first portions thereof. Heat abstracted from the leading edge by air flow in said first portions is transferred by the flow to the second portions and from there through the blade material to the flow in the straight passage.

Abstract: A bladed rotor assembly for a turbomachine includes a hub having a circumferential blade retaining groove and a plurality of blades which have rhombic shaped platforms. The blade retaining groove is continuous and the blades are loaded into the groove by inserting them with the generally axially directed faces of the platforms directed circumferentially. Each blade is then turned to its final operating position when the axially directed faces are axially directed and then further turned to reduce the amount of room the blade platform takes up in the circumferential direction to allow enough room for all the blades to be loaded. After all the blades are in the groove they are all turned finally to their operating positions and locked in place. A method for assembling a turbomachine rotor assembly is disclosed, also.

Abstract: The invention is particularly concerned with the control of blade tip clearance in high pressure compressors of gas turbine engines. The compressor comprises a rotor assembly having radially extending rotor blades and a stator assembly. The stator assembly comprises an inner casing and a cylindrical wall member spaced radially from the inner casing to form a chamber. The axial ends of the cylindrical wall member seal with and are moveable radially with respect to the inner casing. The casing has radially inner and radially outer stops to limit radial movement of the cylindrical wall member. The cylindrical wall member carries a ring of shroud segments which are spaced radially from the rotor blades by a clearance.

Abstract: A thrust reverser for a gas turbine engine of the type to power an aircraft, the thrust reverser comprising a first pair of half round doors nesting in a second pair of half round doors and both being pivoted on an axis extending diametrically of a jet pipe for the gas turbine engine. Each of the first pair of half round doors having a downstream edge which has a scarfed portion thereon and each of said second pair of half round doors having a downstream edge which is fully scarfed, the scarfed portion of said first pair of half round doors and the scarfed edges of said second pair of half round doors abutting when said doors are in a thrust reversing position with said second pair of half round doors preventing axial spillage of efflux. When said first pair and said second pair of half round doors are in non-thrust reversing position, said first pair of half round doors forming a greater portion of a nozzle throat and said scarfed portions of the downstream edges thereof reducing side spillage of efflux.

Abstract: A method of moulding composite materials in which prepregs of resin impregnated filaments are interposed between a shaped partially contained polytetrafluoroethylene body and a die. The assembly is heated to cause the polytetrafluoroethylene body to thermally expand and compress the prepregs against the die so that the prepregs are moulded to the shape of the die and the resin in the prepregs cures.

Abstract: A composite material comprising reinforcing carbon fibres in an epoxy resin matrix is manufactured by stacking alternate layers of resin impregnated fibre and polyester release film, and compressing the stack between shaped formers at a temperature below the resin cure temperature. The layers are than separated and the release film discarded. After re-stacking, the fibre layers are autoclave moulded on an appropriately shaped former at a temperature at which the layers consolidate and the resin cures.

Abstract: A vectorable nozzle (17) comprises a fixed first duct (21), a rotatable second duct (22) scarfed at its downstream end, and a rotatable third duct (23) scarfed at its upstream end. The second and third ducts (22, 23) are mounted in bearings (24,26), respectively, and an epicyclic gear train (30) is provided to rotate them in opposite directions while maintaining the direction of gas efflux issuing from the nozzle (17) in a single plane.

Abstract: A nozzle structure for a gas turbine power plant for a V.T.O.L. or S.T.O.L. type aircraft includes a main engine and at least one auxiliary engine; the efflux from the at least one auxiliary engine may be directed downwardly or alternatively horizontally through a convergent/divergent nozzle which may be closed off when the at least one auxiliary engine is inoperative.

Abstract: A method of manufacturing a brush seal is disclosed which comprises producing a plurality of tufts of brush material and securing the tufts between two side plates to form a seal. Each tuft is produced by forming a tape which comprises one or more layers of parallel contiguous lengths of filaments adhesively bonded together to form a self supporting body. The tape is cut into discrete portions and these constitute tufts of bonded filaments. A plurality of tufts are assembled closely adjacent to each other in a support member with a part of each tuft projecting from the support member, and the tufts are subsequently secured in the support member. Finally the adhesive is removed from at least that part of each tuft which projects from the support member.

Abstract: A method is described for the manufacture of articles in defect-immunized materials in which the defects are eliminated, or broken up and oriented in such manner as to minimize their harmful effects on the article.Referring to FIG. 1e, a rotor disc for a gas turbine engine is formed to an approximate shape by stacking together "sticks" 5 of material in an evacuated container and bonding them together by a hot isostatic pressing process. The "sticks" 5 are produced by extruding a starting body, for example, of powder material, to produce an elongation of up to twenty times and then cutting them to the appropriate length. By this means any non-metallic inclusions in the powder are broken up, inspection of the sticks and rejection of defective ones becomes easier, and the sticks can be oriented so that the effects of any remaining defects can be minimized.

Abstract: A rotor or stator blade for an axial flow compressor has one or both of its end portions designed to minimize the interference between the annulus boundary layer and the blade. The basis of the invention is that the aerofoil is designed so that each section of the aerofoil through the boundary layer performs a constant amount of work per unit height.

Abstract: An axial flow turbine comprises an annular array of rotatable turbine blades which are surrounded by a shroud ring. An open honeycomb structure is attached to the shroud ring, the open cells of which honeycomb structure contain an abradable material. The honeycomb structure is totally covered by an impervious coating which contains a ceramic material. The impervious coating resists oxidation and/or erosion of the abradable material.

Abstract: A gas turbine engine has an axial flow compressor, the inner wall of which is provided with a lining adapted to capture and retain any small hard foreign objects impinging upon it. The lining consists of an aluminum alloy sheet having a silicone rubber coating bonded to one side and a stainless steel honeycomb structure bonded to the other side. The honeycomb structure is bonded in turn to the inner wall of the compressor casing. The silicone rubber coating has a plurality of finely divided hollow silica microspheres dispersed within it.

Abstract: A hollow prop-fan aerofoil blade is covered by a woven containment sheath which is bonded to the aerofoil surface of the blade and operationally anchored to the blade root. The sheath is arranged to be of sufficient strength to contain any of the aerofoil cross-section portion of the blade in the event of any structural failure thereof.

Abstract: A gas turbine engine in which a rotating component (such as a turbine shaft or disc) is cooled by directing preswirled air at it by flowing the air along a passageway defined by the rotating component and a static wall. To reduce the viscous drag on the swirling air due to the presence of the static wall, the static wall is transpired to cause the boundary layer of air adjacent the wall to move in the same direction as the swirling air. The wall may be transpired by sucking or blowing air through air passages through the wall.

Abstract: A shroud for a gas turbine engine rotor stage is divided into a number of arcuate segments each of which is a trapezium. The segments are urged together by compression springs which attempts to make the assembly lengthen peripherally, by virtue of the chamfered ends 30 of the segments riding up each other. The completely circular shape of the shroud prevents that movement but as a result, a hoop load is generated which tends to move the segments radially outwards, thus achieving the same effect, along with a consequent increase in diameter. A cam ring and pins are used to control and reverse the movement, as desired. The shroud diameter can thus be altered to suit rotor expansion and contraction.

Abstract: An integral rocket and ramjet engine (16) comprises, in flow series, an intake duct (18) for aerodynamically compressing air, a port cover (30), a combustion chamber (20) and a propelling nozzle (22). The port cover (30) in a first position prevents air from entering the combustion chamber (20) and enables a rocket charge (24) to be burnt in the combustion chamber (20) thereby accelerating the engine to sufficient velocity for ramjet operation. The port cover (30) is movable axially to a second position when the rocket charge (24) is spent to allow compressed air into the combustion chamber (20) and the engine (16) to operate as a ramjet. In the second position the port cover (30) provides a quiet zone (52) which is substantially shielded from the main flow of air into the combustion chamber. Combustion equipment such as primary and main fuel manifolds (42,46) and flame gutters (50) are also movable with the port cover (30).

Abstract: A hollow propfan aerofoil blade is provided with grooves which extend between its root and tip portions. The grooves contain filaments which are enclosed within a resin matrix material and are anchored to the root and tip portions. The filaments are arranged to be of sufficient strength to contain any of aerofoil portion of the blade in the event of any structural failure thereof.