Abstract: A gas turbine blade (10) which has a rounded trailing edge (18), is provided with a row of side by side arranged ceramic fibres (20) along the trailing edge (18). During operation of the turbine blade (10), the rounded shape of trailing edge (18) causes gasflows to break from the rounded edge before reaching the edge extremity. The presence of the fibres (20) prevent the formation of vortices in the gasflow, and thereby improve turbine efficiency.

Abstract: An aircraft gas turbine propulsion engine includes a compressor (10) and ancillary systems (cabin pressurizing pumps and other service) in a box 16. A fuel cell 18 powers electric motors (12) and (14) via a switch (20) for the purpose of either simultaneously driving the compressor (10) and ancillary systems, or the ancillary systems alone.

Abstract: A pipe replacement apparatus includes a housing having upper and lower walls that are sealed together about a pipe section to be replaced; on the interior surfaces of the walls, tracks are provided for clamps with engage the original pipe and also carry the replacement section; cutting devices are provided operable through one of the walls to sever the original pipe and the clamps are gear driven to be shiftable to move the cut portion out of line and move the replacement section into line; sealing collars are carried on the replacement pipe portion and manipulated to engage the pipe portions to be sealed thereto.

Abstract: A shaft and club head separation device for a golf club includes a base having padded shaft gripping members and a separation disk having at least two peripheral recesses into which the end of shaft is inserted; the gripping members are arranged to maintain the club shaft parallel to the axis of the separation disk; a plunger in slidably disposed in a bore in the base and an actuation device is provided to engage the plunger and move the plunger and disk away from the base so that, with the hosel of a club engaging a face of the disk, separation of the shaft and club head will occur.

Abstract: A combustor arrangement (8) for a gas turbine engine (3) comprising a combustion chamber, fuel nozzles(52a,52b), and a bled diffuser (7) located upstream of said combustion chamber to, in use, direct an airflow from an upstream compressor (6) into the combustor (8). The fuel nozzles (52a,52b) arranged in use to supply fuel into the combustion chamber where it is mixed and combusted with the airflow from the compressor (6). The bled diffuser (7) adapted to bleed off a portion of said airflow from a main airflow into the combustion chamber. At least one bleed duct (46) is connected to the bled diffuser (7) to return and direct the air bled from the diffuser (7) to a main gas flow through the engine (3) at a location downstream of the fuel supply means (52a,52b). The bled air from the diffuser (7) preferably providing cooling of a part of the gas turbine engine (3), for example part of the turbine (10) or combustor outlet vane (58), downstream of the combustor (8).

Abstract: A combustion chamber assembly (22) comprises a primary, a secondary and a tertiary fuel and air mixing ducts (54,78,98) to supply fuel and air to primary, secondary and tertiary combustion zones (40,42,44). Each of the primary, secondary and tertiary fuel and air mixing ducts (54,78,98) comprises a pair of axial flow swirlers (56,60,80,84,102,104) arranged coaxially to swirl the air in opposite directions and fuel injectors (62, 86,106) to supply fuel coaxially of the respective axial flow swirlers (56,60,80,84,102, 104). Valves (66,90) are provided to control the supply of air to the primary and the secondary fuel and air mixing ducts (54,78) respectively. A duct (122,116) is arranged to supply cooling air and dilution air to the combustion chamber (22). The amount of air supplied to the primary, secondary and tertiary fuel and air mixing ducts (54,78,98) and the duct (122,116) is measured.

Abstract: A gas turbine engine comprises a centrifugal compressor (4), an air diffuser (8), a heat exchanger (10), combustion apparatus (12), and first and second turbines (14,18). The combustion chamber assembly (22) comprises a primary, a secondary and a tertiary fuel and air mixing ducts (54,78,98). The compressor (4), diffuser (8), primary and secondary fuel and air mixing ducts (54,78) and turbines (14,18) all comprise means (6,8,16,20) for varying the mass flow area at their inlets such that in operation the amount of air mass through each component may be independently variable. Under part power conditions the mass flow is reduced and under full power conditions the mass flow is increased thereby maintaining a substantially constant gas cycle throughout the engine.

Abstract: A fuel injector assembly designed to withstand a predetermined impact load resulting from ingestion of a bird or other foreign object into the combustion stage of a gas turbine engine comprises a body member adapted to be mounted adjacent to and to extend into the combustion chamber and an abutment member disposed downstream of the injector assembly in the direction of air flow through the combustion stage, the arrangement being such that the body member moves into contact with the abutment member when subjected to an impact load in excess of the impact load it is designed to withstand.

Abstract: A component (24) for a gas turbine engine rotor assembly comprises a wall member (26) for at least partially bridging a space between adjacent rotor blades (18) and a root member (42) for insertion into a groove in the rotor disc (18) of the rotor assembly. The root member (42) includes bearing surfaces (46) for bearing against complementary surfaces (48) in the groove on rotation of the disc. The annulus filler (26) comprises a plurality of layers of carbon fiber reinforced plastics material, the layers being oriented substantially perpendicular to the bearing surfaces (46).

Abstract: A gas turbine engine core casing comprises axially adjacent portions (20,21) that are flanged to receive bolts (28) that attach the portions (20,21) to each other. One of the casing portions (20) is provided with flange (23) that is of L-shape cross-section to provide local enhancement of the rigidity of the casing portion (20). Such rigidity enhancement provides improved debris containment in the event of the failure of rotor components within the casing.

Abstract: A high rhenium-containing nickel superalloy article (10) has a multilayer coating (12) comprising a barrier coating (14) and an aluminide coating (16). The aluminide coating (16) is a corrosion and oxidation protective coating for the superalloy article (10). The barrier coating (14) comprises an alloy with a similar composition to the high rhenium-containing nickel superalloy article (10) but with less rhenium. The barrier coating (14) minimises diffusion of elements between the aluminide coating (16) and the superalloy article (10) to minimise the formation of topologically close packed phases at the interface between the superalloy article (10) and the multilayer coating (12). The barrier coating (12) preferably has some rhenium to minimise diffusion of rhenium from the superalloy article (10) to the barrier coating (14).

Abstract: Power generation equipment (10) comprises a compressor (11,12) into which finely dispersed water is directed. Air compressed by the compressor (11,12) is directed into combustion equipment (13) where it is mixed with steam and fuel before combustion takes place. The resultant combustion products pass through two turbines (14,15), the second of which (15) is a power turbine. The exhaust efflux from the power turbine (15) is then cooled to such an extent that its pressure falls to below atmospheric pressure. The carbon dioxide is removed from the efflux by means of a sequestration membrane (33) before the pressure of the efflux is raised to above atmospheric pressure by a further compressor (16) driven by the power turbine (15). The equipment (10) is very efficient and produces low levels of pollution.

Abstract: A catalytic combustion chamber is provided with at least two cataleptic combustion zones arranged in flow series, In a first mode of operation fuel is supplied from first fuel injectors, positioned upstream of the first catalytic combustion zone, into the catalytic combustion chamber and is burnt in the first catalytic combustion zone in order to preheat the subsequent catalytic combustion zones. In the second mode of operation the supply of fuel to the first fuel injectors is reduced and fuel is supplied from second fuel injectors positioned between the first catalytic combustion zone and the second catalytic combustion zone into the space between the first catalytic combustion zone and the second catalytic combustion zone. This prevents the first catalytic zone becoming overheated, and reduces the possibility of the second and third catalytic combustion zones becoming overheated and allows the optimum catalyst to be selected for the first catalytic combustion zone.

Abstract: An apparatus for detecting pressure in a hollow fan blade of a turbofan gas turbine engine comprises a magnetostriction transducer arranged in a sub chamber of the fan blade which is interconnected to a main chamber. A magnetic coil is arranged in the fan casing of the turbofan remote from the fan blade, and an alternating current is supplied from a supply to the magnetic coil to produce an alternating magnetic field. The alternating magnetic field causes the magnetostriction transducer to generate vibrations in the fan blade and the magnetostriction transducer. A magnetic search coil detects changes in the magnetic field, which corresponds to the vibrations, and a processor analyses the vibrations to determine if there has been a change in the resonant frequency of the vibrations which is indicative of a change in the pressure in the fan blade. The processor supplies a signal to a display or to an alarm. The change in the pressure is indicative that the fan blade is cracked and needs replacing.

Abstract: An aerofoil blade damper comprising an elongate member which is inserted within a core passage of the blade and extends within the blade with the damper retained therein at one end which is closest to the blade root with the remainder of the damper free to move relative to and within the passage; such movement generating friction which dissipates any vibration. The damper comprises a plate insert including at least two discrete substantially oppositely directed contact regions which are arranged, in use, to frictionally engage the passage. The discrete contact regions may be provided by protrusions in the form of contact pads or arms extending from the plate or be provided by the corrugations of the plate. A contact load on the contact regions is provided by the resilience of the damper, an interference fit of the damper and/or centrifugal loading of the damper and contact regions during operation.

Abstract: An axial flow turbomachine has at least one circumferential row of aerofoil members in which at least one of the two end walls (33) between successive blades (30) is given a non-axisymmetric profile to modify the boundary layer flow at the wall. In one form of the profile, a convex region (33) adjacent each member pressure surface (35) and a complementary concave region (34) adjacent each member suction surface (34) extend over at least a major part of the blade chord lengths to reduce the transverse pressure gradient and thereby reduce vortical energy losses. In another form of the profile, at least one end wall (33) has complementary convex and concave regions (50,51) extending through the zone of the trailing edges of the members (30) on the suction end pressure surface sides (34,35) respectively of each member, thereby to reduce over turning of the flow. Both forms of profiling can be employed in combination.

Abstract: A flexible coupling between drive and driven shaft members includes a diaphragm member having an offset base and outer periphery that is relatively more flexible than the radially inner base which is formed integrally with one of the shaft members or at least an intermediate coupling sleeve. A transition zone of the diaphragm extends through an arcuate curve to facilitate attachment between the shafts and one that will allow flexing of the diaphragm to accommodate dimensional changes in the shaft members as a result of cooling after assembly of the coupling.

Abstract: A high pressure fuel system includes tapped high pressure flows (22,24) with restrictors (44,46) to provide low pressure flows with which, via a valve (50) to enable either a low pressure flow to bias a main, high pressure fuel cut off valve (16) to close at a steady given rate, or a high pressure flow across a further valve (30), to close the valve (16) at a faster rate.

Abstract: A damping member (27) for positioning between the undersides of the platforms of adjacent turbine blades (20) and the peripheral surface of a disc (19) upon which the turbine blades (20) are mounted. The damping member (27) comprises first, second and third portions (28,31 and 32). The first portion (28) simultaneously engages the adjacent blade platforms (24) whereas the second and third portions (31,32) converge to engage each other at the disc (19) peripheral surface from a spaced apart relationship with each other at their attachment to the first portion (28). Frictional interaction between the second and third portions (31,32) and the disc (19) peripheral surface provides damping of turbine blade (20) vibration.