Abstract: A rotor (30) comprises a disc (50) for rotating about an axis, supporting a plurality of radially extending blades (40) in slots (52) therein. At least one retaining plate (60), having a radially outer edge (62), is received by a channel (48, 58) formed by the disc (50) and at least one blade (40). The blade (40) and disc (50) are contiguous in the channel (48, 58) at first (55) and second (57) distinct parts of the disc (50), and the radially outer edge (62) of the retaining plate (60) has at least one indentation (64) aligned with a first part (55) of the disc (50).

Abstract: A gas turbine aerofoil with at least one internal cavity for conveying cooling fluid through the aerofoil has at least one internal cavity adjacent the aerofoil leading edge, at least one impingement cooling passage extending from, and in flow communication with the cavity, and at least one film cooling hole extending from, and in flow communication with, the cooling passage. The impingement passage extends into the aerofoil wall in the region of the leading edge and opens into the film cooling hole. The cooling hole extends through the aerofoil wall to an external surface of the aerofoil at a position downstream of the leading edge such that, in use, cooling fluid from the cavity exits the cooling passage into the cooling hole as an impingement jet against the internal surface of the cooling hole in the region of the leading edge.

Abstract: A casing arrangement (20) for surrounding a rotary component (16) of a gas turbine engine (10) especially a turbine is disclosed. The arrangement (20) comprises an annular casing member (22A) surrounding a bladed rotary section, such as a turbine stage. An array of gas path liner segments are suspended from the casing at a nominal clearance gap form the tips of the rotary blades. Running clearance is determined by expansion and contraction of the casing which, in turn, is controlled by a supply of heat transfer fluid passed through heat transfer holes in the casing. The casing arrangement described includes a passive pressure bleed effective during engine deceleration to reduce a pressure differential driving the fluid flow in order to avoid too rapid contraction of the casing as a result of overcooling.

Abstract: A stall detection and recovery system (10) for a gas turbine engine (1) comprises surge detection means (50 for detecting an engine surge and providing a surge detection signal (154) for indicating said engine surge. The system also includes stall detection means to provide a stall detection signal (48) relating to a condition of gas flowing through the engine (1). The system (10) also includes stall recovery means (58) to receive the surge detection signal (154) and the stall detection signal. The stall recovery means (58) is arranged to control combustor operating means, whereby when the stall detection signal (48) indicates an engine stall, the stall recovery means controls the combustor operating means to effect recover from the stall.

Abstract: A laminar flow nacelle for an aircraft engine (10) has an outer member (26) defining an aerodynamic shape. The nacelle (24) has an inner member (28) defining a chamber (30) with the outer member (26) of the nacelle (24). The outer member (26) of the nacelle (24) has a porous region (32) at a first region (34) of the outer member (26) and the porous region (32) allows a flow of fluid into the chamber (30). A duct (36) connects the chamber (30) to an aperture (38) in the outer member (26) at a second region (40) of the outer member (26) downstream of the first region (34). In operation the static pressure at the first region (34) is greater than the static pressure at the second region (40) such that the boundary layer of the fluid flows through the porous region (32) at the first region (34) through the duct (36) to the aperture (38) at the second region (40). The first region (34) extends between 10% and 20% of the chord length of the nacelle (24) from the highlight (42) of the nacelle (24).

Abstract: A heat transfer arrangement comprises a target member (28 or 30), and an impingement member (32). The impingement member (32) defines a fluid path (38 or 40) there through to direct fluid onto the target member. The arrangement includes a fluid directing formation (42) on the impingement member (32). The fluid path (38 or 40) extends through the fluid directory formation (42) such that the fluid path directs fluid to exit there from at an exit angle that is substantially orthogonal to the fluid directing formation.

Abstract: A component such as a turbine blade of a gas turbine engine includes a film cooling arrangement which is optimised by a process in which a component is manufactured having a film cooling arrangement of an initial design and evaluation of the performance of the film cooling arrangement is conducted, for example on the basis of the blowing rate of cooling holes present in the initial design. The configuration of a cooling hole of the initial design is subsequently modified in order to improve the performance of the film cooling arrangement and a component is manufactured in accordance with the modified design. By way of example, the initial design may provide a single cooling hole 128 and the modification may comprise the provision of additional cooling holes 126, 130, each intersecting the initial cooling hole 128 adjacent its inlet end which opens into a source of cooling fluid 120.

Abstract: A combustion chamber inner casing (24) has a flange (28) on its downstream end. A retaining seal ring (40) is bolted to the flange (28) and traps a flanged seal ring (46) against the flange (28) in a manner that allows the seal ring (46) to move by expansion, relative to seal ring (40).

Abstract: An arrangement 10 for providing thermal energy to a portion 16 of a component 14 of a gas turbine engine while the component 14 is forming, wherein the component 14 is substantially elongate in a first direction 28 and the portion 16 extends from the component 14 in a second direction 38, substantially perpendicular to the first direction 28, and comprises a first surface area component 32 oriented in the first direction 28. The arrangement 10 further comprising an elongate member 18, connected to the portion 16 for heating the portion 16, wherein the elongate member 18 comprises a second surface area component 36, oriented in the second direction 38, wherein the second surface area component 36 of the elongate member 18 is greater than the first surface area component 32 of the portion 16.

Abstract: Turbine blades used in jet engines generally require cooling in order to achieve desired engine performance. Cooling of blade tip areas is difficult due to the limited space available such that thickening tip areas in order to allow passages and holes to be incorporated adds to weight and therefore stressing along with causing additional manufacturing costs. The present cooling arrangement includes coolant release passages 5 which present coolant flow to coolant entrainment elements or fins 2 such that the coolant flow is entrained close to the blade tip surface 12. Thus, turbulent air flow caused by adjacent shrouds and edges are inhibited from diluting the coolant flow and therefore reducing thermal efficiency.

Abstract: An investment mould process for forming a cast member is described. The process comprises forming a precursor casting (12) from a master mould (10). The master mould (10) has an internal surface (16) defining a space (14) in which the precursor casting (12) can be formed. The surface (16) defines a location indentation (20) to provide a corresponding location projection (3) on the precursor casting (12). A holding member (22) can be inserted into the location projection (3) to hold an internal member (17A, 17B).

Abstract: An investment mould process for forming a cast member is described. The process comprises forming a precursor casting (12) from a master mould (10). The master mould (10) has an internal surface (16) defining a space (14) in which the precursor casting (12) can be formed. The surface (16) defines a location indentation (20) to provide a corresponding location projection (3) on the precursor casting (12). A holding member (22) can be inserted into the location projection (3) to hold an internal member (17A, 17B).

Abstract: An investment mould process for forming a cast member is described. The process comprises forming a precursor casting (12) from a master mould (10). The master mould (10) has an internal surface (16) defining a space (14) in which the precursor casting (12) can be formed. The surface (16) defines a location indentation (20) to provide a corresponding location projection (3) on the precursor casting (12). A holding member (22) can be inserted into the location projection (3) to hold an internal member (17A, 17B).

Abstract: An instrumentation system including a plurality of sensors attached to a test piece, each of the sensors provided with an electrical signal generating device and an electrical signal measurement device. Each sensor is located remotely from its electrical signal measurement device, and connected thereto by connectors. A method of testing connectivity of the instrumentation system including: inducing an electrical signal in a first sensor; and checking that the electric signal is received in the corresponding signal measurement device.

Abstract: A combustion device, such as an afterburner 4 for a gas turbine engine, comprises primary and secondary fuel delivery means 12, 20. Pressure fluctuations in the afterburner 4 are damped by reducing the supply of fuel through the primary fuel delivery means 12 during pressure peaks. This is achieved by means of a spill valve 18, controlled by a controller 22 in response to pressure in the afterburner 4 as sensed by a transducer 24. The spill valve 18 diverts a proportion of the fuel to the secondary fuel delivery means 20. The secondary fuel delivery means 20 is positioned upstream of the primary fuel delivery means 12 by a distance X, such that the secondary fuel delivery means serves to supply additional fuel to the combustion zone 26 during pressure troughs.

Abstract: A sealing element for a turbine of a gas turbine engine includes a radially inner surface region provided with an integrally formed seal structure comprising a plurality of radially inwardly projecting walls. The walls may be abradable and may define cells for receiving an abradable sealing material.

Abstract: A method and device for monitoring the condition of an electrode in an electro-discharge machine are disclosed. The device includes an electrode detector operable to produce a position signal indicative of the position of an end region of the electrode in the nose-guide, and a signal processing unit connected to receive a position signal from the electrode detector and operable to determine from the position signal when retraction of the electrode into the nose-guide has occurred. The method monitors the position of the end region of the electrode and determines when retraction of the electrode has occurred.

Abstract: A sealing arrangement between annular components 2, 4, for example in a gas turbine engine, comprises an L-shaped sealing ring 20. The sealing ring 20 has a first limb 24 received in a groove 28 in the first component 2, and a second limb 26 having a sealing face 30 which is maintained in contact with an abutment surface 32 on the second component 4 under the resilient action of the sealing ring 20. The second limb 26 is exposed to the pressure in a chamber 34 which assists the resilience of the sealing ring 20 in maintaining sealing contact between the sealing face 30 and the abutment surface 32. Bleed holes 38 are provided.

Abstract: An irreversible temperature indicating paint comprises 0.7 wt % to 0.9 wt % toluidine red, 7 wt % to 32 wt % cobalt aluminium spinel and glass, 7 wt % to 32 wt % cobalt chromite spinel, alumina, purple gold and frit, 37 wt % to 41 wt % acrylic resin, 20 wt % to 23 wt % silicone resin excluding solvent. The solvent comprises a mixture of 80% 1-methoxy-2-propanol and 20% dipropylene glycol monomethyl ether. A particular irreversible temperature indicating paint comprises 0.78 wt % toluidine red, 19.55 wt % cobalt aluminium spinel and glass, 19.55 wt % cobalt chromite spinel, alumina, purple gold and frit, 39.1 wt % acrylic resin, 21.1 wt % silicone resin excluding solvent. The irreversible temperature indicating paint is used to determine the temperatures to which various parts of turbine blades, turbine vanes or other components are subjected in operation of a gas turbine engine.

Abstract: A Ship Propulsion Arrangement A ship propulsion arrangement is provided in which a mechanical drive shaft (3) extends through a hull (9) of a ship (1) to an external body (7). The external body incorporates an electric motor. The mechanical drive shaft (3) is driven when required by either a prime mover (2) located within the ship (1) or by the electric motor in the external body (7). This arrangement provides the benefits of hybrid operation without necessary accommodation of the electric motors within the limited hull space available in a ship (1).