Abstract: The present invention relates to control of engine variable of a gas turbine engine to regulate the surge margins of at least two compressors. A controller (20) receives data measured from engine sensors (22, 23) and uses said data to determine an indication of surge margin for each of at least two compressors (6, 7) of the gas turbine engine. The controller (20) uses the indications of surge margin for each of the compressors to determine a control strategy that balances the requirements of each compressor. In one embodiment a surge margin operating map divided into different control domains (40, 43, 43) is used. The indication of surge margin determined for each compressor is plotted to determine which control domains the current operating point on the surge margin operating map falls within.

Abstract: A valveless pulse combustor having a combustion chamber with a closed first end and an open second end, the combustor also having a tailpipe in fluid communication with the open second end of the combustion chamber, the combustor further having an inlet pipe in fluid communication with the open second end of the combustion chamber, the inlet pipe and the tailpipe being arranged such that one is located within the other.

Abstract: A control assembly operable in the event of overheating of an engine electronic controller on an aircraft. The assembly including one or more temperature sensors located in the vicinity of the aircraft engine and connected to a control unit. The control unit is arranged to measure the temperature or temperatures detected by the sensors, and to determine whether in view of detected increased temperatures, a situation constitutes an emergency situation or a controlled situation. The control unit communicates a controlled situation to a pilot of the aircraft, and automatically causes an override action to occur to the aircraft engine when an emergency situation is determined.

Abstract: An aerofoil blade or vane (1) suitable for the turbine of a gas turbine engine includes a longitudinally extending aerofoil portion (7) having facing wall parts (20, 22). The wall parts being interconnected by a generally longitudinally extending divider member (17) to partially define first and second cooling fluid passage portions (11, 15) disposed in side-by-side generally longitudinally extending relationship. The first and second passage portions being interconnected in series fluid flow relationship by a bend passage portion (13). The first passage portion is adapted to direct cooling fluid to the bend portion and the second passage portion being adapted to exhaust cooling fluid from the bend portion. The divider member has a first local thickening (33) in the region of the bend portion to provide a localised contraction of the downstream end of the first passage portion to accelerate the cooling fluid flow before it enters the bend passage portion.

Abstract: A method for depositing material at an intersection region (116) of at least two surfaces (102,104,106,108,110) of a component (100). The method comprises a first deposition process providing material at said intersection and directing an energy beam (114) toward the intersection region to join the material to the component. The beam is maintained at an obtuse angle relative to at least two of the surfaces such that substantially all of the working area of the beam is incident upon the intersection region.

Abstract: A variable stator vane assembly for a gas turbine engine has a projection protruding from the upper stem of the vane and which engages an aperture in the vane lever to provide a visual indication of correct alignment between the lever and the vane.

Abstract: The present invention relates to an aerofoil having an erosion resistant leading edge. The aerofoil comprises an aerofoil main body portion and a leading edge portion. The leading edge portion comprises a plurality of functionally distinct parts, including at least; a first part for resisting abrasion of the leading edge portion, and a second part for resisting a bending force applied to the leading edge portion, and located behind the first part.

Abstract: There is disclosed a self-regulating electric heater assembly arranged to heat an electrically conductive substrate such as a flexible substrate of shape memory alloy for use as an actuator. The heater assembly comprises a plurality of substantially rigid PTC elements arranged in spaced-apart relation to one another in a flexible array, said PTC elements each having a contact surface arranged in contact with the substrate and being urged against said substrate so as to remain in contact with the substrate upon flexure of the substrate. The substrate serves as one of a pair of electrical conductors electrically connected to the PTC elements for the supply of electric current to the PTC elements.

Abstract: A method of inspecting a hole in a component situated within an engine. The engine comprises a chamber configured to receive fluid from a fluid supply, and the component is in flow communication with the fluid supply via the chamber. The method comprises the steps of (i) delivering a test fluid via the chamber to the component to induce flow of the test fluid through the hole; (ii) deploying a sensor panel adjacent the component in the path of test fluid issuing from the hole, the sensor panel being sensitive to the impingement of the test fluid; and (iii) assessing the condition of the hole in dependence on the response of the sensor panel at the location corresponding to the hole.

Abstract: An aerofoil blade assembly having a rotor disc having a radially outer surface provided with an axial slot and a radially inner surface provided with a circumferential slot, wherein the lower region of the axial slot intersects the upper region of the circumferential slot; an aerofoil blade having a blade root which is received by the axial slot and which is provided with a locating slot which corresponds to the circumferential slot; a retention key which engages with the circumferential slot and the locating slot to restrict axial displacement of the aerofoil blade with respect to the rotor disc, the retention key being provided with socket which, when the retention key is engaged in the circumferential slot and the locating slot, is substantially aligned with the axial slot.

Abstract: A method of detecting an ice shedding event in a gas turbine engine, the engine having a rotor comprising a compressor drivingly connected via a shaft to a turbine; the method for detecting an ice shedding event comprising the steps of measuring the temperature at regular intervals and where a temperature drop of at least 20 degrees per second is recorded, producing a signal indicative of an ice shedding event and sending the signal to an indicator device. Advantageously, the engine may then be inspected for damage associated to an ice impact rather than other foreign object ingestion event or fuel flow irregularity. Alternatively, instead of temperature, pressure may be measured and a pressure drop of at least 20 kPa per second is indicative of an ice shedding event.

Abstract: A method of manufacturing an aerofoil structure to have a portion with an increased erosion resistance, the method comprising: providing one or more titanium elements (2, 4, 6) and a beta-stabilizing material (8); wherein the one or more elements (2, 4, 6) have an alpha-beta microstructure; assembling the one or more elements (2, 4, 6) and the beta-stabilizing material (8) such that the beta-stabilizing material (8) is adjacent to the one or more elements (2, 4, 6); and heating the assembly such that the beta-stabilizing material (8) diffuses into an adjacent portion of the one or more elements (2, 4, 6), causing the adjacent portion of the one or more elements (2, 4, 6) to have a beta microstructure which provides an increased erosion resistance.

Abstract: A fluidfoil tip vortex disruption arrangement. There is a fluid inlet on a pressure side of the fluidfoil; and a fluid outlet in a low pressure region at or near the tip of the fluidfoil. Fluid exiting the fluid outlet inserts instability into a tip vortex to disrupt or destroy the vortex.

Abstract: A component, for example an aerofoil vane or other gas path structure in a gas turbine machine such as a gas turbine engine, in which an insulated electrical conductor path (2) is embedded structurally integrally, and in which electromagnetic shielding (1, 2, 20, 22), for the embedded electrical conductor path is provided structurally integrally.

Abstract: A liner for an interior surface of an engine casing (18), comprises a number of liner panels (32) disposed radially outwardly of a rotatable component (56) of the engine, such as a fan blade, and mounting means (46) arranged to mount the panels on the interior surface of the casing, wherein each panel is individually adjustable, such that the clearance between the panel and the rotatable component of the engine may be varied individually for each panel.

Abstract: A bypass turbofan gas turbine engine is started by means of electric starter motor mounted directly about a downstream end or upstream of the low pressure spool of the engine. This causes air to be driven by a fan through a bypass duct around the engine casing. Closures close to substantially seal an outlet of the bypass duct, and the air is directed into a combustion chamber of the engine and through the turbines, causing the high pressure spool to pick up speed for starting.

Abstract: A turbomachine casing assembly comprises a first casing portion for at least partially encasing one or more rotating aerofoil structures of a turbomachine and a second casing portion; the first casing portion being movable with respect to the second casing portion. One or more resilient elements are arranged so as to resist movement of the first casing portion with respect to the second casing portion.

Abstract: The present invention relates to a test specimen (11) for testing the through-thickness properties of a fibre-reinforced composite material. The specimen comprises a first loading portion (10) having a first longitudinal axis in a first plane and a second loading portion (20) having a second longitudinal axis in a second plane. The first and second planes are substantially parallel and the first and second axes are oblique. The first and second loading portions (10, 20) partially overlap to form a test gauge portion (30) that is integral with the first and second loading portions (10, 20). The first loading portion (10) extends beyond the test gauge portion (30) in opposite directions along the first axis and the second loading portion (20) extends beyond the gauge portion (30) in opposite directions along the second axis. This allows a force be applied to the test gauge portion through the first and second loading portions.

Abstract: In prop fan and contra fan gas turbine engines particularly utilized with regard to pusher configurations towards the aft of an aircraft, there is a significant problem with regard to debris and bird strikes resulting in particulate matter which may damage the engine thrust core. By providing an inlet rotor which fragments and creates a centrifugal aspect to the inlet flow, particulate matter can be directed to a bypass chute normally in the form of an annular section extending into a number of chute exits. In such circumstances, the particulate matter is removed continuously from the inlet flow and the chutes are normally angled to project any particulate matter towards propellers of the engine at robust positions.

Abstract: A panel assembly of the kind found in the air intake of a gas turbine engine or a fan duct is designed to be demountable for maintenance purposes. Consequently there exists between the panels a gap to be sealed. Thus, the assembly has a deformable seal for such a purpose and has an elongated, hollow resilient member mounted on a flange at the edge of one of the panels and arranged to be pressed against a flange at the confronting edge of the adjacent panel. The flanges are shaped such that the hollow member when pressed between the flanges effectively seals the gap between the panels and a portion of the seal member lies flush between the panel edges.