Abstract: In a method for closing open ends of tubes, an open tube end is closed using spin-tubing methods known in the art. Following the tube end closing, a concavity is formed in the tube end in a tube end forming machine. The tube end is brazed with the tube end facing upright, so that the braze alloy pools in the concavity, strengthening the tube end.

Abstract: A fan blade is provided with an aerofoil portion for which, for cross-sections through the aerofoil portion at radii between 15% and 25% of the blade span from the root radius, the average leading edge thickness is greater than the leading edge thickness at the tip. The geometry of the fan blade may result in a lower susceptibility to flutter.

Abstract: A gas turbine engine has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The bypass flow rate at the optimum bypass efficiency is appreciably lower than the maximum bypass flow rate at the given conditions. This results in increased design flexibility and improved overall engine performance.

Abstract: A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Abstract: A gas turbine engine system has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The fan bypass inlet mass flow rate at the reference operating point is appreciably higher than the mass flow rate through the bypass duct at the peak bypass efficiency at a given fan reference rotational speed and cruise conditions. This results in increased design flexibility and improved overall engine performance.

Abstract: A gas turbine engine has a cycle operability parameter ? in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter ? may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Abstract: A fan blade is provided with an aerofoil portion for which, at radii between 20% and 40% of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than 70% of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Abstract: A system for automatically brazing joints in a manifold has a loading station, a brazing station, and a cooling station. The brazing station has a plurality of brazing torches moveable to a joint in the manifold to braze the joint. First, second and third fixture frames extend from a common rotatable platform. The platform rotates each of the fixture frames to each of the loading station, brazing station, and cooling station in turn. The fixture frames support manifolds with joints requiring brazing. The torches are disposed on a lifting platform that lifts the torches up to a desired joint. The lifting platform is disposed on a sliding platform that slides the torches horizontally to the desired joint. The torches surround the joint and braze it from all sides simultaneously. While brazing is being performed at the brazing station, loading and unloading of manifolds may be done at the loading station, and cooling of already-brazed manifolds may take place at the cooling station.

Abstract: A gas turbine engine 10 is provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P102) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P104), is no greater than a certain value. The gas turbine engine 10 may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Abstract: A gas turbine engine 10 is provided in which a fan root pressure ratio is no greater than a given value at cruise conditions. The gas turbine engine may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades in which the camber distribution along the span allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Abstract: A gas turbine engine comprising a compressor, a compressor case surrounding the compressor and a compressor tip injector system is disclosed. The compressor tip injector system comprises a cabin blower system comprising a cabin blower compressor arranged in use to compress air used in a cabin of an aircraft and by the compressor tip injector system. The compressor case comprises one or more injectors of the compressor tip injector system through which in use air from the cabin blower compressor is injected towards blade tip ends of blades of the compressor as they rotate.

Abstract: In a method for closing open ends of tubes, an open tube end is closed using spin-tubing methods known in the art. Following the tube end closing, a concavity is formed in the tube end in a tube end forming machine. The tube end is brazed with the tube end facing upright, so that the braze alloy pools in the concavity, strengthening the tube end.

Abstract: A fan blade for a gas turbine engine is arranged such that for any two points on its leading edge that are in the radially outer 40% of the blade span and are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point. The radius of the leading edge of a given fan blade at the hub divided by the radius of the leading edge of the fan blade at the tip is less than or equal to 0.3. Such an arrangement may result in an improved operability range.

Abstract: A fan blade for a gas turbine engine is provided with forward axial lean. The fan blade may have a substantially straight leading edge. The geometry of the fan blade results in a lower susceptibility to flutter, thereby allowing a gas turbine engine comprising such a fan blade to operate over a wider range of operating conditions.

Abstract: A fan blade is provided with an aerofoil portion for which, for cross-sections through the aerofoil portion at radii between 15% and 25% of the blade span from the root radius, the average leading edge thickness is greater than the leading edge thickness at the tip. The geometry of the fan blade may result in a lower susceptibility to flutter.

Abstract: A fan blade is provided with an aerofoil portion for which, at radii between 20% and 40% of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than 70% of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

Abstract: A fan outlet assembly comprises a fan discharge nozzle having a radially inner platform, a radially outer platform arranged coaxially and in radial alignment with the radially inner platform and a circumferential array of outlet guide vanes (110) spanning an annulus defined by the radially inner platform and radially outer platform. An annular duct extends downstream from the circumferential array, the duct having a radially outer wall (112) contiguous with the radially outer platform and a radially inner wall (111) contiguous with the radially inner platform. In a region extending downstream from the circumferential array, the radially inner and outer platforms and/or the radially inner and outer walls having a non-axisymmetric surface bounding the duct. A non-axisymmetric surface may be applied to either or both of the radially outer and radially inner contiguous surfaces and may extend upstream as well as downstream of the circumferential array of outlet guide vanes.