Abstract: A shear coupling can be used with reciprocating and rotary pumps. The shear coupling uses a shear nut and connecting member as the shear element. The connecting member and shear nut are located inside the axial bore of the shear coupling to protect from corrosive wellbore elements. The coupling members may be connected by matching polygonal shaft and hub profiles on the coupling members to transmit torque. The shaft and hub may have octagonal profiles for better torque transmission capacity. The shear nut may be isolated from large alternating stresses due to pretensioning of the connecting member. The length of the connecting member, especially when pretensioned, and a reduced cross-section that may be formed on the outer surface of the shear coupling increase the overall flexibility of the shear coupling, thereby increasing the resistance to bending stress.

Abstract: Systems, apparatuses, and methods to secure a film to a container are provided. An example sealing device utilizes film from a supply of film to seal a lid onto a container. Various sizes of containers are usable with some example sealing devices. Additional features, such as printing on the film and piercing the film for ventilation and/or insertion of a straw are contemplated. One or more markings along the film may be utilized for confirming that an approved film has been loaded into the sealing device. In response, various components or features of the sealing device may be appropriately enabled or disabled. The one or more markings may also be utilized to convey data to the sealing device regarding the installed film, such as for improved operation thereof.

Abstract: A wheel fastener alarm assembly is provided having a fastener body with a first portion defining a wrenching surface and a cavity, and a second portion with a threaded portion to attach and detach from a wheel of a vehicle. A sensor array is disposed in the cavity of fastener body to detect an attribute of the fastener body and generate an output signal based on the attribute of the fastener body. An antenna connected to the sensor array to transmit the signal to a remote location. A cap is secured to the first portion of the fastener body and covers the wrenching surface and the cavity opening to define a capped fastener body.

Abstract: An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric Stip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition Mrel, wherein Mrel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and Stip is from ?1 to 0.1.

Abstract: An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition Mrel, wherein Mrel is between 0.4 and 0.93, and L/D is between 0.2 and 0.45.

Abstract: A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (?3??1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

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: An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric Stip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition Mrel, wherein Mrel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and Stip is from ?1 to 0.1.

Abstract: An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition Mrel, wherein Mrel is between 0.4 and 0.93, and L/D is between 0.2 and 0.45.

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 fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (?3??1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, compressor, and core shaft connecting turbine to the compressor; a fan located upstream of engine core, the fan containing a plurality of fan blades mounted for rotation about an engine axis, each blade of the plurality of fan blades having a leading edge and trailing edge extending across a span of an airflow duct from blade root to tip; and a gearbox that receives an input from core shaft and outputs drive to fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the trailing edge of each blade is characterized by a metric M defined as a rate of change of an angle of trailing edge between 0.4 and 0.8 of the span divided by an area averaged trailing edge angle, M being not less than around 4.

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 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 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 for an aircraft including: an engine core including a turbine, compressor, and core shaft connecting turbine to the compressor; a fan located upstream of engine core, the fan containing a plurality of fan blades mounted for rotation about an engine axis, each blade of the plurality of fan blades having a leading edge and trailing edge extending across a span of an airflow duct from blade root to tip; and a gearbox that receives an input from core shaft and outputs drive to fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the trailing edge of each blade is characterized by a metric M defined as a rate of change of an angle of trailing edge between 0.4 and 0.8 of the span divided by an area averaged trailing edge angle, M being not less than around 4.

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 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 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.