A redundant oil and fuel pumping system for use with a gas turbine engine. The pumping system includes a plurality of power sources, a fuel system and an oil system. The fuel system pump being driven by electric motors controlled via variable frequency drives powered by the plurality of power sources. The oil system pump being driven by electric motors controlled via variable frequency drives powered by the plurality of power sources.

A gear box assembly for an engine, having a gear box for transmitting a torque, at least one first, static part, at least one second, rotating part, which is mounted so as to be rotatable relative to the first, static part and on which at least one element of the gear box is provided, and a conduit system for conveying a fluid to elements of the gear box,
wherein the conduit system has at least one first supply line and one second supply line for supplying fluid to at least one element of the gear box.

A leakage recovery facility is provided, by means of which at least a proportion of a leakage flow of fluid which originates from a first duct portion of the first supply line and which flows across at least one seal can be conducted to the second supply line.

Methods and systems of starting a gas turbine engine are provided. During startup, a fuel pressure associated with a primary fuel supply of the gas turbine engine is monitored. A low-pressure event for the primary fuel supply is detected when the fuel pressure falls below a predetermined threshold. Responsive to detecting the low pressure event, an electric backup boost pump is activated by an engine controller to provide fuel to the gas turbine engine.

A heat management system for turbomachinery is provided. The heat management system includes a thermal delivery system configured to providing heating, cooling, or a combination thereof, to a turbomachinery. The thermal delivery system includes a first heat transfer conduit and a second heat transfer conduit. The second heat transfer conduit is disposed on the turbomachinery adjacent to the first heat transfer conduit. The heat management system additionally includes a controller operatively coupled to the thermal delivery system and configured to control the heating, the cooling, or the combination thereof, of the turbomachinery via the thermal delivery system.

A method, apparatus, system, and computer program product for managing pumps in an aircraft. Flight information about an operation of the pumps in a pump package in the aircraft is received by a computer system. The flight information is received from the aircraft. A number of times that an abnormal switching occurred for the pumps within a window of consecutive flights is determined by the computer system when the abnormal switching is identified from the flight information. A set of actions is performed by the computer system when the abnormal switching occurred a number of times for the pumps within the window of consecutive flights that exceeds a set of thresholds for the abnormal switching that is considered healthy for the pumps.

A system for controlling the geometry of a variable geometry compressor. The system having: a mechanical linkage operable to vary the compressor geometry; a first fluid-powered actuator arranged to operate the linkage and configured to be powered by a pressurised supply of an incompressible fluid; and a second fluid-powered actuator arranged to operate the linkage and configured to be powered by a pressurised supply of a compressible fluid. Wherein the first and second actuators are further arranged to operate in combination on the mechanical linkage such that a first actuator force exerted by the first actuator on the mechanical linkage sums with a second actuator force exerted by the second actuator on the mechanical linkage.

Methods and systems of starting a gas turbine engine are provided. During startup, a fuel pressure associated with a primary fuel supply of the gas turbine engine is monitored. A low-pressure event for the primary fuel supply is detected when the fuel pressure falls below a predetermined threshold. Responsive to detecting the low pressure event, an electric backup boost pump is activated by an engine controller to provide fuel to the gas turbine engine.

An assembly for an aircraft turbine engine includes a fan drive reduction gear and a lubrication system including: a reduction gear housing; a lubricant tank; a lubricant supply circuit including a feed pump; and a lubricant recovery circuit including a pump for recovering lubricant from the reduction gear housing. The recovery circuit includes a lubricant distributor, including: a lubricant inlet communicating with a lubricant outlet of the housing; an air inlet; and a distributor outlet, the distributor being able to adopt a lubricant recovery configuration and a configuration for retaining the lubricant in the housing.

The subject of the disclosure is an aircraft comprising at least one left motor generator, at least one right motor generator, a control device for a jet pipe nozzle with variable cross-section of the aircraft comprising at least one control member powered by a first airplane electrical power supply linked to at least one right or left motor generator, the control member powered by a second electrical power supply linked to an electrical power source of the aircraft independent of the first airplane electrical power supply.

A protector for an electric submersible pump includes a plurality of chambers, with adjacent or consecutive pairs of chambers coupled, e.g., fluidly coupled, via a series or parallel body. The body defines two separate hydraulic circuits, which allows for the inclusion of a shaft seal, even when the body couples consecutive chambers in a parallel configuration or couples to a thrust bearing chamber. The shaft tube circuit can include a compensator and/or a relief valve.