Abstract: 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.

Abstract: An apparatus is provided for a gas turbine engine. This apparatus includes a vane array and a heat exchanger integrated with the vane array. The vane array includes an inner platform, an outer platform and a plurality of vanes. The inner platform extends circumferentially about a centerline and forms an inner peripheral boundary of a flowpath through the vane array. The outer platform extends circumferentially about the centerline and forms an outer peripheral boundary of the flowpath through the vane array. The vanes extend across the flowpath between the inner platform and the outer platform. The heat exchanger includes a passage following a tortuous trajectory along a first wall of the vane array.

Abstract: A system is provided for a gas turbine engine. This gas turbine engine system includes a flowpath wall and a deflector. The flowpath wall includes a surface and an opening to a blind aperture. The surface forms a peripheral boundary of an internal engine flowpath. The opening is disposed in the surface. The blind aperture extends vertically into the flowpath wall from the opening. The deflector projects vertically out from the flowpath wall into the internal engine flowpath. The deflector is configured to deflect gas flowing within the internal engine flowpath over the opening.

Abstract: A combustor includes a liner defining a combustion chamber. An air and fuel mixing body is received within the liner and upstream of the combustion chamber. The mixing body has a center axis and includes a bluff-body. A plurality of fuel injection ports on the bluff-body communicate with a central fuel supply such that fuel passes from the fuel supply passage and into a mixing chamber with a component in an axially downstream direction and a radially outward direction relative to said central axis. A plurality of inner air swirlers provide air into the mixing chamber with a component in an axially downstream direction, a radially outward direction, and with a circumferential component due to swirler structure. The fuel injection ports are downstream of an outlet of the inner air swirlers.

Abstract: A portable attitude test stand (PATS) for an aircraft engine, has: a support frame mountable on a trailer of a road vehicle; a test cell supported by the support frame and sized to receive the aircraft engine, the test cell operable to rotate the aircraft engine about a pitch axis and a roll axis of the aircraft engine; and an actuator operatively connected to the test cell and to the support frame, the actuator operable to lower and raise the test cell relative to the support frame between a transport configuration in which the test cell has a transport height and a test configuration in which the test cell has a test height, the test height being greater than the transport height.

Abstract: A seal assembly for an aircraft engine includes a first seal having an upstream end exposed to a first pressure area, and a downstream end exposed to a second pressure area, the first seal at least partially defining an intermediate pressure area and a chamber being fluidly connected to the second pressure area and to the intermediate pressure area through passages defined in the first seal, the chamber allowing for a mixing of a first portion of a stream of air with air from the second pressure area, and a second seal connected to the first seal, the second seal biasing the first portion of the stream of air toward the passages, and allowing a second portion of the stream of air from flowing therethrough toward the second pressure area. A method of flowing air through an aircraft seal assembly is also described.

Abstract: A system is provided for an aircraft. This aircraft system includes a gas turbine engine and a sensor system. The gas turbine engine includes an inlet and a compressor section. A flowpath projects radially inward into the gas turbine engine from the inlet and extends through the compressor section. The sensor system includes a plurality of static pressure sensors at least partially within the flowpath. The sensor system is configured to determine a total pressure characteristic within the flowpath using the plurality of static pressure sensors.

Abstract: A system and method for detecting fixed pitch operation of a variable pitch propeller of an engine are provided. A command signal for maintaining a rotational speed of the propeller at a reference speed is output. An actual value of at least one of the rotational speed and a blade angle of the propeller is obtained. From the actual value, it is assessed whether an expected change in the at least one of the rotational speed and the blade angle of the propeller has occurred in response to the command signal. Responsive to determining that the expected change in the at least one of the rotational speed and the blade angle of the propeller has not occurred in response to the command signal, operation of the propeller at fixed pitch is detected and an alert output accordingly.

Abstract: The gas turbine engine can have an engine core; a core output shaft drivable by the engine core; a power output shaft; an auxiliary power shaft; and a reduction gearbox having gears, the gears drivingly connecting the core output shaft to the auxiliary power shaft. The gears can include an epicyclic gearing drivingly connecting the core output shaft and the auxiliary power shaft to the power output shaft. The gas turbine engine can further have a second auxiliary power shaft interconnected to the auxiliary power shaft, the power output shaft, and the core output shaft by the gears.

Abstract: In an asymmetric operating regime, a first engine is operating in an active mode to provide motive power to an aircraft while a second engine is operating in a standby mode and de-clutched from a gearbox of the aircraft. In response to an emergency exit request, the second engine's speed is increased, at a maximum permissible rate, to a re-clutching speed while increasing the first engine's power output at a maximum permissible rate. When the re-clutching speed is reached, the second engine's power output is increased at a maximum permissible rate. In response to a normal exit request, the second engine's speed is increased to the re-clutching speed at a rate lower than the maximum permissible rate. When the re-clutching speed is reached, the second engine's power output is increased at a rate lower than the maximum permissible rate.

Abstract: A tool and an associated process for removing a press-fit component from an associated engine component, wherein the engine component is pushed in a first axial direction while the press-fit component is restrained from axially moving in the first axial direction.

Abstract: An assembly includes a rotor housing, a first rotor, a lubrication system, a first vibration sensor, and an engine control system. The rotor housing forms a first rotor cavity. The first rotor is configured for rotation within the first rotor cavity. The first rotor includes the plurality of apex seals. The lubrication system is configured to supply a lubrication flow for lubrication of the plurality of apex seals. The first vibration sensor is on the rotor housing. The first vibration sensor is configured to generate a vibration measurement signal. The engine control system includes a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: identify that the vibration measurement signal exceeds a first vibration threshold, and increase a flow rate of the lubrication flow based on an identification of the vibration measurement signal exceeding the first vibration threshold.

Abstract: A deaeration system for an engine lubrication system, the deaeration system comprising: a deaeration rotor rotatable about an axis and including: a rotor inlet extending circumferentially around the axis, a first and a second rotor outlet, a first rotor passage in fluid communication between the rotor inlet and the first rotor outlet, and a second rotor passage in fluid communication between the rotor inlet and the second rotor outlet in parallel to the first rotor passage; and a deaeration conduit including: a conduit inlet, a splitter downstream of the conduit inlet relative to a flow of lubricant through the deaeration conduit, a first conduit outlet and a second conduit outlet downstream of the splitter, the first conduit outlet in fluid communication with the rotor inlet, the conduit inlet having a curved portion extending away from the splitter.

Abstract: A combustor includes a liner defining a combustion chamber, and receiving an air and fuel mixing body. The mixing body includes a forward face facing into the combustion chamber. A plurality of circumferentially spaced columns about a central axis of the combustor each have at least two fuel supply passages and two air supply passages extending to the forward face. At least one of the two air supply passages and one of the fuel supply passages extending along a direction with a circumferentially extending component in a first circumferential direction. The other of the two air supply passages and the fuel supply passages extend with a circumferential component in a second opposed circumferential direction. A gas turbine engine is also disclosed.

Abstract: A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

Abstract: A fuel manifold adapter for a fuel system of an aircraft engine, the fuel manifold adapter comprising: a body having a body-output interface defining a downstream end of a body passage including a body bore about a bore axis, the body-output interface movably and fluidly connectable to a first component of the fuel system mounted to a first mounting point of the engine, and a body-input interface defining an upstream end of the body passage, the body-input interface rigidly and fluidly connectable to a second component of the fuel system mounted to a second mounting point of the engine, and a transfer tube having an upstream-tube end slidably engaged with the body along the bore axis via the body bore, the transfer tube having a downstream-tube end opposite the upstream-tube end slidably engageable along the bore axis with the first component, the downstream-tube end defining a downstream end of the fuel manifold adapter relative to fuel flow through the fuel manifold adapter.

Abstract: A combustor for a gas turbine engine includes a liner surrounding a fuel and air mixing body. A gaseous fuel supply passage delivers gaseous fuel into the mixing body. A wall of the mixing body has air openings to communicate air into mixing passages. At least one fuel opening in each of the mixing passages allows fuel to flow into the mixing passages and mix with the air. There are passage sections downstream of a location of the fuel openings, such that the mixed air and fuel travel downstream of the location and into a combustion chamber. The passage sections have baffles that create a tortuous path. A gas turbine engine is also disclosed.

Abstract: Aircraft power plants and associated methods are provided. A method for driving a load on an aircraft includes: transferring motive power from an internal combustion (IC) engine to the load; discharging a flow of first exhaust gas from the IC engine when transferring motive power from the IC engine to the load; receiving the flow of first exhaust gas from the IC engine into a combustor; mixing fuel with the first exhaust gas in the combustor and igniting the fuel to generate a flow of second exhaust gas; receiving the flow of second exhaust gas at a turbine and driving the turbine with the flow of second exhaust gas from the combustor; and transferring motive power from the turbine to the load.

Abstract: Methods and systems for starting an aircraft gas turbine engine are described. The method comprises, in a first phase of a startup upon receipt of a start request, modifying a first set of engine control parameters to cause light-up; in a second phase of the startup, modifying a second set of engine control parameters to set conditions for light-around; and in a third phase of the startup, modifying a third set of engine control parameters to propagate a flame around a combustor of the gas turbine engine.

Abstract: A method of determining an inlet total air pressure includes determining a first parameter indicative of a first inlet total air pressure. The method includes executing a sequence that includes: determining a mass air flow passing through the air inlet based on the first parameter, determining a Mach number of air passing through the air inlet based on the mass air flow, determining a static air pressure at the air inlet, determining an air pressure ratio based on the Mach number, generating a subsequent parameter indicative of the revised inlet total air pressure based on the air pressure ratio and the static air pressure, and substituting the subsequent parameter for the first parameter. The method includes executing at least one additional instance of the sequence with the subsequent parameter, and outputting the subsequent parameter as the inlet total air pressure.