Abstract: An electronic control system (30) for a turbopropeller engine (12) having a gas turbine (20) and a propeller assembly (13) coupled to the gas turbine (20), controls propeller operation based on a pilot input request, via generation of a driving quantity (Ip) for an actuation assembly (29) designed to adjust a pitch angle (?) of propeller blades (2) of the propeller assembly (13).

Abstract: Aspects of the technology relate to lateral propulsion systems in lighter-than-air (LTA) platforms configured to operate in the stratosphere. One or more motor assemblies are used to actuate the lateral propulsion system and to make directional changes, for instance using one or more propellers. This can include a pointing axis motor assembly for orienting the lateral propulsion system along a particular heading, and a drive motor assembly for causing a propeller assembly or other propulsion mechanism to turn on and off Corrective actions may be necessary to adjust the alignment of the lateral propulsion system. A stepper motor control module may be used to control operation of the pointing axis motor assembly, for instance by causing it to rotate in a clockwise (or counterclockwise) direction. A motor current control approach may be used, in which the motor voltage is adjusted until a measured motor current reaches a selected current level.

Abstract: A combustion tube for a gas turbine including an outlet section having a cross-section of an annular sector-shape. The outlet section includes an outer wall forming an outer peripheral boundary of the annular sector-shape, an inner wall forming an inner peripheral boundary of the annular sector-shape, and a pair of side walls forming boundaries on both sides of the annular sector-shape in a circumferential direction, respectively. The outer wall extends obliquely with respect to the inner wall such that a height of the annular sector-shape decreases toward an outlet opening of the combustion tube. A first side wall extends obliquely with respect to a second side such that a perimeter of the annular sector-shape increases toward the outlet opening of the combustion tube. An inclination angle ?1 of the first side wall with respect to the second side wall satisfies 0<?1?56.

Abstract: A nozzle for a combustor burning a hydrogen-containing fuel is provided. The nozzle includes a first tube through which the fuel flows and having a fuel injection hole on a front side to inject the fuel therethrough, a second tube surrounding the first tube and having a premixing injection hole through which the fuel and air are mixed and discharged, and a third tube surrounding the second tube and through which the fuel flows, wherein the second tube includes a plurality of fine injection holes to inject the fuel from the third tube to form fine flames, and the fine injection holes are spaced apart from each other in a circumferential direction of the second tube.

Abstract: The aircraft includes a fuselage and at least one wing extending from the fuselage. The wing includes first and second original portions and a plug portion positioned between the first and second original portions. A propulsion system is positioned on the at least one wing. The propulsion system includes at least one electric powerplant and at least one combustion powerplant. Each powerplant delivers power to a respective air mover for propelling the aircraft. The electric powerplant and/or the combustion powerplant is positioned outboard from the plug portion. A method for retrofitting an aircraft includes segmenting a wing of an aircraft into two original portions, positioning a plug portion between the two original portions, and connecting the plug portion to one of the two original portions, the first nacelle and/or the second nacelle.

Abstract: A drive unit for air vehicle, which allows building the vertical take-off and landing vehicles, intended for use, for instance, in the production of flying taxis, as well as in the model-making branch and in the toy industry. The drive unit is composed of the air channel, in the form of a straight segment of a tube with circular section, which has fans with engines fixed on its both ends. The vertical draft force outlet-inlet nozzle opening is located between fixed fans of the drive unit.

Abstract: A method of distributing power within a gas turbine engine is disclosed. In various embodiments, the method includes driving a high pressure turbine having a first stage and a second stage with an exhaust stream from a combustor, the first stage connected to a high pressure turbine first stage spool and the second stage connected to a high pressure turbine second stage spool; driving a high pressure compressor connected to a high pressure compressor spool via a differential system, the differential system having a first stage input gear connected to the high pressure turbine first stage spool, a second stage input gear connected to the high pressure turbine second stage spool and an output gear assembly connected to the high pressure compressor spool; and selectively applying an auxiliary input power into at least one of the high pressure compressor spool and the high pressure turbine.

Abstract: A gas turbine engine including a compressor section and a combustion section in serial flow arrangement along an engine centerline, the combustion section having a combustor liner, a dome wall coupled to the combustor liner, and a dome inlet located in the dome wall, a fuel injector fluidly coupled to the dome inlet, a combustion chamber fluidly coupled to the fuel injector and defined at least in part by the combustor liner and the dome wall, and at least one set of dilution openings located in the dome wall and fluidly coupled to the combustion chamber.

Abstract: A flow path assembly for a gas turbine engine is provided. The flow path assembly may include an outer casing comprising a metal material having a first coefficient of thermal expansion, a ceramic structure comprising a ceramic material having a second coefficient of thermal expansion, and a mounting component attached on a first end to the outer casing and attached on a second end to the ceramic structure. The mounting component may be constructed from at least two materials transitioning from the first end to the second end such that the coefficient of thermal expansion is different at the first end than the second end.

Abstract: A maintenance tool for gas turbine engine includes a rail system having a plurality of rail segments insertable through one or more inspection holes of the gas turbine engine for assembly within a core air flowpath of the gas turbine engine. The maintenance tool additionally includes a maintenance head movable along the plurality of rail segments of the rail system for performing maintenance operations within the core air flowpath.

Abstract: A power module includes a turbine arranged along a rotation axis, an interconnect shaft fixed in rotation relative to the turbine, and a compressor with a regenerative compressor wheel. The regenerative compressor wheel is fixed in rotation relative to the interconnect shaft supported for rotation with the turbine about the rotation axis. Generator arrangements, unmanned aerial vehicles, and methods of generating electrical power are also described.

Abstract: A dual wall liner for a gas turbine engine may comprise a shell having a first side and a second side, a panel contacting the shell, the panel at least partially defining a hot gas path through which a hot gas flows, wherein the first side of the shell faces the panel, wherein the shell includes a thermal barrier coating (TBC) disposed on the first side of the shell. The TBC may thermally protect the shell from heat from a hot gas path.

Abstract: A multi-engine aircraft that includes two or more powerplants, each including an electric motor, and an electric power system controlling electrical distribution and operatively connected to the electric motors. The electric power system includes primary battery packs operatively connected to the electric motor of a respective one of the two or more powerplants, a reserve battery pack, and a switch interconnecting the reserve battery pack and the electric motors. The reserve battery pack is shared by the electric motors by the switch and configured such that the reserve battery pack provides electric power to a selected one of the electric motors.

Abstract: A combustor liner for a gas turbine includes a liner that at least partially defines a combustion chamber, and that a plurality of dilution openings therethrough. Each dilution opening includes an outer wall defining an outer perimeter of the dilution opening and defining a dilution opening centerline axis through the dilution opening. A plurality of swirl vanes extend from the outer wall into a dilution airflow passage that extends through the dilution opening. Each of the plurality of swirl vanes extends from the outer wall into the dilution airflow passage at a respective swirl vane angle with respect to the outer wall. The plurality of swirl vanes are arranged in a successive arrangement about the outer wall, and successive respective ones of the plurality of swirl vanes extend from the outer wall at a different swirl vane angle.

Abstract: The present invention relates to a control method in case of engine failure of a hybrid helicopter having a power plant connected to at least one lift rotor and to at least one propeller, said lift rotor having a plurality of first blades and said at least one propeller having a plurality of second blades. The method comprises the following steps: (i) measuring a forward speed of the hybrid helicopter, (ii) on condition that said forward speed is greater than a first speed threshold and that each engine has failed, automatically implementing a first emergency piloting mode comprising a step for automatic reduction by an automatic piloting system of a pitch of said second blades toward an objective pitch making said at least one propeller produce a motive power which is transmitted to the lift rotor.

Abstract: A system is provided for multi-engine coordination of gas turbine engine motoring in an aircraft. The system includes a controller operable to determine a motoring mode as a selection between a single engine dry motoring mode and a multi-engine dry motoring mode based on at least one temperature of a plurality of gas turbine engines and initiate dry motoring based on the motoring mode.

Abstract: A waste heat capture system that can be used with at least a gas turbine engine. The system includes: an air scoop connected to a first component, the air scoop configured to direct air from a first duct to an interior compartment of the first component; a second duct along an exterior of the first component; and a thermoelectric material connected to an interior surface of the first component. The interior compartment of the first component is on a first side of the thermoelectric material and the exterior of the first component is on a second side of the thermoelectric material. The first duct is configured to receive air having a first temperature range, and the second duct is configured to receive air having a second temperature range, wherein the second temperature range is an order of magnitude higher than the first temperature range.

Abstract: Embodiments are directed to boosting aircraft engine performance for takeoff and critical mission segments by reducing airflow used for cooling exhaust gases. The airflow is reduced by stopping an accessory blower or by closing an external air vent Eliminating the cooling airflow to the exhaust has the effect of lowering the backpressure on the engine, which thereby increases maximum engine power.

Abstract: An apparatus is provided for a turbine engine. This turbine engine apparatus includes a monolithic body. The monolithic body includes a splash plate and a fuel nozzle. The splash plate includes a splash plate surface. The fuel nozzle includes a nozzle orifice. The fuel nozzle is configured to direct fuel out of the nozzle orifice to impinge against the splash plate surface.

Abstract: Techniques for determining a fan spool speed value associated with a hybrid electric gas engine (“engine”) are described herein. For example, a method includes generating a first fan spool speed value based on a base extraction amount. The method further includes generating a change in the first fan spool speed value based on a flight condition and a change in extraction amount. The method further includes generating a second fan spool speed value based on the extraction amount and an augmentation value for intersection of a combustion instability and burner pressure limit. The method further includes generating a second change in the second fan spool speed value based on the flight condition and the change in extraction amount. The method further includes determining a final fan spool speed value and operating the engine at the final fan spool speed value.