Rtx Corporation

Raytheon Technologies Corporation

United Technologies Corporation

Martin Richard Amari

Timothy J Crowley

Zubair Ahmed Baig

Joshua Adams

Manuj Dhingra

Richard P Meisner

Michael Winter

David Lei Ma

Geoffrey T Blackwell

Sorin Bengea

David Gelwan

Kevin Hendricks

Daniel Bernard Kupratis

Marc J Muldoon

Coy Bruce Wood

Jonathan Gilson

Michael D Greenberg

William S Pratt

Steven D Roberts

Tyler W Hayes

Timothy J Gaudet

James K Berry

Nancy Poisson

Ryan M Stanley

Andrew B Thompson

Xiao D He

Matt Dushek

Methods and systems for monitoring operation of hybrid electric engines of aircraft. The methods include monitoring a motor condition of an electrical power system associated with an engine condition using a motor sensor, wherein the electrical power system comprises an electric machine operably coupled to at least one shaft of an engine core, wherein the electric machine is configured to at least one of add power to the at least one shaft and extract power from the at least one shaft, receiving motor data from the motor sensor at a motor controller, wherein the motor controller is configured to control operation of, at least, the electric machine, analyzing the motor data to determine the presence of a fault in the engine core, and, when a fault is detected, performing a fault response action.

Electric machines for aircraft engine fault detection

A system of a hybrid aircraft includes a first gas turbine engine, a second gas turbine engine, and a controller. The first gas turbine engine includes a first low spool electric machine and a first high spool electric machine. The second gas turbine engine includes a second low spool electric machine and a second high spool electric machine. The controller is operable to determine an operating mode of the hybrid aircraft and control power extraction from either or both of the first low spool electric machine and the first high spool electric machine while a single engine descent mode is active. Electric power is provided to either or both of the second low spool electric machine and the second high spool electric machine while the single engine descent mode is active to balance thrust between the first gas turbine engine and the second gas turbine engine.

Hybrid electric single engine descent power extraction control

A distributed hybrid propulsion system is provided and includes an engine assembly that includes an engine, which is non-propulsive, an energy storage element, a motor-generator and a common bus to which the engine assembly, the energy storage element, the motor-generator and the controller are connected. The common bus includes bi-directional power flow control elements, which are controllable to manage bi-directional power flow between the engine assembly, the energy storage element and the motor-generator.

Power and energy management for distributed hybrid propulsion

A system of a hybrid aircraft includes a first gas turbine engine, a second gas turbine engine, a power source, and a controller. The first gas turbine engine includes a first electric machine. The second gas turbine engine includes a second electric machine. The controller is operable to determine an operating condition of the first gas turbine engine and the second gas turbine engine and to detect a change in a thrust command for the hybrid aircraft. The controller is further operable to determine an adjustment to the second electric machine to compensate for the change in the thrust command while the first gas turbine engine is operating in a fuel-driven mode and the second gas turbine engine is operating in an electrically-driven mode. At least a portion of electric power to perform the adjustment to the second electric machine is extracted from the power source.

Hybrid electric single engine descent energy management

A hybrid-electric aircraft system is provided and includes first and second hybrid-electric engines, each of which includes an electric motor to drive operations thereof, and a supplemental power unit (SPU). The SPU is configured as a thermal engine paired with a generator and is configured to generate electrical power. The first and second hybrid-electric engines are operable normally and off, respectively, with electrical power generated by the SPU being diverted to the electric motor of the second hybrid-electric engine.

Hybrid-electric single engine descent failure management

A hybrid electric propulsion system includes a gas turbine engine having at least one compressor section and at least one turbine section operably coupled to a shaft. The hybrid electric propulsion system includes an electric motor configured to augment rotational power of the shaft of the gas turbine engine. A controller is operable to determine hybrid electric propulsion system parameters based on a composite system model and sensor data, determine a prediction based on the hybrid electric propulsion system parameters and the composite system model, determine a model predictive control optimization for a plurality of hybrid electric system control effectors based on the prediction using a plurality of reduced-order partitions of the composite system model, and actuate the hybrid electric system control effectors based on the model predictive control optimization.

Adaptive model predictive control for hybrid electric propulsion

A hybrid electric propulsion (HEP) system included in an aircraft includes a generator configured to output a first power, an energy storage system configured to output a second power, a propulsion system configured to generate thrust based on at least one of the first power and the second power, and an HEP controller in signal communication with the generator, the battery, and the HEP system. The HEP controller is configured to detect loss of the first power output from the generator, to determine an altitude of the aircraft and to actively control delivery of the second power to the propulsion system based on the altitude during the loss of the first power.

Aircraft hybrid electric propulsion architecture with generator-motor fail mode

A system for bleeding air from a core flow path of a gas turbine engine includes a bleed valve in a bleed air duct configured to receive bleed air from a first entrance point to the core flow path into the bleed air duct; a pressurized air valve in a pressurized air duct configured to receive pressurized air from a second entrance point to the core flow path, the pressurized air at a pressure greater than that received into the first entrance point; an eductor outlet from the pressurized air duct located in the bleed air duct; and a control system operable to control operation of the bleed valve and the pressurized air valve.

Assisted engine start bleed system

An engine system of an aircraft includes an energy storage system, a gas turbine engine, and a controller. The gas turbine engine includes a low spool, a high spool, a low-spool generator operably coupled to the low spool, and a high-spool electric motor operably coupled to the high spool. The controller is configured to detect a braking condition of the aircraft, transfer power from the low-spool generator to the energy storage system based on the storage capacity state of the energy storage system, and transfer power to the high spool through the high-spool electric motor to support combustion in the gas turbine engine while a rotational speed of the low spool is reduced responsive to the low-spool generator extracting energy from the low spool.

Hybrid electric idle and braking for an aircraft

A system for bleeding air from a core flow path of a gas turbine engine is disclosed. In various embodiments, the system includes a bleed valve having a bleed valve inlet configured to receive a bleed air from a first access point to the core flow path and a bleed valve outlet; and an air motor having a first air motor inlet configured to receive the bleed air from the bleed valve outlet and a first air motor outlet configured to exhaust the bleed air, the air motor configured to pump the bleed air from the core flow path of the gas turbine engine.

Power assisted engine start bleed system

An engine system of an aircraft includes a gas turbine engine comprising at least one spool and at least one electric machine operably coupled with the at least one spool. A controller is configured to detect if the at least one spool of the gas turbine engine is in or is approaching an overspeed condition and apply a load to the at least one spool via the at least one electric machine.

Turbine over-speed brake for hybrid electric gas turbine engine

A hybrid electric propulsion system includes a gas turbine engine having at least one compressor section and at least one turbine section operably coupled to a shaft. The hybrid electric propulsion system includes an electric motor configured to augment rotational power of the shaft of the gas turbine engine. A controller is operable to determine an estimate of hybrid electric propulsion system parameters based on a composite system model and sensor data, determine a model predictive control state and a prediction based on the hybrid electric propulsion system parameters and the composite system model, determine a model predictive control optimization for a plurality of hybrid electric system control effectors based on the model predictive control state and the prediction using a plurality of reduced-order partitions of the composite system model, and actuate the hybrid electric system control effectors based on the model predictive control optimization.

A variable vane actuation system of a gas turbine engine is provided. The variable vane actuation system including: a variable vane; a vane stem operably associated with the variable vane, wherein the variable vane is configured to rotate with the vane stem; a vane arm having vane stem end and a vane pin end opposite the vane stem end, the vane arm being operably connected to the vane stem at the vane stem end; and a rotational variable differential transformer operably connected to the vane stem, the rotational variable differential transformer configured to detect an amount of rotation of the vane stem.

Nested direct vane angle measurement shaft

A system for controlling a plurality of hydraulic effectors operably connected to an engine to control engine parameters. The system also includes a plurality of sensors operably connected to measure a state or parameter of each effector, a pump configured to supply fluid to the plurality of effectors, and a controller operably connected to the plurality of sensors, the plurality of effectors, and the pump. The controller executes a method for an adaptive model-based control for controlling each effector, The method includes receiving a request indicative of a desired state for each effector, receiving a weighting associated each request, obtaining information about a current state of each effector, and updating an adaptive model based control (MBC) based upon the information. The method also includes generating a control command for an effector based upon the adaptive MBC and commanding the effector based upon the control command.

Growing community of inventors

Glastonbury, CT, United States of America

Martin Richard Amari