United Technologies Corporation

Rolls-royce Corporation

General Electric Company

Rolls-royce North American Technologies Inc.

Syed J Khalid

David Marion Ostdiek

Craig W Irwin

Stefan Joseph Cafaro

William Joseph Solomon

Michael Scott McCambridge

Steven Mark Schirle

Gary M Perkins

Robert Wayne Cedoz

Trevor Wayne Goerig

Drew Michael Capps

Juan A Marcos

Brian A Robideau

Samy Baghdadi

Robert R Pollak

Timothy M Kasberg

Mark J Gritton

Logan Daniel Fritz

Jason Holden

Matthew Truong

Matthew Vincent Gettemy

Matthew J Drodofsky

James V French

An engine with sound-based active controls is provided. The engine includes a sound sensor coupled to the engine system configured to capture sound from the engine system, one or more thrust effector devices, and an engine controller. The engine controller is configured to: determine a sound profile based on signals from the sound sensor, determine a modified control parameter for at least one of the one or more thrust effector devices based on the sound profile and a nominal schedule; and control the at least one of the one or more thrust effector devices according to the modified control parameter, deviating from the nominal schedule.

Turbine engine with sound-based active control

An engine for aeromechanical instability abatement includes a sensor configured to capture data from rotating blades of the engine system, a airflow effector device, and an engine controller. The engine controller is configured to control the airflow effector device according to a nominal schedule, detect, based on a signal from the sensor indicating a vibration amplitude of the rotating blades within a frequency band, an incipient instability condition, in response to the incipient instability condition being present, determine a modified control parameter for at least one of the airflow effector device, and control the airflow effector device according to the modified control parameter, deviating from the nominal schedule.

Aircraft engine aeromechanical instability detection

A method for reducing a clearance gap between a plurality of rotor blades and a shroud assembly of an engine includes determining, with a flight control system, that an airplane is in a first flight condition. The method also includes adjusting the clearance gap to a first clearance gap distance associated with the first flight condition. Further, the method includes receiving, with the flight control system, a demand for a second flight condition. During the second flight condition, the method includes adjusting at least two independently controllable parameters, the at least two independently controllable parameters comprising, at least, a first parameter for optimizing the clearance gap and a second parameter for satisfying a thrust demand of the engine, the first parameter having a first impact on the clearance gap, the second parameter having a second impact on the clearance gap, the first impact being greater than the second impact.

System and method for reducing a clearance gap in an engine

Embodiments of the present invention include unique gas turbine engines. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Gas turbine engine with variable speed turbines

A gas turbine engine control system is disclosed having a model and an observer that together can be used to adjust a command issued to the gas turbine engine or associated equipment to improve performance. In one form the control system includes a nominal model that is adjusted to real time conditions. The adjusted model is used with a Kalman filter and is ultimately used to determine a perturbation to a control signal. In one form the perturbation can be to a legacy controller.

Gas turbine engine performance seeking control

An aircraft having a moveable ejector member for assisting in alleviating a boundary layer flowing along an aircraft surface is disclosed. The moveable ejector member is capable of being placed at a variety of positions between a fully open position and a nested position to entrain the boundary layer with another fluid flow. The ejector member can take the form of an ejector shroud used with a nacelle. In some forms, a gas turbine engine is used to provide an ejector flow to entrain the boundary layer through the flow path.

Positionable ejector member for ejector enhanced boundary layer alleviation

A turbine engine includes a compressor, and high and low pressure turbines. The configuration includes a mid-compression station which can be, in the case of a single compressor in the middle of that compressor, or at the exit of the first compressor in the case of two compressors. Also, there is an exit pressure station at the exit of the compression system. A first gas flow line is interposed between the mid-compression station of the compressor and the low pressure turbine, and a second gas flow line is interposed between the exit pressure station of the compression system and the high pressure turbine. A first valve is coupled to the first gas flow line and modulates a low pressure flow rate of coolant in the first gas flow line, and a second valve is coupled to the second gas flow line and modulates a high pressure flow rate of coolant in the second coolant flow line. A controller is configured to operate the first and second valves based on an operating condition of the turbine engine so as to improve the specific fuel consumption while ensuring acceptable stall margin.

Modulated cooling flow scheduling for both SFC improvement and stall margin increase

Disclosed is a boundary layer ejector fluidically connecting boundary layer bleed slots from an external surface of an aircraft to reduce aircraft/nacelle/pylon drag, reduce jet noise and decrease thrust specific fuel consumption. A boundary layer withdrawn through the boundary layer bleed slots is entrained with an exhaust flow of a gas turbine engine. In another embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with a flow stream internal to the gas turbine engine, such as a fan stream of a turbofan. A moveable shroud can be used to open and close a passage of an ejector which can be used to assist in withdrawing a boundary layer or entrain an ambient air. A lobed mixer can be used in some embodiments to effect mixing between the boundary layer and a primary fluid of the ejector.

Gas turbine engine with ejector

The present inventions include a boundary layer ejector fluidically connecting boundary layer bleed slots from an external surface of an aircraft to reduce aircraft/nacelle/pylon drag, reduce jet noise and decrease thrust specific fuel consumption. In one embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with an exhaust flow of a gas turbine engine. In another embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with a flow stream internal to the gas turbine engine, such as a fan stream of a turbofan. Members can be provided near an outlet of a passageway conveying the withdrawn boundary layer air to locally reduce the pressure of the fluid in which the withdrawn boundary layer air is to be entrained. A lobed mixer can be used in some embodiments to effect mixing between the boundary layer and a primary fluid of the ejector.

A method and an apparatus for determining the clearance between the rotor blades of a rotor assembly and a shroud disposed radially outside of the rotor assembly is provided that calculates steady-state operating conditions for a given power engine setting and utilizes those steady-state conditions to determine a steady-state clearance at the given power setting. The method and apparatus further calculate instantaneous thermal conditions for the rotor disk, rotor blades, and shroud. The instantaneous thermal conditions are subsequently used to determine the amount of instantaneous thermal expansion of the rotor disk, rotor blades, and shroud. A clearance transient overshoot is determined using the calculated instantaneous thermal expansions. The actual clearance is determined using the steady-state clearance and the clearance transient overshoot.

Method and apparatus for use in control of clearances in a gas turbine engine

A control system for preventing a compressor stall in a gas turbine engine includes means for sensing an aerodynamic signal indicative of an impending stall condition. The weak stages of the compressor amplify an upstream disturbance originating in the low pressure compressor or fan, resulting in a discrete frequency disturbance, which is an effective stall precursor. The control system also includes means, responsive to the sensed signal, for processing the signal by isolating the relevant frequencies of the sensed signal, calculating the magnitude of the relevant portions of the sensed signal, comparing the sensed signal to a predetermined threshold indicative of a healthy compressor, and for providing a processed signal indicative of an impending stall condition. The control system further includes means, responsive to the processed signal, for providing an output to initiate corrective action to prevent the impending stall condition. The engine's electronic control system, in response to the output signal, increases the stall margin by either adjusting the compressor variable vanes or modulating the compressor bleed.

Control system for preventing compressor stall

A control system for controlling the compressor stall margin during acceleration in a gas turbine engine includes means for sensing signals indicative of the gas flow temperature and gas pressure. The control system further includes signal processing means, responsive to the sensed signals, for synthesizing and providing a processed signal indicative of a measure of compressor destabilization due to heat transfer effects. The control system also includes means, responsive to the processed signal, for providing an output to initiate corrective action to increase compressor stall margin if needed. The engine control means, which is a part of the control system, increases the compressor stall margin by either adjusting the compressor variable vanes, reducing fuel flow or modulating the compressor bleed.

Controlling stall margin in a gas turbine engine during acceleration

A method and an apparatus for use with a gas turbine engine receive a signal indicative of an engine operating condition, generate signals representative of thermal conditions of a rotor, blades, and a case, and in response to each of the signals above, determine a signal indicative of a difference between a instantaneous clearance for the thermal conditions and a steady state clearance for the engine operating condition. The determination includes effects related to a temporary difference that results from a difference between the steady state clearance for the engine operating condition and a steady state clearance for a preceding engine operating condition, but does not require computation of the actual temperatures or the steady state temperatures of the rotor, the blades, and the case. A signal indicative of the difference between a instantaneous clearance for the thermal conditions and a steady state clearance for the engine operating condition may be provided to various augmentation schedules.

Method and apparatus for use in control and compensation of clearances

A variable cycle gas turbine engine (10) includes a core unit (30) having a guide vane array (36), and a hybrid stage (38) with a rotor blade array (40) and a stator vane array (42). The engine is operable in at least two modes which channel different quantities of working medium air to the core unit, and each mode corresponds to a distinct thermodynamic cycle. The vanes (54) of the hybrid stage stator vane array are bivariable so that the leading segment (140) and trailing segment (146) of each vane are separately adjustable over a range of pitch angles. The bivariable character of the vanes ensures aerodynamic stability of the engine in each of its modes of operation.

Variable cycle engine with enhanced stability

The compressor section of an insert gas turbine engine contains a insert installed around the compressor blades that includes cells in a honeycomb configuration. Each cell is inclined at a compound angle to the blade tip to energize the tip air flow as the tip passes over the cell as the blade rotates, improving stall margin with minimum efficiency loss. Each cell is oriented in the direction of the blade chord and faces the advancing blades. As the blade rotates, it sweeps by each cell and high pressure airflow is first captured in the cell from the high pressure side of the blade and released to the low pressure side as the blade passes the cell in the form of a transient energizing jet of high velocity flow in the direction of the airflow across the blade thereby providing effective mixing of the endwall flows and enhancing the tip flow streamwise momentum.

Growing community of inventors

Palm Beach Gardens, FL, United States of America

Syed J Khalid