United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Matthew P Forcier

William K Ackermann

Jonathan Ortiz

Brian J Schuler

John C DiTomasso

Paul J Hiester

Frederick M Schwarz

Jordan T Wall

Joseph Brent Staubach

Matthew E Bintz

Andrew J Murphy

John C Ditomasso

A bleed air cooling system for a gas turbine engine includes a bleed port located at an axial location of the gas turbine engine to divert a bleed airflow from a gas turbine engine flowpath, a bleed outlet located at a cooling location of the gas turbine engine and a bleed duct in fluid communication with the bleed port and the configured to convey the bleed airflow from the bleed port to the bleed outlet. A modulating valve is located at the bleed duct and is movable between a fully open position and a fully closed position to regulate the bleed airflow through the bleed duct based on one or more operating conditions of the gas turbine engine.

Active bleed flow modulation

A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade.

Tandem blade rotor disk

A gas turbine engine includes a compressor section and a compressor case with a low pressure compressor (LPC) and a high pressure compressor (HPC). The HPC is aft of the LPC. The compressor case defines a centerline axis. The compressor section also includes a rotor disk defined between the compressor case and the centerline axis. A plurality of stages are defined radially inward relative to the compressor case. The plurality of stages include at least one tandem blade stage. The tandem blade stage includes a plurality of blade pairs. Each blade pair is circumferentially spaced apart from the other blade pairs, and is operatively connected to the rotor disk. Each blade pair includes a forward blade and an aft blade. The aft blade is configured to further condition air flow with respect to the forward blade without an intervening stator vane stage shrouded cavity therebetween.

Tandem rotor blades

A bleed air cooling system for a gas turbine engine includes one or more bleed ports located at one or more axial locations of the gas turbine engine to divert a bleed airflow from a gas turbine engine flowpath, a bleed outlet located at a cooling location of the gas turbine engine and a bleed duct in fluid communication with the bleed port and the configured to convey the bleed airflow from the bleed port to the bleed outlet. One or more safety sensors are configured to sense operational characteristics of the gas turbine engine, and a controller is operably connected to the one or more safety sensors and configured to evaluate the sensed operational characteristics for anomalies in operation of the bleed air cooling system.

Bleed flow safety system

A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade. The tandem rotor disk apparatus may include two separate rotor disk bodies.

A bleed air cooling system for a gas turbine engine includes one or more bleed flowpaths operably connected to a bleed outlet to divert a bleed airflow from a gas turbine engine flowpath. Each bleed flowpath includes two or more bleed ports to divert a bleed airflow from a gas turbine engine flowpath, and a bleed duct in fluid communication with the bleed ports and configured to convey the bleed airflow from the two or more bleed ports to the bleed outlet. A valve is located at each bleed port of and is configured to move between an opened position and a closed position, and one or more sensors are located along the bleed flowpath to sense one or more conditions of the bleed air cooling system. The valve at a particular bleed port is moved to the opened position based on the sensed one or more conditions.

Actively-driven bleed source switching

A graphene heat pipe for a gas turbine engine includes a body of graphene. The body has a hot side to accept heat from the gas turbine engine, a cold side to reject heat from the body, and an adiabatic portion to flow heat within the body between the hot side and the cold side.

Gas turbine engine with graphene heat pipe

A gas-circulation system for conditioning a disk of an aircraft may comprise a first takeoff port configured to extract a combusted gas and a second takeoff port configured to extract an uncombusted gas. A first valve may comprise an inlet in fluid communication with the first and second takeoff ports and an outlet of the first valve in fluid communication with the disk.

High temperature disk conditioning system

A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

High pressure compressor rotor thermal conditioning using outer diameter gas extraction

An integrally bladed rotor includes a monolithic rotor body that has a bore portion, a rim, and a web that joins the bore portion and the rim. A plurality of blades extends outwardly from the rim. An arm that extends axially off of the rim, and there is a pocket on a radially inner side of the arm.

Integrally bladed rotor having axial arm and pocket

A bleed air cooling system for a gas turbine engine includes one or more bleed flowpaths operably connected to a bleed outlet to divert a bleed airflow from a gas turbine engine flowpath. Each bleed flowpath includes two or more bleed ports to divert a bleed airflow from a gas turbine engine flowpath, a bleed duct configured to convey the bleed airflow from the two or more bleed ports to the bleed outlet, and a delta-pressure valve located at each bleed port of the two or more bleed ports configured to move between an opened position and a closed position in response to a difference between a first pressure upstream of the delta-pressure valve and a second pressure downstream of the delta pressure valve. The bleed airflow is selectably conveyed through a bleed port of the two or more bleed ports depending on the operation of the associated delta-pressure valve.

Passively-driven bleed source switching

A rotor includes a rotor hub that is rotatable about an axis. The rotor hub includes a bore portion and a rim. An arm extends axially and radially inwardly from the rim. The arm has a radially inner side, a radially outer side, and a protruding ramp on the radially outer side.

Rotor with axial arm having protruding ramp

A rotor disk for a turbomachine includes a disk body, a rim configured to connect to or include a rotor blade disposed on a radially outward portion of the disk body, a bore defined in a radially inward portion of the disk body and configured to be radially adjacent to a shaft, and a thermal regulation aperture defined in the disk body, radially inward of the rim, for allowing flow to pass axially through the disk body when disposed in a stage of a turbomachine, and a thermal regulation aperture defined in the disk body for allowing flow to pass through the disk body when disposed in a stage of a turbomachine.

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South Windsor, CT, United States of America

Matthew P Forcier