Rolls-royce North American Technologies Inc.

Rolls-royce Corporation

Douglas J Snyder

Patrick C Sweeney

Kenneth M Pesyna

Timothy Unton

Igor Vaisman

Douglas David Dierksmeier

Kerry J Lighty

William Barry Bryan

Paul O'Meallie

Graham Burkholder

Brian Thomas Spangler

Douglas Boyd

Michael Head

Bryan H Lerg

Eric Sean Donovan

Andrew Mark Bollman

Daniel G Edwards

Christopher Banham

Anthony F Pierluissi

Aric Acius

Jeffery Forest Rhodes

Ronald Alan Hall

Scott Schwenker

Ken S Rubel

Nilesh Shah

J Edward Pope

Ken Rubel

A thermal energy system for use with an aircraft includes a cooling loop and a cooler. The cooling loop includes a fluid conduit and a pump configured to move fluid through the fluid conduit to transfer heat from a heat source to the fluid in the fluid conduit to cool the heat source. The cooler includes an air-stream heat exchanger located in a duct and is in thermal communication with the fluid conduit to transfer heat between the fluid in the cooling loop and the air passing through the duct.

Aircraft with thermal energy storage system

A cooling system is provided including a two-phase pump loop and a vapor compression system. The two-phase pump loop cools a thermal load with a first coolant. The vapor compression system is configured to circulate a second coolant. The vapor compression system includes a liquid vapor separator which separates the second coolant into a liquid portion and a gaseous portion. The liquid vapor separator is a thermal energy storage for the two-phase pump loop. A condenser of the two-phase pump loop transfers heat from the first coolant to the liquid portion of the second coolant in the liquid-vapor separator.

Thermal energy storage and heat rejection system

A heat exchanger assembly for a gas turbine engine includes an access panel configured to be removably coupled with an outer wall of a bypass duct arranged circumferentially around a central axis of the gas turbine engine. A heat exchanger is adapted to receive cooling fluid therein and is coupled with the access panel. A shroud extends between the access panel and the heat exchanger to collect a first portion of a flow of air through the bypass duct and direct the first portion through the heat exchanger.

Bypass duct heat exchanger with access panel for gas turbine engines

A bypass duct assembly for a gas turbine engine includes a bypass duct, a heat exchanger assembly, and an inlet cowl. The bypass duct is configured to direct bypass air around an engine core of the gas turbine engine. The heat exchanger assembly includes a heat exchanger located in the bypass duct and configured to transfer heat to the bypass air. The inlet cowl is coupled with the bypass duct and the heat exchanger assembly. The inlet cowl includes a cowl duct configured to collect a first portion of the bypass air and conduct the first portion of the bypass air into the heat exchanger and a flow diverter that guides a second portion of the bypass air around the heat exchanger.

Inlets for gas turbine engine bypass duct heat exchangers

Aircraft with thermal energy storage system with bypass control

Cooling systems and methods of operation are provided. The cooling system may include a two-phase pumped loop (TPPL). The two-phase pumped loop may include, a receiver, a pump downstream from the receiver, a heat load downstream from the pump, a TPPL tee downstream from the heat load, a TPPL check valve downstream from the TPPL tee, and a heat exchanger downstream from the TPPL check valve and upstream from the receiver. The cooling system may further include a vapor cycle system (VCS) loop. The vapor cycle system loop may include the receiver, a compressor downstream from a vapor outlet of the receiver, a compressor check valve downstream from the compressor and upstream of the heat exchanger, the heat exchanger, and the heat load downstream from a liquid outlet of the receiver.

Bimodal cooling system

An example thermal management system may include a first heat exchanger including a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air. A turbine including a turbine inlet may be fluidically coupled to the bleed air outlet. The turbine may be configured to drive a shaft mechanically coupled to the turbine in response to expansion of the cooled bleed air through the turbine. A second heat exchanger may include an expanded bleed air inlet fluidically coupled to a turbine outlet of the turbine. The second heat exchanger may be configured to extract heat from at least one heat source using the expanded bleed air.

Bleed expander cooling with turbine

Thermal management systems for cooling high-power, low-duty-cycle thermal loads by rejecting heat from the thermal loads to the ambient environment are provided. The thermal management systems include a two-phase pump loop in fluid communication with a vapor compression system loop, evaporators disposed in parallel between the two-phase pump loop and the vapor compression system loop, and a thermal energy storage loop including a cold-temperature tank and a warm-temperature tank thermally coupled to the two-phase pump loop and the vapor-compression system loop. Methods of transferring heat from one or more thermal loads to an ambient environment are also provided.

Thermal management system with dual-use serial thermal energy storage for system size reduction

Aircraft with thermal energy storage system for multiple heat loads

A base plate for cooling a power electronics device is provided, the base plate comprising cooling fins, the base plate configured to receive the power electronics device directly above the cooling fins, the cooling fins integral to the base plate, the base plate configured to conduct a liquid coolant past the cooling fins.

High heat flux power electronics cooling design

A system includes a housing for a gas turbine engine, and a fan disposed in the housing to rotate coaxially with a gas turbine included in the housing. The system also includes an inlet guide vane disposed in the housing in axial alignment with the fan and configured to have an open position where a first flow of air is received by the fan through the inlet guide vane, and a closed position where airflow through the inlet guide vane is obstructed. The system further includes a heat exchanger disposed in a supply passage in fluid communication with a second flow of air received by the fan. The second flow of air is received by the fan via the supply passage with the inlet guide vane in the open position or in the closed position.

Inlet guide vane draw heat exchanger system

A gas turbine engine system includes a gas turbine engine having a fan bypass duct, a heat exchanger system, and a flow transfer unit configured to selectively allow or preclude flow between the fan bypass duct and the heat exchanger system. A controller is coupled to and configured to vary the state of one or more of a fan exit nozzle, a heat exchanger exit nozzle, and a flow control valve controlling flow between the fan bypass duct and the heat exchanger system.

Gas turbine engine with reversible heat exchanger

A method for cooling a hybrid electric aircraft propulsion system comprises transporting a coolant through a two-phase pumped loop (TPPL) in thermal contact with an electrical machine and a plurality of power modules to be cooled, where the TPPL includes: a parallel arrangement of cold plates; an evaporator; a condenser; a first control valve; a liquid receiver; and a pump. A sensor positioned upstream of the cold plates, and in some cases upstream of the liquid receiver, measures pressure and/or temperature of a return stream of the coolant and transmits measurement data to a first controller electrically connected to the first control valve. The first controller regulates flow of a first liquid stream through the first control valve based on the pressure and/or temperature measured by the sensor, thereby keeping the return stream at a temperature within a predetermined temperature range.

Thermal management system and method for cooling a hybrid electric aircraft propulsion system

A gas turbine engine includes an electrical device and a microchannel cooling system in communication with the electrical device to remove heat.

Gas turbine engine with microchannel cooled electric device

A thermal management system for tightly controlling temperature of a thermal load (e.g., onboard an aircraft) includes: a pressurized tank for storing an expendable coolant; a control valve downstream of the pressurized tank for controlling a flow rate of the expendable coolant; a heat exchanger downstream of the control valve for transferring heat from a thermal load to the expendable coolant at a predetermined temperature; a back pressure regulator (BPR) downstream of the heat exchanger, the BPR having a set point controlled to maintain the expendable coolant at the predetermined temperature in the heat exchanger; optionally, a sensor or orifice downstream of the heat exchanger for determining vapor quality at an exit of the heat exchanger; and a system exit downstream of the BPR for removing some or all of the expendable coolant from the thermal management system after transferring heat from the thermal load to the expendable coolant.

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Carmel, IN, United States of America

Douglas J Snyder