GB2587687A - Systems and methods for aircraft - Google Patents

Systems and methods for aircraft Download PDF

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Publication number
GB2587687A
GB2587687A GB2003511.9A GB202003511A GB2587687A GB 2587687 A GB2587687 A GB 2587687A GB 202003511 A GB202003511 A GB 202003511A GB 2587687 A GB2587687 A GB 2587687A
Authority
GB
United Kingdom
Prior art keywords
thermal management
management system
nacelle
aircraft
profile
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
GB2003511.9A
Other versions
GB202003511D0 (en
Inventor
Iqbal Kamran
Wood Norman
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Advanced Mobility Research and Development Ltd
Original Assignee
Advanced Mobility Research and Development Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from GB1914220.7A external-priority patent/GB2587637A/en
Priority claimed from GB1914224.9A external-priority patent/GB2589300A/en
Priority claimed from GB1914219.9A external-priority patent/GB2587636A/en
Priority claimed from GB1914227.2A external-priority patent/GB2593417A/en
Priority claimed from GB1914223.1A external-priority patent/GB2589299A/en
Priority claimed from GB1914225.6A external-priority patent/GB2593416A/en
Application filed by Advanced Mobility Research and Development Ltd filed Critical Advanced Mobility Research and Development Ltd
Publication of GB202003511D0 publication Critical patent/GB202003511D0/en
Priority to PCT/GB2020/052382 priority Critical patent/WO2021064383A2/en
Priority to EP20866962.2A priority patent/EP4061707A2/en
Priority to PCT/GB2020/052375 priority patent/WO2021064376A2/en
Priority to EP20866954.9A priority patent/EP4061720A2/en
Priority to EP20866959.8A priority patent/EP4058358A2/en
Priority to EP20864298.3A priority patent/EP4066225A2/en
Priority to EP20864295.9A priority patent/EP4058349A2/en
Priority to EP20786032.1A priority patent/EP4061724A1/en
Priority to PCT/GB2020/052389 priority patent/WO2021064390A2/en
Priority to PCT/GB2020/052391 priority patent/WO2021064392A2/en
Priority to EP20866956.4A priority patent/EP4058355A2/en
Priority to PCT/GB2020/052393 priority patent/WO2021064394A2/en
Priority to PCT/GB2020/052374 priority patent/WO2021064375A2/en
Priority to EP20866963.0A priority patent/EP4061721A2/en
Priority to EP20864297.5A priority patent/EP4061719A2/en
Priority to PCT/GB2020/052379 priority patent/WO2021064380A2/en
Priority to PCT/GB2020/052383 priority patent/WO2021064384A2/en
Priority to EP20866960.6A priority patent/EP4058356A2/en
Priority to PCT/GB2020/052388 priority patent/WO2021064389A2/en
Priority to PCT/GB2020/052376 priority patent/WO2021064377A2/en
Priority to EP20864291.8A priority patent/EP4061705A2/en
Priority to PCT/GB2020/052384 priority patent/WO2021064385A2/en
Priority to PCT/GB2020/052377 priority patent/WO2021064378A2/en
Priority to PCT/GB2020/052395 priority patent/WO2021064396A1/en
Priority to PCT/GB2020/052373 priority patent/WO2021064374A2/en
Priority to PCT/GB2020/052387 priority patent/WO2021064388A2/en
Priority to EP20864296.7A priority patent/EP4061725A2/en
Priority to EP20866964.8A priority patent/EP4059004A2/en
Priority to PCT/GB2020/052381 priority patent/WO2021064382A2/en
Priority to PCT/GB2020/052385 priority patent/WO2021064386A2/en
Priority to PCT/GB2020/052378 priority patent/WO2021064379A2/en
Priority to PCT/GB2020/052394 priority patent/WO2021064395A2/en
Priority to PCT/GB2020/052396 priority patent/WO2021064397A2/en
Priority to EP20866982.0A priority patent/EP4061722A2/en
Priority to PCT/GB2020/052386 priority patent/WO2021064387A2/en
Priority to EP20866961.4A priority patent/EP4058352A2/en
Priority to EP20866957.2A priority patent/EP4061706A2/en
Priority to EP20866958.0A priority patent/EP4061717A2/en
Priority to EP20866955.6A priority patent/EP4061715A2/en
Priority to EP20864293.4A priority patent/EP4058360A2/en
Priority to PCT/GB2020/052390 priority patent/WO2021064391A2/en
Priority to EP20866953.1A priority patent/EP4061714A2/en
Priority to EP20864300.7A priority patent/EP4058348A2/en
Priority to EP20864299.1A priority patent/EP4058361A2/en
Priority to EP20864292.6A priority patent/EP4058347A2/en
Priority to PCT/GB2020/052392 priority patent/WO2021064393A2/en
Priority to PCT/GB2020/052380 priority patent/WO2021064381A2/en
Priority to EP20864294.2A priority patent/EP4058359A2/en
Publication of GB2587687A publication Critical patent/GB2587687A/en
Pending legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C25/00—Alighting gear
    • B64C25/32—Alighting gear characterised by elements which contact the ground or similar surface 
    • B64C25/405—Powered wheels, e.g. for taxing
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L1/00—Supplying electric power to auxiliary equipment of vehicles
    • B60L1/02—Supplying electric power to auxiliary equipment of vehicles to electric heating circuits
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L1/00—Supplying electric power to auxiliary equipment of vehicles
    • B60L1/20—Energy regeneration from auxiliary equipment
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L50/00—Electric propulsion with power supplied within the vehicle
    • B60L50/50—Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells
    • B60L50/60—Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells using power supplied by batteries
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L50/00—Electric propulsion with power supplied within the vehicle
    • B60L50/50—Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells
    • B60L50/60—Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells using power supplied by batteries
    • B60L50/61—Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells using power supplied by batteries by batteries charged by engine-driven generators, e.g. series hybrid electric vehicles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B60—VEHICLES IN GENERAL
    • B60L—PROPULSION OF ELECTRICALLY-PROPELLED VEHICLES; SUPPLYING ELECTRIC POWER FOR AUXILIARY EQUIPMENT OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRODYNAMIC BRAKE SYSTEMS FOR VEHICLES IN GENERAL; MAGNETIC SUSPENSION OR LEVITATION FOR VEHICLES; MONITORING OPERATING VARIABLES OF ELECTRICALLY-PROPELLED VEHICLES; ELECTRIC SAFETY DEVICES FOR ELECTRICALLY-PROPELLED VEHICLES
    • B60L58/00—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles
    • B60L58/10—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
    • B60L58/12—Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries responding to state of charge [SoC]
    • B60L58/15—Preventing overcharging
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
    • B64C1/12—Construction or attachment of skin panels
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B64C11/002—Braking propellers, e.g. for measuring the power output of an engine
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B64C11/02—Hub construction
    • B64C11/14—Spinners
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B64C11/30—Blade pitch-changing mechanisms
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
    • B64C13/24—Transmitting means
    • B64C13/38—Transmitting means with power amplification
    • B64C13/50—Transmitting means with power amplification using electrical energy
    • B64C13/504—Transmitting means with power amplification using electrical energy using electro-hydrostatic actuators [EHA's]
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C19/00—Aircraft control not otherwise provided for
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C25/00—Alighting gear
    • B64C25/02—Undercarriages
    • B64C25/04—Arrangement or disposition on aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C25/00—Alighting gear
    • B64C25/32—Alighting gear characterised by elements which contact the ground or similar surface 
    • B64C25/34—Alighting gear characterised by elements which contact the ground or similar surface  wheeled type, e.g. multi-wheeled bogies
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C25/00—Alighting gear
    • B64C25/32—Alighting gear characterised by elements which contact the ground or similar surface 
    • B64C25/42—Arrangement or adaptation of brakes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C3/00—Wings
    • B64C3/18—Spars; Ribs; Stringers
    • B64C3/185—Spars
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C3/00—Wings
    • B64C3/34—Tanks constructed integrally with wings, e.g. for fuel or water
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C7/00—Structures or fairings not otherwise provided for
    • B64C7/02—Nacelles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D13/00—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft
    • B64D13/006—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft the air being used to cool structural parts of the aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • B64D15/02—De-icing or preventing icing on exterior surfaces of aircraft by ducted hot gas or liquid
    • B64D15/04—Hot gas application
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/026—Aircraft characterised by the type or position of power plants comprising different types of power plants, e.g. combination of a piston engine and a gas-turbine
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/10—Aircraft characterised by the type or position of power plants of gas-turbine type 
    • B64D27/12—Aircraft characterised by the type or position of power plants of gas-turbine type  within, or attached to, wings
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/24—Aircraft characterised by the type or position of power plants using steam or spring force
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D29/00—Power-plant nacelles, fairings, or cowlings
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D31/00—Power plant control systems; Arrangement of power plant control systems in aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D31/00—Power plant control systems; Arrangement of power plant control systems in aircraft
    • B64D31/02—Initiating means
    • B64D31/06—Initiating means actuated automatically
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/08—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D35/00—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions
    • B64D35/02—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions specially adapted for specific power plants
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D37/00—Arrangements in connection with fuel supply for power plant
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D37/00—Arrangements in connection with fuel supply for power plant
    • B64D37/30—Fuel systems for specific fuels
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D37/00—Arrangements in connection with fuel supply for power plant
    • B64D37/32—Safety measures not otherwise provided for, e.g. preventing explosive conditions
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D41/00—Power installations for auxiliary purposes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D41/00—Power installations for auxiliary purposes
    • B64D41/007—Ram air turbines
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D47/00—Equipment not otherwise provided for
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D15/00—Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby
    • F01D15/10—Adaptations for driving, or combinations with, electric generators
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/047—Heating to prevent icing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
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Abstract

A thermal management system 100,200 has a skin portion 208 having a profile, and an energy storage device, such as a battery pack 202, with an outer profile 214 shaped to conform to the profile of the skin portion. Preferably the system is used in an aircraft, for example in a fuselage, nacelle, wing, over-wing fairing, belly fairing or tail cone 204. The skin may have an inner 24,210 and an outer profile 26,212, the profile of the batteries may be shaped to match either of the skin’s profiles, and be close to, preferably touching, the inner profile of the skin, alternatively the battery may itself form a surface of the skin. The skin may have a plate to conduct heat away from the battery along a conductive path toward a cooling fluid temperature regulation region, such as an external airflow. The heat transfer plate may extend partially, or completely, through the skin towards the battery. The system may allow for passive cooling of aircraft batteries in a compact, space-saving manner.

Description

SYSTEMS AND METHODS FOR AIRCRAFT
[0001] The present invention relates to systems and methods for aircraft. In particular, the present invention relates to systems such as thermal management systems, and associated methods.
BACKGROUND
[0002] An aircraft flight includes the following phases: taxiing to the runway; takeoff (which includes the takeoff roll); climb; cruise; descent; final approach; and landing (which includes the landing roll). In the taxiing phase, the aircraft moves to the runway under its own power. The aircraft is positioned on the runway ready for the takeoff roll. In takeoff, thrust from the aircraft propellers or jet engines accelerate the aircraft up to speed during a takeoff roll, and once a suitable speed is achieved the nose of the aircraft is raised to increase lift from the wings and effect take off. In the climb phase, the aircraft climbs to cruise altitude and the aircraft speed is gradually increased. In cruise, the aircraft flies at a required altitude and, typically, at a constant speed. In the descent phase, the aircraft altitude is reduced by increasing drag. In the final approach, the aircraft is aligned with the runway before landing. In the landing phase, the aircraft lands on the runway and performs the landing roll, in which the aircraft is decelerated.
[0003] Hybrid and electric aircraft have been proposed, to make one or more phases of aircraft flight, or general management, more efficient. Hybrid and electric aircraft are driven at least in part by electric components, including storage devices, electric motors and power electronics. In such aircraft, it is beneficial to monitor and maintain component health during operation, which may involve managing and monitoring the component temperature. Storage devices, electric motors and power electronics typically require cooling. It is, however, advantageous not to use large amounts of energy providing cooling to such components. This is because power from batteries or other energy accumulators in electric aircraft will need to be diverted to power cooling systems, which could otherwise be used to propel the aircraft.
[0004] Large loads are placed on energy storage devices during operation, causing them to heat. If unmanaged, the energy storage device can reach temperatures which are dangerous and detrimental both to the function and longevity of the energy storage device, and also to the operation of other nearby components. Commercially available cooling systems are complex and energy intensive.
[0005] Ice protection is necessary to protect the aircraft skin from build-up of ice, which can damage the structure of the aircraft. Ice protection for aircraft wings is of particular importance.
Known systems impose a loss of power on the aircraft, by bleeding air from the aircraft engine, resulting in a loss of thrust. Other known systems are complex and unreliable.
[0006] A more holistic approach to hybrid, or fully electric, aircraft manufacture is required, instead or simply trying to use or adapt existing systems. Nevertheless, adapting existing systems can be advantageous where safety and reliability of aircraft function is ensured. In addition, perhaps more generally, there is a drive to reduce pollution and other negative environmental impacts of aircraft and the aerospace industry as a whole.
[0007] "Turnaround time" is the time between landing and a subsequent takeoff, wherein the aircraft is not in operation. Turnaround time is increased by components which are above a suitable temperature for subsequent takeoff and must be cooled before the aircraft can be allowed to take off. Hybrid and electric aircraft are powered at least in part by batteries. Refuelling such aircraft involves charging batteries. This conventionally takes place on the ground, which also increases turnaround time. Decreasing turnaround time is a high priority for airlines.
[0008] It is an object of the present invention to provide improved systems for aircraft and/or to address one or more of the problems discussed above, or discussed elsewhere, or to at least provide alternative systems.
SUMMARY OF THE INVENTION
[0009] According to the present invention there is provided a thermal management system, and associated method, as set forth in the appended claims. Other features of the invention will be apparent from the dependent claims, and the description which follows.
[0010] According to a first aspect of the present invention there is provided a thermal management system for managing the temperature of one or more components in an aircraft nacelle, the thermal management system comprising a nacelle having a profile and one or more electric components having an outer profile, the outer profile of the electric components being shaped to be conformal with the profile of the nacelle.
[0011] A related aspect might relate to a method of thermal management of one or more components in an aircraft nacelle.
[0012] The electric components may comprise an electric motor and/or power electronics. The nacelle may have an inner profile and an outer profile. The outer profile of the electric components may be shaped to be conformal with the inner or outer profile of the nacelle. The outer profile of the electric components and the nacelle profile may be substantially contiguous.
The outer profile of the electric components may be proximal to the inner profile of the nacelle. The electric components may be in contact with the nacelle.
[0013] The thermal management system may further comprise thermally conductive interface members for interfacing between the electric components and the nacelle, the interface members defining a thermally conductive path such that, when the components are in use, heat produced by the components is conducted away from the components along the conductive path toward the nacelle.
[0014] The interface members may substantially surround the components to form the nacelle. The interface members may be integrally formed with parts of the electric components. The interface members and parts of the electric components may be integrally formed by 3D printing or casting. Heat produced by the electric components may be conducted away from the components along the conductive path toward a temperature regulating region which comprises an external airflow.
[0015] According to a second aspect of the present invention there is provided a nacelle comprising a thermal management system according to the first aspect.
[0016] The nacelle may be comprised in an aircraft wing. The nacelle may be mounted to the aircraft by pylons.
[0017] According to a third aspect of the present invention there is provided an aircraft comprising a thermal management system or a nacelle as claimed in the first or second aspects.
[0018] According to a fourth aspect of the present invention there is provided a thermal management system for managing the temperature of one or more components in an aircraft nacelle, the thermal management system comprising: an inlet opening into the nacelle, an outlet opening out of the nacelle, a gas flow path extending between the inlet and outlet and at least proximal to the components to manage the component temperature, and a flow control system for selectively controlling a mass flow along the gas flow path.
[0019] The flow control system may be operable to control the mass flow along the gas flow path as a function of the component temperature. The flow control system may be operable to control the mass flow leaving the outlet to mitigate the aerodynamic interference caused by the presence of the nacelle.
[0020] The flow control system may comprise a transducer assembly provided in the flow path.
The transducer assembly may be arranged to convert energy of the gas flow into electrical energy. The transducer assembly may comprise a turbine.
[0021] The thermal management system may be arranged to direct external airflow flowing past the nacelle into the inlet. The inlet may comprise a filter to reduce ingress of particulates into the system. The inlet may comprise a louvre to reduce ingress of particulates into the system.
[0022] The one or more components may be comprised in an aircraft propulsion system. The aircraft propulsion system may comprise a plurality of propellers. The one or more components may comprise an electric motor and/or power electronics.
[0023] According to a fifth aspect of the present invention there is provided a nacelle comprising a thermal management system according to the fourth aspect. The nacelle may be comprised in an aircraft wing.
[0024] According to a sixth aspect of the present invention there is provided an aircraft comprising a thermal management system or a nacelle according to the fourth or fifth aspect.
[0025] According to a seventh aspect of the present invention there is provided a method of thermal management of one or more components in an aircraft nacelle, the method comprising the steps of: providing a thermal management system comprising an inlet opening into the nacelle, an outlet opening out of the nacelle, a gas flow path extending between the inlet and outlet and at least proximal to the components; and controlling a mass flow along the gas flow path to control the temperature of the components.
[0026] According to an eighth aspect of the present invention there is provided a thermal management system for managing the temperature of one or more components in an aircraft nacelle, the thermal management system comprising: an airflow generator provided within the nacelle, the airflow generator configured to generate an airflow past components within the nacelle.
[0027] The one or more components may be comprised in an aircraft propulsion system. The one or more components may comprise an electric motor and/or power electronics. The airflow generator may comprise a fan. The thermal management system may comprise a rotatable drive shaft and the airflow generator may be fixed to the drive shaft.
[0028] The drive shaft may be comprised in an aircraft propulsion system. The drive shaft may be a drive shaft for a propeller. The airflow generator may compress the airflow.
[0029] The thermal management system may comprise an inlet opening into the nacelle, an outlet opening out of the nacelle, a gas flow path extending between the inlet and outlet. In use, the airflow generator may draw gas into the nacelle from outside of the nacelle. The gas flow path may pass through the airflow generator, and the airflow generator may be operable to control the mass flow along the gas flow path.
[0030] According to a ninth aspect of the present invention there is provided a nacelle comprising a thermal management system according to the eight aspect.
[0031] The nacelle may be comprised in an aircraft wing.
[0032] According to a tenth aspect of the present invention there is provided an aircraft comprising a thermal management system or a nacelle according to the eighth or ninth aspects.
[0033] According to an eleventh aspect of the present invention there is provided a method of thermal management of one or more components in an aircraft nacelle, the method comprising the steps of: providing a thermal management system comprising an airflow generator provided within the nacelle; and generating an airflow within the nacelle to control the temperature of the one or more components in the nacelle.
[0034] According to a twelfth aspect of the present invention there is provided a thermal management system for managing the temperature of one or more components in an aircraft nacelle, the thermal management system comprising: a rotatable hub of an aircraft propulsion system, the hub comprising an aperture for facilitating airflow into the nacelle via the aperture.
[0035] The hub may comprise a hub duct extending through the hub and the aperture may open into the duct. The aircraft propulsion system may comprise one or more propellers and the hub is a propeller hub.
[0036] The thermal management system may further comprise a fairing provided over the hub, the fairing comprising an aperture for facilitating airflow through the rotatable hub aperture via the fairing aperture. The fairing may comprise a fairing duct extending through the fairing and the aperture opens into the duct.
[0037] The thermal management system may comprise an airflow generator for drawing air into the nacelle through the hub aperture. The fairing duct may comprise channels which, when the fairing is rotated, draws air into the nacelle. The rotating channels may increase the pressure of the airflow. The fairing may be 3D printed and/or cast with the channels integrally formed.
[0038] The thermal management system may comprise an outlet opening out of the nacelle, a gas flow path extending between the and outlet. The one or more components may comprise an electric motor and/or power electronics [0039] According to a thirteenth aspect of the present invention there is provided a nacelle comprising a thermal management system according to the twelfth aspect.
[0040] The nacelle may be comprised in an aircraft wing.
[0041] According to a fourteenth aspect of the present invention there is provided an aircraft comprising a thermal management system or a nacelle according to the eleventh or twelfth aspects.
[0042] According to a fifteenth aspect of the present invention there is provided a method of thermal management of one or more components in an aircraft nacelle, the method comprising the steps of: providing a thermal management system comprising a rotatable hub of an aircraft propulsion system, the rotatable hub comprising an aperture for facilitating airflow into the nacelle via the aperture; and generating an airflow within the nacelle via the aperture to control the temperature of the one or more components in the nacelle.
[0043] According to a sixteenth aspect of the present invention there is provided a thermal management system for managing the temperature of an energy storage device in an aircraft, the thermal management system comprising a skin portion having a profile and the energy storage device having an outer profile, the outer profile of the energy storage device being shaped to be conformal with the profile of the skin portion.
[0044] In one example, the energy storage device comprises a battery pack. In one example the skin portion has an inner profile and an outer profile, and the outer profile of the energy storage device is shaped to be conformal with the inner or outer profile of the skin portion. In one example the outer profile of the energy storage device is proximal to the inner profile of the skin portion. In one example the outer profile of the energy storage device and the skin portion profile are substantially contiguous. In one example the energy storage device is in contact with the skin portion. In one example the skin portion comprises a heat transfer member arranged to conduct heat away from the energy storage device along a conductive path. In one example the heat transfer member is arranged to conduct heat away from the energy storage device toward a temperature regulating region which comprises a cooling fluid flow. In one example the cooling fluid flow of the temperature regulating region is an external airflow. In one example the heat transfer member is integrated into the skin portion profile to extend part way through and/or completely through the skin portion. In one example the heat transfer member extends through the skin portion from the temperature regulating region to the energy storage device. In one example the heat transfer member is a conductive plate. In one example the thermal management system is comprised in an aircraft fuselage, nacelle, wing, over-wing fairing, belly fairing or tail cone.
[0045] According to a seventeenth aspect of the present invention there is provided a skin portion for an aircraft comprising a thermal management system according to the sixteenth aspect.
[0046] According to an eighteenth aspect of the present invention there is provided an aircraft comprising one or more thermal management systems and/or skin portions according to the sixteenth or seventeenth aspects.
[0047] According to a nineteenth aspect of the present invention there is provided a thermal management system for managing the temperature of an aircraft wing, the thermal management system comprising: a gas flow path comprising a first flow path portion and a second flow path portion, the first flow path portion arranged to direct a flow of gas at least proximal to an aircraft component for managing the component temperature, the second flow path portion arranged to receive the flow of gas from the first flow path portion and to direct the flow of gas to the wing for managing the temperature of the wing.
[0048] In one example the thermal management system comprises an arrangement of one or more conduits defining the gas flow path. In one example the second flow path portion is arranged to extend through the leading edge of the wing. In one example the first flow path portion comprises an inlet opening into the first flow path portion. In one example the first flow path portion comprises an internally opening outlet, opening into the second flow path portion, such that the flow of gas is receivable from the first flow path portion by the second flow path portion to direct the flow of gas to the wing. In one example the first flow path portion comprises an externally opening outlet, opening out of the first flow path portion, such that the flow of gas can be discharged from the first flow path portion and the thermal management system. In one example the second flow path portion comprises an externally opening outlet, opening out of the second flow path portion, such that the flow of gas can be discharged from the second flow path portion and the thermal management system. In one example the externally opening outlet of the second flow path portion opens out of the second flow path portion at a location commensurate with the heat profile required on the wing, optionally opening out of the second flow path portion proximal to the wing tip. In one example the thermal management system further comprises a flow control system arranged to selectively control a mass flow of gas along the second flow path portion. In one example the thermal management system further comprises one or more temperature sensors for sensing the temperature of the wing, the flow control system being operable to selectively control the mass flow along the second flow path portion as a function of the sensed temperature. In one example the flow control system is arranged to selectively divert the flow of gas toward the externally opening outlet of the first flow path portion or toward the internally opening outlet of the first flow path portion, thereby to selectively control the mass flow of gas along the second flow path portion. In one example the first flow path portion extends through a nacelle and the component is housed in the nacelle. In one example the component is an electric motor and/or power electronics.
[0049] According to a twentieth aspect of the present invention there is provided an aircraft wing or aircraft comprising a thermal management system according to the nineteenth aspect.
B
[0050] According to a twenty first aspect of the present invention there is provided a method of managing the temperature of an aircraft wing, the method comprising the steps of: providing a gas flow path comprising a first flow path portion and a second flow path portion; directing a flow of gas along the first flow path portion at least proximal to an aircraft component for managing the component temperature; and arranging the second flow path portion to receive the flow of gas from the first flow path portion and directing the flow of gas to the wing for managing the temperature of the wing.
[0051] Although some exemplary embodiments of the present invention are shown and described, it will be appreciated by those skilled in the art that various changes and modifications might be made without departing from the scope of the invention, as defined in the appended claims.
[0052] Additionally, it will be appreciated that the various aspects and embodiments are closely related in terms of concept and technical implementation, and as a result various features of those aspects and embodiments are clearly combinable with one another, and/or may replace one another, unless such combination or replacement would be understood by the skilled person
reading this disclosure to be mutually exclusive.
BRIEF DESCRIPTION OF THE DRAWINGS
[0053] For a better understanding of the invention, and to show how embodiments of the same may be carried into effect, reference will now be made, by way of example only, to the accompanying diagrammatic drawings in which: Figure 1 shows an aircraft according to an embodiment of the present invention; Figure 2 shows an aircraft nacelle comprising a thermal management system according to an embodiment of the present invention; Figure 3 shows a general thermal management system according to an embodiment of the present invention; Figure 4 shows a general thermal management system according to an embodiment of the present invention; Figure 5 shows general methodology principles associated with a thermal management system according to an embodiment of the present invention; Figure 6 shows an aircraft nacelle comprising a thermal management system according to an embodiment of the present invention; Figure 7 shows a general thermal management system according to an embodiment of the present invention; Figure 8 shows general methodology principles associated with a thermal management system according to an embodiment of the present invention; Figure 9 shows an aircraft nacelle comprising a thermal management system according to an embodiment of the present invention; Figure 10 shows a general thermal management system according to an embodiment of the present invention; Figure 11 shows general methodology principles associated with a thermal management system according to an embodiment of the present invention; Figure 12 shows a thermal management system according to an embodiment of the present invention; Figure 13 shows a thermal management system according to an embodiment of the present invention; and Figure 14 shows general methodology principles associated with a thermal management system of the present invention.
DETAILED DESCRIPTION
[0054] Referring to Figure 1, an aircraft 1 comprises a fuselage 2, wings 4, an empennage or tail assembly 6 and landing gear comprising wheels 8. The aircraft further comprises a propulsion system comprising an air propulsion system 10. The air propulsion system 10 comprises propellers 12 and air propulsion system motors 14 arranged to drive the propellers 12. The air propulsion system 10 is connected to a power supply 16 comprising an energy accumulator, which in this exemplary embodiment, comprises a battery 18. The battery 18 is configured to supply power to the air propulsion system motors 14. The aircraft is therefore an electric, or at least hybrid, aircraft.
[0055] In this exemplary embodiment the air propulsion system 10 comprises four propellers 12. Nevertheless, it will be appreciated by the person skilled in the art that the air propulsion system 10 may comprise any suitable number of propellers 12 and associated air propulsion system motors 14. For example, the air propulsion system 10 may comprise one, two, three, four, five or six propellers 12. It is highly advantageous for propellers 12 to be positioned such that the force produced by the air propulsion system 10 is substantially symmetric either side of the aircraft fuselage 2.
[0056] Referring to Figure 2, each air propulsion system motor 14 is housed, or received, in a nacelle 20. Power electronics 22 relating to each motor 14 is also housed in the nacelle 20 alongside the motor 14. Whilst the power electronics 22 can alternatively be housed elsewhere in the aircraft 1, it is advantageous to house them alongside the motor 14 in the nacelle 20 so that the temperature of the power electronics 22 and motor can be managed using thermal management systems provided for both components, as will be described in greater detail below.
[0057] A nacelle is perhaps typically understood to relate to or be a streamlined casing (or pod) on the outside of an aircraft, especially one housing an aircraft engine or motor, or other (e.g. propulsion system) components. However, a nacelle is alternatively or additionally described or defined as a housing, or streamlined casing, for an aircraft, in general, especially one housing an aircraft engine or motor, or other (e.g. propulsion system) components. That is, a nacelle does not need to be 'outside' of an aircraft or separate to an aircraft fuselage. For example, a nacelle may be located in a nose of an aircraft, in a tail of an aircraft, or in a wing, tail or fin of an aircraft. For example, a nacelle may be integral to the main body or fuselage of an aircraft, such as in the tail assembly or nose cone, and hold or house propulsion components, motors, power supplies, electronics, or other equipment on an aircraft including motors and power electronics (i.e. electronic components, or components in general associated with management or propulsion of the aircraft).
[0058] A thermal management system 100 is devised for managing the temperature of components housed, or received, in the nacelle 20. The thermal management system 100 functions to manage the temperature of components in dependence upon the ambient temperature external to the aircraft 1. Such a construction is advantageous as it may eliminate the need for additional thermal management systems for the aircraft motors 14 and power electronics 22. Moreover, a cooling flow of air will be constantly present during flight, or motion in general, thus allowing the motors 14 to be cooled in a manner which does not require additional energy input. In exemplary embodiments, the thermal management system 100 is a cooling system for cooling one or more components in an aircraft nacelle 20.
[0059] Depending on ambient temperatures, the thermal management might involve heating of components, for example when ambient temperatures are higher than temperatures in the nacelle, or temperate of components in the nacelle.
[0060] Each nacelle 20 has an inner profile and an outer profile defined by inner surfaces 24 and outer surfaces 26 of the nacelle 20. Each air propulsion system motor 14 and power electronics unit 22 has an outer profile defined by outer surfaces 28 of the motor 14 and outer surfaces 29 of the power electronics unit 22. In the thermal management system 100, the outer profile, or outer surfaces 28, of each motor 14 is shaped to be conformal with the outer profile, or outer surface 26 of its respective nacelle 20. The outer profile, or outer surfaces 29, of each power electronics unit 22 is shaped to be conformal with the outer profile, or outer surface 26, of its respective nacelle 20. In another exemplary embodiment, the outer profile, or outer surfaces 28, of each motor 14 is shaped to be conformal with the inner profile, or inner surface 24 of its respective nacelle 20. The outer profile, or outer surfaces 29, of each power electronics unit 22 is shaped to be conformal with the inner profile, or inner surface 24, of its respective nacelle 20.
[0061] As explained below, conformal means, in general, that the two respective components take the same general shape. The components might be in contact, or close proximity, or contiguous, sufficient for thermal transfer to be effective.
[0062] A conformal motor 14 and power electronics unit 22 is advantageous as it not only reduces the spatial footprint of the nacelle 20, thus improving aerodynamics, but also ensures that the motor 14 and power electronics unit 22 are close to, or at, or even forming, a surface of the nacelle 20, thus improving transfer of heat from the motor 14 and electronics unit 22 to the surrounding air. A thermal management system 100 can therefore be provided wherein the motor 14 and power electronics unit 22 are exposed to an external flow of air. An additional advantage is that the close fitting of motor 14, power electronics unit 22 and nacelle 20 facilitates easier access to the motor 14 for maintenance and inspection.
[0063] The exterior profile of the nacelle 20 is defined by the nacelle outer mould line 26. In an exemplary embodiment, the outer profile, or exterior surface 28, of each motor 14 is shaped to be conformal with the exterior surface 26 of its respective nacelle 20, such that the outer profiles are contiguous, and the nacelle 20 has a continuous outer skin. This may also be described as the motor 14 being at least partially at the nacelle outersurface 26, or forming that surface. Such a construction is advantageous as part of the motor 14 is directly in contact with ambient airflow around the nacelle 20. In another exemplary embodiment, the outer profile of each motor 14 is shaped to be conformal with an interior profile of the nacelle 20 to ensure that the motor 14 is proximal to the nacelle 20 which substantially surrounds the motor 14. That is, the outer profile of each motor 14 closely follows the interior profile of the nacelle 20. Benefits are obtained by the motor 14 being proximal to, or closely following, the nacelle 20 in that conductive transfer of heat is improved. In another exemplary embodiment, the outer profile of each motor 14 is shaped to ensure that the motor 14 is in contact with the nacelle 20, at least at the points wherein significant heat is generated. The points of contact facilitate the conduction of heat.
[0064] In this exemplary embodiment, the outer profile of each motor 14 is defined by thermally conductive interface members 30 which are 3D printed manifolds extending from various parts of the motor 14. Interface members 30 also extend from various parts of the power electronics 22. The interface members 30 are thus integrally formed with the motor 14 and power electronics 22. The interface members 30 substantially surround the motor 14 and power electronics 22. Whilst 3D printed manifolds are comprised in this exemplary embodiment, casting could also be used to manufacture integrally formed components and interface members 30.
[0065] The interface members 30 are conformal with the exterior surface 26 of the nacelle 20, such that the interface members 30 are contiguous with the exterior profile of the nacelle 20, thereby forming pad of the nacelle outer mould line 26. The interface members 30 therefore provide a thermally conductive path from the motor 14 to an external airflow around the nacelle 20. When the motor 14 is in use, the heat produced by the motor 14 is conducted away from the motor 14 along the conductive path. The external shape of the motor 14 is therefore adapted to form the external skin of the nacelle 20. The interface members 30 facilitate a weight saving, as well as a cooling benefit, as a separate nacelle 20 (or nacelle region or portion) need not be provided in addition to the interface members 30 to form a conductive path between the motor 14 and nacelle 20. That is, in an exemplary embodiment, the motor 14, in effect, is the nacelle 20.
[0066] In an exemplary embodiment, an exterior surface 32 at the rear of the nacelle 20 is comprised (i.e. located) in the aircraft wing 4. The exterior surface 28 of each air propulsion system motor 14 is shaped to be conformal (e.g. contiguous) with the profile of the wing in the region of the exterior surface 32 at the rear of the nacelle 20. The conformal, or matching, shape of the outer profile of the motors 14 and wings 4 reduces drag and provides cooling due to airflow over the wings 4, and by extension, the nacelles 20. In another exemplary embodiment, the nacelles are mounted to the aircraft by pylons extending below the wings, or from the fuselage 2 or tail assembly 6.
[0067] In Figure 3, a general thermal management system 100 according to an embodiment of the present invention is shown. The thermal management system 100 is for managing the temperature of one or more components in an aircraft nacelle 20. The thermal management system 100 comprises a nacelle 20 having a profile and one or more electric components 14, 22 having an outer profile, the outer profile of the electric components 14,22 being shaped to be conformal with the profile of the nacelle 20.
[0068] Referring back to Figure 2, the thermal management system 100 comprises an inlet 40, opening into internal volume of the nacelle 20, and an outlet 42. In an exemplary embodiment, the inlet 40 and outlet 42 are of the type developed by the National Advisory Committee for Aeronautics (NACA). A gas flow path extends between the inlet 40 and outlet 42 through the nacelle and passes over and around the motor 14 and power electronics 22 housed within the nacelle 20. In this example the inlet 40 is shaped to provide an axisymmetric flow, which ensures that shear forces are not applied to the nacelle 20. For example, inlets 40 and/or outlets 42 might be distributed evenly or uniformly, or symmetrically, about the nacelle 20. Additionally, or alternatively, the inlet 40 can be shaped to provide cooling to specific components, or to maximise the positive pressure at the inlet 40. The inlet 40 and outlet 42 are arranged to reduce ingress of particulates into the nacelle 20. The inlet 40 and outlet 42 are positioned on an upper surface of the nacelle 20. The inlet 40 comprises a louvre filter to reduce ingress of particulates into nacelle 20. The outlet 42 might have similar functionality. The inlet 40 also comprises a recirculation system for improved protection from freezing and ice build-up, which is typically an electric recirculation system, but may be a gas recirculation system. The motor 14 and power electronics 22 are provided with temperature sensors. The motor 14 and power electronics 22 are provided downstream in the gas flow path, so that the gas (e.g. air) can flow through the inlet 40 and over or around the motor 14 and power electronics 22, before leaving the nacelle 20 via outlet 42. Such a thermal management system is advantageous in ensuring the motor and power electronics provided in the nacelle do not overheat.
[0069] A flow control system 44 is built into the outlet 42. The flow control system 44 is operable to selectively control a mass flow along the gas flow path. In an exemplary embodiment, the flow control system 44 comprises an autonomous aperture control to throttle the mass flow to achieve a desired motor and/or power electronics temperature. Opening the aperture allows a greater mass flow through the nacelle 20, thus cooling the components therein. Conversely, closing the aperture reduces mass flow through the nacelle 20, allows the components housed therein to heat. Such "throttling" of the air flow allows the system to ensure that component temperature is always optimum. In a related example, such a flow control system could also, or alternatively, be located at the inlet 40.
[0070] In this example, the flow control system 44 comprises turbines comprising turbine motors. The turbines are provided in addition to, or in place of, the autonomous aperture control. The turbines rotate creating suction through the nacelle 20, which draws air through the nacelle 20 at a greater rate. The rate of airflow through the nacelle 20 is thus a function of the rotational velocity of the turbines. By increasing the rate of airflow, the motor 14 and power electronics 22 can be cooled more rapidly, and by decreasing the rate of airflow, they can be cooled less rapidly, the temperature maintained, or the components 14, 22 allowed to heat. The temperature of the components 14,22 is monitored, and an appropriate load (e.g. electrical or mechanical) is placed on the turbine motors to increase or decrease the rotational velocity of the turbines, to thereby increase or decrease the rate of airflow through the nacelle 20 as necessary or desired. The turbine motors are, in this example, connected to a battery, such that a greater motor load produces an increased rate of battery charging and a reduced motor load produces a reduced rate of battery charging. Energy recovery is thus facilitated. Whilst in this exemplary embodiment Li-Ion batteries are employed, supercapacitors and flywheels are also a feasible alternative energy accumulators or energy storage means.
[0071] Exhaust air from the outlet 42 opening out of the nacelle 20 is ejected upwardly out of the nacelle 20 onto the wing upper surface. The aerodynamic interference caused by the presence of the nacelle 20 can thus be controlled or offset, using the air that emerges from the nacelle outlet 42. The low pressure on the upper surface of the wing 4 will augment the flow rate and provide the maximum pressure difference to induce airflow through the nacelle 20. At low aircraft forward speed, induced flow over the wing is used to relieve flow separations around and behind the nacelle. At high aircraft forward speed, induced flow over the wing can be managed by shaping of the nacelle, such as conformal shaping.
[0072] In Figure 4, a general thermal management system 110 according to an embodiment of the present invention is shown. The Figures are used to describe different general thermal management system and methods, but it will be appreciated that these might relate to the same or closely related system or method. Different reference numerals have been used, simply to denote different general functions.
[0073] The thermal management system 110 is for managing the temperature of one or more components 14, 22 in an aircraft nacelle 20. The thermal management system comprises an inlet 40 opening into the nacelle 20, an outlet 42 opening out of the nacelle 20, a gas flow path extending between the inlet 40 and outlet 42 and at least proximal to the components to manage the component temperature, and a flow control system 44 for selectively controlling a mass flow along the gas flow path.
[0074] In Figure 5, general methodology principles associated with a thermal management system 110 according to an embodiment of the present invention are shown. Step 46 comprises providing a thermal management system comprising an inlet 40 opening into the nacelle 20, an outlet 42 opening out of the nacelle 20, a gas flow path extending between the inlet 40 and outlet 42 and at least proximal to the components 14, 22. Step 48 comprises controlling a mass flow along the gas flow path to control the temperature of the components 14,22.
[0075] Referring to Figure 6, an airflow generator 50 is provided in the nacelle 20. The airflow generator 50 is positioned forward of the motor 14 and power electronics in the nacelle 20. The airflow generator 50 is configured to generate an airflow past components within the nacelle 20. In this exemplary embodiment, the airflow generator 50 is positioned forward of the components and is therefore configured to provide an airflow past the components, the airflow being toward a rear of the nacelle 20. Nevertheless, it will be appreciated that the airflow generator 50 could be positioned rearward of the components, and therefore configured to draw airflow past the components from a different position.
[0076] The airflow generator 50 comprises a fan 52 operably connected to the rotating drive 35 shaft 54 of the motor 14. The drive shaft 54 extends through the central hub of the fan 52 such that the fan 52 and drive shaft 54 are coaxial, and the fan 52 is fixed to the drive shaft 54 such that rotation of the drive shaft 54 induces equivalent rotation of the fan 52. In other words, the airflow generator 50 is directly driven by the drive shaft 54. This means that there is synergy, in that driving of the propellers 12 will cause heating of the motor 14, but such driving also causes rotate of fan 52, which serves to cool the motor 12. The inlets and outlets of the previous embodiment may be useful for facilitating such an airflow, or an improved airflow.
[0077] In an exemplary embodiment, the thermal management system 110 comprises an airflow generator 50, an inlet 40 opening into the nacelle 20 and an outlet 42 opening out of the nacelle 20 defining a gas flow path, as described above. The airflow generator 50 is positioned upstream of the motor 14 and power electronics 22 in the gas flow path. Operation of the airflow generator 50 creates a region of lower pressure within the nacelle 20, which draws air through the inlet 40 into the nacelle 20 from outside the nacelle 20. In an exemplary embodiment, the airflow generator 50 comprises a radial compressor operably connected to the drive shaft 54. Cooling of the motor 14 and power electronics 22 is critical to performance, and such a system can induce improved air flow through the nacelle when compared with a simpler inlet/outlet, or "ram air", system. Rotational velocity of the airflow generator 50 is controllable to control the mass flow of air through the nacelle 20.
[0078] In one example, the airflow generator could take the form of an internal surface (that is, facing into the nacelle) of a hub for the propeller, such that a dedicated, separate component (e.g. a fan) is not required. The internal surface could be shaped or textured to provide or drive such air flow.
[0079] In Figure 7, a general thermal management system 120 according to an embodiment of the present invention is shown. The thermal management system 100 is for managing the temperature of one or more components 14, 22 in an aircraft nacelle. The thermal management system 100 comprises an airflow generator 50 provided within the nacelle 20, the airflow generator 50 configured to provide an airflow past components 14,22 within the nacelle 20.
[0080] In Figure 8, general methodology principles associated with a thermal management system 120 according to an embodiment of the present invention are shown. Step 56 comprises providing a thermal management system comprising an airflow generator provided within the nacelle. Step 58 comprises generating an airflow within the nacelle to control the temperature of the one or more components in the nacelle.
[0081] Referring to Figure 9, the air propulsion system 6 comprises propellers 12 and air propulsion system motors 14 arranged to drive the propellers 12. The propellers 12 are connected to the motors 14 via a drive shaft 54 and the central hub 60 of the propeller 12 is fixed to the drive shaft 54 such that rotation of the drive shaft 54 induces propeller rotation. A substantially conical fairing 62, or nose cone, is provided over the hub 60. The fairing tip is the foremost point of the propeller/nacelle construction. Thus, in flight, it experiences maximum positive pressure or "stagnation pressure".
[0082] The fairing 62 and hub 60 comprise apertures 64 for facilitating airflow into the nacelle 20 via the apertures 64. The apertures 64 open into flow paths through the hub 60 and fairing 62. The hub 60 comprises an inlet opening into a hub duct defining a flow path formed within the hub, and an outlet opening out of the duct into the nacelle 20. The fairing 62 comprises an inlet opening into a fairing duct defining a flow path formed within the fairing 62, and an outlet opening out of the duct into the hub aperture. The fairing 62 is 3D printed to form a fairing 62 that is hollow and comprises airflow channels 66 integrally formed within the ducting. The channels 66 allow high pressure air to flow through the fairing 62, around the bearings of the hub 60, and into the nacelle 20 to provide a cooling airflow. In an exemplary embodiment, and as alluded to above, the internal channels 66, or ducting, is shaped to form an axial compressor. The fairing 62 is fixed over the propeller hub 60 and rotation of the fairing energises the airflow. Whilst in this exemplary embodiment the fairing 62 is 3D printed, it could be cast to form the fairing comprising integrally formed airflow channels 66.
[0083] It is notable that such a construction enables cooling air flow into the nacelle 20 even when the aircraft 1 is not in flight, or in motion. During a taxiing phase of flight, the aircraft 1 is often required to wait until space becomes available on the runway for takeoff. The aircraft 1 sits with the propellers 12 rotating, but the brakes applied such that the aircraft 1 has no forward velocity. By the provision of a propeller fairing 62 comprises channels 66 for drawing air into the nacelle, rotation of the propellers 12 creates a cooling airflow through the nacelle 20 even when the aircraft 1 has no forward velocity.
[0084] Whilst in this exemplary embodiment both the fairing 62 and hub 60 comprise apertures 64, it will be appreciated by the person skilled in the art that a fairing 62 need not be provided, and the hub 60 may be shaped, or comprise integrally formed and appropriately shaped manifolds, to form a hub 60 with a conical frontward portion.
[0085] In Figure 10, a general thermal management system 130 according to an embodiment of the present invention is shown. The thermal management system 130 is for managing the temperature of one or more components 14, 22 in an aircraft nacelle 20. The thermal management system comprises a rotatable hub 60 of an aircraft propulsion system 6, the hub comprising an aperture 64 for facilitating airflow into the nacelle 20 via the aperture 64.
[0086] In Figure 11, general methodology principles associated with a thermal management system 130 according to an embodiment of the present invention are shown. Step 68 comprises providing a thermal management system comprising a rotatable hub of an aircraft propulsion system, the rotatable hub comprising an aperture for facilitating airflow into the nacelle 20 via the aperture 64. Step 70 comprises generating an airflow within the nacelle via the aperture to control the temperature of the one or more components in the nacelle. This could be via general airflow, airflow due to movement of the aircraft, or driven airflow (e.g. using an airflow generator).
[0087] Turning now to Figure 12, and with reference to Figure 1, a thermal management system is shown. The thermal management system 200 is for managing the temperature of energy storage devices, which in this exemplary embodiment is a battery pack 202 located in the tail cone 204 (only a portion of a cross-section of which is shown in Figure 12). The energy storage device may additionally or alternatively be a fuel cell 201. The thermal management system 200 functions to manage the temperature of the battery packs 202 in dependence upon the ambient temperature external to the aircraft 1. This construction is advantageous as it may eliminate the need for additional thermal management systems for the energy storage devices 202, for example additional thermal management systems that draw power from on-board power supplies. Moreover, a cooling flow of air is constantly present during flight, or motion in general, thus allowing the batteries 202 to be cooled in a manner which does not require additional energy input.
[0088] The thermal management system 200 comprises a skin portion 208 of fuselage 2 and tail cone 204. As illustrated in Figure 12, the skin portion208 has an inner profile and an outer profile defined by inner surfaces 210 and outer surfaces 212 of the skin portion 208. Each battery 202 has an outer profile defined by outer surfaces 214 of the batteries 18, 202. In the thermal management system 200, the outer profile, or outer surfaces 214, of the battery 202 is shaped to be conformal with the profile of the skin portion. In one exemplary embodiment, the outer surface is shaped to be conformal with the inner profile, or inner surface, of the skin portion 208. In another exemplary embodiment, the outer surface of the battery 202 is shaped to be conformal with the outer profile, or outer surface, of the skin portion. The thermal management system 200 does not require a power source, and passive cooling of battery packs 202 is thus enabled. Shape conformity might also allow for a more compact overall design, or space can be saved for other components.
[0089] Whilst in this exemplary embodiment, the thermal management system 200 is for managing the temperature of battery pack 202 located in tail cone 204, it will be understood by the person skilled in the art that the thermal management system 200 may be used to manage the temperature of battery pack 18 installed in fuselage 2, and conformal to skin portion 206, as shown in Figure 1. That is, the outer profile, or outer surfaces, of battery 18 can be shaped to be conformal with the outer profile of skin portion 16. Further modifications and adaptations are consistent with this disclosure. Moreover, in a further advantageous embodiment, the thermal management system 200 is for managing the temperature of battery pack 18 installed in the wing. That is, the thermal management system 200 may be employed in the belly fairing skin contour, or in the over wing fairing 4 for high wing configurations. It will become clear from the description that installing and shaping the battery pack 18 to the belly fairing skin contour or over-wing fairing facilitates efficient cooling.
[0090] As explained herein, and above, conformal means, in general, that the two respective components take the same general shape. The components might be in contact, or close proximity, or contiguous, for sufficient for thermal transfer to be effective. Conformal may also mean that the external shape of the component is continuous (up to and including the second derivative) with the outer shape or profile of the aircraft skin.
[0091] A conformal energy storage device is advantageous as it not only reduces the spatial footprint of the fuselage 2 and tail cone 204, thus improving aerodynamics, but also ensures that the battery pack 202 is close to, or at, or even forming, a surface of the skin portion, thus improving transfer of heat from the battery pack 202 to the surrounding air. In examples where the fuselage 2 shape is restricted, however, the conformal energy storage device remains advantageous by virtue of the improved cooling. A thermal management system 200 can therefore be provided wherein the battery pack 202 are exposed to an external flow of air. An additional advantage is that the close fitting of battery pack 202 and skin portion facilitates easy access to the battery pack202 for maintenance and inspection. For clarification, the battery pack 202 "forming" a surface of the skin portion allows the battery pack to be referred to as a structural battery. That is, the battery pack forms part of the aircraft structure.
[0092] The exterior profile of the fuselage 2 and tail cone 204 is defined by the outer mould line. In an exemplary embodiment, the outer profile, or exterior surface, of the battery pack202 is shaped to be conformal with the exterior surface of the skin portion where it is located, such that the outer profiles are contiguous, or continuous, and the fuselage 2 or tail cone 204 has a continuous outer skin. This may also be described as the battery being at least partially at the skin portion outer surface, or forming that surface. Such a construction is advantageous as part of the battery is directly in contact with ambient airflow around the outer mould line. In another exemplary embodiment, the outer profile of each battery pack 202 is shaped to be conformal with an interior profile of the skin portion to ensure that the battery is proximal to the skin portion which substantially surrounds the battery. That is, the outer profile of each battery closely follows the interior profile of the skin portion. Benefits are obtained by the battery pack 202 being proximal to, or closely following, the nacelle in that conductive transfer of heat is improved. In another exemplary embodiment, the outer profile of each battery is shaped such that the battery pack 202 is in contact with the skin portion, at least at the regions wherein significant heat is generated. The regions of contact facilitate conduction of heat from the battery pack 202 to the skin portion, and thus to the external airflow. In another exemplary embodiment, an airflow may be provided by cooperation with an internal airflow system, which may be an energy recovery system.
[0093] A heat transfer member 216 in the form of a conductive plate may be arranged between the battery 202 and the skin portion 208, or form part of the outer surface of the battery 202. The heat transfer member is arranged to conduct heat away from the battery 202 toward the skin portion 208. The heat transfer member 216 defines a conductive path to facilitate this. The heat transfer member 216 is conformal to the outer profile of the battery 202 and the inner profile of the skin portion 208. In an exemplary embodiment, the heat transfer member 216 is integrated into the skin portion profile to extend part way through the skin portion 208. That is, the heat transfer member 216 is embedded in the skin portion 208. Such a construction maximises contact with the skin portion, ensuring good thermal contact between the battery pack202 and skin portion 208. Efficient transfer of heat is thus facilitated. In another exemplary embodiment, the heat transfer member 216 is integrated in the skin portion 208 to extend completely through the skin portion 208, and may be contiguous to the skin portion outer surface. Direct contact with the external cooling airflow results, advantageously improving heat transfer. In another example, the heat transfer member 216 may be integrated into, the battery 202.
[0094] The concept might at first appear simple. However, it is powerful and advantageous, and is a good demonstration of the more holistic approach to hybrid, or fully electric, aircraft design and manufacture disclosed herein, as opposed to simply forcing electric principles into combustion-focussed designs.
[0095] Turning now to Figure 13, a thermal management system 220 is shown. The thermal management system 220 is for managing the temperature of the aircraft wing 4. Managing the temperature of the wing may involve selectively heating or cooling the wing. Ultimately, in this example, thermal management system 220 is operable to, in use, manage the temperature of the wing to prevent build up of ice.
[0096] As can be seen from the figure, the system 220 is substantially symmetrical either side of nacelle 20. Although reference will be made to the drawing in the singular, it will be appreciated by the skilled person from the description and drawings that there may be a plurality of systems either side of the nacelle 20 to manage the temperature of wing 4 portions either side of the nacelle 20.
[0097] The thermal management system 220 comprises an arrangement of conduits 222 defining a gas flow path. A first conduit 222a extends through the nacelle 20, for example from front to back. The first conduit 222a defines a first gas flow path portion 224a. A second conduit 222b is in fluid connection with the first conduit 222a. The second conduit 222b extends from the nacelle 20, and through the aircraft wing 4. The second conduit 222b defines a second gas flow path portion 224b.
[0098] As described above in relation to Figure 2, air propulsion system motors 14 and power electronics 22 are received in a nacelle 20. The first gas flow path portion 224a extending through the nacelle is arranged to direct a flow of gas across and around the motor 14 and power electronics 22. A transfer of heat is thus facilitated between the gas flowing through the first conduit 222a and the components 14, 22. This is advantageous as it enables thermal management of the motor 14 and power electronics 22 by cooling the components 14, 22, and, advantageously heating the gas for subsequent use in thermal management of the wing 4.
[0099] In one exemplary embodiment, the gas flow is created by a turbine (not shown) located in the first conduit 222a, which is drivable to induce a flow of gas through the conduits 222a, 222b. In another exemplary embodiment, the first gas flow path portion 224a comprises an inlet 226 opening into the first gas flow path portion 224a from a frontward (that is, with respect to aircraft travel direction) surface. That is, alternatively, the inlet may be inlet 40 described in relation to Figure 2, or may be aperture 64 and channels 66 opening into the nacelle 20 through the propeller hub 60, as described in relation to Figure 9. This construction is highly advantageous as air can be directed into the nacelle 20 and a gas flow facilitated through the conduits 222a, 222b with minimal or no additional energy input to induce gas flow.
[00100] A first outlet 228 opens from the first gas flow path portion 224a (that is, from the first conduit 222a) into the second gas flow path portion 224b (that is, into the second conduit). Such an outlet may therefore be described as an "internally" opening outlet 228, because it opens internal to the conduit arrangement. Thus, the second gas flow path portion 224b is in fluid connection with the first gas flow path portion 224a, and the flow of gas is receivable from the first flow path portion 224a by the second gas flow path portion 224b to direct the flow of gas to the wing. The second gas flow path portion 224b connected to the first gas flow path portion 224a downstream of the region in which the first gas flow path portion 224a passes the motor 14 and power electronics 22, such that gas entering the second gas flow path portion 224b is gas which has been heated by transfer of heat between the components and the gas flow upstream in the first conduit 222a. This is advantageous as the heated gas can be directed to the wing 4 for thermal management of the wing 4, in particular low-pressure regions.
[00101] The second conduit 222b extends through the leading edge of the wing 4. The second gas flow path portion 224b is thus configured to direct a flow of heated gas through the leading edge of the wing 4. Advantageously, this causes heating of the leading edge of the wing 4, on which it is highly desirable to prevent build-up of ice, or to remove ice (e.g. prior to flight). The leading edge of the wing 4 may also house power cables (not shown), and the gas flow along the second gas flow path portion 224b may also manage the temperature of such cables, by cooling or heating (depending on relative temperatures between the gas flow and cables).
[00102] An outlet 230 opens out of the second flow path portion 224b (that is, from the second conduit 222b) to facilitate the expulsion or discharge of gas from the second conduit 222b to the environment, e.g. from the wing. Such an outlet may therefore be described as an "externally" opening outlet 230, because it opens external to the conduit arrangement 222.
Advantageously, the heated gas which passes along the second conduit 222b is used to heat the wing 4 On particular the leading edge), the gas transferring its heat to the wing 4, and the cooled or cooler gas can be discharged from the wing 4 via the outlet opening from the second conduit 222b. In one example, the outlet 230 is provided at a downstream end of the second conduit 222b, distal the first outlet 228. That is, outlet 230 may be proximal to the wing tip 232.
Advantageously, this allows the heated gas to travel the full length of the wing 4, ensuring thermal management of the full wing length. The outlet 230 may be located such that a pressure difference is created along the second conduit 222b, to maximise the draw of gas from the first conduit 222a to the second conduit 222b. Advantageously, this construction is robust and reliable, and may mean that no active (e.g. turbine) propulsion of the air flow is required.
Furthermore, discharging gas at the wing tip is advantageous in mitigating negative aerodynamic effects or turbulent flows near the aircraft. That is, gas may be discharged at relevant locations along the aircraft wing commensurate with the heat profile required on the wing, in particular on the wing leading edge. Additionally or alternatively, outlets 230 may specifically be provided at regions of low pressure to facilitate airflow through the conduits.
[00103] An additional, second, outlet 234 opens from the first gas flow path 224a (that is, from the first conduit 222a) to facilitate the expulsion or discharge of gas from the first conduit 222a to the environment, e.g. from the nacelle 20. Such an outlet 234 may therefore be described as externally opening outlet 234, opening out of the first flow path portion. The second outlet allows a flow of gas to be discharged from the first flow path portion 224a (and from the thermal management system 220), without the gas propagating through the second flow path portion 224b. The second outlet 234 is formed in a rearward surface of the nacelle 20.
[00104] The thermal management system 220 further comprises a flow control system arranged to selectively control a mass flow of gas along the second gas flow path portion 224b. The flow control system comprises a valve 240, or other flow regulator, positioned at the junction between the first conduit 222a and second conduit 222b. The valve 240 is controllable to selectively control a mass flow of gas which is allowed to enter the second conduit 222b to propagate along the second flow path portion 224b. In one exemplary embodiment, an arrangement of temperature sensors 242 may be provided on the wing leading edge, and the valve 240 may be controllable in dependence upon, or as a function of, the sensed temperature.
For example, when a temperature below a threshold temperature is sensed by the sensors 242, the valve 240 may be controlled to allow an appropriate mass flow of gas along the second conduit 222b to facilitate heating of the wing, in particular, the leading edge of the wing. On the contrary, when a temperature above a threshold temperature is sensed by the sensors 242, the valve may be controlled to cause a mass flow of gas to pass straight through the first conduit 222a and discharged from the second outlet 234. Thus, no further wing heating is caused. Advantageously, this selective thermal management ensures that ice build up is prevented, and that the wing surfaces are maintained at a safe operating temperature. In another exemplary embodiment, the valve 240 may be operable under the influence of pressure. Advantageously, this construction is robust and reliable, and does not rely on feedback from other systems.
[00105] The flow control system is therefore arranged to divert gas from the first conduit 222a into the second conduit 222b. That is, the flow control system is arranged to selectively divert the flow of gas toward the externally opening outlet 234 of the first flow path portion 224a, or to allow a flow of gas through the internally opening outlet 228 of the first flow path portion 224a. This allows selective control of the mass flow of gas along the second flow path portion, and thus allows thermal management of the wing to be performed. That is, heating of the wing 4 is controllable using a gas flow which has been used to cool the motor 14, power electronics 22, and generators, thus heating the gas, and then the gas can be selectively allowed to pass into the second conduit 222b for heating the leading edge of the wing 4. The cooled air is then discharged from the wing tip, which advantageously minimises aerodynamic disruption of the thermal management system 220. The system 220 is reliable, and ensures that the wing is maintained at a safe operating temperature. Furthermore, it does not result in a loss of motor thrust or a loss of power generation during operation. The system enables a high motor efficiency to be maintained throughout operation, and can eliminate the need for additional flow control systems.
[00106] Figure 14 shows general methodology principles associated with a thermal management system. The method is a method of managing the temperature of an aircraft wing. Step 250 comprises providing a gas flow path comprising a first flow path portion and a second flow path portion. Step 252 comprises directing a flow of gas along the first flow path portion at least proximal to an aircraft component for managing the component temperature. Step 254 comprises arranging the second flow path portion to receive the flow of gas from the first flow path portion and directing the flow of gas to the wing for managing the temperature of the wing.
[00107] Of course, the gas flow management approaches discussed herein are synergistic, in that ambient air can be used to cool components that are too hot, and then resultant heated air can be used to heat components that are too cold.
[00108] In summary, we provide systems for managing the temperature of components, in particular electric components, provided in aircraft nacelles. Such systems facilitate cooling, both whilst the aircraft is in flight, or stationary. Exemplary embodiments of these systems are energy efficient and require little or no electrical power to operate. Other exemplary embodiments facilitate energy recovery whilst also providing cooling. Accurate control of component temperature is facilitated by the provision of flow control systems, airflow generation within the nacelle and propeller hub and fairing design. Systems provided are mechanically simple, robust and reliable. Described systems ensure safe operating temperatures of electronic components and the aircraft body, in particular the wings, and do not impose a loss of thrust or electrical generation. Finally, the systems can be provided in new aircraft or can be retrofitted in existing aircraft to provide improved performance during the various stages of aircraft flight.
[00109] At least some of the example embodiments described herein may be constructed, partially or wholly, using dedicated special-purpose hardware. Terms such as 'component', 'module' or 'unit' used herein may include, but are not limited to, a hardware device, such as circuitry in the form of discrete or integrated components, a Field Programmable Gate Array (FPGA) or Application Specific Integrated Circuit (ASIC), which performs certain tasks or provides the associated functionality. In some embodiments, the described elements may be configured to reside on a tangible, persistent, addressable storage medium and may be configured to execute on one or more processors. These functional elements may in some embodiments include, by way of example, components, such as software components, object-oriented software components, class components and task components, processes, functions, attributes, procedures, subroutines, segments of program code, drivers, firmware, microcode, circuitry, data, databases, data structures, tables, arrays, and variables. Although the example embodiments have been described with reference to the components, modules and units discussed herein, such functional elements may be combined into fewer elements or separated into additional elements. Various combinations of optional features have been described herein, and it will be appreciated that described features may be combined in any suitable combination. In particular, the features of any one example embodiment may be combined with features of any other embodiment, as appropriate, except where such combinations are mutually exclusive.
Throughout this specification, the term "comprising" or "comprises" means including the component(s) specified but not to the exclusion of the presence of others.
[00110] Although a few preferred embodiments have been shown and described, it will be appreciated by those skilled in the art that various changes and modifications might be made without departing from the scope of the invention, as defined in the appended claims.
[00111] Attention is directed to all papers and documents which are filed concurrently with or previous to this specification in connection with this application and which are open to public inspection with this specification, and the contents of all such papers and documents are incorporated herein by reference.
[00112] All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive.
[00113] Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features.
[00114] The invention is not restricted to the details of the foregoing embodiment(s). The invention extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.

Claims (15)

  1. CLAIMS1. A thermal management system for managing the temperature of an energy storage device in an aircraft, the thermal management system comprising a skin portion having a profile and the energy storage device having an outer profile, the outer profile of the energy storage device being shaped to be conformal with the profile of the skin portion.
  2. 2. A thermal management system as claimed in claim 1 wherein the energy storage device comprises a battery pack.
  3. 3. A thermal management system as claimed in either of claims 1 or 2 wherein the skin portion has an inner profile and an outer profile, and the outer profile of the energy storage device is shaped to be conformal with the inner or outer profile of the skin portion.
  4. 4. A thermal management system as claimed in any preceding claim wherein the outer profile of the energy storage device is proximal to the inner profile of the skin portion.
  5. 5. A thermal management system as claimed in any preceding claim wherein the outer profile of the energy storage device and the skin portion profile are substantially contiguous.
  6. 6. A thermal management system as claimed in any preceding claim wherein the energy storage device is in contact with the skin portion.
  7. 7. A thermal management system as claimed in any preceding claim wherein the skin portion comprises a heat transfer member arranged to conduct heat away from the energy storage device along a conductive path.
  8. 8. A thermal management system as claimed in claim 7 wherein the heat transfer member is arranged to conduct heat away from the energy storage device toward a temperature regulating region which comprises a cooling fluid flow.
  9. 9. A thermal management system as claimed in claim 8 wherein the cooling fluid flow of the temperature regulating region is an external airflow.
  10. 10. A thermal management system as claimed in any of claims 7 to 9 wherein the heat transfer member is integrated into the skin portion profile to extend part way through and/or completely through the skin portion.
  11. 11. A thermal management system as claimed in any of claims 7 to 10 wherein the heat transfer member extends through the skin portion from the temperature regulating region to the energy storage device.
  12. 12. A thermal management system as claimed in any of claims 7 to 11 wherein the heat transfer member is a conductive plate.
  13. 13. A thermal management system as claimed in any preceding claim wherein the thermal management system is comprised in an aircraft fuselage, nacelle, wing, over-wing fairing, belly fairing or tail cone.
  14. 14. A skin portion for an aircraft comprising a thermal management system as claimed in any preceding claim.
  15. 15. An aircraft comprising one or more thermal management systems and/or skin portions as claimed in any preceding claim.
GB2003511.9A 2019-10-02 2020-03-11 Systems and methods for aircraft Pending GB2587687A (en)

Priority Applications (48)

Application Number Priority Date Filing Date Title
EP20864292.6A EP4058347A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
EP20864299.1A EP4058361A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
EP20864291.8A EP4061705A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052387 WO2021064388A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052375 WO2021064376A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
EP20866954.9A EP4061720A2 (en) 2019-10-02 2020-10-01 Electric aircraft propulsion system
EP20866959.8A EP4058358A2 (en) 2019-10-02 2020-10-01 Aircraft electrical power source monitoring system
EP20864298.3A EP4066225A2 (en) 2019-10-02 2020-10-01 Power management system for managing electrical power sources in aircraft
EP20864295.9A EP4058349A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
EP20786032.1A EP4061724A1 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052389 WO2021064390A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052391 WO2021064392A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
EP20866956.4A EP4058355A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052393 WO2021064394A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052374 WO2021064375A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
EP20866963.0A EP4061721A2 (en) 2019-10-02 2020-10-01 Aircraft electrical power management system
EP20864297.5A EP4061719A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052379 WO2021064380A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052383 WO2021064384A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
EP20866960.6A EP4058356A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052388 WO2021064389A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052376 WO2021064377A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052382 WO2021064383A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052384 WO2021064385A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052377 WO2021064378A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052395 WO2021064396A1 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052373 WO2021064374A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
EP20866962.2A EP4061707A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
EP20864296.7A EP4061725A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
EP20866964.8A EP4059004A2 (en) 2019-10-02 2020-10-01 Power management system for managing level of energy storage device in aircraft
PCT/GB2020/052381 WO2021064382A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052385 WO2021064386A2 (en) 2019-10-02 2020-10-01 Systems and methods for aircraft
PCT/GB2020/052378 WO2021064379A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052394 WO2021064395A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
PCT/GB2020/052396 WO2021064397A2 (en) 2019-10-02 2020-10-01 Systems, arrangements, structures and methods for aircraft
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