A nacelle inlet structure is provided for an aircraft propulsion system. This inlet structure includes an inlet lip, a bulkhead, a nozzle and a plurality of heat transfer augmentation features. The inlet lip includes an inner lip skin and an outer lip skin. The bulkhead is configured with the inlet lip to form a cavity axially between a forward end of the inlet lip and the bulkhead and radially between the inner lip skin and the outer lip skin. The annular cavity extends along a curvilinear centerline within the inlet lip. The nozzle is configured to inject fluid approximately tangentially into the annular cavity. The heat transfer augmentation features are configured with the inner lip skin and operable to interact with the fluid flow within the cavity in order to promote heat transfer between the inner lip skin and the fluid within the cavity.

A system comprises a heat source. The system also comprises a bladder comprising opposing thin-walled sheets and a fluid flow conduit defined between the opposing thin-walled sheets. The fluid flow conduit comprises an inlet and an outlet. The system further comprises a first fluid line coupled to the heat source and the inlet of the bladder. The system additionally comprises a second fluid line coupled to the heat source and the outlet of the bladder. The system also comprises fluid flowable through the first fluid line from the heat source to the inlet, from the inlet through the fluid flow conduit to the outlet, and through the second fluid line from the outlet to the heat source.

A system comprises a heat source. The system also comprises a bladder comprising opposing thin-walled sheets and a fluid flow conduit defined between the opposing thin-walled sheets. The fluid flow conduit comprises an inlet and an outlet. The system further comprises a first fluid line coupled to the heat source and the inlet of the bladder. The system additionally comprises a second fluid line coupled to the heat source and the outlet of the bladder. The system also comprises fluid flowable through the first fluid line from the heat source to the inlet, from the inlet through the fluid flow conduit to the outlet, and through the second fluid line from the outlet to the heat source.

An anti-icing system for an aircraft structure includes a cavity that has an exterior surface subject to ice accretion and a bleed source that is configured to provide a fluid to the cavity via a duct. The system includes multiple temperature sensors, disposed at the aircraft structure, with each temperature sensor configured to detect a temperature associated with an aircraft structure location. A controller is in communication with the temperature sensors. The controller is programmed to compare outputs of the temperature sensors and to determine a temperature sensor fault condition.

An anti-icing system for an aircraft structure includes a cavity that has an exterior surface subject to ice accretion and a bleed source that is configured to provide a fluid to the cavity via a duct. The system includes multiple temperature sensors, disposed at the aircraft structure, with each temperature sensor configured to detect a temperature associated with an aircraft structure location. A controller is in communication with the temperature sensors. The controller is programmed to compare outputs of the temperature sensors and to determine a temperature sensor fault condition.

A system for generating electrical energy from thermal waste energy and removing thermal waste energy in an aircraft is provided. The system includes a thermal management circuit configured to channel a flow of coolant fluid that collects thermal waste energy from a thermal energy generating component. A heat exchanger coupled in fluid communication with the thermal management circuit and downstream from the thermal energy generating component. The heat exchanger is coupled in thermal contact to an exterior skin wall and adapted to transfer thermal energy from the flow of coolant fluid to the exterior skin wall. A thermoelectric generator is configured to convert heat flux between the flow of coolant fluid within the heat exchanger and atmosphere surrounding an outside surface of the exterior skin wall into electrical energy. The thermoelectric generator includes thermo-electric interface material extending between an inside surface of the exterior skin wall and the heat exchanger.

A device for deicing a wall of an aircraft, comprising a closed circuit. The closed circuit comprises at least one condenser, positioned in the environment of the wall that is to be deiced, and in which a heat-transfer fluid condenses, generating energy in the form of latent heat which is transmitted to the wall that is to be deiced, at least one evaporator positioned in the environment of a heat source sited remotely with respect to the wall, and in which the heat-transfer fluid evaporates, absorbing energy in the form of latent heat coming from the heat source. At least part of the closed circuit is facing, in contact with, or positioned in, the wall that is to be deiced, being made of a material transparent to electromagnetic fields.

A system and method for providing bleed air compensation for a continuous power assurance analysis of a gas turbine engine includes estimating bleed air flow rate from the gas turbine engine, estimating a shift in power turbine inlet temperature based on the estimated bleed air flow rate, and applying the estimated shift in power turbine inlet temperature to the continuous power assurance analysis of the gas turbine engine.

An electrically powered aerial vehicle includes at least one motor where each motor includes a stator and a rotor, a motor housing having an inlet opening and a discharge opening for airflow, a plurality of rotor blades rotatable by the rotor, each of the plurality of rotor blades having a cavity running from a proximal end of the rotor blade towards a distal end of the rotor blade, and a blade hub coupled to the rotor blades at the proximal end of each rotor blade and coupled to the motor housing at the discharge opening. A chamber is defined in the blade hub and is in fluid communication with the discharge opening of the motor housing and the cavity of each rotor blade. The airflow is centrifugally drawn in from the motor housing through the discharge opening and transported through the chamber and into the cavities of the rotor blades when the rotor blades are rotating.

The present disclosure provides a tape comprising: a) a single layer comprising a crosslinked polymer, selected from the group consisting of crosslinked polyurethane, crosslinked polyurea, and crosslinked mixed polyurethane/polyurea polymer, and b) an adhesive layer. In another embodiment, the tape comprises at least one layer comprising a semi-interpenetrating polymer network of a crosslinked acrylate and an uncrosslinked polymer selected from the group consisting of polyurethane, polyurea, and mixed polyurethane/polyurea polymer.