A deicing system for an aircraft may supply heat to an aircraft component in pulses. A first series of pulses may melt ice built up on the aircraft component. A second series of pulses may prevent ice from forming on the aircraft component. The length of each of the pulses in the first series of pulses may be longer than the length of each of the pulses in the second series of pulses. The pulses may be supplied by a pneumatic deicing system or an electrical deicing system.

The engine (10) includes a low spool (16) disposed aft of an air inlet (12) and a high spool (34) disposed aft of the low spool (16). An intake reverse-duct (44) is disposed radially outward of the high spool (34) and reverses direction of low pressure compressed air from the low spool (16) into a forward-flow high pressure compressor (40) of the high spool (34). A discharge reverse-manifold (48) directs flow of an exhaust gas stream (50) from a forward-flow low pressure turbine (20) into a rearward-flow direction and into at least one pulse detonation firing tube (54). An annular bypass air duct (72) directs cooling air along the engine (10)The at least, one firing tube is positioned radially outward of the high spool (34), overlies the high spool (34) and is also positioned within the bypass air duct (72).

An aero damping measurement system is provided. The system includes a shroud defining a tunnel, a hub disposed within the tunnel, and a plurality of blades coupled to the hub. The blades may rotate about the hub. A gas pressure probe may have a tip extending to the tunnel to deliver a pressurized burst into the tunnel. An aeromechanical identification system may include a pressurized gas source, a valve in fluid communication with the pressurized gas source, and the gas pressure probe may be in fluid communication with the valve. The valve may control a flow of a pressurized gas from the pressurized gas source into the gas pressure probe. A pressure sensor may be coupled to the gas pressure probe and configured to measure a pressure within the gas pressure probe.

New exemplary heat exchange configurations that incorporate internal or external surfaces equipped with perturbators, for changing the thermal behavior of the system, or for modulating the surface temperature distribution of the flow surfaces. This is achieved by applying an acoustic wave to the fluid flow in a heat exchange passage, and selecting the frequency of the acoustic exciting wave to be the same as the acoustic resonance frequency of the heat exchange passage itself. As the traveling waves interact with the boundaries confining the heat exchange passages, constructive interference of the incident and reflected waves give rise to a standing wave. Thus, the heat exchange passages act as a resonator, and by superimposing this standing wave on the separating and reattaching fluid flow, significant heat transfer improvement can be achieved. This is accomplished without the need to significantly increase the pressure required to achieve the desired through flow.

Methods and apparatus for reducing flow distortion at engine fans of nacelles are disclosed. An example apparatus for reducing flow distortion at an engine fan of a nacelle includes a plurality of nozzles radially spaced about an inner wall of the nacelle. In some examples, respective ones of the nozzles are positioned to eject corresponding respective jets of fluid adjacent the inner wall in a downstream direction toward the engine fan. The example apparatus further includes a controller to selectively activate the respective ones of the nozzles according to a time-based sequence. In some examples, the time-based sequence corresponds to a directional sequence that moves in an arcuate direction along a circumference of the inner wall.

A power source is configured to apply a first alternating electric excitation signal to an oscillation blade device at a first excitation frequency causing a blade of the oscillation blade device to oscillate at a first oscillation frequency. A current detector is configured to measure amplitude values of the current supplied by the power source to the oscillation blade device. A processor is configured to assess a plurality of successive peak values of the measured amplitudes, determine a second oscillation frequency for the blade if variation in the successive peak values is detected and send a command to the power source to apply a second alternating electric excitation signal to the oscillation blade device at a second excitation frequency which matches the determined second oscillation frequency.

A method of transferring a fluid flow from a static component to a rotor of a gas turbine engine and a modulated flow transfer system are provided. The modulated flow transfer system includes an annular inducer configured to accelerate the fluid flow in a substantially circumferential direction in a direction of rotation of the rotor. The system further includes a first fluid flow supply including a compressor bleed connection, a feed manifold formed of bendable tubing, and a feed header extending between the compressor bleed connection and the feed manifold. The feed header includes a modulating valve configured to control an amount of fluid flow into the feed manifold. The system also includes a flow supply tube that extends between the feed manifold and the inducer and is couplable to at least one of the plurality of first fluid flow inlet openings through a sliding piston seal.

A deicing system for an aircraft may supply heat to an aircraft component in pulses. A first series of pulses may melt ice built up on the aircraft component. A second series of pulses may prevent ice from forming on the aircraft component. The length of each of the pulses in the first series of pulses may be longer than the length of each of the pulses in the second series of pulses. The pulses may be supplied by a pneumatic deicing system or an electrical deicing system.

The engine (10) includes a low spool (16) disposed aft of an air inlet (12) and a high spool (34) disposed aft of the low spool (16). An intake reverse-duct (44) is disposed radially outward of the high spool (34) and reverses direction of low pressure compressed air from the low spool (16) into a forward-flow high pressure compressor (40) of the high spool (34). A discharge reverse-manifold (48) directs flow of an exhaust gas stream (50} from a forward-flow low pressure turbine (20) into a rearward-flow direction and into at least one pulse detonation firing tube (54). An annular bypass air duct (72) directs cooling air along the engine (10)The at least, one firing tube is positioned radially outward of the high spool (34), overlies the high spool (34) and is also positioned within the bypass air duct (72).

Systems and methods for monitoring aerostructures are provided. In various embodiments, a method for monitoring an aerostructure may include: receiving a signal from a pressure sensor, the pressure sensor located downstream from the aerostructure; performing a time frequency analysis on the signal to calculate a power level over a range of frequencies; monitoring the power level over the range of frequencies; and determining a susceptibility to a flutter condition based on the monitoring the power level.