A hydraulic system 300 for an aircraft including a backup hydraulic pressure source 216 to provide hydraulic pressure to a brake 222 in the event of a failure condition of a primary hydraulic brake pressure source. The hydraulic system 300 also includes a landing gear backup system to provide hydraulic pressure to enable extension and/or retraction of landing gear 100. The backup hydraulic pressure source 216 is arranged to provide hydraulic pressure to the landing gear backup system in the event of a failure condition of a primary landing gear hydraulic pressure source.

A hydraulic system 300 for an aircraft including a backup hydraulic pressure source 216 to provide hydraulic pressure to a brake 222 in the event of a failure condition of a primary hydraulic brake pressure source. The hydraulic system 300 also includes a landing gear backup system to provide hydraulic pressure to enable extension and/or retraction of landing gear 100. The backup hydraulic pressure source 216 is arranged to provide hydraulic pressure to the landing gear backup system in the event of a failure condition of a primary landing gear hydraulic pressure source.

The present disclosure provides an aircraft thermal management system. The aircraft thermal management system includes a first hydraulic system for circulating a first hydraulic fluid, a second hydraulic system for circulating a second hydraulic fluid, and a third hydraulic system for circulating a third hydraulic fluid. The aircraft thermal management system also includes a controller configured to: (i) determine a temperature of the first hydraulic fluid, a temperature of the second hydraulic fluid, and a temperature of the third hydraulic fluid, and (ii) based on the determined temperature of the first hydraulic fluid, utilize the second hydraulic fluid and/or the third hydraulic fluid to modify an operational temperature of the first hydraulic fluid.

A rod end of an actuator assembly is disclosed. The rod end may comprise a rod piston end having an outer profile, and a rod joint end opposite the rod piston end, wherein the outer profile comprises an elliptical shape comprising a major diameter and a minor diameter, wherein the major diameter may be larger than the minor diameter in length.

An emission-capturing apparatus includes a tank having an inlet and an outlet. The apparatus further includes a muffler fluidly coupled with the outlet to intercept fluid exiting the outlet, permit passage of gas through the muffler, and inhibit passage of liquid through the muffler. The apparatus further includes a hose having a tank end and a coupler end. The tank end is coupled to the inlet. The apparatus further includes a coupler coupled to the coupler end of the hose. The coupler has a drain hole. The coupler is configured to be coupled to a surface having an ejection port, to completely cover the ejection port when the coupler is coupled to the surface and to support the coupler end of the hose in a position to receive an emission when the drain hole is aligned with the ejection port.

An emission-capturing apparatus includes a tank having an inlet and an outlet. The apparatus further includes a muffler fluidly coupled with the outlet to intercept fluid exiting the outlet, permit passage of gas through the muffler, and inhibit passage of liquid through the muffler. The apparatus further includes a hose having a tank end and a coupler end. The tank end is coupled to the inlet. The apparatus further includes a coupler coupled to the coupler end of the hose. The coupler has a drain hole. The coupler is configured to be coupled to a surface having an ejection port, to completely cover the ejection port when the coupler is coupled to the surface and to support the coupler end of the hose in a position to receive an emission when the drain hole is aligned with the ejection port.

An aircraft assembly having: a first part; a second part, the second part being movably mounted with respect to the first part; an electro-hydraulic actuator coupled between the second part and a first anchor point, the actuator comprising a cylinder defining a bore and a piston and rod assembly slidably mounted within the bore and an active chamber within which an increase in fluid pressure causes the actuator to change during a first phase between first and second extension states to move the second part relative to the first part. The electro-hydraulic actuator further includes a hydraulic fluid supply circuit comprising a piezo-electric pump operable to supply pressurised fluid to the active chamber to change the actuator between first and second extension states.

An aircraft assembly having: a first part; a second part, the second part being movably mounted with respect to the first part; an electro-hydraulic actuator coupled between the second part and a first anchor point, the actuator comprising a cylinder defining a bore and a piston and rod assembly slidably mounted within the bore and an active chamber within which an increase in fluid pressure causes the actuator to change during a first phase between first and second extension states to move the second part relative to the first part. The electro-hydraulic actuator further includes a hydraulic fluid supply circuit comprising a piezo-electric pump operable to supply pressurised fluid to the active chamber to change the actuator between first and second extension states.

A power control unit configured to supply hydraulic pressure and mechanical torque. To this end, the power control unit includes an electric motor coupled to a differential output transmission and a hydraulic pump. A switchable transmission device allows the torque of the motor to be redirected either to the output transmission, e.g., to drive high-lift devices, or to the hydraulic pump, e.g., to extend or retract landing gears.

The landing gear systems, methods and apparatuses disclosed herein may comprise a shrink pump and a shrink valve that are capable of shrinking a landing gear by up to 40% of its available stroke, or more depending on the air spring configuration. The shrink pump may be configured to pump fluid (e.g., a hydraulic fluid) between an oil chamber, where hydraulic fluid is likely present and a shrink chamber to shrink the landing gear. Moreover, the shrink pump and shrink valve may be part of a strut shrink system.