A system and method for mitigating thermal loading between engine exhaust structures having different coefficients of thermal expansion. The engine exhaust structure comprises a metallic duct portion, a ceramic duct portion, and a double bipod fitting joining the metallic duct portion to the ceramic duct portion. The double bipod fitting is capable of flexing and taking up the thermal expansion differences between the joined metallic and ceramic ducts across the full temperature spectrum that an engine exhaust structure will experience in service.

A labyrinth barrier is disclosed and comprises two or more members each defining respective vertical axes. One or more of the members are constructed at least in part of a shape memory material having a first energy state and a second energy state. The members are oriented relative to one another by their respective vertical axes in an original state to create a flow pathway that restricts fluid flow in a direction transverse to the respective vertical axes. The members are urged towards one another to further restrict the flow pathway when the shape memory material transitions from the first energy state to the second energy state.

An assembly is provided for an aircraft propulsion system. This assembly includes a nozzle extending axially along and circumferentially about an axial centerline. The nozzle includes a nozzle panel and a nozzle fairing fixedly connected to the nozzle panel at an axial end of the nozzle. The nozzle is configured with an internal flow passage radially between the nozzle panel and the nozzle fairing. The internal flow passage extends axially within the nozzle to an outlet between the nozzle panel and the nozzle fairing at the axial end of the nozzle.

A highly adaptive seal for use in advanced aircraft propulsion and missile systems is provided. The seal is capable of effectively limiting invasive thermal energy and corrosive airflow from entering the internal systems and hardware of the aircraft by managing and controlling the incoming air flow. The seal is configured to be disposed between component parts of the aircraft and create a tortuous path for air flow. The seal maintains contact with the component parts between which it seals by allowing for significant radial and axial displacement freedom of the component parts. Accordingly, hot gas intrusion may be limited in a wide variety of dynamic environments and component displacement conditions. Limiting such hot gas intrusion has a substantial impact on component survivability over the course of free flight of an aircraft.

An assembly is provided for an aircraft propulsion system. This assembly includes a nozzle extending axially along and circumferentially about an axial centerline. The nozzle includes a nozzle panel and a nozzle fairing fixedly connected to the nozzle panel at an axial end of the nozzle. The nozzle is configured with an internal flow passage radially between the nozzle panel and the nozzle fairing. The internal flow passage extends axially within the nozzle to an outlet between the nozzle panel and the nozzle fairing at the axial end of the nozzle.

An apparatus is provided for an aircraft propulsion system. This apparatus includes a variable area nozzle apparatus. The variable area nozzle apparatus includes a plurality of panels, a plurality of inter-panel members and an actuation system. The panels are arranged circumferentially about an axial centerline. The panels include a first panel and a second panel. The inter-panel members are arranged circumferentially about the axial centerline. The inter-panel members include a first inter-panel member between and connected to the first panel and the second panel. The first inter-panel member is configured from or otherwise includes an elastomeric material. The actuation system is configured to move the panels and the inter-panel members between a restricted flow arrangement and an unrestricted flow arrangement. The actuation system includes a first linkage and a second linkage. The first linkage is coupled to the first panel. The second linkage is coupled to the second panel.

Apparatuses and methods are provided herein useful to sealing a dynamic gap between a moveable flap and a sidewall. The apparatus may be a seal assembly that includes a spring body, a clip coupled to a distal end of the spring body, and a wear shoe coupled to the clip. The spring body includes a flap arm adjacent the flap and a wall arm adjacent the sidewall. The flap arm and the wall arm bias away from one another when under compression in the gap. A distal end of the wall arm includes a first knuckle and a second knuckle that are pivotally coupled to the clip to create a hinge feature. The wear shoe is pivotally coupled to the clip to create another hinge feature. A distal end of the flap arm includes a skirt that is received by the clip seal an interior space of the spring body.

The present invention relates to an assembly for a turbomachine turbine extending along an axis (X), comprising: —an ejection cone (100) comprising a radially outer annular wall (102) defining a flow duct for a flow of hot gases and a sound box radially arranged inside the outer annular wall (102), the sound box comprising a radially inner annular wall (104), —a connecting member (106) intended to be axially inserted between the exhaust housing and the ejection cone (100), the connecting member (106) comprising an upstream annular flange (108) intended to be attached to the exhaust housing and a plurality of downstream securing tabs (110) connected to the inner annular wall (104), —an annular sealing shroud (112) comprising an upstream portion surrounding the securing tabs (110) of the connecting member (106) so as to cover the spaces circumferentially located between the securing tabs (110) and axially located between the upstream annular flange (108) of the connecting member (106) and the radially inner annular wall (104).

An exterior nozzle member for a turbomachine extending along an axis oriented from upstream to downstream and comprising a primary duct configured to conduct a primary flow, an annular cavity of axis and a turbine, said exterior nozzle member comprising: an acoustic attenuation panel formed by an acoustic attenuation structure on which there are mounted an interior wall and an exterior wall facing the primary duct and the annular cavity, respectively, a nozzle flange connected to the panel and comprising an upstream end configured to be connected to a turbine casing flange and a sealing member connected to the panel and comprising an interior longitudinal branch extending radially inside the nozzle flange in order to thermally protect it from the primary flow, the nozzle flange being made of a first material and the sealing member being made of a second material, the first material having a coefficient of thermal expansion greater than that of the second material.

An exterior nozzle member for a turbomachine extending along an axis oriented from upstream to downstream and comprising a primary duct configured to conduct a primary flow, an annular cavity of axis and a turbine, said exterior nozzle member comprising: an acoustic attenuation panel formed by an acoustic attenuation structure on which there are mounted an interior wall and an exterior wall facing the primary duct and the annular cavity, respectively, a nozzle flange connected to the panel and comprising an upstream end configured to be connected to a turbine casing flange and a sealing member connected to the panel and comprising an interior longitudinal branch extending radially inside the nozzle flange in order to thermally protect it from the primary flow, the nozzle flange being made of a first material and the sealing member being made of a second material, the first material having a coefficient of thermal expansion greater than that of the second material.