Passive reduction of internal jet engine component temperature in supersonic and hypersonic vehicles results from use of nanocomposite optical ceramic materials between the heat-generating portions of each jet engine and the ambient environment, allowing heat dissipation from the jet engine components directly to the ambient environment. A propulsion-airframe integrated scramjet aircraft includes a jet engine and an airframe supporting the jet engine, with at least a portion of the airframe between a heat-generating portion of the jet engine and an ambient environment comprising a nanocomposite optical ceramic material in the form of a panel or a grid of windows each supported within a frame. The nanocomposite optical ceramic material portion of the airframe disposed between the heat-generating portion of the jet engine and the ambient environment is infrared-transparent, and may transmit at least 75% of heat energy from the heat-generating portion of the jet engine to the ambient environment.

According to one aspect, a system for alignment of vanes to suppress infrared detection in a gas turbine engine is provided. The system includes a first vane disposed on a first component, and a second vane disposed on a second component. The second vane is configured to engage the first vane such that the second component is capable of being positioned proximal to the first component.

Devices, systems, and methods of a casing for a heat suppression system of a gas turbine engine exhaust include a heat shield and an insulation layer for arrangement between the heat shield and an outer skin.

A gas turbine exhaust assembly includes an exhaust flow path configured to receive an exhaust flow from a gas turbine engine, the exhaust flow path defined by an inner hub and a radially outer wall. The gas turbine exhaust assembly also includes a plurality of vanes circumferentially spaced from each other and operatively coupled to the radially outer wall of the exhaust flow path, each of the plurality of vanes extending only partially toward the inner hub and terminating at an inner end of the vanes, the inner end defining an open portion.

A serpentine baffle for a liner assembly of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a wall that defines a serpentine shape and a multiple of tabs which extend from the wall. In a further embodiment of the foregoing embodiment, the wall and the multiple of tabs are formed from a single cross-shaped piece. In a further embodiment of any of the foregoing embodiments, wherein at least one of the multiple of tabs is located adjacent to a distal end of the wall. In a further embodiment of any of the foregoing embodiments, the multiple of tabs are perpendicular to the wall.

Embodiments are directed to systems and methods for reducing the infrared signature of a vehicle, specifically to the use of mixed flow or centrifugal blowers to reduce the amount of infrared radiation being emitted from the engine exhaust. A blower and mixer are used to cool the exhaust gas from an aircraft engine. The blower and mixer use a tertiary exhaust duct to swirl cool air from an engine particle separator (EPS) and/or oil cooler, for example, with the hot engine exhaust gas. Existing blowers in the EPS or oil cooler may be used to motivate air flow so that a unique or dedicated blower is not required. In addition to reducing exposure to hostile forces, other benefits of cooling the exhaust gas include greatly reduced ground-impingement temperatures and safer personnel working zones because the hot plume is mostly eliminated.

A seal assembly can accommodate deflection between two components through a cartridge that is slidingly engaged with a slide plate.

An exhaust system for reducing infrared emissions of a rotary wing aircraft includes a duct assembly having an inlet portion and an outlet portion; the inlet portion configured to receive exhaust from an engine of the aircraft; and the outlet portion coupled to the inlet portion, the outlet portion having an outlet duct with an outlet opening, the outlet duct configured expel an emission containing engine exhaust proximate to a tail fairing of the rotary wing aircraft.

According to one aspect, an infrared suppression system that utilizes a primary airflow and a secondary air flow for a gas turbine engine is provided. The infrared suppression system includes an exit flap configured to control a flow of exhaust air out of the gas turbine engine and a secondary flow door that is configured to selectively control the secondary air flow into the gas turbine engine. The mixture of the secondary air flow and the primary air flow from the gas turbine engine suppresses an infrared signature produced by the gas turbine engine.

A thermal panel having a high temperature side located nearest a source of heat and a low temperature side. The thermal panel includes a corrugated composite material core having hot side ridges and cold side ridges. A hot side skin is attached to the hot side ridges to form a plurality of first cells. A cold side skin is attached to the cold side ridges to form a plurality of second cells. The first cells and second cells are substantially filled with an insulating material that has a thermal conductivity which is lower than the thermal conductivity of the composite material.