U.S. Pat. No. 2,938,333 shows a typical gas turbine engine that is used for aircraft propulsion and includes a hot gas flow section, following the turbine, that is contained in an air cooled liner "assembly" having an impingement sheet surrounding a liner sheet. The hot engine exhaust is constrained within the liner, and air is forced through the the impingement sheet and the liner, cooling the liner. The cooling air exits at the exhaust end of the engine.
The liner assembly is exposed to considerable heating, and, as would be expected, liner heating can be very uneven, creating local thermal stresses. For one thing, the upstream area is much hotter than the down stream area. The liner, which must be as light as possible, is supported on the engine, but in such a way that buckling is minimized, as buckling can lead to local stresses and premature liner fatigue. Support is typically provided by brackets, as demonstrated in U.S. Pat. No. 2,938,333.
The conventional liner, such as the liner in U.S. Pat. No. 2,938,333, is made of sheet metal and the impingement and liner sheets are attached by welding or by fasteners, a design, though common, is not particularly rigid, increasing the tendency for temperature induced liner warpage. These designs also do not lend themselves to modulating cooling airflow so that the hotter areas receive more airflow than the cooler areas. Additionally, the flexible cupped membrane configuration used in such liners, as shown in U.S. Pat. No. 2,938,333, is not compatible with the smoothness and coating retention requirements for low observable advanced military engine exhaust systems.