Acoustic structure for gas turbine engine
A bypass air passage can be provided between a radially-outer wall of a bypass duct and an engine core, the radially-outer wall having a first, radially inner skin apertured and exposed to the bypass air passage, and a second skin radially-outwardly offset from the first skin, the second skin having an aperture blocked by a thermal fuse, wherein the thermal fuse is configured to melt upon reaching a threshold temperature, thereby opening a gas path from the bypass air passage across the second skin, via the aperture.
1. A gas turbine engine comprising a bypass duct, a bypass air passage between a radially-outer wall of the bypass duct and an engine core, the radially-outer wall having a first, radially inner skin apertured and exposed to the bypass air passage, and a second skin radially-outwardly offset from the first skin, the second skin having an aperture blocked by a thermal fuse, wherein the thermal fuse is configured to melt upon reaching a threshold temperature, thereby opening a gas path from the bypass air passage across the second skin, via the aperture.
2. The gas turbine engine of claim 1 wherein the first skin and the second skin are secured to one another, further comprising a porous acoustic media sandwiched between the first skin and the second skin.
3. The gas turbine engine of claim 2 wherein the acoustic media is made of wood, metal, foam material, or of a fiber matrix composite of carbon glass, ceramic, boron or kevlar with a thermoset or thermoplastic matrix.
4. The gas turbine engine of claim 2 wherein the acoustic media is a honeycomb having open cells extending normal to the first skin.
5. The gas turbine engine of claim 1 wherein the second skin is exposed to a nacelle cavity.
6. The gas turbine engine of claim 1 wherein the second skin is a pressure resisting skin and prevents through fluid flow unless the thermal fuse has melted.
7. The gas turbine engine of claim 1 wherein the second skin has a plurality of apertures blocked by a corresponding plurality of thermal fuses.
8. The gas turbine engine of claim 1 further comprising a third skin radially-outwardly offset from the second skin, and a porous fire retardant media between the third skin and the second skin, the third skin being apertured.
9. The gas turbine engine of claim 8 further comprising a plurality of thermal fuses blocking a corresponding plurality of apertures in the third skin.
10. The gas turbine engine of claim 1 further comprising a third skin between the second skin and the first skin, and a porous fire retardant media between the third skin and the second skin, the third skin being apertured.
11. The gas turbine engine of claim 1 wherein the thermal fuse is a thermoplastic blind rivet.
12. The gas turbine engine of claim 1 wherein the first skin and the second skin are made of metal or fiber matrix composite of carbon, glass, ceramic, boron and/or kevlar with a thermoset or thermoplastic matrix.
13. An acoustic structure for an aircraft, the acoustic structure having a first skin exposed to a first fluidic environment, the first skin being apertured, a second skin between the first skin and a second fluidic environment, an porous acoustic media sandwiched between the first skin and the second skin, the second skin having an aperture, and a thermal fuse blocking the aperture in the second skin, the thermal fuse being configured to melt upon reaching a target temperature to allow a flow from the first fluidic environment to the second fluidic environment across the aperture.
14. The acoustic structure of claim 13 wherein the first skin and the second skin are secured to one another.
15. The acoustic structure of claim 13 wherein the second skin is a pressure resisting skin and prevents through fluid flow unless the thermal fuse has melted.
16. The acoustic structure of claim 13 wherein the second skin has a plurality of apertures blocked by a corresponding plurality of thermal fuses.
17. The gas turbine engine of claim 1 further comprising a third skin offset from the second skin, and a porous fire retardant media between the third skin and the second skin, the third skin being apertured.
18. The gas turbine engine of claim 17 further comprising a plurality of thermal fuses blocking a corresponding plurality of apertures in the third skin.
19. The gas turbine engine of claim 18 wherein the thermal fuse is a thermoplastic blind rivet.
20. A method of operating a gas turbine engine including
pressurizing air in a bypass air passage between a radially-outer wall of a bypass duct and an engine core, the radially-outer wall having a first, radially inner skin apertured and exposed to the bypass air passage, and a second skin radially-outwardly offset from the first skin,
maintaining, with the second skin, a pressure difference between the pressurized air in the bypass air passage and a nacelle cavity radially external to the second skin, the second skin having an aperture blocked by a thermal fuse;
the thermal fuse melting upon reaching a predetermined temperature, opening a cooling path across the second skin, and
cooling the second skin by circulation of cooling air flowing across the second skin.