Liner attaching scheme
A gas turbine engine includes a liner disposed around a flowpath. The liner has a forward end, a radially outer surface, and a radially inner surface. A hole extends axially into the forward end of the liner between the radially outer surface and the radially inner surface, and an engagement member is partially disposed in the hole and extends axially forward from the forward end of the liner.
1. An assembly for a gas turbine engine, the assembly comprising:
a liner disposed around a flowpath, the liner having a forward end, a radially outer surface, and a radially inner surface, wherein the liner further comprises:
a radially outer sheet that forms the radially outer surface of the liner;
a radially inner perforated sheet that forms the radially inner surface of the liner; and
a honeycomb core disposed between the radially outer sheet and the radially inner perforated sheet;
a hole extending axially into the forward end of the liner between the radially outer surface and the radially inner surface, wherein the hole extends axially into the honeycomb core at the forward end of the liner; and
an engagement member at least partially disposed in the hole and extending axially forward from the forward end of the liner.
2. The assembly of claim 1 further comprising:
a bracket disposed on the outer surface of the liner proximate an aft end of the liner.
3. The assembly of claim 2 , wherein the outer surface of the liner comprises a reduced diameter portion proximate the aft end of the liner and the bracket is disposed on the reduced diameter portion of the outer surface of the liner.
4. The assembly of claim 1 , wherein a bushing is disposed in the hole and around the engagement member.
5. The assembly of claim 4 , wherein the radially outer sheet, the radially inner perforated sheet, the honeycomb core, and the bushing are all made of the same metal.
6. The assembly of claim 4 , wherein the radially outer sheet, the radially inner perforated sheet, the honeycomb core, and the bushing are all made of aluminum.
7. A method of assembling a fan liner in a gas turbine engine, the method comprising:
positioning the fan liner at least partially radially inward from a fan case and around a flowpath, the fan liner having a forward end, a radially outer surface, a radially inner surface, a honeycomb core disposed between the radially outer surface and the radially inner surface, and a hole disposed in the forward end between the radially outer surface and the radially inner surface, wherein the hole extends axially into the honeycomb core at the forward end of the fan liner;
inserting an engagement member into the hole such that a portion of the engagement member extends forward from the hole; and
positioning a fan inlet cowl on the engagement member and forward from the fan liner such that the engagement member extends into a first flange of the fan inlet cowl, wherein the first flange is disposed forward from the forward end of the fan liner.
8. The method of claim 7 , the method further comprising:
connecting a second flange of the fan inlet cowl to a forward end of the fan case, wherein the second flange is disposed aftward from the first flange and the forward end of the fan liner, and forward from the fan case and an aft end of the fan liner, and at least partially radially outward from the first flange.
9. A gas turbine engine comprising:
a fan inlet cowl;
a fan case disposed aftward from the fan inlet cowl;
a fan liner disposed at least partially radially inward from the fan case, the fan liner having a forward end disposed aftward from the fan inlet cowl, a radially outer surface facing the fan case, and a radially inner surface facing a gas flowpath bounded by the fan case and fan inlet cowl, wherein the fan liner further comprises a radially outer sheet that forms the radially outer surface of the fan liner, a radially inner perforated sheet that forms the radially inner surface of the fan liner, and a honeycomb core disposed between the radially outer sheet and the radially inner perforated sheet; and
at least one engagement member extending axially aftward from the fan inlet cowl into the forward end and the honeycomb core of the fan liner.
10. The gas turbine engine of claim 9 further comprising:
a first bracket and a second bracket disposed on the outer surface of the fan liner; and
a block extending radially inward from the fan case and between the first bracket and the second bracket, the block contacting the first bracket and the second bracket.
11. The gas turbine engine of claim 10 , wherein the first bracket and the second bracket are each substantially L-shaped.
12. The gas turbine engine of claim 9 , wherein the fan inlet cowl further comprises:
a first flange disposed forward from the forward end of the fan liner; and
a second flange disposed aftward from the first flange and the forward end of the fan liner, and forward from the fan case and an aft end of the fan liner, wherein the second flange is disposed at least partially radially outward from the first flange.
13. The gas turbine engine of claim 12 , wherein the at least one engagement member extends aftward from the first flange into the forward end of the fan liner.
14. The gas turbine engine of claim 12 , wherein the second flange is connected to a forward end of the fan case.
15. The gas turbine engine of claim 12 , wherein a gap is disposed between the first flange and the forward end of the fan liner and a compliant seal is disposed in the gap between the first flange and the forward end of the fan liner.
16. The gas turbine engine of claim 15 , wherein a second gap is disposed between the radially outer surface of the fan liner and the case.
17. The gas turbine engine of claim 9 , wherein the fan case comprises a composite material and the fan liner comprises aluminum.
18. The gas turbine engine of claim 9 , wherein the at least one engagement member is selected from the group comprising pins, studs, flat metal biscuits, and rods.