Aerodynamic track fairing for a gas turbine engine fan nacelle
A gas turbine engine includes an aerodynamic track fairing adjacent to a convergent-divergent nozzle, the aerodynamic track fairing including a localized curvature along an outside edge. The aerodynamic track fairing is configured to offset a circumferential pressure gradient otherwise introduced in part by a transition between the convergent-divergent nozzle with the aerodynamic track fairing.
1. A gas turbine engine comprising:
a convergent-divergent nozzle including an outlet having a trailing edge; and
an aerodynamic track fairing adjacent to said convergent-divergent nozzle, said aerodynamic track fairing defining a compound edge extending aft from the trailing edge, the aerodynamic fairing including an aft portion that extends toward an aft end of said aerodynamic track fairing to define a primary curvature and a forward portion extending between the trailing edge and said aft portion, the forward portion of the compound edge defining a substantially reverse curvature relative to said primary curvature thereby minimizing a pressure gradient between said convergent-divergent nozzle and said aerodynamic track fairing;
wherein a width of the aerodynamic fairing at the forward portion is less than a width of the aerodynamic track fairing at the aft portion.
2. The gas turbine engine as recited in claim 1 , wherein said aerodynamic track fairing is adjacent to an engine pylon.
3. The gas turbine engine as recited in claim 2 , wherein said aft portion defining said reverse curvature extends toward said engine pylon.
4. The gas turbine engine as recited in claim 1 , wherein said aerodynamic track fairing is adjacent to a Bi-Fi splitter.
5. The gas turbine engine as recited in claim 4 , wherein said aft portion defining said reverse curvature extends toward said Bi-Fi splitter.
6. The gas turbine engine as recited in claim 1 , further comprising a thrust reverser system upstream of said aerodynamic track fairing.
7. The gas turbine engine as recited in claim 1 , wherein said reverse curvature decreases a magnitude of said circumferential pressure gradient to reduce said non-axial component of flow adjacent to said fan nozzle exit plane.
8. The gas turbine engine as recited in claim 1 , wherein said reverse curvature initiates at the trailing edge of a fan nacelle.
9. The gas turbine engine as recited in claim 1 , wherein said reverse curvature initiates at the trailing edge of a fan nacelle adjacent said convergent-divergent nozzle.
10. The gas turbine engine as recited in claim 1 , wherein said primary curvature is convex and said reverse curvature is concave.
11. The gas turbine engine of claim 1 , wherein
the convergent-divergent nozzle comprises a convergent portion and a divergent portion;
the convergent portion extends radially inward as the convergent portion extends axially, in a downstream direction, to the divergent portion; and
the divergent portion extends radially outward as the divergent portion extends axially, in the downstream direction, away from the convergent portion.
12. The gas turbine engine of claim 1 , wherein the convergent-divergent nozzle comprises a convergent-divergent surface that forms an outer peripheral boundary portion of a bypass flowpath.