Fan blade removal feature for a gas turbine engine
AssigneeUnited Technologies Corporation
InventorDavid A. Welch
Life of this patent
Estimating the term…
Abstract
A gas turbine engine includes a fan forward of a primary flowpath inlet. The fan includes multiple fan blades distributed radially about, and connected to, a hub. A fan nacelle is positioned radially outward of the fan and includes an inner diameter. The inner diameter is sloped relative to an engine axis such that a forward portion of the fan nacelle has a smaller inner diameter than an aft portion of the fan nacelle. A fan blade spacer is disposed between a radially inward facing surface of at least one fan blade and a radially outward facing surface of the hub. The fan blade spacer has a radial thickness at least equal to a slope drop of the fan nacelle.
Claims18
A gas turbine engine comprising:
- a fan forward of a primary flowpath inlet;
- the fan including a plurality of fan blades distributed radially about, and connected to, a hub;
- a fan nacelle positioned radially outward of the fan and including an inner diameter, the inner diameter being sloped relative to an engine axis such that a forward portion of the fan nacelle has a smaller inner diameter than an aft portion of the fan nacelle;
- and
- a ring shaped fan blade spacer including a center hole coaxial with an axis defined by the ring shaped fan blade spacer, the ring shaped fan blade spacer being disposed between a radially inward facing surface of at least one fan blade and a radially outward facing surface of said hub, the fan blade spacer having a radial thickness at least equal to a slope drop of the fan nacelle.
A method for removing a fan blade from an aircraft engine comprising:
- at least partially removing a ring shaped tan blade spacer, wherein the ring shaped fan blade spacer includes center hole coaxial with an axis defined by the ring shaped fan blade spacer;
- shifting a fan blade radially inward, at least a distance equal to a slope drop of a fan nacelle;
- and
- sliding said fan blade axially forward.
A fan blade spacer for a gas turbine engine comprising:
- a ring shaped body including a center hole, coaxial with an axis defined by the fan blade spacer, the center hole protruding axially through the ring shaped body.
Dependencies are inferred from the claim text.