Splitter assembly for a gas turbine engine of an aircraft propulsion system

A gas turbine engine for an aircraft propulsion system, the gas turbine engine comprising:

• an engine core extending along an engine axis of the gas turbine engine, the engine core including a compressor section, a combustor section, and a turbine section, the engine core forming a core flow path extending through the compressor section, the combustor section, and the turbine section;
• a bypass flow path disposed radially outward of the engine core;
• and
• a splitter assembly disposed radially between the core flow path and the bypass flow path, the splitter assembly including:
• a splitter structure including an inner structure member, an outer structure member, and a splitter housing, the inner structure member, the outer structure member, and the splitter housing extending circumferentially about an engine axis, the inner structure member forming the core flow path, the outer structure member forming the bypass flow path, the splitter housing disposed between the inner structure member and the outer structure member at an upstream end of the splitter structure;
• a splitter band including a plurality of band segments, the plurality of band segments arranged circumferentially about the engine axis within the splitter housing, each of the plurality of band segments translatable within the splitter housing between an extended position and a retracted position, the plurality of band segments in the extended position extending a least partially outside of the splitter structure at the upstream end;
• a plurality of first cams disposed within the splitter housing, each of the plurality of first cams contacting at least one of the plurality of band segments, said each of the plurality of first cams pivotable about a respective first pivot axis, the plurality of first cams configured to effect translation of the plurality of band segments between the extended position and the retracted position by pivoting about the first pivot axis;
• and
• a plurality of variable inlet guide vanes disposed within the core flow path, the plurality of variable inlet guide vanes arranged circumferentially about the engine axis, each of the plurality of variable inlet guide vanes pivotably mounted to and pivotable with a respective one of the plurality of first cams.

A method for operating a splitter assembly of a gas turbine engine for an aircraft propulsion system, the method comprising:

• directing an air flow through a fan section and into a core flow path and a bypass flow path by rotating a bladed fan rotor, directing the air flow into the core flow path including directing the air flow through a plurality of variable inlet guide vanes;
• separating the air flow into the core flow path and the bypass flow path with the splitter assembly;
• operating the splitter assembly to position a splitter band of the splitter assembly, the splitter band including a plurality of band segments arranged circumferentially about an engine axis of the gas turbine engine, operating the splitter assembly to position the splitter band including translating each of the plurality of band segments along a translation axis between an extended position and a retracted position, the plurality of band segments in the extended position forming a leading edge of the splitter assembly, the plurality of band segments in the retracted position disposed radially inward and axially downstream of the plurality of band segments in the extended position, and
• controlling each of the plurality of variable guide vanes, with the splitter assembly, to pivot about a pivot axis as said each of the plurality of band segments translates along the translation axis.