US 12,473,863 B2
Gas turbine engine having composite fan blades
Arthur William Sibbach, Boxford, MA (US); and Gary Willard Bryant, Jr., Loveland, OH (US)
Assigned to General Electric Company, Evendale, OH (US)
Filed by General Electric Company, Schenectady, NY (US)
Filed on Oct. 8, 2024, as Appl. No. 18/909,259.
Application 18/909,259 is a continuation in part of application No. 18/603,773, filed on Mar. 13, 2024.
Prior Publication US 2025/0290453 A1, Sep. 18, 2025
This patent is subject to a terminal disclaimer.
Int. Cl. F02C 7/36 (2006.01)
CPC F02C 7/36 (2013.01) [F05D 2220/36 (2013.01); F05D 2260/4031 (2013.01)] 23 Claims
OG exemplary drawing
 
1. A gas turbine engine defining a radial direction, the gas turbine engine comprising:
a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath;
a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio equal to a mass flowrate of an airflow from the fan over the turbomachine to a mass flowrate of an airflow from the fan through the inlet to the working gas flowpath during operation of the gas turbine engine in a cruise operating mode, the bypass ratio being greater than or equal to 10 and less than or equal to 100; and
a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan;
wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8, the FLTCF being equal to:

OG Complex Work Unit Math