US 11,781,505 B2
Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
Wesley K. Lord, South Glastonbury, CT (US); Robert E. Malecki, Storrs, CT (US); Yuan J. Qiu, Glastonbury, CT (US); Becky E. Rose, Colchester, CT (US); and Jonathan Gilson, W. Hartford, CT (US)
Assigned to RTX CORPORATION, Farmington, CT (US)
Filed by RAYTHEON TECHNOLOGIES CORPORATION, Farmington, CT (US)
Filed on Mar. 11, 2021, as Appl. No. 17/198,651.
Application 17/198,651 is a continuation of application No. 15/887,183, filed on Feb. 2, 2018, granted, now 11,015,550.
Application 15/887,183 is a continuation of application No. 14/091,862, filed on Nov. 27, 2013, granted, now 9,932,933, issued on Apr. 3, 2018.
Application 14/091,862 is a continuation in part of application No. 13/721,095, filed on Dec. 20, 2012, granted, now 9,920,653, issued on Mar. 20, 2018.
Claims priority of provisional application 61/884,325, filed on Sep. 30, 2013.
Prior Publication US 2021/0285403 A1, Sep. 16, 2021
This patent is subject to a terminal disclaimer.
Int. Cl. F02K 3/068 (2006.01); F01D 25/24 (2006.01); F02C 7/04 (2006.01)
CPC F02K 3/068 (2013.01) [F01D 25/24 (2013.01); F02C 7/04 (2013.01); F05D 2220/36 (2013.01); F05D 2240/303 (2013.01); F05D 2260/40311 (2013.01); F05D 2260/96 (2013.01); Y02T 50/60 (2013.01)] 30 Claims
OG exemplary drawing
 
1. A gas turbine engine assembly, comprising:
a fan section including a fan and a fan case surrounding the fan to define a bypass duct, the fan including a plurality of fan blades having circumferentially outermost edges, a diameter of the fan having a dimension D extending between the circumferentially outermost edges of the fan blades, each fan blade having a leading edge, a forward-most portion on the leading edges of the fan blades in a first reference plane, wherein a portion of the fan case is forward of the leading edges of the fan blades relative to an engine central longitudinal axis, and each of the fan blades is a swept fan blade;
a geared architecture;
a compressor section including a first compressor and a second compressor axially aft of the first compressor relative to the engine central longitudinal axis;
a turbine section including a first turbine and a fan drive turbine, wherein each of the first compressor and the fan drive turbine includes a greater number of stages than the first turbine;
a nacelle including an inlet portion forward of the fan relative to the engine central longitudinal axis, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along the engine central longitudinal axis between the first reference plane and the second reference plane;
wherein the fan delivers a portion of air into the compressor section and a portion of air into the bypass duct, and a bypass ratio of greater than 10, the bypass ratio being a ratio of a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section;
a first shaft concentric with a second shaft, the first and second shafts rotatable about the engine central longitudinal axis, wherein the first shaft connects the fan and the first compressor to the fan drive turbine, the second shaft connects the second compressor to the first turbine, and the first shaft drives the fan through the geared architecture; and
wherein a dimensional relationship of L/D is between 0.30 and 0.40.