公开(公告)号：US20240426249A1

公开(公告)日：2024-12-26

申请号：US18824100

申请日：2024-09-04

Applicant: General Electric Company

Inventor： Daniel John Oehrle ,  Randy M. Vondrell

IPC: F02C7/32 ,  B64D27/12

Abstract: A gas turbine engine defines an axial direction and a radial direction and comprises a turbomachine having an unducted primary fan, a core engine a combustor casing enclosing a combustor and defining an outer surface, a core cowl surrounding at least a portion of the core engine. The outer surface of the core cowl defines a peak cowl diameter (D) in the radial direction, and the outer surface of the combustor casing defines a maximum combustor casing diameter (d) along the radial direction. The core engine defines an overall core axial length (L) along the axial direction and an under-core cowl axial length (L1) along the axial direction. The gas turbine engine defines a core cowl diameter ratio (CDR) equal to the peak cowl diameter (D) divided by the maximum combustor casing diameter (d) and a core cowl length ratio (CLR) equal to the under-core cowl axial length (L1) divided by the overall core axial length (L). The CDR is between 2.7 and 3.5 and the CLR is between 0.25 and 0.50.

公开(公告)号：US12085027B2

公开(公告)日：2024-09-10

申请号：US17470261

申请日：2021-09-09

Applicant: General Electric Company

Inventor： John Carl Glessner ,  Daniel John Oehrle ,  Mark Edward Linz

IPC: F02C9/18

CPC classification number: F02C9/18 ,  F05D2220/32

Abstract: A gas turbine engine includes a turbomachine defining a core flowpath extending through a first compressor and a second compressor, wherein a first compressor frame defines a compressor bypass passage extending from the core flowpath at the first compressor frame; and a forward compressor frame positioned between the first compressor and the second compressor, wherein the forward compressor frame defines at least in part a second portion of the core flowpath at a location downstream of the first portion of the core flowpath, wherein the forward compressor frame defines a compressor bleed passage extending from the core flowpath at a location downstream of the first compressor and upstream of the second compressor to egress a flow of air away from the core flowpath.

公开(公告)号：US20240229719A9

公开(公告)日：2024-07-11

申请号：US17972720

申请日：2022-10-25

Applicant: General Electric Company

Inventor： Daniel John Oehrle ,  Randy M. Vondrell

IPC: F02C7/32 ,  B64D27/12

CPC classification number: F02C7/32 ,  B64D27/12

Abstract: A gas turbine engine defines an axial direction and a radial direction and comprises a turbomachine having an unducted primary fan, a core engine a combustor casing enclosing a combustor and defining an outer surface, a core cowl surrounding at least a portion of the core engine. The outer surface of the core cowl defines a peak cowl diameter (D) in the radial direction, and the outer surface of the combustor casing defines a maximum combustor casing diameter (d) along the radial direction. The core engine defines an overall core axial length (L) along the axial direction and an under-core cowl axial length (L1) along the axial direction. The gas turbine engine defines a core cowl diameter ratio (CDR) equal to the peak cowl diameter (D) divided by the maximum combustor casing diameter (d) and a core cowl length ratio (CLR) equal to the under-core cowl axial length (L1) divided by the overall core axial length (L). The CDR is between 2.7 and 3.5 and the CLR is between 0.25 and 0.50.

公开(公告)号：US12104539B2

公开(公告)日：2024-10-01

申请号：US17972720

申请日：2022-10-25

Applicant: General Electric Company

Inventor： Daniel John Oehrle ,  Randy M. Vondrell

IPC: F02C7/32 ,  B64D27/12

CPC classification number: F02C7/32 ,  B64D27/12

Abstract: A gas turbine engine defines an axial direction and a radial direction and comprises a turbomachine having an unducted primary fan, a core engine a combustor casing enclosing a combustor and defining an outer surface, a core cowl surrounding at least a portion of the core engine. The outer surface of the core cowl defines a peak cowl diameter (D) in the radial direction, and the outer surface of the combustor casing defines a maximum combustor casing diameter (d) along the radial direction. The core engine defines an overall core axial length (L) along the axial direction and an under-core cowl axial length (L1) along the axial direction. The gas turbine engine defines a core cowl diameter ratio (CDR) equal to the peak cowl diameter (D) divided by the maximum combustor casing diameter (d) and a core cowl length ratio (CLR) equal to the under-core cowl axial length (L1) divided by the overall core axial length (L). The CDR is between 2.7 and 3.5 and the CLR is between 0.25 and 0.50.

公开(公告)号：US20240133345A1

公开(公告)日：2024-04-25

申请号：US17972720

申请日：2022-10-24

Applicant: General Electric Company

Inventor： Daniel John Oehrle ,  Randy M. Vondrell

IPC: F02C7/32 ,  B64D27/12

CPC classification number: F02C7/32 ,  B64D27/12

Abstract: A gas turbine engine defines an axial direction and a radial direction and comprises a turbomachine having an unducted primary fan, a core engine a combustor casing enclosing a combustor and defining an outer surface, a core cowl surrounding at least a portion of the core engine. The outer surface of the core cowl defines a peak cowl diameter (D) in the radial direction, and the outer surface of the combustor casing defines a maximum combustor casing diameter (d) along the radial direction. The core engine defines an overall core axial length (L) along the axial direction and an under-core cowl axial length (L1) along the axial direction. The gas turbine engine defines a core cowl diameter ratio (CDR) equal to the peak cowl diameter (D) divided by the maximum combustor casing diameter (d) and a core cowl length ratio (CLR) equal to the under-core cowl axial length (L1) divided by the overall core axial length (L). The CDR is between 2.7 and 3.5 and the CLR is between 0.25 and 0.50.

公开(公告)号：US20230073647A1

公开(公告)日：2023-03-09

申请号：US17470261

申请日：2021-09-09

Applicant: General Electric Company

Inventor： John Carl Glessner ,  Daniel John Oehrle ,  Mark Edward Linz

IPC: F02C9/18

Abstract: A gas turbine engine includes a turbomachine defining a core flowpath extending through a first compressor and a second compressor, wherein a first compressor frame defines a compressor bypass passage extending from the core flowpath at the first compressor frame; and a forward compressor frame positioned between the first compressor and the second compressor, wherein the forward compressor frame defines at least in part a second portion of the core flowpath at a location downstream of the first portion of the core flowpath, wherein the forward compressor frame defines a compressor bleed passage extending from the core flowpath at a location downstream of the first compressor and upstream of the second compressor to egress a flow of air away from the core flowpath.