US 11,840,967 B2
Gas turbine engine
Peter Daniel Silkowski, Scotia, NY (US)
Assigned to GENERAL ELECTRIC COMPANY, Schenectady, NY (US)
Filed by General Electric Company, Schenectady, NY (US)
Filed on Jan. 13, 2022, as Appl. No. 17/647,855.
Application 17/647,855 is a continuation of application No. 16/295,542, filed on Mar. 7, 2019, granted, now 11,434,831.
Claims priority of provisional application 62/675,284, filed on May 23, 2018.
Prior Publication US 2022/0290618 A1, Sep. 15, 2022
Int. Cl. F02C 9/22 (2006.01); F02C 6/00 (2006.01); F02C 3/04 (2006.01); F02C 3/14 (2006.01); F23R 3/02 (2006.01)
CPC F02C 9/22 (2013.01) [F02C 3/04 (2013.01); F02C 3/14 (2013.01); F02C 6/003 (2013.01); F23R 3/02 (2013.01); F05D 2220/32 (2013.01); F05D 2240/127 (2013.01); F05D 2240/128 (2013.01)] 19 Claims
OG exemplary drawing
 
1. A gas turbine engine comprising:
a compressor section, the compressor section comprising a compressor mean radius;
an open swirl combustor section having a flow path therethrough and fluidly coupled downstream of the compressor section, the open swirl combustor section comprising:
a combustor mean radius;
a combustor inlet defined by a combustor inlet inner wall and a combustor inlet outer wall;
a combustor exit defined by a combustor exit inner wall and a combustor exit outer wall; and
a heat addition section between the combustor inlet and the combustor exit; and
a turbine section fluidly coupled downstream of the open swirl combustor section, the turbine section comprising a first stage turbine rotor and a turbine mean radius, wherein the first stage turbine rotor is located immediately downstream of the open swirl combustor section,
wherein the gas turbine engine is free from a last stage compressor vane and a first stage turbine vane between a last compressor rotor and the first stage turbine rotor,
wherein the combustor mean radius is greater than the compressor mean radius, the combustor mean radius is between about 1.5 times and about five times greater than the turbine mid-span radius, and the turbine mean radius is greater than or equal to the compressor mean radius,
wherein the compressor section, the open swirl combustor section, the turbine section have a flow successively therethrough, the flow having a non-zero flow angle, and wherein the flow angle is reduced due to a heat addition in the heat addition section of the open swirl combustor section, and
wherein the flow path of the open swirl combustor section is contoured in a radial direction from the combustor inlet to the combustor exit such that the combustor inlet inner wall and the combustor inlet outer wall curve radially outward from the compressor section to the heat addition section and the combustor exit inner wall and the combustor exit outer wall curve radially inward from the heat addition section to the turbine section and such that the combustor exit is radially outward of the combustor inlet and radially inward of the heat addition section, and
wherein a span of the open swirl combustor section is varied such that, in combination with the contoured flow path and the reduced flow angle due to the heat addition, the flow has a predetermined flow angle at an inlet of the first stage turbine rotor.