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US 2022/0074352 A1Published application
Turbine section of gas turbine engine
AssigneeRaytheon Technologies Corporation
InventorsPaul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
Life of this patent
Estimating the term…
Abstract
A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.
Claims30
A gas turbine engine comprising:
- a propulsor including a circumferential array of blades;
- a compressor in fluid communication with the propulsor, the compressor including a high pressure compressor section and a low pressure compressor section, and the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area;
- a combustor in fluid communication with the compressor;
- a shaft assembly having a first portion and a second portion;
- a two-stage high pressure turbine section coupled to the first portion of the shaft assembly to drive the high pressure compressor section, and a low pressure turbine section coupled to the second portion of the shaft assembly, each of the high pressure turbine section and low pressure turbine section including blades and vanes, and a low pressure turbine section airfoil count defined as the numerical count of all of the blades and vanes in the low pressure turbine section, wherein the low pressure turbine section airfoil count is below 1600, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and the low pressure compressor section and the low pressure turbine section have an equal number of stages;
- a speed reduction mechanism coupled to the propulsor and rotatable by the low pressure turbine section through the second portion of the shaft assembly to drive the propulsor;
- and
- wherein the low pressure turbine section further includes a maximum gas path radius, the blades of the propulsor include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades of the propulsor is equal to or greater than 0.35, and is less than 0.55.
5 more elements
A gas turbine engine comprising:
- a propulsor including a circumferential array of blades;
- a compressor in fluid communication with the propulsor, the compressor including a nine-stage high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and the low pressure compressor section including four stages;
- a combustor in fluid communication with the compressor;
- a shaft assembly having a first portion and a second portion;
- a two-stage high pressure turbine section coupled to the first portion of the shaft assembly to drive the high pressure compressor section, and a four-stage low pressure turbine section coupled to the second portion of the shaft assembly, each of the high pressure turbine section and low pressure turbine section including blades and vanes, and a low pressure turbine section airfoil count defined as the numerical count of all of the blades and vanes in the low pressure turbine section, and wherein the low pressure turbine section airfoil count is below 1600;
- and
- a speed reduction mechanism coupled to the propulsor and rotatable by the low pressure turbine section through the second portion of the shaft assembly to drive the propulsor;
- and
- wherein the low pressure turbine section further includes a maximum gas path radius, the blades of the propulsor include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades of the propulsor is equal to or greater than 0.35, and is less than 0.55.
6 more elements
Dependencies are inferred from the claim text.