Gas turbine engine

Patent metadata

Field	Value
Publication number	US-11566562-B2
Application number	US-202217723736-A
Country	US
Kind code	B2
Filing date	Apr 19, 2022
Priority date	Oct 23, 2019
Publication date	Jan 31, 2023
Grant date	Jan 31, 2023

Abstract

A gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator compressor is an axi-centrifugal compressor comprising a plurality of axial compression stages followed by a single centrifugal compression stage, wherein the International Standard Atmosphere, sea-level static (hereinafter ISA SLS) design point pressure ratio of the axi-centrifugal compressor is from 12 to 16, and a ratio of the ISA SLS pressure rise across the axial compression stages to the ISA SLS pressure rise across the centrifugal compression stage is from 0.75 to 1.

First claim

The invention claimed is:

1. A gas turbine engine comprising, in fluid flow series:
a gas-generator compressor;
a combustor,
a single stage gas-generator turbine, which comprises an uncooled, subsonic throughflow rotor and a supersonic throughflow stator; and
a free power turbine,
wherein the gas-generator turbine has a specific work of from 200 to 350 kilojoules per kilogram, the specific work being the power output of the gas-generator turbine per unit mass flow therethrough,
the gas-generator compressor is an axi-centrifugal compressor comprising a plurality of axial compression stages followed by a single centrifugal compression stage,
an International Standard Atmosphere, sea-level static (hereinafter ISA SLS) design point pressure ratio of the axi-centrifugal compressor is from 12 to 16, and
a ratio of ISA SLS pressure rise across the axial compression stages to ISA SLS pressure rise across the centrifugal compression stage is from 0.75 to 1.

2. The gas turbine engine of claim 1 , wherein the axi-centrifugal compressor comprises from 3 to 5 axial compression stages.

3. The gas turbine engine of claim 1 , wherein the axi-centrifugal compressor comprises 4 axial compression stages.

4. The gas turbine engine of claim 1 , further comprising a fixed-geometry inlet guide vane for the compressor.

5. The gas turbine engine of claim 1 , wherein the engine has a core size, defined as an ISA SLS design point non-dimensional flow function in which mass flow W is measured at an entry to the compressor (station 2) and temperature T and pressure P are measured at an exit of the compressor (station 3) such that the core size is equal to W 2 √T 3 /P 3 .

6. The gas turbine engine of claim 5 , wherein the core size is from 0.5 to 1.5.

7. The gas turbine engine of claim 5 , wherein the core size is from 0.7 to 0.9.

8. The gas turbine engine of claim 5 , wherein the core size is from 1.0 to 1.5.

9. The gas turbine engine of claim 1 , wherein the free power turbine is one of:
a single-stage turbine; and
a two-stage turbine.

10. The gas turbine engine of claim 1 , wherein the gas-generator turbine and the free power turbine are arranged to contra-rotate.

Assignees

• Rolls Royce Plc
• Rolls Royce Corp

Inventors

• Stieger Rory D
• Swain Daniel

Classifications

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F02C3/08 Primary

  the compressor comprising at least one radial stage (F02C3/10 takes precedence) · CPC title

• F02C3/10 Primary

  with another turbine driving an output shaft but not driving the compressor · CPC title

• F02C3/09

  of the centripetal type · CPC title