Geared turbofan with integrally bladed rotor

Patent metadata

Field	Value
Publication number	US-10060282-B2
Application number	US-201514709573-A
Country	US
Kind code	B2
Filing date	May 12, 2015
Priority date	Jun 10, 2014
Publication date	Aug 28, 2018
Grant date	Aug 28, 2018

Abstract

A gas turbine comprises a compressor module, with a lower pressure compressor section including a plurality of stages, with at least one of the plurality of stages being an integrally bladed rotor. A higher pressure compressor section includes a plurality of stages with at least one of the plurality of stages being an integrally bladed rotor. A fan drive turbine shaft drives a fan rotor through a gear reduction. The fan rotor delivers a portion of air into a bypass duct, and a portion of air into the compressor module. A bypass ratio defined by the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor module is greater than or equal to about 6.0.

First claim

The invention claimed is:

1. A gas turbine comprising:
a compressor module, including:
a lower pressure compressor section including a plurality of stages, with at least one of said plurality of stages being an integrally bladed rotor;
a higher pressure compressor section, including a plurality of stages with at least one of said plurality of stages being an integrally bladed rotor;
a fan drive turbine shaft driving a fan rotor through a gear reduction, said fan rotor delivering a portion of air into a bypass duct, and a portion of air into said compressor module;
wherein at least one stage in said higher pressure compressor section being provided by a compressor hub having removable blades, said compressor hub having slots and said removable blades having dovetails received in said slots; and
at least one stage having removable blades being downstream of said at least one integrally bladed rotor in said higher pressure compressor section.

2. The gas turbine engine as set forth in claim 1 , wherein each of said plurality of stages in said lower pressure compressor section are integrally bladed rotors.

3. The gas turbine engine as set forth in claim 1 , wherein a final compressor stage in said higher pressure compressor section is said at least one stage with said removable blades.

4. The gas turbine engine as set forth in claim 3 , wherein said higher pressure compressor section having at least six compressor stages.

5. The gas turbine engine as set forth in claim 1 , wherein a common turbine drives said lower pressure compressor section and said fan rotor.

Assignees

• United Technologies Corp

Inventors

• Schwarz Frederick M

Classifications

• F05D2260/4031

  as in toothed gearing · CPC title

• F01D5/34

  Rotor-blade aggregates of unitary construction {, e.g. formed of sheet laminae; (discs formed of sheet laminae F01D5/028; ceramic materials F01D5/284, composite materials F01D5/282)} · CPC title

• F05D2220/32

  in gas turbines · CPC title

• Y02T50/671

  Cross-Sectional Technologies · mapped topic

• F01D15/12 Primary

  Combinations with mechanical gearing (driven by multiple engines F01D13/00) · CPC title