Low hub-to-tip ratio fan for a turbofan gas turbine engine

Patent metadata

Field	Value
Publication number	US-9303589-B2
Application number	US-201213687540-A
Country	US
Kind code	B2
Filing date	Nov 28, 2012
Priority date	Nov 28, 2012
Publication date	Apr 5, 2016
Grant date	Apr 5, 2016

Abstract

A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (R HUB ) is the radius of the leading edge at the hub relative to a centerline of the fan. A tip radius (R TIP ) is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (R HUB /R TIP ) is at least less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than or equal to 0.25.

First claim

The invention claimed is:

1. A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of fan blades arranged in a single blade row on the rotor hub, the fan blades radially extending from and being integral with the hub to form an integrally bladed rotor, each of the fan blades of the single blade row having a leading edge, a hub radius (R HUB ) which is the radius of the leading edge at the hub relative to a centerline of the fan, and a tip radius (R TIP ) which is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan, and wherein the ratio of the hub radius to the tip radius (R HUB /R TIP ) is than 0.29.

2. The fan as defined in claim 1 , wherein the ratio of the hub radius to the tip radius (R HUB /R TIP ) is less than about 0.25.

3. The fan as defined in claim 1 , wherein the ratio of the hub radius to the tip radius (R HUB /R TIP ) is between 0.25 and 0.29.

4. The fan as defined in claim 1 , wherein the fan preform includes root stubs disposed on the hub at positions corresponding to at least alternate ones of said fan blades, the root stubs being first formed on the hub prior to blades being fastened thereto.

5. The fan as defined in claim 4 wherein the root stubs formed on the preform have airfoils welded thereto to provide the integrally bladed rotor.

6. The fan as defined in claim 5 wherein the airfoils are linear-friction-welded to the respective root stubs.

7. The fan as defined in claim 4 wherein all of the fan blades are first formed as root stubs on the hub.

8. A method of manufacturing an integrally bladed rotor fan for a turbofan gas turbine engine, comprising: forming a rotor hub preform defining a hub radius and having at least a number of root stubs circumferentially spaced apart on a periphery of the rotor hub preform and aligned in a single blade row thereon; providing blade airfoils having a length such that a ratio of the hub radius to a tip radius of the blade airfoils, once mounted to the hub, is less than 0.29; and subsequently fastening the blade airfoils to the root stubs to form fan blades integrally formed with the hub—resulting in an integrally bladed rotor fan having the fan blades thereof arranged in the single blade row and having a hub to tip radius ratio of less than 0.29.

9. The method as defined in claim 8 , wherein the step of selecting further comprises selecting the length of the blade airfoils to define a ratio of the hub radius to the tip radius of the blade airfoils that is between 0.25 and 0.29.

10. The method as defined in claim 9 , wherein the step of selecting further comprises selecting the length of the blade airfoils to define a ratio of the hub radius to the tip radius of the blade airfoils that is less than about 0.25.

11. The method as defined in claim 8 , wherein circumferentially alternate ones of said fan blades are formed integrally with the hub perform without root stubs, leaving alternate root stubs on the hub preform to provide access for machine tools between the circumferentially alternate ones of said fan blades.

12. The method as defined in claim 8 , wherein the step of fastening further comprises welding the blade airfoils to the root stubs using Linear Friction Welding.

13. A turbofan gas turbine engine comprising a fan upstream of at least one compressor, the fan having a rotor hub and a plurality of substantially radially extending fan blades integral with the rotor hub to form an integrated bladed rotor, the fan blades being arranged in a single blade row on the rotor hub, each said fan blade of the single blade row having an airfoil defining a leading edge and defining a tip radius (R TIP ) which is the radius of a tip of the fan blade at the leading edge, the rotor hub defining a hub radius (R HUB ) which is the radius of the hub at the blade leading edge, and wherein a ratio of the hub radius to the tip radius (R HUB /R TIP ) is less than 0.29.

14. The turbofan gas turbine engine as defined in claim 13 , wherein the ratio of the hub radius to the tip radius (R HUB /R TIP ) is less than about 0.25.

15. The turbofan gas turbine engine as defined in claim 13 , wherein the ratio of the hub radius to the tip radius (R HUB /R TIP ) is between 0.25 and 0.29.

Assignees

• Pratt & Whitney Canada

Inventors

• Heikurinen Kari
• Townsend Peter

Classifications

• 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

• B23K2101/001

  Turbines · CPC title

• F04D29/321

  for axial flow compressors · CPC title

• F04D29/329

  Details of the hub · CPC title

• F04D29/384 Primary

  characterised by form · CPC title