Compressor rotor for supersonic flutter and/or resonant stress mitigation

Patent metadata

Field	Value
Publication number	US-10670041-B2
Application number	US-201715436091-A
Country	US
Kind code	B2
Filing date	Feb 17, 2017
Priority date	Feb 19, 2016
Publication date	Jun 2, 2020
Grant date	Jun 2, 2020

Abstract

A compressor rotor, such as a fan, for a gas turbine engine is described which includes alternating at least first and second blade types. The leading edge of the second blade types includes a leading edge tip cutback extending to the blade tip thereof. The leading edge tip cutback of the second blade type defines a chord length at the blade tip of the second blade types that is less than that of the first blades types. The first and second blade types generate different shock patterns when the fan or compressor rotor operates in supersonic flow regimes.

First claim

The invention claimed is:

1. A gas turbine compressor for an aircraft gas turbine engine, the gas turbine engine compressor comprising a compressor rotor having compressor blades circumferentially distributed around a hub and each extending radially outward from the hub to a blade tip, the compressor blades including a first compressor blade and a second compressor blade alternating continuously around a circumference of the compressor rotor, each first compressor blade and each second compressor blade having an airfoil with a pressure side and a suction side, the pressure side and suction side extending on opposed sides of the airfoil between a leading edge and a trailing edge, the second compressor blade including at least one tip cutback, the first compressor blade and the second compressor blade being identical but for the at least one tip cutback on the second compressor blade, the at least one tip cutback including a leading edge tip cutback formed in the leading edge of the second compressor blade and extending to the blade tip, wherein the leading edge tip cutback defines a chord length at the blade tip of the second compressor blade that is less than a chord length at the blade tip of the first compressor blade, and the leading edge tip cutback is disposed within a radially outermost 15% of the total span length of the second compressor blade.

2. The gas turbine compressor of claim 1 , wherein the chord length at the blade tip of the second compressor blade is greater than 75% and less than 100% of the chord length at the blade tip of the first compressor blade.

3. The gas turbine compressor of claim 2 , wherein the chord length at the blade tip of the second compressor blade is greater than 80% of the chord length at the blade tip of the first compressor blade.

4. The gas turbine compressor of claim 1 , wherein the leading edge tip cutback has a span-wise length and a chord-wise length, the span-wise length of the leading edge tip cutback is greater than the chord-wise length of the leading edge tip cutback.

5. The gas turbine compressor of claim 1 , wherein a chord-wise length of the leading edge tip cutback on the second compressor blade is less than 25% of the chord length at the blade tip of the first compressor blade.

6. The gas turbine compressor of claim 5 , wherein the chord-wise length of the leading edge tip cutback on the second compressor blade is less than 20% of the chord length at the blade tip of the first compressor blade.

7. The gas turbine compressor of claim 1 , wherein the first compressor blade and the second compressor blade are aerodynamically mistuned to generate different shock patterns when the compressor operates in supersonic flow regimes.

8. The gas turbine compressor of claim 1 , wherein the leading edge tip cutback defines a tip portion of the leading edge of the second compressor blade that extends linearly between an upstream inflection point and a downstream inflection point, the upstream inflection point located at a junction between the leading edge of the airfoil and the tip portion, and the downstream inflection point located at a junction between the tip portion and an outer edge of the blade tip.

9. The gas turbine compressor of claim 1 , wherein the first compressor blade and the second compressor blade generate different shock patterns and/or aerodynamic instabilities when the compressor operates in supersonic flow regimes, the different shock patterns and/or aerodynamic instabilities mitigating at least one of supersonic flutter and resonant stresses in the compressor blades.

10. The gas turbine compressor of claim 1 , wherein the at least one tip cutback of the second compressor blade further includes a pressure side tip pocket disposed at the blade tip of the second compressor blade and extending radially inwardly from the blade tip on the pressure side of the airfoil of the second compressor blade.

11. The gas turbine compressor of claim 1 , wherein the first compressor blade is free of leading edge tip cutbacks and free of tip pockets.

12. The gas turbine engine compressor of claim 1 , wherein the first compressor blade includes an axial tip projection thereon, the axial tip projection extending axially forwardly relative to a baseline leading edge of the majority of the airfoil of the first compressor blade.

13. The gas turbine compressor of claim 1 , wherein the compressor rotor is a fan of a turbofan engine.

14. A gas turbine compressor for an aircraft engine, the compressor comprising a compressor rotor having a hub from which a plurality of airfoil blades extend to outer blade tips, the airfoil blades each having an airfoil selected from at least first and second airfoil types and arranged on the hub as alternating with one another around a circumference of the rotor, the second airfoil types including a leading edge having a leading edge tip cutback extending to the outer blade tip thereof, the first airfoil types and the second airfoil types being identical but for the leading edge tip cutback on the outer blade tip of the second airfoil types, wherein the leading edge tip cutback defines a chord length at the blade tip of the second airfoil types that is less than a chord length at the blade tip of the first airfoil types, the first and second airfoil types generating different shock patterns and/or aerodynamic instabilities when the compressor rotor operates in supersonic flow regimes, and wherein the chord length at the blade tip of the second airfoil types is greater than 75% and less than 100% of the chord length at the blade tip of the first airfoil types.

15. The gas turbine compressor of claim 14 , wherein the leading edge tip cutback is disposed within a radially outermost 15% of a total span length of the second airfoil types.

16. The gas turbine compressor of claim 14 , wherein the leading edge tip cutback defines a tip portion of the leading edge of the second airfoil types that extends between an upstream inflection point and a downstream inflection point, the upstream inflection point located at a junction between the leading edge of the second airfoil types and the tip portion thereof, and the downstream inflection point located at a junction between the tip portion and an outer edge of the blade tips.

Assignees

• Pratt & Whitney Canada

Inventors

• Theratil Ignatius
• Balike Krishna Prasad
• Abate Aldo
• Kelly Sean
• Townsend Peter
• Dutton Ronald
• Brown Anthony
• Stone Paul
• Klein Myron
• Kojovic Milica
• Venditti Robert

Classifications

• F04D19/002

  Axial flow fans · CPC title

• F05D2240/302

  characteristics related to shock waves, transonic or supersonic flow · CPC title

• F05D2240/303

  related to the leading edge of a rotor blade · CPC title

• F05D2240/307

  related to the tip of a rotor blade · CPC title

• F04D29/666

  by means of rotor construction or layout, e.g. unequal distribution of blades or vanes · CPC title