COMPRESSOR ROTOR WITH COATED BLADES

• US 20190085708A1
• Filed: 09/18/2017
• Published: 03/21/2019
• Est. Priority Date: 09/18/2017
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: VEITCH, Thomas, ABRARI, Farid, ADIQUE, Ernest, AITCHISON, Paul, FUDGE, Daniel, HEIKURINEN, Kari, STONE, Paul, URAC, Tibor
• Granted Patent: US 10,443,411 B2
• Time in Patent Office: Days
• CPC Class Codes: F01D 5/141 Shape, i.e. outer, aerodynamic form F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147
• CPC Class Codes: F01D 5/16 for counteracting blade vibration
• CPC Class Codes: F01D 5/26 Antivibration means not restricted to blade form or construction or to blade-to-blade connections or to the use of particular materials
• CPC Class Codes: F01D 5/288 Protective coatings for blades
• CPC Class Codes: F05D 2220/3216 for a special compressor stage
• CPC Class Codes: F05D 2230/90 Coating; Surface treatment manufacture with deposition of material F05D2230/30
• CPC Class Codes: F05D 2240/301 Cross-sectional characteristics
• CPC Class Codes: F05D 2260/961 by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape
• CPC Class Codes: F05D 2300/133 Titanium
• CPC Class Codes: F05D 2300/143 Platinum group metals, i.e. Os, Ir, Pt, Ru, Rh, Pd
• CPC Class Codes: F05D 2300/16 Other metals not provided for in groups F05D2300/11 - F05D2300/15
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

Claims

1. A compressor rotor for a gas turbine engine, the rotor comprising blades circumferentially distributed around and extending a span length from a central hub, the blades including alternating first and second blades having airfoils with a leading edge, a trailing edge, a root, a tip and a mid-span region midway between the root and the tip along the span, the airfoils of the first and second blades having corresponding geometric profiles, the airfoil of the first blades having a coating defining a first coating structure, the coating being provided on at least a portion of the first blade adjacent the root and having a root coating thickness, the mid-span region of the first blade having a mid-span thickness, the coating being provided on a portion adjacent the tip of the first blade and having a tip coating thickness, the root coating thickness being greater than at least one of the tip coating thickness and a coating thickness of the airfoil of the first blade at the mid-span region , the first coating structure of the first blade selected to provide the first blade with a first natural vibration frequency different from a second natural vibration frequency of the second blade.

2. The compressor rotor as defined in claim 1, wherein the airfoil of the second blade is free of coating.

3. The compressor rotor as defined in claim 1, wherein the airfoil of the second blade has a coating defining a second coating structure, the coating being provided on at least a portion adjacent the root of the second blade and having a root coating thickness, the coating being provided on a portion adjacent the tip of the second blade and having a tip coating thickness, the root coating thickness being greater than at least the tip coating thickness and a coating thickness of the airfoil of the second blade at the mid-span region, the second coating structure of the second blade leading to the second natural vibration frequency that is different from the first natural vibration frequency.

4. The compressor rotor as defined in claim 1, wherein the root coating thickness of the first blade is greater than the tip coating thickness of the first blade.

5. The compressor rotor as defined in claim 1, wherein the root coating thickness and the tip coating thickness on the trailing and the leading edges of the first blade is greater than the coating thickness of the airfoil at the mid-span region on the trailing and the leading edges of the first blade.

6. The compressor rotor as defined in claim 1, wherein the first and the second airfoil comprise a titanium based substrate or platinum based substrate.

7. The compressor rotor as defined in claim 1, wherein the coating is a nano-coating.

8. The compressor rotor as defined in claim 7, wherein the nano-coating is a nickel nano-coating.

9. The compressor rotor as defined in claim 7, wherein the rotor is a fan.

10. A method of manufacturing a compressor rotor of a gas turbine engine, the rotor having a plurality of blades circumferentially distributed around and extending a span length from a central hub, the method comprising the steps of:

providing first and second blades respectively having first and second airfoils with corresponding geometric profiles, a leading edge, a trailing edge, a root, a tip, and a mid-span region midway between the root and the tip along the span; and

applying a coating on an outer surface of the first airfoil to form a first coating structure, including applying the coating on a portion of the first airfoil adjacent the root so that the first blade has a root coating thickness and applying the coating on a portion adjacent the tip so that the first blade has a tip coating thickness, the root coating thickness being greater than at least one of the tip coating thickness and a coating thickness of the airfoil of the first blade at the mid-span region, wherein the first coating structure of the first blade is selected to provide a first natural vibration frequency different from a second natural vibration frequency of the second blade.

11. The method as defined in claim 11, wherein the coating is applied only on the first airfoil.

12. The method as defined in claim 11, wherein the coating is applied on an outer surface of the second airfoil to form a second coating structure, the coating being applied on at least a portion adjacent the root so that the second blade has a root coating thickness and a portion adjacent the tip so that the second blade has a tip coating thickness, the root coating thickness being greater than at least the tip coating thickness and a coating thickness of the airfoil of the second blade at the mid-span region, wherein the second coating structure of the second blade is selected to provide the second natural vibration frequency that is different from the first natural vibration frequency.

13. The method as defined in claim 11, wherein the coating is applied so that root coating thickness of the first blade is greater than the tip coating thickness of the first blade.

14. The method as defined in claim 11, wherein the coating is applied so that the root coating thickness and the tip coating thickness on the trailing and the leading edges of the first fan blade is greater than the coating thickness of the airfoil at the mid-span region on the trailing and the leading edges of the first blade.

15. A compressor rotor for a gas turbine engine, the mistuned compressor rotor comprising blades circumferentially distributed around and extending a span length from a central hub, the blades including alternating first and second blades having corresponding geometric profiles, the first blade having airfoil with a coating thereon within one or more portions thereof and defining a first coating structure, the one or more portions of the airfoil including a radially inner portion of the airfoil adjacent a blade root of the first blade and having a root coating thickness and a radially outer portion of the airfoil adjacent a blade tip of the first blade and having a tip coating thickness, the root coating thickness being greater than at least one of the tip coating thickness and a mid-span coating thickness at a mid-span region of the airfoil of the first blade, the first coating structure of the first blade selected to provide a first natural vibration frequency different from a second natural vibration frequency of the second blade.

16. The compressor rotor as defined in claim 15, wherein the airfoil of the second blade is free of coating.

17. The compressor rotor as defined in claim 15, wherein the airfoil of the second blade has a coating defining a second coating structure, the coating being located on at least a portion adjacent the root of the second blade and having a root coating thickness and a portion adjacent the tip of the second blade and having a tip coating thickness, the root coating thickness being greater than at least the tip coting thickness and a coating thickness of the airfoil of the second blade at the mid-span region, the second coating structure of the second blade selected to provide the second natural vibration frequency that is different from the first natural vibration frequency.

18. The compressor rotor as defined in claim 15, wherein the root coating thickness of the first blade is greater than the tip coating thickness of the first blade.

19. The compressor rotor as defined in claim 15, wherein the root coating thickness and the tip coating thickness on the trailing and the leading edges of the first blade is greater than the coating thickness of the airfoil at the mid-span region on the trailing and the leading edges of the first blade.

20. A rotor blade for a compressor rotor of a gas turbine engine, the compressor rotor having alternating blades having corresponding geometric profiles but different coating structures selected to provide different natural vibration frequencies, the rotor blade comprising an airfoil having a blade root, a blade tip, a coating on one or more portions of the airfoil and defining a first coating structure, the one or more portions of the airfoil including a radially inner portion adjacent the blade root and having a root coating thickness and a radially outer portion adjacent the blade tip and having a tip coating thickness, the root coating thickness being greater than at least one of the tip coating thickness and a mid-span coating thickness defined at a mid-span region midway between the blade root and the blade tip.