Compressor rotor with coated blades
First Claim
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.
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Accused Products
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.
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Citations
20 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.
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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:
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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. - View Dependent Claims (11, 12, 13, 14)
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- 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.
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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.
Specification