# US 10,865,806 B2 · Mistuned rotor for gas turbine engine

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- Number: US 10,865,806 B2
- Document ID: US10865806B2
- Status: Granted
- Kind: B2
- Application number: 15/706,247
- Filed: 2017-09-15
- Published: 2019-03-21
- Granted: 2020-12-15
- Earliest claimed: 2017-09-15
- Assignees: Pratt & Whitney Canada Corp.
- Applicants: Pratt & Whitney Canada Corp.
- Inventors: Thomas Veitch; Farid Abrari; Ernest Adique; Daniel Fudge; Kari Heikurinen; Paul Stone; Ignatius Theratil; Peter Townsend; Tibor Urac
- Web page: https://patentgrep.com/patent/US10865806B2

## CPC

- F01D 5/16: for counteracting blade vibration
- F04D 29/327: {with non identical blades}
- F01D 5/141: {Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147)}
- F04D 29/666: {by means of rotor construction or layout, e.g. unequal distribution of blades or vanes}
- F04D 29/668: {damping or preventing mechanical vibrations}
- F05D 2240/301: Cross-sectional characteristics
- F05D 2260/961: by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape
- Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F04D: NON-POSITIVE-DISPLACEMENT PUMPS (engine fuel-injection pumps F02M; ion pumps H01J41/12; electrodynamic pumps H02K44/02)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## Abstract

A rotor for a gas turbine engine. The rotor includes blades circumferentially distributed around a hub. The blades have airfoils with a span defined between a root and tip, a chord defined between a leading edge and a trailing edge, and a thickness defined between a pressure side surface and suction side surface. The blades include first blades and second blades. The airfoil of the first blades has a first thickness distribution defining a first natural vibration frequency of the airfoils of the first blades. The airfoil of the second blades has a second thickness distribution defining a second natural vibration frequency different than the first natural vibration frequency. The first thickness distribution is different than the second thickness distribution along a radially-inner half of the span, and the first thickness distribution matches the second thickness distribution along a radially-outer half of the span.

## Claims

### Claim 1

A rotor for a gas turbine engine, the rotor comprising blades circumferentially distributed around a hub, the blades having airfoils with a span defined between a root and tip of the airfoils, the airfoils having a chord defined between a leading edge and a trailing edge of the airfoils, the airfoils having a thickness defined between a pressure side surface and suction side surface of the airfoils, the blades including first blades and second blades interleaved about the rotor, the airfoil of the first blades having a first thickness distribution along the span defining a first natural vibration frequency of the airfoils of the first blades, the airfoil of the second blades having a second thickness distribution along the span defining a second natural vibration frequency different than the first natural vibration frequency, the first thickness distribution being different than the second thickness distribution along a radially-inner portion of the span, and the first thickness distribution matching the second thickness distribution along a radially-outer portion of the span extending from a 50% span position to a 100% span position, the first and second thickness distributions having a maximum thickness (t-max′) at the root of the airfoils located at a first position on the chord, and the first and second thickness distributions having a maximum thickness (t-max″) at the tip of the airfoils located at a second position on the chord different than the first position, wherein Δt-max is the difference between the maximum thickness of the airfoils of the first blades and the maximum thickness of the airfoils of the second blades, and the Δt-max is largest at a 0% span position of the airfoils of the first and second blades.

### Claim 2 (depends on 1)

The rotor as defined in claim 1, wherein the second position on the chord is closer to the trailing edge than the first position on the chord.

### Claim 3 (depends on 2)

The rotor as defined in claim 2, wherein the first position on the chord is 35% of the chord, and the second position on the chord is 55% of the chord.

### Claim 4 (depends on 1)

The rotor as defined in claim 1, wherein the Δt-max has a value greater than 0 between a 0% span position and a 45% span position of the span of the airfoil of the first and second blades.

### Claim 5 (depends on 4)

The rotor as defined in claim 4, wherein the Δt-max has a value greater than 0 between the 0% span position and a 25% span position of the span of the airfoil of the first and second blades.

### Claim 6 (depends on 1)

The rotor as defined in claim 1, wherein the Δt-max has a value substantially equal to 0 between a 25% span position and the 100% span position of the span of the airfoil of the first and second blades.

### Claim 7 (depends on 1)

The rotor as defined in claim 1, wherein the radially-inner portion of the span of the airfoils of the first and second blades extends between the 0% span position and a 45% span position.

### Claim 8 (depends on 1)

The rotor as defined in claim 1, wherein the first position on the chord of the maximum thickness (t-max′) at the root of the airfoils is the same for the first and second blades.

### Claim 9

A fan for a gas turbine engine, the fan comprising blades circumferentially distributed around a hub, the blades having airfoils with a span defined between a root and tip of the airfoils, the airfoils having a chord defined between a leading edge and trailing edge of the airfoils, the airfoils having a thickness defined between a pressure side surface and suction side surface of the airfoils, the blades including first blades and second blades interleaved about the rotor, the airfoil of the first blades having a first thickness distribution along the span defining a first natural vibration frequency of the airfoils of the first blades, the airfoil of the second blades having a second thickness distribution along the span of the airfoil defining a second natural vibration frequency different than the first natural vibration frequency, the first thickness distribution being different than the second thickness distribution along a radially-inner portion of the span, the first thickness distribution matching the second thickness distribution along a radially-outer portion of the span extending from a 50% span position to a 100% span position, the first and second thickness distributions having a maximum thickness (t-max′) at the root of the airfoils located at a first position on the chord, and the first and second thickness distributions having a maximum thickness (t-max″) at the tip of the airfoils located at a second position on the chord different than the first position, wherein a Δt-max is the difference between the maximum thickness of the airfoil of the first blades and the maximum thickness of the airfoil of the second blades, and the Δt-max is largest at a 0% span position of the airfoils of the first and second blades.

### Claim 10 (depends on 9)

The rotor as defined in claim 9, wherein the maximum thickness (t-max′) at the root of the airfoil is located at 35% of the chord, and the maximum thickness (t-max″) at the tip of the airfoil is located at 55% of the chord.

### Claim 11 (depends on 9)

The fan as defined in claim 9, wherein the Δt-max has a value greater than 0 between a 0% span position and a 45% span position of the span of the airfoil of the first and second blades.

### Claim 12 (depends on 9)

The fan as defined in claim 9, wherein the fan is an integrally-bladed fan.

### Claim 13

A method of forming a rotor of a gas turbine engine, the method comprising:
providing first blades and second blades, the first blades having a first thickness distribution defining a first natural vibration frequency, the second blades having a second thickness distribution defining a second natural vibration frequency different than the first natural vibration frequency, the first thickness distribution being different than the second thickness distribution along a radially-inner portion of the first and second blades, the first thickness distribution matching the second thickness distribution along a radially-outer portion of the first and second blades extending from a 50% span position to a 100% span position, the first and second thickness distributions having a maximum thickness (t-max′) at a root of the first and second blades located at a first position on a chord of the first and second blades, and the first and second thickness distributions having a maximum thickness (t-max″) at a tip of the first and second blades located at a second position on the chord different than the first position, and defining a difference between the maximum thickness of the first blade thickness distribution and the maximum thickness of the second blade thickness distribution being defined as Δt-max, the Δt-max being largest at a 0% span position of the airfoils of the first and second blades;
positioning at least one of the second blades relative to a hub of the rotor to be circumferentially between two of the first blades; and
fastening the first and second blades to the hub.

### Claim 14 (depends on 13)

The method as defined in claim 13, wherein providing the first blades and the second blades includes adjusting a difference between the first and second natural vibration frequencies by increasing the difference between the maximum thickness of the first blades and the maximum thickness of the second blades along the radially-inner portion of the first and second blades.