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A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibr

A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor.

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1. A mistuned fan for a gas turbine engine, the fan comprising fan blades circumferentially distributed around and extending a total span length from a central hub, the fan blades including successively alternating first and second fan blades each having airfoil with a pressure side and a suction si

1. A mistuned fan for a gas turbine engine, the fan comprising fan blades circumferentially distributed around and extending a total span length from a central hub, the fan blades including successively alternating first and second fan blades each having airfoil with a pressure side and a suction side disposed on opposed sides of a median chord line, the pressure side and suction side extending on opposed sides of the airfoils between a leading edge and a trailing edge, the first and second fan blades respectively having different first and second airfoil thickness distributions, the first airfoil thickness distribution including a first baseline thickness and a first frequency modifier on the pressure side, the first frequency modifier defining an airfoil thickness differential relative to the first baseline thickness and being located at a first span distance away from the central hub, and the second airfoil thickness distribution including a second baseline thickness and a second frequency modifier on the pressure side, the second frequency modifier defining an airfoil thickness differential relative to the second baseline thickness and being located at a second span distance away from the central hub, the second span distance being different from the first span distance, both the first span distance and the second span distance being between 60% and 100% of the total span length of the first and second fan blades, the first and second frequency modifiers generating different natural vibration frequencies for each of the first and second fan blades, wherein a thickness of the airfoil of the first fan blade at the first span distance is less than the thickness of the second fan blade at the first span distance, and a thickness of the airfoil of the second fan blade at the second span distance is less than the thickness of the first fan blade at the second span distance, and the first span distance corresponds to a span-wise location of high strain energy and the second span distance corresponds to a span-wise location of low strain energy. 2. The mistuned fan according to claim 1, wherein the first natural vibration frequency of the first fan blades is less than a baseline frequency and the second natural vibration frequency of the second fan blades is greater than the baseline frequency, wherein the baseline frequency is defined as the natural vibration frequency of a fan blade having corresponding size and shape but absent said first and second frequency modifiers. 3. The mistuned fan according to claim 1, wherein the first and second frequency modifiers comprise at least one of a local region of reduced thickness relative to the first and second baseline thicknesses, respectively, and a local region of increased thickness relative to said first and second baseline thicknesses. 4. The mistuned fan according to claim 3, wherein the first and second frequency modifiers both comprise regions of reduced thickness relative to their respective baseline thicknesses, the thickness of the first and second fan blades at the respective first and second span distances are both less than their respective baseline thicknesses. 5. The mistuned fan according to claim 4, wherein the airfoil thickness of the first fan blades at the first span distance is between 40% and 60% of said baseline thickness of the pressure side. 6. The mistuned fan according to claim 4, wherein the thickness of the second fan blades at the second span distance is between 40% and 60% of said baseline thickness of the pressure side. 7. The mistuned fan according to claim 1, wherein a frequency separation between the first and second natural vibration frequencies is between 3 and 10%. 8. The mistuned fan according to claim 7, wherein the frequency separation between the first and second natural vibration frequencies is greater than or equal to 5%. 9. The mistuned fan according to claim 1, wherein the second span distance is greater than the first span distance. 10. The mistuned fan according to claim 1, wherein the first span distance of the first frequency modifier is disposed between 65% and 100% of the total span length, and the second span distance of the second frequency modifier is disposed between 80% and 100% of the total span length. 11. The mistuned fan according to claim 10, wherein the first span distance of the first frequency modifier is disposed between 65% and 90% of the total span length, and the second span distance of the second frequency modifier is disposed between 90% and 100% of the total span length. 12. The mistuned fan according to claim 1, wherein the first frequency modifier extends over a greater chord-wise extent of the pressure side of the first fan blades than does the second frequency modifier of the second fan blades. 13. The mistuned fan according to claim 12, wherein the first frequency modifier extends in a chord-wise direction substantially the entire cord-wise width of the first fan blades, and the second frequency modifier extends in the chord-wise direction from 10% to 90% of the entire chord-wise width of the second fan blades. 14. A mistuned compressor rotor assembly for a gas turbine engine, the mistuned compressor rotor assembly comprising a hub to which a plurality of airfoil blades are mounted, the airfoil blades having a full span length extending from the hub to tips of the airfoil blades, the airfoil blades each having an airfoil selected from at least first and second airfoil types and arranged as generally alternating with one another around the circumference of the rotor, the first airfoils having an airfoil thickness less than an airfoil thickness of the second airfoils at a first selected span of the respective blades, and the second airfoils having an airfoil thickness less than an airfoil thickness of the first airfoil at a second selected span of the respective blades different from the first selected span, both the first selected span and the second selected span being located within a radially outermost 40% of the full span length of airfoil blades. 15. The mistuned compressor rotor assembly of claim 14, wherein the first and second airfoils have substantially identical thickness distribution profiles but for in regions immediately adjacent the first and second selected spans. 16. The mistuned compressor rotor assembly of claim 14, wherein the first airfoil thickness at the first selected span at least partially provides the first airfoil blade with a lower natural vibration frequency than the second airfoil blade. 17. The mistuned compressor rotor assembly of claim 16, wherein the second airfoil thickness at the second selected span at least partially provides the second airfoil blade with a higher natural vibration frequency than the first airfoil blade. 18. The mistuned compressor rotor assembly of claim 17, wherein the second selected span corresponds in use to a span of a region of strain energy in the second airfoil blade lower than an average strain energy in the second airfoil blade. 19. The mistuned compressor rotor assembly of claim 14, wherein the first selected span corresponds in use to a span of a region of strain energy in the first airfoil blade higher than an average strain energy in the first airfoil blade. 20. The mistuned compressor rotor assembly of claim 19, wherein the second selected span corresponds in use to a span of a region of strain energy in the second airfoil blade lower than an average strain energy in the second airfoil blade. 21. The mistuned compressor rotor assembly of claim 14, wherein the rotor is a fan. 22. The mistuned compressor rotor assembly of claim 14, wherein the airfoil thickness of the first and second airfoils at the first and second selected span locations provide a natural vibration frequency difference between the first and second blade airfoil types of greater than 3%. 23. The mistuned compressor rotor assembly of claim 22, wherein the airfoil thickness of the first and second airfoils at the first and second selected span locations provide a natural vibration frequency difference between the first and second blade airfoil types of between 3% and 10%. 24. The mistuned compressor rotor assembly of claim 14, wherein the first selected span location is associated with a region of high strain energy and the second selected span location is associated with a region of low strain energy. 25. The mistuned compressor rotor assembly of claim 24, wherein the airfoil thickness of the first airfoils is less than the airfoil thickness of the second airfoils at the first selected span location. 26. The mistuned compressor rotor assembly of claim 25, wherein the airfoil thickness of the second airfoils is less than the airfoil thickness of the first airfoils at the second selected span location. 27. The mistuned compressor rotor assembly of claim 14, wherein the first selected span is disposed between 65% and 100% of the full span length, and the second selected span is disposed between 80% and 100% of the full span length. 28. The mistuned compressor rotor assembly of claim 27, wherein the first selected span is disposed between 65% and 90% of the full span length, and the second selected span is disposed between 90% and 100% of the full span length. 29. A compressor for a gas turbine engine, the compressor comprising: a plurality of first blades having a first airfoil thickness distribution defining a first natural vibration frequency;a plurality of second blades having a second airfoil thickness distribution different from the first airfoil thickness distribution and defining a second natural vibration frequency different from the first natural vibration frequency;the first airfoil thickness distribution including a first frequency modifier on the pressure side of the first blades at a first span distance away from the central hub and the second airfoil thickness distribution defining a second first frequency modifier on the pressure side of the second blades at a second span distance away from the central hub, the second span distance different from the first span distance, both the first span distance and the second span distance being between 60% and 100% of a full span length of the respective first and second blades, wherein first and second pressure side airfoil thicknesses are respectively defined by the first and second first frequency modifiers, wherein the first pressure side airfoil thickness of the first blades is less than a thickness of the second blades at the first span distance, and the second pressure side airfoil thickness of the second blades is less than a thickness of the first blades at the second span distance, and the first span distance corresponds to a span-wise location of high strain energy and the second span distance corresponds to a span-wise location of low strain energy. 30. The compressor of claim 29, wherein the first span distance of the first frequency modifier is disposed between 65% and 100% of the full span length, and the second span distance of the second frequency modifier is disposed between 80% and 100% of the full span length. 31. The compressor of claim 30, wherein the first span distance of the first frequency modifier is disposed between 65% and 90% of the full span length, and the second span distance of the second frequency modifier is disposed between 90% and 100% of the total span length. 32. A method of mitigating supersonic flutter in a compressor rotor, the rotor having a plurality of circumferentially disposed blades, the method comprising the steps of: providing a nominal airfoil having a nominal airfoil definition;determining a first span location associated with a region of high strain energy expected on the airfoil while in use on the compressor rotor, the first span location being disposed within an outermost 40% of the total span of the blade;determining a second span location associated with a region of low strain energy expected on the airfoil while in use on the compressor rotor, the second span location being different than the first span location, and both the first span location and the second span location being disposed within the outermost 40% of a total span of the blade;providing a first blade airfoil definition substantially the same as the nominal airfoil definition but having a different thickness at the first span location associated with the region of high strain energy;providing a second blade airfoil definition substantially the same as the nominal airfoil definition but having a different thickness at the second span location associated with the region of low strain energy; andproviding the compressor rotor where the blades are providing with the first and second blade airfoil definitions in an alternating fashion around the circumference of the rotor. 33. The method according to claim 32, wherein the region of high strain energy is the highest strain energy expected in the airfoil, and the region of low strain energy is the lowest strain energy expected in the airfoil. 34. The method according to claim 32, wherein the first blade airfoil definition is thinner than the nominal profile at the first span location associated with the region of high strain energy, and the second blade airfoil definition is thinner than the nominal profile at the second span location associated with the region of low strain energy. 35. The method according to claim 32, further comprising selecting relative airfoil thicknesses at the first and second span locations to provide a natural vibration frequency difference between the first and second blade airfoil definitions of greater than 3%. 36. The method according to claim 35, further comprising selecting the relative airfoil thicknesses at the first and second span locations to provide a natural vibration frequency difference between the first and second blade airfoil definitions of between 3% and 10%. 37. The method according to claim 32, wherein, at the first span location associated with the region of high strain energy, the airfoil thickness of the first blade airfoil definition is less than the airfoil thickness of the second blade definition at the same first span location. 38. The method according to claim 37, wherein, at the second span location associated with the region of low strain energy, the airfoil thickness of the second blade airfoil definition is less than the airfoil thickness of the first blade definition at the same second span location. 39. A method of mitigating supersonic flutter for a fan of a turbofan gas turbine engine, the method comprising: providing the fan with a plurality of fan blades, the fan blades circumferentially distributed around and extending a total span length away from a central hub, the fan blades composed of a plurality of pairs of circumferentially alternating first and second fan blades each having a different airfoil thickness distribution on a pressure side of the fan blades, the airfoil thickness distributions creating different natural vibrational frequencies of each of the first and second fan blades;selecting a desired frequency separation between natural vibration frequencies of the first and second fan blades in use, the frequency separation selected to mistune the pairs of fan blades to reduce the occurrence of supersonic flutter of the fan blades;determining respective first and second airfoil thickness distributions of the first and second fan blades to provide said desired frequency separation; andproviding the first airfoil thickness distribution on the pressure side of the first fan blade and providing the second airfoil thickness distribution on the pressure side of the second fan blade, wherein the first airfoil thickness distribution includes a first frequency modifier at a first span distance on the first fan blades, and the second airfoil thickness distribution including a second frequency modifier located on the second fan blades at a second span distance different from the first span distance, both the first span distance and the second span distance being between 60% and 100% of the total span length of the first and second span blade, and selecting the first span distance to correspond to a span-wise location of high strain energy and selecting the second span distance to correspond to a span-wise location of low strain energy. 40. The method according to claim 39, further comprising forming a thickness of the airfoil of the first fan blade at the first span distance to be less than the thickness of the second fan blade at the first span distance, and forming a thickness of the airfoil of the second fan blade at the second span distance which is less than the thickness of the first fan blade at the second span distance.