BLADE ROOT STRESS RELIEF

1 . A turbine engine section comprising:

a disk having:

fore and aft faces;

a periphery; and

a circumferential array of slots, each slot:

extending between the first and second faces; and

open to the periphery;

a plurality of blades, each having:

an airfoil; and

an attachment root, the attachment root captured in an associated one of the slots and having a cross-sectional profile interfitting with a cross-sectional profile of the associated slot to lock the blade to the disk against radial extraction,

wherein:

along at least a first circumferential side of the root and slot there is a static engagement; and

along at least a portion of said first circumferential side, there is a slot-wise tapering gap having a length of at least 10% of a local length along the slot and a height of at least 0.001 inch.

2 . The engine section of claim 1 wherein

the gap length is at least 0.25 inch.

3 . The engine section of claim 2 wherein

the gap height is at least 0.002 inch.

4 . The engine section of claim 1 wherein

the gap height is 0.004-0.006 inch.

5 . The engine section of claim 1 wherein

the gap is formed by a chamfer along a portion of the root, an adjacent portion of the slot being unchamfered.

6 . The engine section of claim 1 wherein:

the gap is an essentially constant chamfer along at least two protuberances/lobes of the attachment root and a fillet area therebetween.

7 . The engine section of claim 1 wherein:

the gap is formed by a chamfer extending continuously along an outboard face of a head of the root, outboard through at least one lobe and to an outboard lobe.

8 . A gas turbine engine blade comprising:

an airfoil having:

a leading edge;

a trailing edge;

a pressure side;

a suction side;

an inboard end; and

a tip;

a platform at the inboard end;

an attachment root depending from the platform and having:

leading and trailing end faces; and

first and second circumferential sides, generally respectively to the suction side and pressure of the airfoil,

wherein:

along at least a portion of at least one of said first and circumferential sides, the is a chamfer having a span of at least 0.25 inch and an end depth of at least 0.002 inch.

9 . The blade of claim 8 wherein:

the chamfer extends to the trailing end face and is along the second circumferential side.

10 . The blade of claim 8 wherein:

there is no corresponding chamfer directly on the opposite side of the root.

11 . The blade of claim 8 wherein:

there is a corresponding chamfer directly on the opposite side of the root but of lesser depth.

12 . The blade of claim 8 wherein:

the span is at least 0.6 inch.

13 . The blade of claim 8 wherein:

the depth is at least 0.004 inch.

14 . The blade of claim 8 wherein:

the chamfer is along an outboard protuberance/lobe of the root and a next protuberance/lobe inboard thereof and along a fillet area therebetween.

15 . A method for reengineering a configuration of a gas turbine engine from a baseline configuration to a reengineered configuration comprising:

providing said baseline configuration having:

a turbine rotor stack comprising:

a plurality of disks; and

a plurality of stages of blades, each stage carried by an associated disk of the plurality of disks; and

a plurality of stator vane stages, interspersed with the blade stages; and reengineering so as to:

provide a longitudinally-varying relief on at least one adjacent surface of the blade roots and disk slots of at least one of the disks.

16 . The method of claim 15 wherein:

the relief is only on an aft pressure side portion of the root.

17 . The method of claim 15 wherein:

in the baseline configuration, the blade center of gravity is ahead of an I max axis; and

in the reengineered configuration, the I max axis is forward of the baseline configuration I max axis and closer to the blade center of gravity.

18 . The method of claim 17 wherein:

the forward shift of the I max axis from the baseline configuration to the reengineered configuration is by at least 30% of the baseline configuration spacing of the blade center of gravity ahead of the I max axis.