A turbine blade for a gas turbine engine comprises an airfoil having a pressure side, a suction side, a span direction and a chord-wise direction. The airfoil has an airfoil span on a pressure line being a projection of the stacking line onto the pressure side. The airfoil has a plurality of chords
A turbine blade for a gas turbine engine comprises an airfoil having a pressure side, a suction side, a span direction and a chord-wise direction. The airfoil has an airfoil span on a pressure line being a projection of the stacking line onto the pressure side. The airfoil has a plurality of chords extending between a leading edge and a trailing edge of the airfoil. A generally round dimple is disposed on the pressure side. The dimple is contained in an area extending on the stacking line between 0% and 23% of the airfoil span from the inner end, and in the chord-wise direction between 0% of a first chord and 82% of a second chord from the leading edge. The dimple is configured to initiate fracture of the blade at a predetermined speed of rotation. A method of preventing rupture of a disk of a turbine rotor is also presented.
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1. A turbine rotor for a gas turbine engine, the turbine rotor comprising: a turbine disk having a rupture rotational speed at which the turbine disk ruptures;a plurality of blades receivable in the turbine disk each having an airfoil, the airfoil having a pressure side and a suction side, the airfo
1. A turbine rotor for a gas turbine engine, the turbine rotor comprising: a turbine disk having a rupture rotational speed at which the turbine disk ruptures;a plurality of blades receivable in the turbine disk each having an airfoil, the airfoil having a pressure side and a suction side, the airfoil having a span direction and a chord-wise direction, the airfoil extending from an inner end to an outer end in the span direction and from a leading edge to a trailing edge in the chord-wise direction, the airfoil having an airfoil span on a pressure line being a projection of a stacking line onto the pressure side, the airfoil having a plurality of chords extending between the leading edge and the trailing edge; anda generally round dimple on the pressure side of the airfoil, the dimple being contained in an area extending on the stacking line between 0% and 23% of the airfoil span from the inner end, and in the chord-wise direction between 0% of a first chord and 82% of a second chord from the leading edge, the dimple corresponding to material removed from the airfoil, the material removed being greater than 15% and up to a maximum of 39% of un-dimpled nominal airfoil thickness of the airfoil, the dimple being configured to initiate fracture at the dimple of each one of the plurality of blades at a predetermined rotational speed being less than the rupture rotational speed of the turbine disk. 2. The turbine rotor as defined in claim 1, wherein the dimple is configured to cause the corresponding blade to fracture at an abnormal predetermined speed of rotation, and wherein the corresponding blade is a low stress blade that would not by itself fracture at the abnormal predetermined speed of rotation. 3. The turbine rotor as defined in claim 2, wherein the abnormal predetermined speed of rotation is determined as a function of a maximum allowed rotation speed of the turbine disk. 4. The turbine rotor as defined in claim 1, wherein the predetermined rotational speed is about 10% less than the rupture rotational speed of the turbine disk. 5. The turbine rotor as defined in claim 2, wherein the corresponding blade is a power turbine blade of a first stage power turbine rotor. 6. The turbine rotor as defined in claim 1, wherein the first chord and the second chord are distinct from each other. 7. The turbine rotor as defined in claim 1, wherein the dimple extends on the stacking line between 0% and 21% of the airfoil span from the inner end. 8. The turbine rotor as defined in claim 1, wherein the dimple extends on the stacking line between 0% and 19% of the airfoil span from the inner end. 9. The turbine rotor as defined in claim 1, wherein the dimple extends in the chord-wise direction between 0% of the first chord and 71% of the second chord from the leading edge. 10. The turbine rotor as defined in claim 1, wherein the dimple extends in the chord-wise direction from 5% of the first chord to 76% of the second chord from the leading edge. 11. The turbine rotor as defined in claim 1, wherein the dimple extends in the chord-wise direction from 9% of the first chord to 82% of the second chord from the leading edge. 12. The turbine rotor as defined in claim 1, wherein the dimple corresponds to material removed from the airfoil corresponding up to 35% of the un-dimpled nominal airfoil thickness. 13. A low stress turbine blade for a gas turbine engine, the low stress turbine blade comprising: a root portion adapted to be received in a turbine disk of the gas turbine engine, the turbine disk having a rupture rotational speed at which the turbine disk ruptures;an airfoil extending from the root portion having a pressure side and a suction side, the airfoil extending radially from an inner end to an outer end thereof along a span of the airfoil; anda weakened area of reduced airfoil cross-section disposed on the pressure side adjacent to a leading edge of the airfoil and the inner end, the weakened area extending on a blade pressure line from 0% up to 23% of the span from the inner end, the weakened area extending in a chord-wise direction from 5% of a first chord up to 82% of a second chord from the leading edge, the weakened area configured to initiate fracture at the weakened area of the blade at a predetermined rotational speed being less than the rupture rotational speed of the turbine disk. 14. The low stress turbine blade as defined in claim 13, wherein the weakened area comprises a generally round dimple on the pressure side of the blade. 15. The low stress turbine blade as defined in claim 13, wherein the weakened area comprises a dimple, the dimple having a maximum thickness of 39% of a local thickness of the airfoil and a minimum thickness greater than 15% of the local thickness of the airfoil. 16. The low stress turbine blade as defined in claim 13, wherein the blade has a 10% burst margin. 17. The low stress turbine blade as defined in claim 13, wherein the blade is a power turbine blade of a first stage power turbine rotor. 18. A method of preventing rupture of a turbine disk of a turbine rotor, the turbine disk having a rupture rotational speed at which the turbine disk ruptures, the method comprising: driving the disk in rotation, the disk carrying a set of circumferentially spaced-apart airfoil shaped blades, each of the blades having a generally round dimple disposed on a pressure side of the blade next to the disk, the dimple being contained in an area extending on a stacking line from 0% up to 23% of an airfoil span on the stacking line from an inner end of the blade, and chord-wise from the leading edge from 5% of a first chord up to 82% of a second chord; andwhen rotating the disk above a predetermined rotational speed being less than the rupture rotational speed, locally increasing stresses on the dimple of at least one blade and causing the at least one blade to break at the dimple, thereby preventing rupture of the disk. 19. The method as defined in claim 18, wherein driving the turbine disk in rotation further comprises driving in rotation the plurality of blades with the dimple having a maximum thickness of 39% of a local thickness of the airfoil and a minimum thickness greater than 15% of the local thickness of the airfoil. 20. The method as defined in claim 18, wherein the predetermined rotational speed is about 10% less than the rupture rotational speed.
De Moura, Raul Fernando; Jean, Pierrick Bernard; Le Hong, Son; Lombard, Jean-Pierre Francois, Blade made of composite material comprising a damping device.
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