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EP-4286650-B1 - ROTOR OF AN AIRCRAFT ENGINE COMPRISING A BLADE WITH A RIB INFLUENCING CRACK PROPAGATION

EP4286650B1EP 4286650 B1EP4286650 B1EP 4286650B1EP-4286650-B1

Inventors

  • AITCHISON, PAUL
  • STONE, PAUL
  • MANGARDICH, Dikran

Dates

Publication Date
20260513
Application Date
20230602

Claims (13)

  1. A rotor (20) of an aircraft engine (10), the rotor (20) comprising: a disc (30) having an outer rim surface (34) extending circumferentially about a rotation axis (A R ) and circumscribed by an outer rim diameter; and a plurality of blades (40) extending to radially outward of the outer rim surface (34) relative to the rotation axis (A R ), at least one blade (40) of the plurality of blades (40) including: an airfoil (46) spaced radially outward from the outer rim surface (34) relative to the rotation axis (A R ); a root (42) extending from the outer rim surface (34) to the airfoil (46), the root (42) corresponding to a fillet being radially bound between an inner transition radius and an outer transition radius of the blade (40), a difference between the outer and the inner transition radii defining a maximum radial height (R H ) of the fillet; and a tip (44) radially outward of the airfoil (46); at least one crack-mitigating rib (48) extending chordwise along the airfoil (46), the at least one crack-mitigating rib (48) being radially closer to the root (42) than to the tip (44), the at least one crack-mitigating rib (48) extending radially outwardly relative to the inner transition radius by no more than three times the maximum radial height (R H ) of the fillet, characterized in that the at least one crack-mitigating rib (48) has a cross-section defining an arcuate convex crest portion.
  2. The rotor of claim 1, wherein the at least one crack-mitigating rib (48) projects from the airfoil (46) by a rib depth (D) and extends radially by a rib height (H), the rib depth (D) being less than the rib height (H).
  3. The rotor of claim 2, wherein the rib depth (D) and the rib height (H) are defined such that a depth ratio (D/H) of the rib depth (D) over the rib height (H) is between 0.01 and 0.5.
  4. The rotor of claim 2 or 3, wherein the at least one crack-mitigating rib (48) has a cross-section including a concave transition portion (R F ) and a convex crest portion between the airfoil (46) and the concave transition portion (R F ), the rib height (H) being defined exclusive of the concave transition portion (R F ).
  5. The rotor of any one of claims 1 to 4, wherein the at least one crack-mitigating rib (48) includes a first rib (48 I ) and a second rib (48 II ) spaced radially from one another relative to the rotation axis (A R ).
  6. The rotor of claim 5, wherein the first and second ribs (48 I , 48 II ) are spaced from one another by a rib spacing (S I-II ) and respectively extend radially by a first rib height (H I ) and a second rib height (H II ), and the rib spacing (S I-II ), the first rib height (H I ) and the second rib height (H II ) are defined such that a spacing ratio (S I-II /H I +H II ) of the rib spacing (S I-II ) over a sum of the first and second rib heights (H I , H II ) is between 0.25 and 5.
  7. The rotor of any one of claims 1 to 6, wherein the at least one rib (48) includes a suction side rib (48 A ) and a pressure side rib (48 B ) respectively projecting from a suction side (46A) and a pressure side (46B) of the airfoil (46) by a suction side depth (D A ) and a pressure side depth (48 B ) greater than the suction side depth (D A ).
  8. The rotor of claim 7, wherein the suction side rib (48 A ) and the pressure side rib (48 B ) are portions of a same rib (48).
  9. The rotor of any one of claims 1 to 8, wherein the airfoil (46) defines a leading edge (E L ) and a trailing edge (E T ) and extends chordwise therebetween, and the at least one crack-mitigating rib (48) has a sloped end (48 E ) at a chordwise location of the airfoil (46) between the leading and trailing edges (E L , E T ).
  10. The rotor of any one of claims 1 to 9, wherein a radial distance (R RA , R RB ) between the at least one crack-mitigating rib (48) and the root (42) varies chordwise.
  11. The rotor of any one of claims 1 to 10, wherein the disc (30) and the plurality of blades (40) are parts of a monolithic bladed rotor (20), and wherein the at least one crack-mitigating rib (48) extends chordwise along the airfoil (46), the at least one crack-mitigating rib (48) having a cross-section defining an arcuate convex crest portion.
  12. A turbine engine (10) comprising: an axial compressor (14A) including the rotor (20) as defined in any of claim 1 to 11, and a rotor shroud defining a radially outer boundary of the axial compressor (14A) around the rotor (20).
  13. The turbine engine of claim 12, wherein the at least one crack-mitigating rib (48), the airfoil (46) and the root (42) of each blade (40) have tangential continuity with a rim (32) of the disk (30).

Description

TECHNICAL FIELD The disclosure relates generally to rotors and, more particularly, to rotor blades. BACKGROUND Rotors are typically used in turbine engine applications, and include a hub from which a plurality of circumferentially arranged rotor blades radially extend. The rotor blades may be subjected to stress fields during engine operation, which may extend into the rotor hub from which the blades extend. Such phenomenon may be accentuated in integrally bladed rotors (IBRs), whose rotor hub and blades form a unitary structure. US 2019/120064 and US 2019/024673 A1 disclose a prior art integrally bladed rotor having a double fillet. GB2251897A discloses a prior art rotor as set forth in the preamble of claim 1. SUMMARY In accordance with an aspect of the present disclosure, there is provided a rotor of an aircraft engine as recited in claim 1. There is also provided a turbine engine as recited in claim 12. Features of embodiments are set forth in the dependent claims. In examples of the above, the at least one crack-mitigating rib, the airfoil and the root of each blade have tangential continuity with the rim. BRIEF DESCRIPTION OF THE DRAWINGS Reference is now made to the accompanying figures in which: Fig. 1 is a schematic cross-sectional view of a turbine engine;Fig. 2 is a perspective view of an integrally bladed rotor having blades each provided with a crack-mitigating rib;Fig. 3 is an elevation view of a portion of the rotor of Fig. 2;Fig. 4 is a cross-section view of the portion of the bladed rotor taken along the line 4-4 of Fig. 3;Fig. 5 is a perspective view of a portion of a bladed rotor having blades each provided with a plurality of crack-mitigating ribs;Fig. 6 is a cross-section view of the portion of the bladed rotor taken along the line 6-6 of Fig. 5;Fig. 7 is a perspective view of a portion of a bladed rotor having blades each provided with a crack-mitigating rib having an end;Fig. 8 is a cross-section view of the portion of the bladed rotor taken along the line 8-8 of Fig. 3;Fig. 9 is a perspective view of a portion of a bladed rotor having blades each provided with a crack-mitigating rib having a pair of ends;Fig. 10A is a schematic radial cross-section view of a portion of an exemplary bladed rotor without crack-mitigating rib(s); andFig. 10B is a schematic radial cross-section view of a portion of an exemplary bladed rotor having blades each provided with a crack-mitigating rib. DETAILED DESCRIPTION The present disclosure relates to technologies for mitigating crack propagation in bladed rotors. In some embodiments, the mitigation of crack propagation in bladed rotors is achieved by way of a rib formed on an outer surface of an airfoil of one or more blades of the bladed rotor. The rib is configured to influence crack propagation to reduce the risk of a large and uncontained fragment of the bladed rotor being released from the bladed rotor due to fracture ultimately resulting from crack propagation during operation of the turbine engine. Fig. 1 illustrates a turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled, a compressor section 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases. Depending on the embodiment, the compressor section 14 includes one or more bladed rotors 20. The compressor section 14 thus includes one or more axial compressors 14A or compressor stages, each having a suitable rotor 20. The rotor 20 is rotatable about a rotation axis AR (Fig. 2) during operation of engine 10. In some embodiments of engine 10, the rotation axis AR may correspond to a central axis AC of engine 10. The rotor 20 may be part of a high-pressure spool or of a low-pressure spool of the engine 10. In some embodiments of the engine 10, the fan 12 may instead or in addition also be a rotor 20 as described herein. Although the engine 10 depicted in Fig. 1 is of the turbofan type, it is understood that aspects of the present disclosure are also applicable, mutatis mutandis, to other types (e.g., turboshaft, turboprop) of turbine engines, including hybrid aircraft engines. The compressor 14 defines a gas path P of the engine 10. The gas path P may be defined by and be disposed between a radially inner shroud and a radially outer shroud of the compressor 14. The gas path P may have an annular configuration and may surround the central axis AC. Lengthwise, the gas path P may extend principally axially relative to the central axis AC at the location of the rotor 20. The rotor 20 may be used as an airfoil-based axial compressor in the engine 10 and may compress and convey the air toward the combustor 16 during operation of the engine 10. The air being compressed through the gas path P in the region of the ro