Patent Application
20240117748
The disclosure relates generally to aircraft engines and, more particularly, to rotors used in such aircraft engines.
Aircraft engines, such as gas turbine engines, include rotors in compressor and/or turbines which usually include circumferentially spaced blades extending radially outwardly from a rotor disc and mounted thereto. The blades of such rotors are disposed within an air passage and typically face an upstream flow, such as pressurized air and/or hot combustion gases, that may infiltrate interstitial spaces between attached components of the rotors. Secondary air at a lower temperature may also infiltrate these interstitial spaces between attached components of the rotors. The presence of such colder secondary air may have a positive impact on the performance and/or durability of the rotor discs, seals and/or blades of rotors. However, secondary air ingested in such interstitial spaces may leak out via air leakage paths, which can limit the performance of rotor discs, seals and/or blades of such rotors.
In one aspect, there is provided a rotor assembly for an aircraft engine, comprising: blades circumferentially distributed about a central axis, the blades having airfoils and roots protruding from opposite sides of platform segments; a rotor disc having a peripheral face and slots extending radially inward from the peripheral face toward the central axis, the peripheral face extending from a first axial face of the rotor disc to a second axial face of the rotor disc, the roots of the blades being received within the slots along a blade insertion direction extending from the first axial face to the second axial face, the peripheral face defining recesses proximate the second axial face, a recess of the recesses located between two adjacent ones of the slots, the recess bounded by a step, the recess located axially between the step and the second axial face relative to the central axis; feather seals located radially between the peripheral face of the rotor disc and the platform segments of the blades, a feather seal of the feather seals having a core axially extending from a trailing end proximate the first axial face to a leading end proximate the second axial face, the core circumferentially overlapping a gap defined between two adjacent ones of the platform segments and tabs protruding from the core, the tabs including: trailing tabs proximate the trailing end of the core and positioned axially outside the recess; and
The rotor assembly described above may include any of the following features, in any combinations.
In some embodiments, the step has a height taken in a radial direction relative to the central axis, the fillet having a radius greater than the height.
In some embodiments, the radius of the fillet is at least 1.5 times the height of the step.
In some embodiments, the radius is about two times the height of the step.
In some embodiments, the radius is at most a width of the leading tab taken along the blade insertion direction.
In some embodiments, the leading tab is axially aligned with the recess and defines the fillet.
In some embodiments, the leading tab is a first lateral leading tab protruding from the core transversally to the blade insertion direction.
In some embodiments, the leading tabs includes a second lateral leading tab protruding from the core transversally to the blade insertion direction and away from the first lateral leading tab, the second lateral leading tab being axially offset from the first lateral leading tab.
In some embodiments, the second lateral leading tab is axially offset from the recess.
In some embodiments, the second lateral leading tab defines a second fillet, a second radius of the second fillet being at least 1.5 times a height of the step taken in a radial direction relative to the central axis.
In some embodiments, the core defines a dimple between the first lateral leading tab and a trailing tab of the trailing tabs, the dimple matingly engaged by a bump of a segment of the platform segments.
In some embodiments, the leading tabs include a longitudinal leading tab protruding from the core, the trailing tabs including a longitudinal trailing tab protruding from the core, the longitudinal leading tab and the longitudinal trailing tab extending away from one another, the longitudinal trailing tab positioned axially outside the recess, the longitudinal leading tab axially aligned with the recess, the longitudinal leading tab located forward of the leading tab relative to the blade insertion direction.
In another aspect, there is provided a turbine section of an aircraft engine comprising a rotor assembly having: blades circumferentially distributed about a central axis, the blades having airfoils and roots protruding from opposite sides of platform segments; a rotor disc having a peripheral face and slots extending radially inward from the peripheral face toward the central axis, the peripheral face extending from a first axial face of the rotor disc to a second axial face of the rotor disc, the roots of the blades being received within the slots along a blade insertion direction, the roots being removable from the slots solely along a direction opposite the blade insertion direction, the peripheral face defining recesses, a recess of the recesses located between two adjacent ones of the slots, the recess bounded by a step, the recess located forward of the step relative to the blade insertion direction; feather seals located radially between the peripheral face of the rotor disc and the platform segments of the blades, a feather seal of the feather seals having a core circumferentially overlapping a gap defined between two adjacent ones of the platform segments and tabs protruding from the core, the tabs including: trailing tabs axially offset from the recess, and leading tabs axially forward of the trailing tabs relative to the blade insertion direction, a leading tab of the leading tabs extending from a root at the core to a tip, the leading tab having one or more of: the tip axially positioned outside of the recess; and a fillet at an intersection between the tip and an edge of the leading tab, the edge extending between the tip and the core, and facing the step.
The turbine section described above may include any of the following features, in any combinations.
In some embodiments, the step has a height taken in a radial direction relative to the central axis, the fillet having a radius greater than the height and/or wherein the radius is at most a width of the leading tab taken along the blade insertion direction.
In some embodiments, the leading tab is axially aligned with the recess and defines the fillet.
In some embodiments, the leading tab is a first lateral leading tab protruding from the core transversally to the blade insertion direction.
In some embodiments, the leading tabs includes a second lateral leading tab protruding from the core transversally to the blade insertion direction and away from the first lateral leading tab, the second lateral leading tab being axially offset from the first lateral leading tab, the second lateral leading tab axially offset from the recess.
In yet another aspect, there is provided a feather seal for a rotor assembly of an aircraft engine, the feather seal comprising: a core extending along a longitudinal axis from a from a leading end to a trailing end, the feather seal having a seal insertion direction extending from the trailing end to the leading end, the feather seal being insertable between blades and a rotor disc solely along the seal insertion direction; tabs protruding from the core from roots at the core to tips, the tips being offset from the roots along a vertical direction normal to the longitudinal axis, the tabs including trailing tabs, and leading tabs axially forward of the trailing tabs relative to the seal insertion direction, a leading tab of the leading tabs defining a fillet at an intersection between a corresponding tip of the tips and an edge of the leading tab, the edge facing a trailing tab of the trailing tabs.
The feather seal described above may include any of the following features, in any combinations.
In some embodiments, the intersection is free of a sharp corner.
In some embodiments, a ratio of a radius of the fillet to a length of the leading tab from the corresponding tip to a corresponding root of the roots ranges from 0.25 to 0.75.
Reference is now made to the accompanying figures in which:
Referring to
Referring to
In embodiments where the rotor assembly 20 is disposed in the turbine section 18 of the engine downstream of the combustor 16, the components of the rotor assembly 20 may have to sustain high pressures and temperatures during operation of the gas turbine engine 10. Such operating conditions may affect the durability of said components. Hot combustion gases and/or air upstream of the rotor assembly 20 may infiltrate interstitial spaces between components connecting/interfacing together in the rotor assembly 20. However, colder air which circulates within the gas turbine engine 10 may reduce the temperature of the components in fluid communication with the hot combustion gases. In operation, such colder air (often referred to as secondary air) flowing upstream of the rotor assembly 20 may be ingested in these interstitial spaces between components connecting/interfacing together in the rotor assembly 20. Increasing said colder air retention in interstitial spaces between components of the rotor assembly 20 may be desirable in order to limit (reduce) the rate at which these components heat up during normal operation of the gas turbine engine 10 and/or so as to limit the negative impacts of infiltration of hot combustion gases through these interstitial spaces on the efficiency of the gas turbine engine 10 and/or limit the negative impacts of excessive secondary air flowing through these interstitial spaces. As discussed below, components of the rotor assembly 20 may be adapted to increase the retention of secondary air at selected locations about the rotor disc 30, more particularly at a disc/blades interface.
In the embodiment shown, the rotor assembly 20 comprises the rotor disc 30 and the rotor blades 40 distributed circumferentially about the central axis 11 and removably connected to the rotor disc 30. Multiple rotor assemblies 20 may be provided, each with an associated stator disposed either downstream (compressor) or upstream (turbine) of the rotor, such as to form multiple compressor or turbine stages as the case may be. These stages may correspond to compression stages or pressure stages in certain embodiments. The blades 40 may be equally circumferentially spaced apart from one another about the disc 30.
As seen in
The rotor disc 30 has a plurality of fixing members 34 defined therein through the peripheral face 33 and circumferentially spaced apart from one another. As in
In an embodiment, the fixing members 34 have a profiled contour which may be, for example, formed by a series of lobes having increasing circumferential widths from the radially outermost lobe (“top lobe”), to the radially innermost lobe (“bottom lobe”), with, in some cases, a radially central lobe (“mid lobe”) disposed therebetween and having an intermediate lobe width. Such a multi-lobed profiled contour is typically referred to as a “firtree” (or “fir tree”), because of this characteristic shape. It is to be understood from the above that the slots 35 may have a complementary firtree shape, as in some embodiments side walls of the slots 35 may define a respective side of the profiled contour of the fixing members 34. Whether or not in the shape of a firtree or lobes, the fixing members 34 and slots 35 define mechanical interferences that form abutments that prevent a radial outward movement of blades 40 connected to the disc 30.
As visible in
Still referring to
Each of the blades 40 has a blade root 41, an airfoil 42 and a platform or platform segments 43 radially disposed between the blade root 41 and the airfoil 42, the platform segments 43 extending laterally to (into opposing relationship with) corresponding platform segments 43 of adjacent ones of the blades 40. Stated differently, the blades 40 have the airfoils 42 and the blade roots 41 protruding from opposite sides of the platform segments 43. These portions of the blade 40 may all merge together to form a single monolithic piece blade, though a multi-piece configuration is also possible.
The blade root 41 of each of the blades 40 may be received within a corresponding one of the slots 35 of the disc 30. The root 41 has a shape and size that dovetail with the shape and size of the corresponding slot 35. The size of the blade roots 41 is slightly smaller than or equal to the size of the slots 35 to allow the blade roots 41 to slide within the slots 35 along a blade insertion direction D1 when connecting the blades 40 to the disc 30. The blade insertion direction D1 extends from the front axial face at the front end portion 31 to the rear axial face at the rear end portion 32. In an alternate embodiment, the blade insertion direction D1 extends from the rear axial face to the front axial face. Once received in the slot 35, the blade root 41 may be secured therein with a retaining member (not shown). The retaining member may be any fastening structure such as a retaining ring, a rivet connector or any other suitable types of retaining member that may secure the blade roots 41 inside respective slots 35 to prevent axial movement between the blade roots 41 and the slots 35 in at least one direction, for instance the direction opposite the insertion direction of the blade root 41 within the slot 35.
The airfoil 42 of the blade 40 extends generally or partially transversally to the direction of the flow path of air/combustion gases in the core flow path 19 (
The platform segment 43 has a curved profile forming a trailing flange 44 protruding rearwardly and a leading flange 45 protruding forwardly. As shown in
As shown in
When used within the turbine section 18, and to ensure that a blade feather seal performs its functions, it may be made from a high-temperature-resistant material, have a shape conforming as closely as possible to the blade under-platform pocket's three-dimensional surface profile and be as light as possible to minimize its centrifugal load contribution on the rotor assembly 20. Feather seals may also have features such as side tabs that prevent them from moving within the pockets 48 and have a shape that may allow it to be made by stamping a pre-cut piece of sheet metal in a forming die.
As shown in
Referring now to
The feather seal 50 has a core 51 circumferentially overlapping one of the gaps 49 (
In the embodiment shown, the tabs include leading tabs and trailing tabs located rearward of the leading tabs relative to the blade insertion direction D1. The blade insertion direction D1 may correspond to a seal insertion direction along which the feather seals 50 and the blades 40 are inserted. These two directions may be parallel to each other. The leading tabs may include three leading tabs, namely, a longitudinal leading tab 52 protruding from the core 51 along the blade insertion direction D1, a first lateral leading tab 53 protruding from the core 51 transversally to the blade insertion direction D1, and a second lateral leading tab 54 protruding from the core 51 transversally to the blade insertion direction D1 and away from the first lateral leading tab 53. The second lateral leading tab 54 may be axially offset from the first lateral leading tab 53. In the embodiment shown, the first lateral leading tab 53 is located forward of the second lateral leading tab 54 relative to the blade insertion direction D1. The trailing tabs may include three trailing tabs, namely, a longitudinal trailing tab 55 protruding from the core 51 along the direction D2 opposite the blade insertion direction D1 and extending away from the longitudinal leading tab 52, a first lateral trailing tab 56 protruding from the core 51 transversally to the blade insertion direction D1, and a second lateral trailing tab 57 protruding from the core 51 transversally to the blade insertion direction D1 and away from the first lateral trailing tab 56.
The leading and trailing tabs 52, 53, 54, 55, 56, 57 extend from roots at the core 51 to tips. The tips of the leading and trailing tabs 52, 53, 54, 55, 56, 57 are in abutment against the peripheral face 33 of the rotor disc 30. The trailing tabs 55, 56, 57 abut the peripheral face 33 outside the recess 33A. In the embodiment shown, at least two of the leading tabs 52, 53, 54 abut the peripheral face 33 within the recess 33A. The longitudinal leading tab 52 has its tip axially aligned with the recess 33A; said tip may thus abut the peripheral face 33 within the recess 33A. The longitudinal trailing tab 55 has its tip axially offset from the recess 33A; said tip may thus abut the peripheral face 33 outside the recess 33A.
To prevent the first and second lateral leading tabs 53, 54 from getting caught on the step 33B while removing the feather seals 50 and the blades 40 along the blade removal direction D2, the first lateral leading tab 53 may have one or more of its tip axially positioned outside the recess 33A, and a fillet at an intersection between its tip and an edge of the first lateral leading tab 53; the edge extending between the tip and the core 51 and facing the step 33B. Similarly, the second lateral leading tab 54 may have one or more of its tip axially positioned outside the recess 33A, and a fillet at an intersection between its tip and an edge of the second lateral leading tab 54; the edge extending between the tip and the core 51 and facing the step 33B. In the present embodiment, the first lateral leading tab 53 is axially aligned with the recess 33A. The first lateral leading tab 53 may thus abut the peripheral face 33 within the recess 33A. The second lateral leading tab 54 is positioned axially outside the recess 33A. The second lateral leading tab 54 is thus axially offset from the recess 33A. The second lateral leading tab 54 may thus abut the peripheral face 33 outside the recess 33A. In an alternate embodiment, both of the first and second lateral leading tabs 53, 54 may be positioned axially outside the recess 33A and may thus abut the peripheral face 33 outside the recess 33A. The first lateral leading tab 53, which is axially aligned with the recess 33A, defines a fillet 53A at an intersection between its tip 53B and an edge 53C that faces the step 33B. This edge 53C faces the first lateral trailing tab 56. A radius of the fillet 53A is greater than a height H1 of the step 33B. The height H1 is taken in a radial direction relative to the central axis 11. Preferably, the radius of the fillet 53A is at least 1.5 times the height H1 of the step 33C. The radius of the fillet 53A may be at least 2 times the height H1 of the step 33C. The radius of the fillet 53A may be at most a width W1 of the first lateral leading tab 53 taken along the blade insertion direction D1. The width W1 may be an average width of the first lateral leading tab 53 since the width may vary from the root to the tip. It may be the width at the root or, alternatively, the width at the tip. The height H1 may be about 0.02 inch whereas the radius may be about 0.06 inch. The expression “about” in the present disclosure encompasses variations by plus or minus 20%. A ratio of the radius of the fillet 53A to a length of the first lateral leading tab 53 from its root 53D to its tip 53B may range from 0.25 to 0.75. In the embodiment shown, the second lateral leading tab 54 also defines a second fillet 54A between its tip and an edge facing the second lateral trailing tab 57. The second fillet 54A may have a second radius at least 1.5 times, preferably 2 times, the height H1 of the step 33B. The second radius may have the same characteristics as the radius of the fillet 53A of the first leading lateral tab 53. The second lateral leading tab 54 may be free of this second fillet 54A since its is located outside the recess 33A and, thus, may be less subjected to be caught on the step 33B. The lateral leading tabs abutting the peripheral face 33 within the recess 33A may be free of a sharp corner to limit chances of these tabs getting caught on the step 33B.
Referring more particularly to
The feather seal 50 described herein includes upstream side tabs having large fillets at their free end. One of these two upstream side tabs is located further downstream than the other to ensure that it always sits on the upper portion of the disc outer diameter step. This may further reduce the risk of back-lock when removing the blades 40 and the feather seals 50 from the disc 30 along the blade removal direction D2. The axial position of the other upstream side tab is located to avoid interaction with the dimple 58 of the core 51 to avoid stamping/manufacturing issues during manufacturing of the feather seal 50. In fact, the feather seal 50 may be cut from sheet metal in a coplanar or two dimensional configuration to create a blank. A three dimensional shape may be imparted to the blank by stamping said blank.
The disclosed feather seal 50 may improve sealing efficiency and may eliminate assembly mistake, which may cause loss of sealing efficiency and high bending loads within the seal itself as it may not sit properly in the pocket 48.
In an alternate embodiment, each of the leading and trailing lateral tabs may be three-tangent at their tips. In other words, tips of the leading and trailing lateral tabs may be circular and my have a radius corresponding to a width of said tabs. Put differently, each of the leading and trailing lateral tabs may be free of sharp corner on all sides.
The embodiments described in this document provide non-limiting examples of possible implementations of the present technology. Upon review of the present disclosure, a person of ordinary skill in the art will recognize that changes may be made to the embodiments described herein without departing from the scope of the present technology. Yet further modifications could be implemented by a person of ordinary skill in the art in view of the present disclosure, which modifications would be within the scope of the present technology.
