WO2018175356A1

WO2018175356A1 - Alternately mistuned blades with modified under-platform dampers

Info

Publication number: WO2018175356A1
Application number: PCT/US2018/023218
Authority: WO
Inventors: Yuekun ZHOU; Jr. Nicholas F. MARTIN
Original assignee: Siemens Aktiengesellschaft
Priority date: 2017-03-22
Filing date: 2018-03-20
Publication date: 2018-09-27

Abstract

The invention is related to a bladed rotor system (10) for a turbomachine including blades (14) mounted circumferentially on a rotor disc (12), and a plurality of under-platform dampers (30). Each damper (30) is located between adjacent blade platforms (24) and frictionally contacts a radially inner surface (24a) of the adjacent blade platforms (24). The plurality of dampers (30) includes a first set (H) of dampers (30) and a second set (L) of dampers (30). The dampers (30) of the first set (H) are distinguished from the dampers (30) of the second set (L) either by a geometry of the damper (30) or by its respective position, that is unique to the respective set (H, L). Dampers (30) of the first (H) and second set (L) are positioned alternately in a periodic fashion in a circumferential direction, thereby providing a frequency mistuning to stabilize flutter of the blades (30).

Description

ALTERNATELY MISTUNED BLADES WITH MODIFIED UNDER- PLATFORM DAMPERS

BACKGROUND

1. Field

[0001] The present invention is relates to rotating blades in a turbomachine, and in particular, to a row of blades with alternate frequency mistuning for improved flutter resistance.

2. Description of the Related Art

[0001] Turbomachines, such as gas turbine engines include multiple stages of flow directing elements along a hot gas path in a turbine section of the gas turbine engine. Each turbine stage comprises a circumferential row of stationary vanes and a circumferential row of rotating blades arranged along an axial direction of the turbine section. Each row of blades may be mounted on a respective rotor disc, with the blades extending radially outward from the rotor disc into the hot gas path. A blade includes an airfoil extending span-wise along the radial direction from a root portion to a tip of the airfoil.

[0002] Typical turbine blades at each stage are designed to be identical aerodynamically and mechanically. These identical blades are assembled together into the rotor disc to form a bladed rotor system. During engine operation, the bladed rotor system vibrates in system modes. This vibration may be more severe in large blades, such as in low pressure turbine stages. An important source of damping in the modes is from aerodynamic forces acting on the blades when the blades vibrate. Under certain conditions, the aerodynamic damping in some of the modes may become negative, which may cause the blades to flutter. When this happens, the vibratory response of the system tends to grow exponentially until the blades either reach a limit cycle or break. Even if the blades achieve a limit cycle, their amplitudes can still be large enough to cause the blades to fail from high cycle fatigue.

[0003] Alternate frequency mistuning can cause system modes to be distorted, so that the resulting new, mistuned system modes are stable, i.e., they all have positive aerodynamic damping. It is therefore desirable to be able to design blades with a certain amount of predetermined alternate mistuning. Alternate mistuning may be implemented in blades by having the blades in the blade row alternate between high and low frequencies in periodic fashion in the circumferential direction. So far, alternate mistuning of blades has been implemented by modifying the mass and/or geometry of the airfoil in alternate blades in a blade row.

[0004] However, there remains a room for improvement to better address the problem of blade vibration.

SUMMARY

[0005] Briefly, aspects of the present invention are directed to a row of blades with modified under-platform dampers to provide alternate frequency mistuning for improved flutter resistance.

[0006] According to a first aspect of the invention, a bladed rotor system for a turbomachine is provided. The bladed rotor system comprises a circumferential row of blades mounted on a rotor disc. Each blade comprises a platform, a root extending radially inward from the platform for mounting the blade to the rotor disc, and an airfoil extending span-wise radially outward from the platform. During operation, platforms of adjacent blades align circumferentially to define an inner diameter boundary for a working fluid flow path. The bladed rotor system further includes a plurality of dampers, each damper being located between adjacent platforms and frictionally contacting a radially inner surface of said adjacent platforms. The plurality of dampers comprise a first set of dampers and a second set of dampers. The dampers of the first set are distinguished from the dampers of the second set by a geometry of the damper that is unique to the respective set. Dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades.

[0007] According to a second aspect of the invention, method for servicing a bladed rotor system is provided. The bladed rotor system comprises a circumferential row of blades mounted on a rotor disc, each blade comprising a platform, a root extending radially inward from the platform for mounting the blade to the rotor disc, and an airfoil extending span-wise radially outward from the platform. The bladed rotor system further comprises a plurality of dampers, each damper being installed between adjacent platforms and frictionally contacting a radially inner surface of said adjacent platforms. The method comprises modifying a geometry of at least a subset of the plurality of installed dampers or providing replacement dampers for at least a subset of the plurality of installed dampers. As a result, first and second sets of dampers are obtained, in which the dampers of the first set are distinguished from the dampers of the second set by a geometry of the damper that is unique to the respective set. The method further comprises installing the modified or replacement dampers, such that dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades.

BRIEF DESCRIPTION OF THE DRAWINGS

[0008] The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.

[0009] FIG. 1 schematically illustrates, in axial view, a portion of a bladed rotor system comprising under-platform dampers;

[0010] FIG. 2 schematically illustrates, in perspective view, an embodiment of the present invention implementing mistuning of under-platform dampers;

[0011] FIG. 3 is a perspective view, looking from the pressure side, illustrating an assembly of an under-platform damper;

[0012] FIG. 4 is a perspective sectional view, along the section IV in FIG. 3;

[0013] FIG. 5 is a radial bottom view of an under-platform damper according an exemplary configuration;

[0014] FIG. 6, 7 and 8 illustrate exemplary embodiments of different damper cross-sections along the section line C-C in FIG. 5; and

[0015] FIG. 9 graphically illustrates alternate mistuning in a row of turbine blades.

DETAILED DESCRIPTION

[0016] In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

[0017] In the drawings, the direction A denotes an axial direction parallel to an axis of the turbine engine, while the directions R and C respectively denote a radial direction and a circumferential direction with respect to said axis of the turbine engine.

[0018] Referring now FIG. 1, a portion of a bladed rotor system 10 is illustrated. The bladed rotor system 10 includes a circumferential row of blades 14 mounted on a rotor disc 12. Each blade 14 comprises an airfoil 16 extending span-wise along a radial direction from a platform 24 to an airfoil tip 20. As known to one skilled in the art, the airfoil 16 may comprise a generally concave pressure side 2 and a generally convex suction side 4, joined at a leading edge 6 and at a trailing edge (not shown). The blade 14 is mounted on the disc 12 via an attachment structure, referred to as a blade root, which extends radially inward from the platform 24. In the illustrated embodiment, the root 18 has a fir-tree shape, which fits into a correspondingly shaped slot 26 in the rotor disk 12. In the context of the illustrated embodiments, it may be assumed that each blade 14 of the blade row has essentially identical fir-tree attachments. Each platform 24 comprises a radially inner surface 24a and a radially outer surface 24b. In operation, the platforms 24 of adjacent blades 14 align circumferentially, without necessarily contacting each other. The circumferential alignment of the radially outer surfaces 24b of neighboring platforms 24 form an inner diameter flow path boundary for a working fluid of the turbomachine. The airfoils 16 extend radially outward into the flow path and extract energy from the working fluid, which causes the blades 14 to rotate about a rotation axis 22.

[0019] As the airfoils 16 extract energy from the working fluid, the working fluid exerts a loading force on the airfoils 16. Variations in the loading force cause the blades 14 to deflect and vibrate. This vibration has a broad spectrum of frequency components, with greatest amplitude at the natural resonant frequency of the blades 14. When the blades 14 are unshrouded, the vibration is primarily tangential to the direction of rotation, i.e. the circumferential direction. There, is also a secondary vibration component in the direction of fluid flow, i.e. the axial direction. The above- mentioned vibrations may be reduced by incorporating under-platform dampers 30. Each damper 30 may be constructed as a rigid element which spans the gap between a pair of adjacent platforms 24, and contacting the radially inner surfaces 24a of the adjacent platforms 24. A friction force is thereby applied by the damper 30 to the platforms 24. This friction force reduces blade to blade vibration and consequently reduces individual blade vibration. Conventionally, the dampers 30 of the blade row were designed to be geometrically identical to each other.

[0020] An underlying idea of the illustrated embodiments involves designing the bladed rotor system 10 to have alternate mistuning of blade frequencies by modifying the geometry of the dampers 30 in an alternating pattern. Modifying the geometry may include, for example, modifying the axial length and/or axial position and/or cross-sectional shape of the dampers 30. Herein, the dampers 30 of the bladed rotor system 10 may be divided into first and second sets of dampers 30, designated respectively as H and L. The dampers 30 of the first set H are distinguished from the dampers 30 of the second set L by a geometry of the damper 30 that is unique to the respective set H or L. Dampers 30 of the first set H and the second set L may be positioned alternately in a periodic fashion in the circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades 14. The term "alternately" may refer to every other damper, or include a continuous group of dampers with similar vibratory characteristics. In the illustrated embodiment, the dampers 30 of the first set H and the second set L alternate in groups of two in a circumferential direction, in a pattern HHLLHH. In further embodiments groups of one or more dampers of the first set H and the second set L may alternate in a periodic fashion along the circumferential direction in the blade row, for example in patterns including HHLLHH, HHHLLHHH, HHHLLLHHH etc.

[0021] Illustrated embodiments of the present invention are directed to freestanding blades. In the context of this specification, a free-standing blade may be understood to be an unshrouded blade, i.e., a rotatable blade comprising an airfoil extending span-wise radially outward from a blade platform to an airfoil tip, without any shroud attached to the airfoil at the tip or at any point between the platform and the airfoil tip. However, the illustrated embodiments are exemplary, and aspects of the present invention may be extended to shrouded blades.

[0022] In a preferred embodiment, the above-described alternate mistuning may be achieved without modifying the geometry of the airfoils. That is, all the airfoils 16 in the circumferential row of blades 14 may have essentially identical cross-sectional geometry about a rotation axis 22. This makes it easier to design the airfoil to have optimum aerodynamic efficiency since a uniform airfoil geometry has to be considered. Moreover, the illustrated embodiments make it possible to employ alternate mistuning for blades with hollow airfoils, for example, containing internal cooling channels. The design of hollow airfoils is more constrained than the design of solid airfoils. The use of mistuned under-platform dampers provide a possibility for implementing alternate mistuning for such hollow blades without compromising the aero-efficiency.

[0023] FIG. 2 schematically illustrates an arrangement of mistuned under-platform dampers 30 according an example embodiment of the invention. In this embodiment, the dampers 30 of the first set H and the dampers 30 of the second set L differ in their axial lengths, i.e., the damper length along the longitudinal axis of the turbine. As shown, the dampers 30 of the first set L have an axial length t, while the dampers 30 of the second set L have a reduced axial length t-Δΐ. Assuming the material and the cross-sectional geometry of the dampers 30 in the first and second sets H and L to be constant, the difference At in the axial length results in different damped frequencies produced by the dampers 30 of the first set H and the dampers 30 of the second set L, which is due to the differences in mass and contact loading between the dampers of the two sets. By installing the dampers 30 of the first and second sets H and L in a defined pattern, the damped frequency of the blades 14 may be changed, so as to achieve alternate mistuning without changing the airfoil geometry.

[0024] As shown in FIG. 3-4, each damper 30 may be installed in an under- platform pocket 50, such that a radially outer surface 30b of the damper 30 contacts the radially inner surface 24a of the platform 24. The damper 30 may be secured in position by axially spaced tabs 52a, 52b provided on the under-side of the platform 24. To this end, as shown in FIG. 5, the radially inner surface 30a of the damper may be provided with axially spaced projections 34a, 34b that engage in corresponding recesses (not shown) in the tabs 52a, 52b. The location of tabs 52a and 52b may provide a "fool proofing" installation safeguard. In an exemplary embodiment, the damper length for the second set of dampers L may be reduced by reducing a free length tei, and/or t_e2 between a projection 34a, 34b and the nearest respective axial end 36a, 36b of the damper 30. In this way, an alternate frequency mistuning may be implemented by minimum modification to the existing damper geometry, for example, as part of a service upgrade. In FIG. 3-4, only one platform 24 is shown. However, it is to be understood that in operation, the radially outer surface 30b of the damper 30 frictionally contacts the radially inner surfaces 24a of the adjacent platforms 24.

[0025] In another embodiment, the dampers 30 of the first set H and the second set L may be positioned at different axial locations in relation to the respective platforms 24. In this case, it may be possible that the dampers 30 of both sets H, L may have the same (reduced) axial length, whereby a frequency mistuning is achieved simply based on the differences in contact loading between the first and second set of dampers.

[0026] It has been recognized that the contact loading of a damper, during operation, is a function of the cross-sectional shape of the damper along the area of contact with the platform. FIG. 5 illustrates a damper 30 with a generally rectangular or trapezoidal cross-sectional shape along the section C-C perpendicular to the axial length of the damper 30. The contact loading may be modified by providing different cross-sectional damper geometries, which may include for example, a circular shape (see FIG. 5), a wedge shape (see FIG. 7), or an asymmetrical shape (see FIG. 8), among many other shapes. In one embodiment, alternate frequency mistuning may be implemented by having different damper cross-sectional shapes between the first and second sets H, L. In this case, for example, the cross-sectional shape of each of the dampers 30 in the first set H and the second set L may be uniform across the entire axial length of the respective damper 30. In an alternate embodiment, the cross- sectional shape of each of the dampers 30 in the first set H, or the second set L, or both, may vary across an axial length of the respective damper 30, such that the variation of the cross-sectional shape along an axial direction is unique to the particular set H, L. By using different uniform cross-sectional shapes, or axial variation of cross-sectional shapes, it is possible to employ dampers of the same axial length in both the first set H and the second set L. Having the same length damper allows for a uniform under-platform geometry for the entire row of blades, as well as straight forward installations. In this case, a frequency mistuning may be achieved on the basis of the differences in contact loading between the first and second sets H, L.

[0027] Aspects of the present invention may be incorporated in a service upgrade method, whereby an intentional alternate mistuning may be introduced in an existing row of blades, to improve flutter resistance of the blades. This may be achieved by modifying the geometry of at least a subset of the existing dampers, or by providing replacement dampers, such that one or more of the inventive concepts described above are realized.

[0028] As an example, to effectively stabilize flutter, the under-platform damper geometries may be modified to achieve a mistuning of about 1.5 - 2 % above manufacturing tolerances. FIG. 9 graphically illustrates alternate mistuning in a row of 40 turbine blades. Herein, the odd number blades have a frequency of 250Hz, while the even numbered blades have a frequency of 255 Hz. In this example, the difference in blade frequencies is 5 Hz. Consequently, the frequency of even numbered blades is 2% than the frequency of odd numbered blades, i.e., the amount of mistuning is 2%.

[0029] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.

Claims

1. A bladed rotor system (10) for a turbomachine, comprising:

a circumferential row of blades (14) mounted on a rotor disc (12), each blade (14) comprising:

a platform (24);

a root (18) extending radially inward from the platform (24) for mounting the blade (14) to the rotor disc (12); and

an airfoil (16) extending span-wise radially outward from the platform

(24);

wherein platforms (24) of adjacent blades align circumferentially to define an inner diameter boundary for a working fluid flow path; and

a plurality of dampers (30), each damper (30) being located between adjacent platforms (24) and frictionally contacting a radially inner surface (24a) of said adjacent platforms (24);

wherein the plurality of dampers (30) comprise a first set (H) of dampers (30) and a second set (L) of dampers (30), wherein the dampers (30) of the first set (H) are distinguished from the dampers (30) of the second set (L) by a geometry of the damper (30) that is unique to the respective set (H, L), and

wherein dampers (30) of the first set (H) and the second set (L) are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades (14).

2. The bladed rotor system (10) according to claim 1, wherein said circumferential row is a row of free-standing blades (14).

3. The bladed rotor system (10) according to claim 1, wherein all the airfoils (16) in the circumferential row of blades (14) have substantially identical cross-sectional geometry about a rotation axis (22).

4. The bladed rotor system (10) according to claim 1, wherein the dampers (30) of the first set (H) and the second (L) set have different axial lengths (t, t - Δΐ), which is unique to the particular set (H, L).

5. The bladed rotor system (10) according to claim 1, wherein the dampers (30) of the first set (H) and the second set (L) are positioned at different axial locations in relation to the respective platforms (24), which is unique to the particular set (H, L).

6. The bladed rotor system (10) according to claim 1, wherein the dampers (30) of the first set (H) and the second set (L) have different cross-sectional shapes, which is unique to the particular set (H, L).

7. The bladed rotor system (10) according to claim 6, wherein the cross- sectional shape of each of the dampers (30) in the first (H) and second (L) sets is uniform across the entire axial length of the respective damper (30).

8. The bladed rotor system (10) according to claim 1, wherein a cross- sectional shape of each of the dampers (30) in the first (H) and/or the second (L) sets varies across an axial length of the respective damper (30), such that the variation of the cross-sectional shape along an axial direction is unique to the particular set (H, L).

9. The bladed rotor system (10) according to any of claims 6 to 8, wherein the dampers (30) in the first (H) and the second (L) sets have equal axial lengths.

10. A method for servicing a bladed rotor system (10):

wherein the bladed rotor system (10) comprises :

a circumferential row of blades (14) mounted on a rotor disc (12), each blade (14) comprising a platform (24), a root (18) extending radially inward from the platform (24) for mounting the blade (14) to the rotor disc (12), and an airfoil (16) extending span-wise radially outward from the platform (24); and

a plurality of dampers (30), each damper (30) being installed between adjacent platforms (24) and frictionally contacting a radially inner surface (24a) of said adjacent platforms (24);

wherein the method comprises:

modifying a geometry of at least a subset of the plurality of installed dampers (30) or providing replacement dampers (30) for at least a subset of the plurality of installed dampers (30),

so as to resultantly obtain a first set (H) of dampers (30) and a second set (L) of dampers (30), wherein the dampers (30) of the first set (H) are distinguished from the dampers (30) of the second set (L) by a geometry of the damper (30) that is unique to the respective set (H, L), and

installing the modified or replacement dampers (30), such that dampers (30) of the first set (H) and the second set (L) are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades (14).