US20140314578A1 - Securing segment for the vibration damping of turbine blades and rotor device - Google Patents

Securing segment for the vibration damping of turbine blades and rotor device Download PDF

Info

Publication number
US20140314578A1
US20140314578A1 US14/343,460 US201214343460A US2014314578A1 US 20140314578 A1 US20140314578 A1 US 20140314578A1 US 201214343460 A US201214343460 A US 201214343460A US 2014314578 A1 US2014314578 A1 US 2014314578A1
Authority
US
United States
Prior art keywords
retaining segment
rotor
region
receiving region
accordance
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
US14/343,460
Other languages
English (en)
Inventor
Jianmin Xu
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: XU, JIANMIN
Publication of US20140314578A1 publication Critical patent/US20140314578A1/en
Abandoned legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/10—Anti- vibration means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/32—Locking, e.g. by final locking blades or keys
    • F01D5/323—Locking of axial insertion type blades by means of a key or the like parallel to the axis of the rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • F01D5/3015—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type with side plates
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention relates to a locking segment for axially securing at least one rotor blade on a disk wheel of a rotor device of a jet engine in accordance with the type defined in more detail in patent Claim 1 and to a rotor device in accordance with the type defined in more detail in patent Claim 12 .
  • Rotor devices of jet engines known from actual practice have disk wheels and rotor blades connected thereto, which are arranged with blade roots in locating rails of the disk wheel running in the axial direction inside the disk wheel and which are secured on either one side or both sides with locking rings formed from locking segments and extending in the circumferential direction of the disk wheel to counter against any unwanted loosening from the disk wheel in the axial direction.
  • the locking segments are each arranged next to one another in the circumferential direction of the rotor devices and are each in contact with one another at their lateral end faces.
  • the blade roots are designed in cross-section in at least some areas at least approximately fir tree-shaped or dovetail-shaped, with the fir tree-shaped or dovetail-shaped areas of the rotor blades being arranged in the locating rails of the disk wheel designed correspondingly thereto.
  • a platform and an airfoil adjoin each of the blade roots of the rotor blades, said platforms each being designed wider than the blade roots between a blade root and an airfoil of a rotor blade in the circumferential direction of the disk wheel.
  • a damping element designed with a roof-shaped cross-section, extending in the axial direction of the disk wheel over a certain length underneath the joint areas between the platforms, contacting the rotor blades in the area of the platforms and also referred to as a damper underneath the platform.
  • the damping of the damping elements results from frictional forces in the contact area between the damping elements and the platforms of the rotor blades.
  • the damping effect of the damping elements is all the higher with a larger quotient from a distance (perpendicular in the radial direction of the disk wheel) between a damping element and a blade root end and the entire length of a rotor blade.
  • the object underlying the present invention is therefore to provide a locking segment for a rotor device of a jet engine and a rotor device, by means of which jet engines can be operated at high speeds in the area of a rotor device, while ensuring the required service life.
  • a locking segment for axially securing at least one rotor blade on a disk wheel of a rotor device of a jet engine that is operatively connectable to at least one rotor blade and to a disk wheel
  • a projecting region is provided in a first peripheral area of the locking segment that is laterally disposed in the circumferential direction in the installed state, and that in a second lateral peripheral area facing away from the first peripheral area, a receiving region is provided, with the projecting region of the locking segment in the installed state being designed for at least partially overlapping a receiving region of a locking segment that is identical in design and arranged adjacently in the circumferential direction, and for performing frictional work with said receiving region.
  • the locking segment in accordance with the invention is, in the installed position, in operative connection with a rotor blade and a disk wheel and is at the same time in frictional engagement with adjacent locking segments, vibrations and oscillations occurring during operation of a rotor device via the locking segment in the area of a rotor blade can be damped in simple manner.
  • the vibrations and oscillations of the rotor blades are here damped in that adjacent locking segments move relative to one another in their common overlapping region due to the vibrations and oscillations of the rotor blades interacting therewith, and the frictional force present between the locking segments counteracts the movement of the locking segments and hence of the rotor blades.
  • This is not possible with locking segments known from actual practice, since the considerably smaller end faces of the locking segments do not form any large-area contact surfaces that cause friction.
  • both the projecting region and the receiving region each have at least one contact surface which, in the installed state of the locking segment, interacts with a contact surface of a projecting region or a receiving region of a locking segment adjacently arranged in the circumferential direction, where a normal of the contact surface of the receiving region and a normal of the contact surface of the projecting region are directed in substantially opposite directions.
  • the normal of the contact surface of the receiving region and the normal of the contact surface of the projecting region in the installed state point substantially in the axial direction of the disk wheel.
  • the normals can, depending on the respective application, also be designed inclined relative to the axial alignment of the disk wheel or form an acute angle, preferably an angle in the single-digit range.
  • the contact surfaces of the projecting region and/or of the receiving region extend, in an advantageous embodiment of a locking segment in accordance with the invention, in the radial direction over an overall width of the locking segment.
  • both the projecting region and the receiving region are designed as a step with a material thickness reduced in the axial direction in the installed state compared to an area in the middle of the locking segment in the circumferential direction.
  • the shapes of the steps in the projecting region and in the receiving region are designed correlating to one another such that in the installed state they interact in a simple way that is favourable in terms of installation space.
  • a material thickness of the locking segments can, starting from the steps of the projecting region and of the receiving region, increase in particular in linear manner in a direction facing away from the contact surfaces, so that the material thickness of the locking segment decreases in particular in linear manner in the direction of the steps, starting from a material thickness applying in a middle area of the locking segments and adapted to a width of the grooves of the disk wheel and of the rotor blades.
  • a locking segment in accordance with the invention withstands stresses in an installed state particularly reliably when a material thickness of the projecting region and of the receiving region in the axial direction is at least approximately half of a material thickness of the locking segment in an area which is in the middle in the circumferential direction, so that a stress arising in the overlapping region in the installed state of the locking segment can be evenly distributed to the respective adjacent locking segments.
  • the total of a wall thickness of the projecting region and of a wall thickness of the receiving region is in particular equal to a wall thickness in a middle area of the locking segment, so that the locking segments can be installed securely into the grooves of the rotor blades and of the disk wheel.
  • both the projecting region and the receiving region have a material thickness which is at least approximately comparable to a material thickness of an area of the locking segment which is in the middle in the circumferential direction.
  • the material thickness here substantially corresponds to half of a width of the grooves of a disk wheel or of rotor blades intended to receive the locking segments.
  • the locking segment is, in this embodiment in particular, designed using a material having a good bending resistance.
  • Two locking segments adjacent in the installed state can each be designed curved in the axial direction in an area adjoining the overlapping region.
  • two locking segments adjacent in the circumferential direction in the installed state can be offset overall to one another in the axial direction, while interacting in the overlapping region.
  • a contour facing outwards in the radial direction in the installed state of the locking segment has at least one contact area for making contact with at least one rotor blade, said contact area extending in the circumferential direction substantially only over part of the length of the locking segment.
  • the contour is here in particular designed such that in the installed state of the locking segment it interacts only with at least one area of the rotor blade which during operation receives oscillation movements with particular sensitivity.
  • the locking segment contour facing at least one rotor blade in the installed state of the locking segment has at least two contact areas which interact in particular with different rotor blades in the installed state of the locking segment.
  • a rotor device for a jet engine with a disk wheel and several rotor blades connected to said disk wheel is proposed, the rotor blades being arranged in each case via a blade root substantially in the axial direction inside recesses of the disk wheel.
  • Several locking segments according to the present invention are provided for axially securing the rotor blades in the recesses of the disk wheel, said locking segments interacting on the one hand with grooves of the rotor blades and on the other hand with at least one groove of the disk wheel, where a projecting region of a locking segment overlaps with a receiving region of a locking segment adjacent in the circumferential direction in order to perform frictional work during operation of the rotor device.
  • the rotor device in accordance with the invention designed with locking segments in accordance with the invention can, in the area of the rotor blades and also in areas of the disk wheel receiving the blade roots, advantageously have smaller dimensions due to the lower stresses even at higher speeds, and hence be designed with a lower dead weight and also manufactured less expensively, without detriment to the service life of known rotor devices, or with a longer service life or better operating performance in comparison with known rotor devices.
  • locking segments are provided for securing the rotor blades in the axial direction on both sides of the rotor blades. If a plurality of and in particular all locking segments are designed with a projecting region and a receiving region and if these overlap with respectively adjacent locking segments in the circumferential direction, a particularly high damping effect of oscillations and vibrations respectively, of the rotor blades occurring during operation can be achieved by the locking segments. Alternatively, locking segments can be arranged exclusively on a front or on a rear side of the rotor blades.
  • a damping device can be provided in an area between platforms of adjacent rotor blades on a side of the platforms facing in the direction of a rotary axis of the rotor device.
  • Both the locking segment in accordance with the invention and the rotor device in accordance with the invention can be used for a variety of engine designs and are applied in particular for any required stages of turbines. Furthermore, the locking segment in accordance with the invention, and the rotor device in accordance with the invention can for example also be used in a compressor or a fan of an engine.
  • FIG. 1 shows a highly schematized longitudinal sectional view of a jet engine featuring a turbine provided with several rotor devices
  • FIG. 2 shows a schematized detail of the sectional view of FIG. 1 , illustrating more clearly two stages of the rotor device, each having rotor blades secured by locking segments in the axial direction inside a disk wheel,
  • FIG. 3 shows a simplified view of three locking segments of a rotor device of FIG. 2 in stand-alone position, with the locking segments being designed in accordance with the present invention and arranged adjacently in the circumferential direction,
  • FIG. 4 shows a simplified representation of a detail of the locking segments of FIG. 3 , illustrating an overlapping region of two adjacent locking segments
  • FIG. 5 shows simplified sectional representations of details of locking segments, each corresponding to FIG. 4 , the locking segments being differently designed in the area of the overlapping region, and
  • FIG. 6 shows a simplified cross-sectional view of a rotor device of FIG. 2 with a damper underneath the platform.
  • FIG. 1 shows a jet engine 1 in a longitudinal sectional view, with the jet engine 1 being designed with a bypass duct 2 and an intake area 3 .
  • a fan 4 adjoins the intake area 3 on the downstream side in a manner known per se.
  • the fluid flow in the jet engine 1 splits into a bypass flow and a core flow, with the bypass flow flowing through the bypass duct 2 and the core flow into an engine core 5 which is in turn designed in a manner known per se with a compressor device 6 , a burner 7 and a turbine device 8 .
  • the turbine device 8 has in this case three rotor devices 9 , 10 , 11 designed in a substantially comparable manner, where the rotor device 9 and the rotor device 10 are illustrated in more detail in FIG. 2 .
  • the rotor device 9 representing a first stage of the turbine device 8 is designed with a centrally arranged disk wheel 13 connected to an engine axis 12 , on which a plurality of rotor blades 14 is arranged in radially outer areas on the circumferential side.
  • the rotor blades 14 each have a blade root 15 designed as a fir-tree root and shown here only in schematic form, via which the rotor blades 14 are arranged in a manner known per se inside recesses 16 of the disk wheel 13 which run substantially in the axial direction in the disk wheel 13 and correlate with the fir-tree roots 15 .
  • a locking ring 18 , 19 with several locking segments 17 is provided on a side of the rotor device 9 facing a flow in a jet engine turbine and on a side of the rotor device 9 facing away from the flow in the jet engine 1 respectively, or on both sides of the rotor blades 14 of the rotor device 9 .
  • the locking rings 18 , 19 are of substantially identical design, where the front locking ring 18 arranged on that side of the rotor blades 14 facing the intake area 3 is described in the following representatively for the rear locking ring 19 arranged on that side of the rotor blades 14 facing away from the intake area 3 .
  • the front locking ring 18 engages with its locking segments 17 in the area of the fir-tree roots 15 of the rotor blades 14 on the one hand in a groove 20 provided continuously in the disk wheel 13 and on the other hand in grooves 22 of the rotor blades 14 arranged underneath platforms 21 or an inner shroud of the rotor blades 14 and extending in the circumferential direction of the disk wheel.
  • the rear locking ring 19 engages with its locking segments 17 in the area of the fir-tree roots 15 of the rotor blades 14 on the one hand in a groove 20 A provided continuously in the disk wheel 13 and on the other hand in grooves 22 A of the rotor blades 14 arranged underneath platforms 21 of the rotor blades 14 and extending in the circumferential direction of the disk wheel.
  • FIG. 3 shows three locking segments 17 A to 17 C, adjacent in the circumferential direction, of the front locking ring 18 , which are all identical in design.
  • the locking segments 17 A to 17 C made using a metallic material have on their side facing the grooves 22 of the rotor blades 14 in the installed state not a concentric surface, but a contour 23 with, in the present case, three contact areas 24 , 25 , 26 , which have a larger radial extent than the areas between the contact areas 24 , 25 , 26 . Between each two contact areas 24 , 25 and 25 , 26 respectively, a radial extent of the locking segments 17 A to 17 C is reduced in the form of a rounded part.
  • the locking segments 17 A to 17 C each extend in the installed state over several rotor blades 14 , in the present case three, where in each case one contact area 24 , 25 , 26 or several contact areas 24 , 25 , 26 of the locking segments 17 A to 17 C interact(s) with a rotor blade 14 .
  • the contact areas 24 , 25 , 26 of the locking segments 17 A to 17 C are arranged here relative to the rotor blades 14 such that they each interact with areas of the rotor blades 14 , in which in each case a curve, varying in the circumferential direction of the rotor blades 14 , of an oscillation amplitude of the displacement components of the oscillations of the rotor blades 14 occurring in all directions during operation of the jet engine 1 is at least approximately at its maximum underneath the platform 21 .
  • the oscillations of the rotor blades 14 are then transmitted to a required extent to the locking segments 17 A to 17 C when the contact areas 24 , 25 , 26 are each arranged directly underneath those areas of the rotor blades 14 in which the maximum of the oscillation movements is underneath the platform 21 .
  • a first peripheral area, laterally disposed in the circumferential direction, of the locking segment 17 A shown in FIG. 4 is designed as a receiving region 27 .
  • a lateral peripheral area 28 facing the first peripheral area 27 is a projecting region.
  • the projecting region 28 of the locking segment 17 B encompasses the receiving region 27 of the locking segment 17 A in the installed state of the locking segments 17 A and 17 B, with the receiving region 27 and the projecting region 28 forming an overlapping region 29 .
  • the peripheral areas 27 and 28 of the locking segments 17 A and 17 B are arranged at least partially one behind the other in the axial direction.
  • the projecting region 28 can be arranged in the axial direction either in front of or behind the receiving region 27 .
  • the receiving region 27 of the left-hand locking segment 17 A has a contact surface 30 which in the installed state of the locking segments 17 A, 17 B interacts with a contact surface 31 of the right-hand locking segment 17 B.
  • the contact surfaces 30 , 31 are arranged substantially parallel to one another, where normals point to the contact surfaces 30 , 31 in directions opposite to one another.
  • the normals to the contact surfaces 30 , 31 face substantially in the axial direction, but can also be inclined relative to this direction in particular about an axis pointing in the radial direction and/or an axis pointing in the circumferential direction.
  • Both the contact surface 30 of the receiving region 27 and the contact surface 31 of the projecting region 28 are arranged in the area of a step 32 , 33 of the respective locking segment 17 A or 17 B, which each have in the axial direction about half of a material thickness of the locking segment 17 otherwise designed with a constant material thickness.
  • An end face 34 of the step 32 of the receiving region 27 of the left-hand locking segment 17 A forms a stop for a shoulder 35 of the step 33 of the projecting region 28 of the right-hand locking segment 17 B.
  • Both the receiving region 27 of the locking segment 17 A and the projecting region 28 of the locking segment 17 B extend in the radial direction over the entire locking segment 17 A and 17 B respectively, where an end face 36 of the step 33 of the projecting region 28 is in the present case curved in the radial direction and is designed least approximately semi-circular.
  • FIG. 5 a shows a section through the two locking segments 17 A and 17 B shown in FIG. 4 and arranged adjacently to one another in the installed state.
  • FIG. 5 b shows two alternatively designed locking segments 17 D, 17 E in which a material thickness decreases, in the present case in linear manner, starting from a thickness adapted to a groove width of the rotor blades 14 and of the disk wheel 13 , in the direction of the steps 32 , 33 with the contact surfaces 30 , 31 , to a thickness which corresponds to around half the maximum material thickness of the locking segments 17 D, 17 E.
  • FIG. 5 c A further alternative embodiment of the locking segments 17 F, 17 G is shown in FIG. 5 c .
  • the locking segments 17 F, 17 G have overall a thickness substantially corresponding to half a groove width of the rotor blades 14 or of the disk wheel 13 , where the locking segments 17 F, 17 G adjacent in the circumferential direction in the installed state together have in the overlapping region 29 a thickness corresponding to the groove width.
  • the locking segments 17 F, 17 G are each designed curved in the axial direction.
  • the locking ring 18 is installed substantially as in known solutions, with the locking ring 18 having pre-curved part sections to provide an annular closure for the entire rotor blade set in the known manner.
  • the locking segments 17 are inserted one after the other into the grooves 20 , 22 of the rotor blades 14 and of the disk wheel 13 in the circumferential direction, with adjacent locking segments 17 each interacting via a receiving region 27 and a projecting region 28 .
  • the contact areas 24 , 25 , 26 are placed with the rotor blades 14 at required positions underneath the blade platform 21 .
  • a locking plate known from actual practice is used to secure the position of the locking segments 17 in the circumferential direction.
  • a locking ring 38 designed comparable to the locking rings 18 , 19 with locking segments 43 designed substantially identical to the previously described locking segments 17 or 17 A to 17 G is provided for axial securing of rotor blades 37 in a disk wheel 39 of the rotor device 10 designed comparable to the rotor device 9 and representing a second stage of the turbine device 8 .
  • the rotor blades 37 are secured against a movement of the rotor blades 37 in the axial direction relative to the disk wheel 39 by the locking ring 38 arranged on a side of the rotor blades 37 facing away from the intake area 3 on the one hand and by a stop arranged on the other side of the rotor blades 37 on the other hand.
  • the design described above of the locking segments 17 and 43 respectively, is used during operation of the jet engine 1 to damp oscillations and/or vibrations of the rotor blades 14 , 37 occurring in particular in the radial direction, which are in particular incited thereto by stator vanes arranged upstream of the rotor blades 14 , 37 .
  • the radial components of maximum oscillation movements occurring during operation of the rotor blades 14 , 37 underneath the platform 21 are here received by the contact areas 24 , 25 , 26 of the locking segments 17 , 43 and transmitted via the contact surfaces 30 , 31 between the locking segments 17 , 43 .
  • Frictional work in the form of metal abrasion between the locking segments 17 and 43 respectively, is performed via the contact surfaces 30 , 31 of the respective receiving regions 27 and projecting regions 28 , with the locking segments 17 , 43 , as a result converting the kinetic energy of the blade oscillations by friction and generated heat and/or by metal wear on the contact surface 30 , 31 into deformation energy.
  • This energy conversion by the locking segments 17 , 43 counteracts the oscillations and vibrations of the rotor blades 14 , 37 , such that the latter are damped.
  • damping devices 41 are provided in the view, in a highly schematized representation in FIG. 6 , of a cross-section through the rotor device 9 of the jet engine 1 , and have damping elements substantially known per se, but which can if necessary be dimensioned smaller.
  • the damping devices 41 are located in intermediate spaces 40 arranged between the fir-tree roots 15 of the rotor blades 14 and the platforms 21 in the area of blade necks 42 , or in intermediate spaces 40 confined by the latter.
  • the locking segments 17 , 43 in accordance with the invention can however also have a damping effect against blade and disk oscillations so great that damping devices 41 can be dispensed with.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US14/343,460 2011-09-12 2012-09-12 Securing segment for the vibration damping of turbine blades and rotor device Abandoned US20140314578A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
DE102011082493.6 2011-09-12
DE102011082493A DE102011082493A1 (de) 2011-09-12 2011-09-12 Sicherungssegment zur Schwingungsdämpfung von Turbinenlauf-schaufeln und Rotorvorrichtung
PCT/EP2012/003815 WO2013037483A1 (fr) 2011-09-12 2012-09-12 Segment de blocage pour amortir les vibrations d'aubes de turbine et système rotor

Publications (1)

Publication Number Publication Date
US20140314578A1 true US20140314578A1 (en) 2014-10-23

Family

ID=47115711

Family Applications (1)

Application Number Title Priority Date Filing Date
US14/343,460 Abandoned US20140314578A1 (en) 2011-09-12 2012-09-12 Securing segment for the vibration damping of turbine blades and rotor device

Country Status (4)

Country Link
US (1) US20140314578A1 (fr)
EP (1) EP2756169A1 (fr)
DE (1) DE102011082493A1 (fr)
WO (1) WO2013037483A1 (fr)

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105386794A (zh) * 2015-12-10 2016-03-09 西安航天动力研究所 涡轮盘刚度自增强的涡轮结构
US20160130977A1 (en) * 2014-11-07 2016-05-12 United Technologies Corporation Turbine rotor segmented sideplates with anti-rotation
US20170321556A1 (en) * 2016-05-03 2017-11-09 Pratt & Whitney Canada Corp. Damper ring
US20170328224A1 (en) * 2016-05-13 2017-11-16 United Technologies Corporation Contoured retaining ring
WO2019040171A1 (fr) * 2017-08-25 2019-02-28 Siemens Aktiengesellschaft Ensemble plaque d'étanchéité
US10450865B2 (en) 2016-05-27 2019-10-22 Pratt & Whitney Canada Corp. Friction damper
US10502061B2 (en) 2016-09-28 2019-12-10 Pratt & Whitney Canada Corp. Damper groove with strain derivative amplifying pockets
CN110925030A (zh) * 2019-12-05 2020-03-27 中国航发四川燃气涡轮研究院 一种带嵌入式叶冠阻尼的低压涡轮模拟叶片
CN114382549A (zh) * 2020-10-21 2022-04-22 中国航发商用航空发动机有限责任公司 涡轮和航空发动机
US11391157B1 (en) 2021-03-23 2022-07-19 Pratt & Whitney Canada Corp. Damped rotor assembly
US11525464B2 (en) 2021-03-23 2022-12-13 Pratt & Whitney Canada Corp. Rotor with centrifugally wedged damper

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112855282B (zh) * 2021-03-01 2022-04-12 杭州汽轮机股份有限公司 一种工业汽轮机调节级锥销装配过盈量控制方法

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3137478A (en) * 1962-07-11 1964-06-16 Gen Electric Cover plate assembly for sealing spaces between turbine buckets
US3572966A (en) * 1969-01-17 1971-03-30 Westinghouse Electric Corp Seal plates for root cooled turbine rotor blades
US3807898A (en) * 1970-03-14 1974-04-30 Secr Defence Bladed rotor assemblies
US4344740A (en) * 1979-09-28 1982-08-17 United Technologies Corporation Rotor assembly
US5211407A (en) * 1992-04-30 1993-05-18 General Electric Company Compressor rotor cross shank leak seal for axial dovetails
US5257909A (en) * 1992-08-17 1993-11-02 General Electric Company Dovetail sealing device for axial dovetail rotor blades
US5338154A (en) * 1993-03-17 1994-08-16 General Electric Company Turbine disk interstage seal axial retaining ring
US20060257262A1 (en) * 2005-05-12 2006-11-16 Itzel Gary M Coated bucket damper pin
EP1944472A1 (fr) * 2007-01-09 2008-07-16 Siemens Aktiengesellschaft Partie axiale d'un rotor de turbine, élément d'étanchéité pour un rotor équipé d'aubes de rotor d'une turbine et rotor de turbine

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3112915A (en) * 1961-12-22 1963-12-03 Gen Electric Rotor assembly air baffle
FR2524933B1 (fr) * 1982-04-13 1987-02-20 Snecma Dispositif de verrouillage axial d'aubes de rotor de turbine ou de compresseur
JP2000186502A (ja) * 1998-12-24 2000-07-04 Hitachi Ltd ガスタービン
US7264448B2 (en) * 2004-10-06 2007-09-04 Siemens Power Corporation, Inc. Remotely accessible locking system for turbine blades
US8128371B2 (en) * 2007-02-15 2012-03-06 General Electric Company Method and apparatus to facilitate increasing turbine rotor efficiency

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3137478A (en) * 1962-07-11 1964-06-16 Gen Electric Cover plate assembly for sealing spaces between turbine buckets
US3572966A (en) * 1969-01-17 1971-03-30 Westinghouse Electric Corp Seal plates for root cooled turbine rotor blades
US3807898A (en) * 1970-03-14 1974-04-30 Secr Defence Bladed rotor assemblies
US4344740A (en) * 1979-09-28 1982-08-17 United Technologies Corporation Rotor assembly
US5211407A (en) * 1992-04-30 1993-05-18 General Electric Company Compressor rotor cross shank leak seal for axial dovetails
US5257909A (en) * 1992-08-17 1993-11-02 General Electric Company Dovetail sealing device for axial dovetail rotor blades
US5338154A (en) * 1993-03-17 1994-08-16 General Electric Company Turbine disk interstage seal axial retaining ring
US20060257262A1 (en) * 2005-05-12 2006-11-16 Itzel Gary M Coated bucket damper pin
EP1944472A1 (fr) * 2007-01-09 2008-07-16 Siemens Aktiengesellschaft Partie axiale d'un rotor de turbine, élément d'étanchéité pour un rotor équipé d'aubes de rotor d'une turbine et rotor de turbine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
EP 1944472, (07-2008), Engle et al. Foreign Copy and Translation Copy *

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10408087B2 (en) * 2014-11-07 2019-09-10 United Technologies Corporation Turbine rotor segmented sideplates with anti-rotation
US20160130977A1 (en) * 2014-11-07 2016-05-12 United Technologies Corporation Turbine rotor segmented sideplates with anti-rotation
CN105386794A (zh) * 2015-12-10 2016-03-09 西安航天动力研究所 涡轮盘刚度自增强的涡轮结构
US20170321556A1 (en) * 2016-05-03 2017-11-09 Pratt & Whitney Canada Corp. Damper ring
US10436032B2 (en) * 2016-05-03 2019-10-08 Pratt & Whitney Canada Corp. Damper ring
US20170328224A1 (en) * 2016-05-13 2017-11-16 United Technologies Corporation Contoured retaining ring
US10215037B2 (en) * 2016-05-13 2019-02-26 United Technologies Corporation Contoured retaining ring
US10450865B2 (en) 2016-05-27 2019-10-22 Pratt & Whitney Canada Corp. Friction damper
US10502061B2 (en) 2016-09-28 2019-12-10 Pratt & Whitney Canada Corp. Damper groove with strain derivative amplifying pockets
US11274556B2 (en) 2016-09-28 2022-03-15 Pratt & Whitney Canada Corp. Damper groove with strain derivative amplifying pockets
WO2019040171A1 (fr) * 2017-08-25 2019-02-28 Siemens Aktiengesellschaft Ensemble plaque d'étanchéité
CN110925030A (zh) * 2019-12-05 2020-03-27 中国航发四川燃气涡轮研究院 一种带嵌入式叶冠阻尼的低压涡轮模拟叶片
CN114382549A (zh) * 2020-10-21 2022-04-22 中国航发商用航空发动机有限责任公司 涡轮和航空发动机
US11391157B1 (en) 2021-03-23 2022-07-19 Pratt & Whitney Canada Corp. Damped rotor assembly
US11525464B2 (en) 2021-03-23 2022-12-13 Pratt & Whitney Canada Corp. Rotor with centrifugally wedged damper

Also Published As

Publication number Publication date
EP2756169A1 (fr) 2014-07-23
DE102011082493A1 (de) 2013-03-14
WO2013037483A1 (fr) 2013-03-21

Similar Documents

Publication Publication Date Title
US20140314578A1 (en) Securing segment for the vibration damping of turbine blades and rotor device
US8790086B2 (en) Turbine blade assembly for retaining sealing and dampening elements
JP7237441B2 (ja) タービン翼の先端シュラウドの冷却用シールレールのためのシステム
US10132169B2 (en) Shrouded turbine rotor blades
EP2204535B1 (fr) Contour de la plate-forme d'une aube de turbine
EP2149674B1 (fr) Rotor aubagé de turbine avec amortisseur de vibrations
EP3187689B1 (fr) Pales de rotor de turbine carénées
US8894368B2 (en) Device and method for aligning tip shrouds
KR102273496B1 (ko) 터빈 버킷 클로져 조립체 및 그 조립 방법
CN106499442B (zh) 用于涡轮叶片的阻尼器销
JP6862128B2 (ja) 隣接するタービンブレードの制振用ダンパピン
US8353672B2 (en) Turbine blade damper arrangement
US8807927B2 (en) Clearance flow control assembly having rail member
JP6557478B2 (ja) タービンバケット及びタービンバケットの先端シュラウドをバランスさせるための方法
RU2647170C2 (ru) Лопаточный аппарат и соответствующий способ изготовления лопаточного аппарата
US20140023506A1 (en) Damper system and a turbine
US20100166562A1 (en) Turbine blade root configurations
US8616838B2 (en) Systems and apparatus relating to compressor operation in turbine engines
CA2673079C (fr) Turbo machine, en particulier turbine a gaz
KR102261350B1 (ko) 터빈 노즐 고정 방법 및 시스템
US10247013B2 (en) Interior cooling configurations in turbine rotor blades
US8721289B2 (en) Flow balancing slot
US10066494B2 (en) Turbine with bucket fixing means
US9080455B2 (en) Rotor device for a jet engine with a disk wheel and several rotor blades
RU2526129C2 (ru) Рабочее колесо осевой турбомашины гтд

Legal Events

Date Code Title Description
AS Assignment

Owner name: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:XU, JIANMIN;REEL/FRAME:032375/0264

Effective date: 20140304

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION