U.S. patent application number 09/844207 was filed with the patent office on 2002-10-31 for methods and apparatus for damping rotor assembly vibrations.
Invention is credited to Crall, David William.
Application Number | 20020159882 09/844207 |
Document ID | / |
Family ID | 25292115 |
Filed Date | 2002-10-31 |
United States Patent
Application |
20020159882 |
Kind Code |
A1 |
Crall, David William |
October 31, 2002 |
METHODS AND APPARATUS FOR DAMPING ROTOR ASSEMBLY VIBRATIONS
Abstract
A multi-stage rotor assembly for a gas turbine engine including
a damper system that facilitates damping vibrations induced to the
rotor assembly is described. The rotor assembly includes a blisk
rotor including a plurality of rotor blades and a radially outer
rim. The rotor blades are integrally formed with the rim and extend
radially outward from the rim. The damper system is attached to
rotor blades within at least one stage of the rotor assembly, and
includes at least one layer of damping material and a cover sheet.
The cover sheet is attached to the rotor blade with adhesive.
Inventors: |
Crall, David William;
(Loveland, OH) |
Correspondence
Address: |
JOHN S. BEULICK
C/O ARMSTRONG TEASDALE LLP
ONE METROPOLITAN SQUARE
SUITE 2600
ST. LOUIS
MO
63102-2740
US
|
Family ID: |
25292115 |
Appl. No.: |
09/844207 |
Filed: |
April 27, 2001 |
Current U.S.
Class: |
415/119 ;
416/229A; 416/500 |
Current CPC
Class: |
F01D 5/16 20130101; Y10S
416/50 20130101 |
Class at
Publication: |
415/119 ;
416/229.00A; 416/500 |
International
Class: |
F01D 005/16 |
Goverment Interests
[0001] The government has rights in this invention pursuant to
Contract No. F33615-96-C-2657 awarded by the Department of the Air
Force.
Claims
What is claimed is:
1. A method of fabricating a rotor assembly for a gas turbine
engine to facilitate damping vibrations induced to the rotor
assembly, the rotor assembly including a radially outer rim and a
plurality of rotor blades that extend radially outward from the
radially outer rim, each of the rotor blades including an airfoil
including a pair of opposing sidewalls, said method comprising the
steps of: forming a cavity in each rotor blade airfoil that extends
radially inward from the airfoil first sidewall towards the airfoil
second sidewall; embedding a first layer of damping material within
the airfoil cavity adjacent the airfoil; and attaching a
constraining layer to the airfoil with adhesive, such that the
constraining layer is adjacent the first layer of damping material;
and attaching a cover sheet to the airfoil with adhesive, such that
the cover sheet extends around a periphery of the airfoil cavity in
sealing contact with the airfoil.
2. A method in accordance with claim 1 wherein said step of forming
a cavity in each rotor blade airfoil further comprises the step of
machining a cavity into each rotor blade airfoil.
3. A method in accordance with claim 1 further comprising the step
of embedding a second layer of damping material within the airfoil
cavity such that the constraining layer is between the first and
second layers of damping material.
4. A method in accordance with claim 1 wherein said step of
embedding a first layer of damping material further comprises the
step of embedding a first layer of visco-elastic material within
the airfoil cavity adjacent the airfoil.
5. A method in accordance with claim 1 wherein said step of
attaching a cover sheet to the airfoil further comprises the step
of attaching a cover sheet fabricated from titanium to the airfoil
with adhesive.
6. A rotor assembly for a gas turbine engine, said rotor assembly
comprising a rotor comprising a radially outer rim and a plurality
of rotor blades extending radially outward from said radially outer
rim, each said rotor blade comprising an airfoil and a damper
system comprising at least one layer of damping material and a
cover sheet, said cover sheet attached to said rotor blade airfoil
with adhesive.
7. A rotor assembly in accordance with claim 6 wherein each said
rotor blade airfoil comprises a first sidewall and a second
sidewall, and a cavity therebetween, said cavity extending
partially from said first sidewall towards said second sidewall,
said damping system cover sheet having an outer perimeter larger
than an outer perimeter of said sidewall cavity.
8. A rotor assembly in accordance with claim 7 wherein said damping
system cover sheet configured affix to said airfoil such that said
sidewall cavity is sealed.
9. A rotor assembly in accordance with claim 7 wherein said damping
material secured within said cavity by said cover sheet.
10. A rotor assembly in accordance with claim 6 wherein said damper
system further comprises a constraining layer affixed to said
airfoil with adhesive.
11. A rotor assembly in accordance with claim 6 wherein said
damping material comprises visco-elastic material, said damping
system comprises at least one constraining layer.
12. A rotor assembly in accordance with claim 11 wherein said
constraining layer between adjacent damping material layers.
13. A gas turbine engine comprising a rotor assembly comprising a
rotor comprising a radially outer rim and a plurality of rotor
blades extending radially outward from said radially outer rim,
each said rotor blade comprising an airfoil and a damper system
comprising at least one layer of damping material and a cover
sheet, said cover sheet attached to said rotor blade airfoil with
adhesive such that said damping material between said airfoil and
said damper system cover sheet, said damper system configured to
damp vibrations induced to said rotor blades.
14. A gas turbine engine in accordance with claim 13 wherein each
said rotor assembly rotor blade airfoil comprises a first sidewall,
a second sidewall, and a cavity extending radially inward from an
exterior surface of said first sidewall, such that said cavity
between said airfoil first and second sidewalls, said damper system
damping material within said cavity.
15. A gas turbine engine in accordance with claim 14 wherein said
rotor assembly damper system cover sheet affixed to said rotor
assembly rotor blade airfoil with adhesive.
16. A gas turbine engine in accordance with claim 14 wherein said
rotor assembly damper system further comprises at least one
constraining layer affixed to said rotor assembly rotor blade
airfoil with adhesive.
17. A gas turbine engine in accordance with claim 16 wherein said
rotor assembly damper system constraining layer within said airfoil
cavity between said damping material and said cover sheet.
18. A gas turbine engine in accordance with claim 16 wherein said
rotor assembly damper system constraining layer within said airfoil
cavity between a first layer of said damping material and a second
layer of said damping material.
19. A gas turbine engine in accordance with claim 14 wherein said
rotor assembly damper system damping material comprises
visco-elastic material.
20. A gas turbine engine in accordance with claim 14 wherein said
rotor assembly damper system cover sheet affixed to said airfoil in
sealing contact around said airfoil cavity.
Description
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to rotor assemblies and,
more particularly, to damper systems for damping vibrations induced
to the rotor assemblies.
[0003] A gas turbine engine typically includes at least one rotor
including a plurality of rotor blades that extend radially
outwardly from a common annular rim. Specifically, in blisk rotors,
the rotor blades are formed integrally with the annular rim rather
than attached to the rim with dovetail joints. An outer surface of
the rim typically defines a radially inner flowpath surface for air
flowing through the rotor assembly.
[0004] Centrifugal forces generated by the rotating blades are
carried by portions of the rims below the rotor blades. The
centrifugal forces generate circumferential rim stress
concentration between the rim and the blades that may be induced
through the blades. Additionally, within blisk rotors, because of
an absence of friction damping created when dovetails and shrouds
contact each other during operation, vibrational stresses may be
induced to the rotor assembly.
[0005] To facilitate vibration damping, rotor assemblies may
include dampers. At least some known rotor assemblies include
sleeve dampers positioned beneath the rim to damp airfoil modes.
The sleeve dampers provide damping to airfoil modes that have
significant rim participation.
[0006] At least some other known rotor assemblies include rotor
blades including pockets formed within the blades. A layer of
damper material is embedded in the pocket and covered with a
titanium constraining layer. The pocket is covered with a titanium
cover that is welded to the rotor blade. During operation, forces
induced within the rotor blade may cause the constraining layer to
separate from the damper material and forcibly contact the cover.
Over time, continued contact between the constraining layer and the
cover sheet may cause the cover sheet to separate from the rotor
blade.
BRIEF SUMMARY OF THE INVENTION
[0007] In an exemplary embodiment, a multi-stage rotor assembly for
a gas turbine engine includes a damper system for facilitating
damping vibrations induced to the rotor assembly. More
specifically, the rotor assembly includes a blisk rotor including a
plurality of rotor blades and a radially outer rim. The rotor
blades are integrally formed with the outer rim and extend radially
outward from the rim. The damper system is attached to the rotor
blades forming at least one stage of the rotor assembly, and
includes at least one layer of damping material and a cover sheet.
The cover sheet is attached to the rotor blade with adhesive to
secure the damping material against the rotor blade.
[0008] During operation, as the rotor assembly rotates, the
adhesive placed between the cover sheets and the rotor blades
carries centrifugal loads induced through the rotor blades.
Vibration damping is facilitated by the damper system. More
specifically, as the rotor assembly rotates, shear strains induced
into the damper material facilitate vibration damping. As a result,
the damper assembly facilitates damping vibrations induced to the
rotor assembly in a reliable and cost-effective manner.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] FIG. 1 is a schematic illustration of a gas turbine
engine;
[0010] FIG. 2 is a partial cross-sectional view of a rotor assembly
including a damper system and that may be used with the gas turbine
engine shown in FIG. 1;
[0011] FIG. 3 is an enlarged front view of a portion of the damper
system shown in FIG. 2; and
[0012] FIG. 4 is a side view of the damper system shown in FIG.
3.
DETAILED DESCRIPTION OF THE INVENTION
[0013] FIG. 1 is a schematic illustration of a gas turbine engine
10 including a low pressure compressor 12, a high pressure
compressor 14, and a combustor 16. Engine 10 also includes a high
pressure turbine 18 and a low pressure turbine 20. Compressor 12
and turbine 20 are coupled by a first shaft 21, and compressor 14
and turbine 18 are coupled by a second shaft 22. In one embodiment,
gas turbine engine 10 is an F 110 engine commercially available
from General Electric Aircraft Engines, Cincinnati, Ohio.
[0014] In operation, air flows through low pressure compressor 12
and compressed air is supplied from low pressure compressor 12 to
high pressure compressor 14. The highly compressed air is delivered
to combustor 16. Airflow from combustor 16 drives turbines 18 and
20 and exits gas turbine engine 10 through a nozzle 24.
[0015] FIG. 2 is a partial cross-sectional view of a rotor assembly
40 that may be used with gas turbine engine 10. Rotor assembly 40
includes a plurality of rotors 44 joined together by couplings 46
co-axially about an axial centerline axis 47. Each rotor 44 is
formed by one or more blisks 48, and each blisk 48 includes an
annular radially outer rim 50, a radially inner hub 52, and an
integral web 54 extending radially therebetween. Each blisk 48 also
includes a plurality of blades 56 extending radially outwardly from
outer rim 50. Blades 56, in the embodiment illustrated in FIG. 2,
are integrally joined with respective rims 50. Alternatively, and
for at least one stage, each rotor blade 56 may be removably joined
to rims 50 in a known manner using blade dovetails (not shown)
which mount in complementary slots (not shown) in a respective rim
50.
[0016] Rotor blades 56 are configured for cooperating with a motive
or working fluid, such as air. In the exemplary embodiment
illustrated in FIG. 2, rotor assembly 40 is a compressor of gas
turbine engine 10, with rotor blades 56 configured for suitably
compressing the motive fluid air in succeeding stages. Outer
surfaces 58 of rotor rims 50 define a radially inner flowpath
surface of the compressor as air is compressed from stage to
stage.
[0017] Blades 56 rotate about the axial centerline axis up to a
specific maximum design rotational speed, and generate centrifugal
loads in rotating components. Centrifugal forces generated by
rotating blades 56 are carried by portions of rims 50 directly
below each rotor blade 56. Rotation of rotor assembly 40 and blades
56 imparts energy into the air which is initially accelerated and
then decelerated by diffusion for recovering energy to pressurize
or compress the air. The radially inner flowpath is bound
circumferentially by adjacent rotor blades 56 and is bound radially
with a shroud (not shown).
[0018] Rotor blades 56 each include a leading edge 60, a trailing
edge 62, and an airfoil 64 extending therebetween. Airfoil 64
includes a suction side 76 and a circumferentially opposite
pressure side 78. Suction and pressure sides 76 and 78,
respectively, extend between axially spaced apart leading and
trailing edges 60 and 62, respectively and extend in radial span
between a rotor blade tip 80 and a rotor blade root 82. A blade
chord 84 is measured between rotor blade trailing and leading edges
62 and 60, respectively.
[0019] Each airfoil 64 also includes a damper system 90. In the
exemplary embodiment, only first stage rotors 44 include damper
system 90. In another embodiment, additional stages of rotors 44
extending through rotor assembly 40 include damper system 90.
During operation, as described in more detail below, damper system
90 damps airfoil modes within rotor assembly 40 to facilitate
damping vibration induced to rotor assembly 40.
[0020] FIG. 3 is an enlarged front view of rotor blade airfoil 64
including damper system 90. FIG. 4 is a side view of airfoil 64 and
damper system 90. Airfoil 64 includes a pocket cavity 100 extending
from an external surface 102 of airfoil body suction side 76
towards airfoil body pressure side 78. In one embodiment, cavity
100 is machined into airfoil 64. More specifically, cavity 100
extends a distance 104 radially inward from airfoil external
surface 102. Cavity depth 104 is less than a thickness (not shown)
of airfoil 64 measured between airfoil suction side 76 and airfoil
pressure side 78.
[0021] Cavity 100 has a width 110 measured from a leading edge 112
to a trailing edge 114. Cavity width 110 is smaller than airfoil
blade chord 84 such that cavity leading and trailing edges 112 and
114, respectively, are each a respective distance 116 and 118 from
airfoil leading and trailing edges 60 and 62. In addition, cavity
100 has a height 120 extending from a bottom edge 122 to a top edge
124 that is less than the radial span of airfoil 64. In the
exemplary embodiment, cavity 100 has a substantially rectangular
shape including rounded corners 126. Alternatively, cavity 100 is
non-rectangular shaped. Cavity leading and trailing edges 112 and
114, respectively, connect with cavity bottom and top edges 122 and
124, respectively, with corners 126, and define an outer periphery
128 of cavity 100.
[0022] Damper system 90 includes a plurality of damper material
layers 130, a constraining layer 132, and a cover sheet 134. In one
embodiment, damping material layers 130 are fabricated from a
visco-elastic material (VEM). A first damper material layer 136 is
embedded into cavity 100 against a back wall 138 of cavity 100.
More specifically, damper material layer 136 is embedded against
cavity back wall 138 a distance 139 from cavity bottom edge 122.
Adhesive material 140 extends between damper material layer 136 and
cavity bottom edge 122.
[0023] Constraining layer 132 is inserted within cavity 100 against
damper material layer 136. In one embodiment, constraining layer
132 is fabricated from titanium. More specifically, constraining
layer 132 extends between cavity top and bottom edges 124 and 122,
respectively, and is held in position against damper material layer
136 with adhesive material 140. In one embodiment, adhesive
material 140 is AF191 commercially available from 3M Bonding
Systems, St. Paul, Minn. 55144. In another embodiment, damper
system 90 includes a plurality of constraining layers 132 stacked
adjacent to each other and held together with adhesive material
140.
[0024] A second damper material layer 144 is embedded into cavity
100 against constraining layer 132. Second damper material layer
144 extends between cavity top and bottom edges 124 and 122,
respectively. Accordingly, constraining layer 132 extends between
damper material layers 130.
[0025] Damper system cover sheet 134 has a width 150 that is wider
than cavity width 110, and is narrower than airfoil blade chord 84
(shown in FIG. 2). In one embodiment, damper system cover sheet 134
is fabricated from titanium. Damper system cover sheet 134 also has
a height 152 that is taller than cavity height 120, and is shorter
than the radial span of airfoil 64. In the exemplary embodiment,
damper system cover sheet 134 has a substantially rectangular
profile and includes rounded lower corners 154. In an alternative
embodiment, damper system cover sheet 134 has a non-rectangular
profile.
[0026] Damper system cover sheet 134 is attached in sealing contact
to rotor blade airfoil 64 with adhesive material 140 extending
around cavity periphery 128. More specifically, damper system cover
sheet 134 is positioned relative to airfoil cavity 100 such that a
distance 160 between a bottom edge 162 of cover sheet 134 and
cavity bottom edge 122 is larger than a distance 164 between a top
edge 166 of cover sheet 134 and cavity top edge 124. Furthermore,
cover sheet 134 is positioned relative to airfoil cavity 100 such
that a distance 170 between each side edge 172 of cover sheet 134
and each respective cavity leading and trailing edge 112 and 114,
is approximately equal, and less than cover sheet distance 160. In
one embodiment, distance 162 is approximately twice as long as
distance 164. Because damper system cover sheet 134 is affixed in
sealing contact to airfoil 64, cover sheet 134 shields damper
material layers 130 from exposure to hot combustion gases flowing
through rotor assembly 40.
[0027] Adhesive material 140 extends between each respective cavity
edge 112, 114, 122, and 124, and each respective cover sheet edge
172, 172, 162, and 166. Accordingly, more adhesive material 140
extends between cavity bottom edge 122 and cover sheet bottom edge
162 than between any other cavity edge 112, 114, and 124, and a
respective cover sheet edge 172, 172, and 166.
[0028] During operation, as rotor assembly 40 rotates, vibration
damping is facilitated by damper material layers 130. More
specifically, vibration damping is facilitated by shear strains
induced within first damper material layer 136 between airfoil 64
and constraining layer 132, and within second damper material layer
144 between constraining layer 132 and cover sheet 134. Adhesive
material 140 placed between cavity bottom edge 122 and cover sheet
bottom edge 162 facilitates carrying centrifugal force loading
induced into airfoil 64, but does not prohibit first damper
material layer 136 from straining during chord-wise bending
vibration.
[0029] Additionally, during operation, damper system cover sheet
134 prevents constraining layer 132 from separating from damper
material layers 130. Further more, because damper system cover
sheet 134 is affixed to airfoil 64 with adhesive material 140,
during rotation of rotor assembly 40, cover sheet 134 induces shear
strains into second damper material layer 144 to facilitate
vibration damping within damper system 90.
[0030] The above-described rotor assembly is cost-effective and
highly reliable. The rotor assembly includes a damper system that
facilitates damping vibrations induced to each rotor blade. More
specifically, the damper system includes at least one layer of
damping material, a constraining layer, and a cover sheet. The
constraining layer is affixed within the airfoil cavity with
adhesive. The cover sheet is also affixed to the airfoil with
adhesive extending around the cavity periphery, such that the cover
sheet is in sealing contact with the airfoil. During operation, the
adhesive material carries the centrifugal force loading induced to
the rotor blade, while shear strains generated within the damping
material damp vibrations. As a result, the damper system
facilitates damping vibrational forces induced to the rotor
assembly.
[0031] While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that
the invention can be practiced with modification within the spirit
and scope of the claims.
* * * * *