U.S. patent application number 13/242297 was filed with the patent office on 2013-03-28 for airfoil air seal assembly.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is Charles P. Gendrich, Christopher R. McNeill, Bradley L. Pike, David J. Pitney. Invention is credited to Charles P. Gendrich, Christopher R. McNeill, Bradley L. Pike, David J. Pitney.
Application Number | 20130078084 13/242297 |
Document ID | / |
Family ID | 46880960 |
Filed Date | 2013-03-28 |
United States Patent
Application |
20130078084 |
Kind Code |
A1 |
Gendrich; Charles P. ; et
al. |
March 28, 2013 |
AIRFOIL AIR SEAL ASSEMBLY
Abstract
An air seal assembly for a gas turbine engine the air seal
comprises a first assembly and a second assembly. One of the first
assembly and the second assembly is rotatable relative to the other
of the first assembly and the second assembly. The second assembly
is aligned annularly with the first assembly and includes a
circumferential surface with an abradable coating disposed
annularly adjacent to the first and second airfoil tips. The first
assembly includes at least one first airfoil with a first tip
having an abrasive coating, and at least one second airfoil with a
second tip absent the abrasive coating, the at least one first
airfoil co-aligned axially and intermingled with a respective at
least one second airfoil around a periphery of the first
assembly.
Inventors: |
Gendrich; Charles P.;
(Middletown, CT) ; Pike; Bradley L.; (Burlington,
CT) ; McNeill; Christopher R.; (Plainville, CT)
; Pitney; David J.; (Moodus, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Gendrich; Charles P.
Pike; Bradley L.
McNeill; Christopher R.
Pitney; David J. |
Middletown
Burlington
Plainville
Moodus |
CT
CT
CT
CT |
US
US
US
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
46880960 |
Appl. No.: |
13/242297 |
Filed: |
September 23, 2011 |
Current U.S.
Class: |
415/173.1 ;
416/241R |
Current CPC
Class: |
F01D 5/20 20130101; F05D
2260/961 20130101; F01D 11/122 20130101; F01D 11/12 20130101 |
Class at
Publication: |
415/173.1 ;
416/241.R |
International
Class: |
F01D 11/08 20060101
F01D011/08; F01D 5/14 20060101 F01D005/14 |
Claims
1. An air seal assembly for a gas turbine engine the air seal
comprising: a first assembly including at least one first airfoil
with a first tip having an abrasive coating, and at least one
second airfoil with a second tip absent the abrasive coating, each
of the at least one first airfoil co-aligned axially and
intermingled with a respective at least one second airfoil around a
periphery of the first assembly; and a second assembly aligned
annularly with the first assembly, the second assembly including a
circumferential surface with an abradable coating disposed
annularly adjacent to the first and second airfoil tips, with one
of the first assembly and the second assembly being rotatable
relative to the other of the first assembly and the second
assembly.
2. The air seal assembly of claim 1, wherein the engine is operable
in at least a run-in mode and an operational mode, the run-in mode
including rotating the rotatable assembly intermittently to
centrifugally expand the rotatable assembly for abrasively forming
a circumferential seal groove into the abradable coating with the
at least one first airfoil tip, and the operational mode including
rotating the rotatable assembly continuously to centrifugally and
thermally expand the rotatable assembly forming an air seal.
3. The blade outer air seal assembly of claim 1, wherein the seal
assembly is installed into a compressor section of the gas turbine
engine.
4. The air seal assembly of claim 1, wherein the at least one
airfoil tip includes at least one squealer tip cut under the
abrasive coating.
5. The air seal assembly of claim 4, wherein the at least one
airfoil tip includes a first squealer tip cut proximate the
pressure surface and a second squealer tip cut proximate the
suction surface.
6. The air seal assembly of claim 1, wherein the abrasive coating
comprises cubic boron nitride (CBN) suspended in a matrix.
7. The air seal assembly of claim 1, wherein the abradable coating
comprises a boro-nitride ceramic.
8. The air seal assembly of claim 1, wherein the at least one first
airfoil and the at least one second airfoil each comprise a
titanium alloy.
9. The air seal assembly of claim 1, wherein the first assembly is
a rotor assembly and the second assembly is a stator assembly.
10. The air seal assembly of claim 9, wherein the at least one
first airfoil is intermingled with the respective at least one
second airfoil based at least in part on results of a Monte Carlo
simulation.
11. The air seal assembly of claim 9, wherein the first and second
airfoils are rotor blades integrally formed with a rotor disc.
12. The air seal assembly of claim 1, wherein the first assembly is
a stator assembly, and the second assembly is a rotor assembly.
13. The air seal assembly of claim 12, wherein the first and second
airfoils are cantilevered stator vanes.
14. A rotor assembly comprising: a rotor disc; at least one first
rotor blade including a first airfoil tip with an abrasive coating;
and at least one second rotor blade including a second airfoil tip
absent an abrasive coating, the at least one first rotor blade
intermingled and co-aligned with a respective at least one second
rotor blade axially around a periphery of the rotor disc.
15. The rotor assembly of claim 14, wherein the at least one first
rotor blade is intermingled with the respective at least one second
rotor blade for rotational balance of the rotor assembly, the
intermingling determined using a Monte Carlo simulation.
16. The rotor assembly of claim 14, wherein a ratio of the at least
one first rotor blade and the at least one second rotor blade is
between about 1:12 and about 1:1.
17. The rotor assembly of claim 16, wherein the ratio of the at
least one first rotor blade and the at least one second rotor blade
is between about 1:3 and about 1:9.
18. The rotor assembly of claim 17, wherein the ratio of the at
least one first rotor blade and the at least one second rotor blade
is about 1:6.
19. The rotor assembly of claim 14, wherein the at least one first
rotor blade and the at least one second rotor blade are integrally
formed with the rotor disc.
20. The rotor assembly of claim 14, wherein the first airfoil tip
includes at least one squealer tip cut under the abrasive coating.
Description
BACKGROUND
[0001] The invention relates generally to air seal assemblies for a
gas turbine engine, and more specifically to rotor assemblies for
air seal assemblies.
[0002] To maximize efficiency and minimize clearances, each
operative section of a gas turbine engine (fan, compressor, and
turbine) includes a variety of seals and coatings. Maintaining
appropriate clearances between moving parts and adjacent stationary
parts is critical to balancing efficiency and improving stability
to limit damage to the engine. Too small of a clearance results in
increased contact severity and frequency between components,
particularly due to maneuver loads, rapid temperature changes, and
other sudden changes in engine operation. Excessive clearance can
cause efficiency losses from lost work embodied in the compressed
gases escaping through gaps between respective rotor and stator
elements. Large clearances also increase the risk and severity of
operational instability such as compressor surge.
[0003] In one relatively simple example, every blade tip on a rotor
is abrasively coated to run a groove into an abradable coating on
the casing to form a seal. In another simple example, stator vanes
include an abradable seal coating on all of the vane tips or
shrouds that are rubbed by an abrasive coating on the rotor land.
However, coating every blade tip or every vane is quite expensive
and time-consuming. The coating process itself is subject to error
which can also result in additional repair or scrapping of the
entire component or assembly.
SUMMARY
[0004] An air seal assembly for a gas turbine engine the air seal
comprises a first assembly and a second assembly. One of the first
assembly and the second assembly is rotatable relative to the other
of the first assembly and the second assembly. The second assembly
is aligned annularly with the first assembly and includes a
circumferential surface with an abradable coating disposed
annularly adjacent to the first and second airfoil tips. The first
assembly includes at least one first airfoil with a first tip
having an abrasive coating, and at least one second airfoil with a
second tip absent the abrasive coating, the at least one first
airfoil co-aligned axially and intermingled with a respective at
least one second airfoil around a periphery of the first
assembly.
[0005] A rotor assembly comprises a rotor disc, at least one first
rotor blade and at least one second rotor blade. The at least one
first rotor blade includes a first airfoil tip with an abrasive
coating. The at least one second rotor blade includes a second
airfoil tip absent an abrasive coating. The at least one first
rotor blade is intermingled and co-aligned with a respective at
least one second rotor blade axially around a periphery of the
rotor disc.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] FIG. 1A schematically depicts a cross-section of a gas
turbine engine.
[0007] FIG. 1B shows a cross-section of a portion of the high
pressure compressor (HPC) section of the engine.
[0008] FIG. 2 is a magnified portion of a blade outer air seal
(BOAS) in the HPC prior to engine run-in.
[0009] FIG. 3A is magnified view of one portion of the BOAS prior
to run-in.
[0010] FIG. 3B shows the magnified portion of the BOAS from FIG. 3A
after run-in.
[0011] FIG. 3C is a magnified first blade tip with an abrasive
coating.
[0012] FIG. 3D is a magnified uncoated second blade tip absent an
abrasive coating.
[0013] FIG. 4 shows a first blade prior to being coated adjacent a
second uncoated blade.
[0014] FIG. 5 is a magnified view of a processed first blade with
squealer tip cuts.
[0015] FIG. 6 is a magnified view of a portion of an inner air
seal.
[0016] FIG. 7 depicts an alternative embodiment of the IBR with
individual first and second blades secured to a traditional rotor
disc.
DETAILED DESCRIPTION
[0017] FIG. 1A shows engine 10, fan section 12, low-pressure
compressor section 14, high-pressure compressor section 16,
combustor section 18, high-pressure turbine section 20,
low-pressure turbine section 22, bypass 24, low-pressure shaft 26,
and high-pressure shaft 28.
[0018] FIG. 1A schematically depicts an axial cross-section of an
example dual-spool turbofan engine 10. A portion of atmospheric air
pulled in by rotation of fan section 12 is directed toward
low-pressure compressor section 14, while the remainder is directed
through bypass section 24. Air entering low-pressure compressor
section 14 is further compressed by high-pressure compressor
section 16. Fuel is added to the compressed air and ignited in
combustor section 18. Blades in turbine sections 20 and 22 capture
a portion of the energy from expanding combustion products. Both
fan section 12 and low-pressure compressor section 14 are rotatably
linked via low-pressure shaft 26 to low-pressure power turbine
section 22. High-pressure compressor section 16 is rotatably
connected to high-pressure turbine section 22 via high-pressure
shaft 28. Thrust is generated in engine 10 by the force of the air
drawn in by fan section 12 and pushed through bypass section 24,
and by the force of exhaust gases exiting from low-pressure turbine
section 22.
[0019] Engine 10 is shown as a dual-spool turbofan engine. However,
this is merely one application of the inventive concepts described
herein. In the example dual-spool engine, the blade, vane, and seal
arrangements described below can be readily adapted to other
compressor sections as well as to fan section 12, low-pressure
compressor 14, and turbine sections 20, 22. The concepts can also
be adapted to other turbofan engines with a single spool or with
multiple spools, as well as to other types of gas turbine engines
including industrial gas turbines (IGTs), turboprops, and
turboshafts.
[0020] FIG. 1B includes low-pressure compressor section 16, IBR
assemblies 30, shrouded stator vane 31, first cantilevered stator
vane 32, second cantilevered stator vane 33, casing 34, first
coated stator vane tip 35. seal ring 36, second uncoated stator
vane tip 37, outer and inner abradable seal coatings 38A, 38B,
rotor lands 39, first blades 40, coated first blade tips 42, second
rotor blades 46, uncoated second blade tips 48, and knife edge
seals 50.
[0021] FIG. 1B is an axial cross-section of a portion of compressor
section 16, which includes multiple compressor stages where each
stage has a rotor assembly 30 and a respective stator assembly.
During rotation, rotor assembly 30 includes blades extending
radially outward adjacent to an adjacent circumferential surface
such as seal ring 36 secured to casing 34. Rotor assembly 30
cooperates with a stator disposed axially downstream to
progressively compress and direct inlet air axially into subsequent
stages toward combustor 18 (shown in FIG. 1A). Stator vanes are
mounted circumferentially and extend radially inward from
compressor casing 34 toward an adjacent circumferential surface
such as rotor lands 39. In this example, compressor section 16
includes shrouded stator vanes 31 at the initial upstream stages,
and cantilevered vanes 32, 33 in the higher downstream stages.
[0022] Here, rotor assemblies 30 are integrally bladed rotors
(IBR's), which are also known in the art as blisks. IBR assemblies
30 include both first blades 40 and second blades 46 integrally
secured thereto. Each IBR 30 has at least one first blade 40 with
abrasively coated airfoil tip 42 and at least one second blade 46
with uncoated airfoil tip 48. Coated first airfoil tips 42 interact
with seal rings 36 to form a groove for a blade outer air seal
(BOAS) in outer abradable coating 38A. Since not every blade 40, 46
is coated on each rotor 30, the example depicted in FIG. 1B
includes both coated first tips 42 and uncoated second airfoil tips
48 at various stages to better illustrate the differences and
similarities therebetween.
[0023] At lower pressure stages, shrouded vanes 31 are mounted
conventionally to casing 34. Shrouded vanes 31 include conventional
labyrinth seal 50 to minimize airflow therepast. In some
embodiments, labyrinth seals 50 are disposed in lower compressor
stages while the upper compressor stages can include the inner air
seal as described below. The particular number and distribution of
inner air seals and labyrinth seals can be determined via
simulation.
[0024] Similar to the BOAS, cantilevered vanes 32, 33 can be
adapted to interact with inner abradable coatings 38B on adjacent
rotor lands 39 to form an inner air seal. Casing 34 includes one
seal ring 36 at each stator stage. At least one first cantilevered
vane 32 is intermingled circumferentially with at least one second
cantilevered vane 33 around each seal ring 36. First vane 32 has
abrasively coated first airfoil tip 35 and second vane has second
airfoil tip 37 absent an abrasive coating.
[0025] Airfoils, which can include one or both of rotor blades and
stator vanes, previously included abrasive coatings on the tip of
each and every airfoil tip for a given stage to facilitate
formation of inner and outer air seals. However, as will be seen in
detail below, not every airfoil includes an abrasively coated tip,
which saves processing time and effort. It should be noted that
coated airfoil tips and numerous other elements have been
exaggerated for clarity. Other useful elements in the various
stator and rotor stages, such as dampers and anti-rotation devices,
have also been omitted for clarity.
[0026] FIG. 2 shows IBR assembly 30, casing 34, outer seal ring 36,
outer abradable seal coating 38A, first blades 40, coated first
blade tips 42, second blades 46, and uncoated second blade tips 48.
FIG. 2 depicts a portion of compressor rotor assembly 30 and casing
34 prior to run-in. IBR assembly 30 is a single rotor stage of
high-pressure compressor 16 (shown in FIGS. 1A and 1B). In this
example, at least one first rotor blade 40 with an abrasively
coated first airfoil tip 42 is intermingled around the periphery of
IBR assembly 30 with second blades 46 absent an abrasive coating on
second airfoil tip 48.
[0027] Casing 34 includes outer seal ring 36 secured annularly
around rotor 30 with outer abradable coating 38 disposed radially
adjacent to coated blade tips 42 and uncoated blade tips 48. These
combine to form a BOAS according to the details below. Depending on
exact tolerances, abrasively coated first airfoil tips 42 closely
approach or actually contact the annularly adjacent abradable
coating 38A. In this example, prior to operating in run-in mode for
the first time, there is a larger gap between the surfaces of
abradable coating 38A and uncoated blade tips 48, as compared to
the gap between abradable coating 38A and abrasively coated tips
42. This shorter length helps ensure the abrasive coated tips 42
contact abradable coating 38A rather than uncoated tips 48 both
during run-in mode and normal operational mode.
[0028] When the engine is cold, newly assembled or refurbished, and
has not yet been operated in a run-in mode, coated blade tips 42
will nearly or actually touch abradable coating 38, while uncoated
tips 48 generally avoid contact. As the engine is run in for the
first time with new or refurbished seal components, the speed is
quickly ramped up and down, on the order of only a few seconds per
cycle, to take advantage of the centrifugal expansion of IBR
assembly 30. This simulates the effects of a quick excursion to
full engine throttle under normal operation, which can occur for
example during takeoff and emergency maneuvers. Centrifugal
expansion caused by quick acceleration of IBR assembly 30 results
in first blades 40 extending spanwise, where abrasively coated
first airfoil tips 42 wear into abradable coating 38A.
[0029] During run-in operation, second blades 46 generally avoid
contact with abradable coating 38. However, when the engine is
operating normally and both rotor 30 and seal ring 36 have
centrifugally and thermally expanded, both coated tips 44 and
uncoated tips 48 form a seal with the groove. This minimizes surge
(backward airflow) through the compressor and eccentricity of rotor
30.
[0030] FIG. 3A is a magnified view of the seal shown in FIG. 2
prior to run-in, while FIG. 3B shows a magnified view of the same
arrangement after run-in. FIGS. 3A and 3B include seal ring 36,
outer abradable coating 38A, first rotor blade 40, coated first
airfoil tip 42, second rotor blade 46, and uncoated second airfoil
tip 48. FIG. 3B also shows outer seal groove 52A.
[0031] As was seen in FIG. 1B and FIG. 2, each IBR assembly 30 has
at least one first blade 40 each with coated first airfoil tips 42
intermingled with at least one second blade 46 each with uncoated
airfoil tips 48. As described with respect to FIG. 2, prior to
run-in, first blade 40 is approximately or actually in contact with
the surface of outer abradable seal coating 38A, while second blade
46 includes a gap between tip 48 and coating 38A. As the engine is
run-in, blades 40, 46 expand spanwise toward seal ring 36, but the
rapid speed transients of run-in mode minimize thermal expansion of
seal ring 36 as shown in FIG. 2. This leaves outer coating 38A in
substantially the same radial position allowing coated first
airfoil tips 42 to form groove 52A.
[0032] Minimizing tip clearances between stages optimizes the seal
to improve efficiency and operational stability of the engine. As
IBR 30 (shown in FIG. 2) spins up and down during run-in, first
blades 40 perform most or all of the work in wearing outer seal
groove 52A into outer abradable coating 38A. Groove 52A is a path
for the thermally and/or centrifugally expanded blades to travel as
they rotate, forming an effective seal around blade tips 42, 48
during the operational life of the engine. When blades 40, 46 are
both thermally and centrifugally expanded into groove 52A, they
together form a seal to minimize airflow past tips 42, 48. Depth of
blades 40, 46 in groove 52A determines the degree of backflow past
blade tips 42, 48, minimizing surge in the compressor. This
arrangement also removes eccentricity between the rotor and stator.
By ensuring both first blades 40 and second blades 46 are long
enough to extend into groove 52A when thermally expanded, most or
all of the sealing effect provided by groove 52A can be
maintained.
[0033] It should be noted that blades 40, 46 including respective
blade tips 44, 48 may have additional features and/or more complex
geometries than shown in FIGS. 2A and 2B. For simplicity, blades
40, 46 are shown without any additional features that further
reduce tip leakage. However, leakage reducing features may be
provided in certain embodiments to the extent they do not adversely
affect run-in or normal operation of the engine. As such, the
particular arrangements may not be precisely as shown in FIGS. 2,
3A, and 3B.
[0034] Abradable coating 38A (as well as inner coating 38B shown in
FIG. 1B) is more abradable than the material used for seal ring 36.
In this example, abradable coatings 38A, 38B are boro-nitride
ceramic. Abradable coatings 38A, 38B can additionally include a
thermal barrier coating such as a ceramic thermal barrier coating
(TBC) like yttria-stabilized zirconia (YSZ) disposed between it and
seal ring 36. The TBC can also be an MCrAlY coating, where M is one
of Ni, Co, or Fe and Y is an oxide like yttria or ytterbia. Some
MCrAlY TBCs are also abradable and can thus also serve in those
instances as both abradable coating 38 and as a thermal coating.
Such coatings are abradable enough to be worn away by contact with
coated tips 42 while still providing sufficient thermal protection
to seal ring 36. Each coating is applied using conventional
techniques appropriate for the particular compounds and substrate
materials. The techniques can include but are not limited to
electroplating, physical or chemical vapor deposition, plasma
spray, and combustion flame spray. In some cases, abradable coating
38 is provided as a strip adhesively or mechanically applied to
seal ring 36.
[0035] As seen here, first blades 40 can remain slightly longer
than second blades 46 both before and after run-in. In this way,
first blades 40 will continue to preferentially contact abradable
coating 38A as opposed to second blades 46. In situations such as
during high maneuver loads, rotor 30 is displaced relative to
casing 34. This could cause both types of blades 40, 46 to
penetrate beyond abradable coating 38, and potentially an
underlying thermal coating strike or seal ring 36. In such a case,
it is preferable that coated first airfoil tips 42 rather than
uncoated tips 48 absorb the majority of the contact forces. When
blades 40, 46 comprise titanium alloys, they are less abradable
than seal ring 36, which is itself less abradable than coating 38.
Continued friction between the components can produce excessive
heat and cause a titanium fire. In the case of nickel or other
superalloy blades, significant wearing of both second blade 46 and
seal ring 36 are likely in the event of inadvertent contact.
[0036] To form a conventional BOAS, each and every blade tip is
coated with abrasive material. This is done because varying the
blade weight increased the risk of rotor imbalance either during or
after run-in. Increased stresses on blade tips was also thought to
increase risk of tip or blade damage caused by fewer tips
performing more of the work in forming the seal groove. Similarly,
it was also required that all vanes in an individual stage have an
abradable coating to form inner seal grooves because leaving some
vane tips uncoated could cause metal buildup around the vane tips
when the abrasively coated rotor lands strike the uncoated metal.
This eventually led to larger clearances and greater efficiency
losses. However, coating each and every airfoil tip with abrasive
material introduces several complications in addition to increased
costs. The coating process weakens the airfoil itself, making it
more susceptible to bending stresses particularly around the tip.
In addition, coating every airfoil tip increases processing time,
effort, and opportunity for blade damage and scrapping.
[0037] IBR assembly 30 can be balanced to help minimize damage due
to uneven rotational moments and wearing in of groove 52A.
Particular intermingling of first blades 40 with second blades 46
is done with the goal of balancing the center of gravity of IBR
assembly 30. The minimum number and distribution of coated blade
tips 42 relative to uncoated tips 48, will depend on a confluence
of factors, including the CTE of airfoil, rotor, coating, and
casing materials, centrifugal expansion of the rotor, operating
temperatures and pressures, rotational speed and harmonics based on
blade and rotor shape, desired and minimum tip clearances, removal
of airfoil material to facilitate coating, among others. Modern
analytic and predictive software tools, such as SIMULIA.RTM.,
available from Dassault Systemes of Paris, France, can be used
model and analyze the combined effects of these and other blade and
operational characteristics in an effort to balance IBR assembly 30
during both run-in mode and normal operational mode.
[0038] These tools operate by using existing CAD definitions of the
various components and modifying variable characteristics of the
simulated components within the required parameters. Simulation
occurs under different operating conditions to identify suitable
and optimal blade distributions. For example, Monte Carlo
simulations can be performed with these or other software tools in
order to identify and analyze suitable blade distributions as well
as to calculate and minimize failure risks given the randomness of
possible operating conditions.
[0039] Using this or similar software, in combination with
empirical testing, suitable and/or optimal intermingling of first
rotor blades 40 with second blades 46 can be determined and tested
over a wide range of normal and abnormal conditions, including bird
strike and blade-off. These analytic tools can also be used to
identify optimal conditions for run-in, including slower rotation,
preheating or cooling of inlet air to expand blades relative to the
casing, etc. In certain embodiments, the ratio of second rotor
blades 46 with uncoated tips 48 to first rotor blades 40 with
coated tips 42 ranges from about 12:1 to about 1:1. In certain of
those embodiments, the ratio of second blades 46 to first blades 40
ranges from about 9:1 to about 3:1. In yet certain of those
embodiments, the ratio of second blades 46 to first blades 40 is
about 6:1. These blades will generally but are not required to be
evenly distributed. The ratio of second blades 46 to first blades
40 tends to decrease with the overall swept diameter of IBR 30
because more energy is absorbed by blade tips to form a larger
groove 52A in abradable coating 38.
[0040] In the example shown in FIG. 2, IBR assembly 30 shows first
blade 40 with abrasive coated tips 42 on either end and five second
blades 46 with uncoated tips 48 therebetween. Thus, in an example
IBR assembly 30 having sixty total blades, where first blades 40
and second blades 46 are intermingled in the manner depicted in
FIG. 2, there will be a total of ten first blades 40 and fifty
second blades 46. Though FIG. 2 shows a single first blade 40
adjacent second blades 46, first blades 40 can alternatively be
arranged in groups of two or three.
[0041] Blades 40, 46 can be welded around the periphery of a rotor
disc or alternatively, the entire IBR assembly 30 can be machined
out of a single block of metal like a titanium or nickel alloy. In
some examples, such as for cold-side applications in fan section 12
and compressor sections 14, 16 (shown in FIG. 1A), IBR 30,
including blades 40, 46, comprises a Ti-6Al-4V alloy. As will be
seen in FIG. 7, IBR assembly 30 can alternatively be a
traditionally bladed rotor with individually formed blades
removably secured around the periphery of a rotor disc. Many
engines combine traditionally bladed rotors and integrally-bladed
rotors in the same engine; however each particular section is often
limited to the same type of rotor in each section (e.g. LPC, HPC,
etc.).
[0042] FIG. 3C shows a magnified view of coated first airfoil tip
42, abrasive tip coating 54, matrix 56, and abrasive particles 58.
FIG. 3D shows uncoated second airfoil tip 48 with tip edge 60. Even
with a slightly larger gap as compared to coated tip 42, abrasive
coating 54 includes a jagged surface made up of abrasive particles
protruding unevenly to various degrees from matrix 56. Comparing
FIGS. 3C and 3D, uncoated blade tips 48 provide a smoother profile
relative to coated tips 42, improving airflow. When a substantial
number of second blades 46 are on IBR 30 (shown in FIG. 2), this
offers overall efficiency and stability comparable to fully coating
every blade tip without the attendant increases in processing time,
expense, and risks.
[0043] In certain embodiments, the clearance between coated tip 42
and the deepest point of seal groove 52A can range from about 1.0
mil (about 0.25 mm) to about 5.0 mils (about 1.3 mm) while the
corresponding clearance for uncoated tips 48 can range from about
2.0 mils (about 0.50 mm) to about 10.0 mils (about 2.5 mm). It will
be noted that first vanes 32 (shown in FIG. 1B) can be similarly
coated to form seal groove 52B on rotor land 39 (shown in FIG.
6).
[0044] Here, tip coating 54 is an abrasive coating, such as a cubic
boron nitride (CBN) based material, while the blade material is a
titanium alloy such as Ti-6Al-4V. Blades in higher compression
stages or in the turbine may require a more temperature resistant
blade or abrasive coating. Coating 54 can be added any time after
the airfoil tips 42 are formed, such as is shown in FIGS. 4 and 5.
For example, due to logistical considerations, blades 40 often will
not be coated until machining of IBR 30 is otherwise completed. In
addition, coating of blades 40 can be coordinated with other
coating processes, including with the addition of leading edge
protective coatings used to prevent thermal, corrosive, and/or
mechanical damage.
[0045] FIG. 4 shows unprocessed first blade 40' with spanwise
length L.sub.1, unprocessed blade tip 42', second blade 46 with
spanwise length L.sub.2, uncoated blade tip 48, blade leading edge
regions 62, blade trailing edge regions 64, and blade suction
surfaces 66. Unprocessed first blade 40' with spanwise length
L.sub.1 is adjacent to second blade 46 with spanwise length
L.sub.2. Unprocessed first blade 40' is shown prior to processing
and adding abrasive coating 54 (shown in FIG. 3C) of tip 42'. Blade
spanwise lengths L.sub.1 and L.sub.2 are determined based on a
number of considerations. In certain embodiments, prior to coating,
blades 40' are machined or otherwise formed to a slightly shorter
spanwise length L.sub.1 than remaining second blades 46, which are
machined to length L.sub.2 (i.e. L.sub.1<L.sub.2). This can be
done to provide room for a thicker coating to be added to blades 40
if needed to balance the weight, rotational moment, and harmonics
of rotor 30 (shown in FIG. 2). In some embodiments, given different
densities of blade material and coating material, blades 40' can
also ne machined or otherwise formed into a slightly different
shape around tip 42' as compared to uncoated tips 48 to
substantially equalize the mass and/or center of gravity of blades
40 and 46. In certain of those embodiments, unprocessed coated tips
42' can then be further processed and coated to form blade 40 seen
in FIGS. 3A and 3B. One example of additional processing is shown
in FIG. 5.
[0046] Alternatively, depending on the results of simulation and
real-life testing, second blades 46 can have substantially the same
underlying spanwise length and other dimensions (within an
acceptable tolerance) as blade 40'. It should be noted that in the
previous example where L.sub.1<L.sub.2, seal ring 36 has a
lesser CTE than the rotor assembly 30 to limit the depth of groove
50 formed by first blades 40, as well as reducing the chances of
second blades 46 striking the interior of seal groove 50 (shown in
FIG. 2B). However, blades 40, 46 can be adapted for rotors and seal
rings with more similar CTE's. In one example, seal ring 36 has
approximately the same CTE as rotor assembly 30. In such an
example, finished first blades 40 (shown in FIG. 3A) are
substantially longer than second blades 46 with uncoated tips 48.
This can be done by starting with blades 40' and 46 having
substantially the same respective lengths L.sub.1 and L.sub.2, then
adding coating 52 (shown in FIG. 3C) to extend the finished length
of blades 40. Thus, in certain alternative embodiments, all blades
are manufactured prior to coating to substantially the same length
(within tolerance limits) as second blades 46. In these
embodiments, L.sub.1 is approximately equal to L.sub.2. This
arrangement can complicate balance of rotor 30 and limit the number
of ways to suitably intermingle blades 40, 46, but the overall
balance of longer finished first blades 40 during both run-in and
normal operation can be accounted for in the blade definitions of
the software package as described above.
[0047] FIG. 5 shows processed blade 40'', processed first airfoil
tip 42'', leading edge region 62, trailing edge region 64, pressure
surface 68, and squealer tip cuts 70.
[0048] The strength of blade material around first airfoil tip 42''
can be weakened by the heat and pressure of the coating process,
making them more susceptible to bending stresses. Thus with all
airfoils on a rotor or stator coated with an abrasive material,
they are susceptible to damage during the run-in mode and during
high maneuver loads. To partially alleviate this risk, titanium
alloy blades sometimes included two 45.degree. chamfer cuts
chordwise along one or both the suction and pressure surfaces
between the leading and trailing edges.
[0049] Replacing chamfers with rounded squealer tip cuts 70 moves
the stress peaks, particularly those resulting from second- and
third-order bending resonances, into the thicker uncoated part of
finished first blade 40. With the stress peaks pushed past the
runout of tip cuts 70 to points adjacent leading edge 62 and
trailing edge 64, first blade 40 can better absorb contact with
abradable coating 38A (shown in FIGS. 3A and 3B). With a more
resilient first airfoil tip 42, it is therefore not necessary that
all first blades 40 be coated when forming groove 50A. Similarly,
first vanes 32 with similar or identical geometries as blades 40,
can include similar squealer tip cuts on the suction and pressure
sides of first coated airfoil tips 35 (shown in FIG. 6). This
similarly enables first vanes 32 to better absorb contact with
abradable coating 38B on rotor land 39 during formation of groove
50B, which in turn allows a number of second vanes 33 to remain
uncoated at tip 37.
[0050] Squealer tip cuts 70 have previously been used on many types
of uncoated airfoils to reduce tip leakage. This is ordinarily
accomplished by allowing for a smaller tip clearance while reducing
the risk and magnitude of damage in the event that the blade
strikes the casing or other outer stationary structure. Such tip
cuts are thus most frequently made at the center of the tip between
the suction and pressure surfaces rather than on the periphery as
seen above.
[0051] Regarding second blades 48 (and second vanes 33 shown in
FIG. 6), neither a chamfer nor a squealer cut is required on
airfoils that have uncoated tips. Since uncoated tips 48 are not
weakened from the coating process, squealer tip cuts 70 need only
be provided on first processed blades 40'' and not on second
uncoated blades 46 (shown in FIG. 3B). Thus, in addition to saving
on coating costs, effort, and the attendant issues therein, coating
only first blade tips 42 eliminates the time and effort needed to
form squealer tip cuts 70 or chamfers on a substantial number of
uncoated blade tips 48. Uncoated airfoil provide a much smoother
profile in the respective seal groove, improving the sealing effect
and reducing backflow into upstream compressor stages.
[0052] Squealer tip cuts 70 also can improve the balance of rotor
30 (shown in FIGS. 2A and 2B) by better balancing weights of
intermingled first blades 40 and second blades 46. With lower
density abrasive coatings replacing blade material around tip cuts
70, overall density of first blade 40 can be less than that of
blade 46, especially at respective airfoil tips 42, 48. By
equalizing weight around tips 42, 48, where imbalance can have the
largest effects, IBR 30 is more likely to remain balanced during
various operational parameters.
[0053] FIG. 6 schematically shows first vane 32, second vane 33,
coated vane tip 35, uncoated vane tip 37, inner abradable seal 38B,
rotor land 39, and inner seal groove 52B. First and second vanes
32, 33 can also be adapted with an alternative version of the
invention described above. FIG. 6 shows adjacent first vane 32 and
second vane 33 with respective coated tip 35 and uncoated tip 37.
The abrasive coating on tip 35 interacts with inner abradable seal
38B on rotor land 39 in a similar manner to form inner seal grooves
50B.
[0054] First and second vanes 32, 33 are analogous to first and
second rotor blades 40, 46 shown in previous figures. The first
airfoils (vanes 32 and blades 40) include abrasive coatings on the
respective first airfoil tips, which rub a seal groove (50A or 50B)
into an adjacent abradable seal coating region (38A or 38B). In
contrast, using an abrasively coated rotor land and uncoated vane
tips led to larger clearances and greater efficiency losses caused
by buildup of metal rubbed from the vane tips.
[0055] By slightly reducing the length of vanes 32 and/or providing
a slightly thicker abrasive coating on vanes 32, some of the vanes
33 can remain uncoated similar to FIG. 4. In both cases, each
abrasive surface is performing more work relative to the prior
case. This causes abrasive first airfoil tips 35 to experience
slightly more wear and during run-in and during high maneuver
loads. Thus they end up shorter than they otherwise would with a
full coating run. These factors can be also evaluated for a
particular design by including the various parameters into a
software package. Over time, first tips 35 and second uncoated tips
37 result in a similar overall clearance and similar sealing
effects in groove 50B as compared to seal groove 50A. In both
cases, each abrasive surface is performing more work and
experiences increased bending stresses relative to the conventional
approach of coating all airfoil tips. Thus applying squealer tip
cuts 70 (shown in FIG. 5) under abrasive coating of first tips 35,
first vanes 32 can similarly withstand these forces with lower risk
of wear and breakage.
[0056] Also similar to first and second blades 40, 46, significant
time and money is saved by only coating first vanes 32 and leaving
second vanes 33 without an abrasive. Generally, only one vane per
cluster need be coated. However, the number of coated vanes and
their relative coating thickness is determined using similar
considerations as for identifying the number of abrasively coated
rotor blades. Thus a greater or lesser number of coated tips 35 may
be required relative to uncoated tips 37. The effects of both inner
and outer seal grooves 50A, 50B (shown respectively in FIG. 3B and
FIG. 6) can be further enhanced by performing a computational fluid
dynamic (CFD) analysis as a parameter of the software
simulation.
[0057] FIG. 7 includes individually bladed rotor assembly 130 with
first blade 140, coated blade tip section 142, rotor disc 144,
second blade 146, uncoated blade tip section 148, blade leading
edges 162, blade trailing edges 164, blade pressure surfaces 168,
platforms 172, and root sections 174. Bladed rotor assembly 130 is
an alternative embodiment of IBR assembly 30, and includes a
plurality of second blades 146 and first blades 140 distributed
around rotor disc 144 in a manner similar to that shown in FIG. 2.
In this alternative embodiment, individual blades 140, 146 are
similar to corresponding IBR blades 40, 46. Blades 140, 146 each
include platforms 172 and root sections 174 for retention by a
traditional rotor disc 144. Blades 140 and 146 can be removably
secured to rotor disc 144 with a ring or other structure
corresponding to root sections 174. It should be noted that other
features that may form a part of rotor disc 144, such as rotor
lands or spacer arms have been omitted for clarity.
[0058] In this alternative embodiment, first blades 140 and second
blades 146 include respective coated blade tips 142 and uncoated
tips 148 (like those shown in FIG. 4). Like coated IBR blades 40,
coated individual blades 140 can be distributed equally around
rotor disc 144 to ensure balance during and after run-in. For
example, depending on whether the rotor is an IBR or a
traditionally bladed rotor, both first blades 40 or 140 can be
distributed in groups of one, two, three, to help balance second
blades 46 or 146. Groupings and the particular coating
characteristics for a particular rotor can then be optimized, for
example, using the software packages described above along side
empirical testing. A Monte Carlo type simulation can also be used
in order to determine suitable blade intermingling in this
configuration (as described with reference to IBR 30), and to
evaluate the likelihood of imbalance and failure under randomized
operational scenarios.
[0059] While the invention has been described with reference to an
exemplary embodiment(s), it will be understood by those skilled in
the art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment(s) disclosed, but that the invention will
include all embodiments falling within the scope of the appended
claims.
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