Title: Counterstagger compressor airfoil
Abstract: A compressor airfoil includes opposite pressure and suction sides joined together at leading and trailing edges and extending in span between a root and tip. The airfoil includes stagger increasing above the root and decreasing above a midspan pitch section thereof.
Patent Number: 6,899,526 Issued on 05/31/2005 to Doloresco, et al.
| Inventors:
|
Doloresco; Bryan Keith (Cincinnati, OH);
Wood; Peter John (Loveland, OH)
|
| Assignee:
|
General Electric Company (Schenectady, NY)
|
| Appl. No.:
|
634545 |
| Filed:
|
August 5, 2003 |
| Current U.S. Class: |
416/238; 416/223A; 416/242; 416/243; 416/DIG.2 |
| Intern'l Class: |
F01D 005/14 |
| Field of Search: |
416/238,223.A,DIG.2,DIG.5,242,243
|
References Cited [Referenced By]
U.S. Patent Documents
| 2415847 | Feb., 1947 | Redding.
| |
| 4585395 | Apr., 1986 | Nourse et al.
| |
| 4682935 | Jul., 1987 | Martin.
| |
| 5088892 | Feb., 1992 | Weingold et al.
| |
| 5167489 | Dec., 1992 | Wadia et al.
| |
| 5249922 | Oct., 1993 | Sato et al.
| |
| 5397215 | Mar., 1995 | Spear et al.
| |
| 5642985 | Jul., 1997 | Spear et al.
| |
| 6071077 | Jun., 2000 | Rowlands.
| |
| 6079948 | Jun., 2000 | Sasaki et al.
| |
| 6290465 | Sep., 2001 | Lammas et al.
| |
| 6299412 | Oct., 2001 | Wood et al.
| |
| 6331100 | Dec., 2001 | Liu et al.
| |
Other References
Smith et al, "Sweep and Dihedral Effects in Axial-Flow Turbomachinery, " ASME
62-WA-102, 1962, pp: 1-14.
|
Primary Examiner: Look; Edward K.
Assistant Examiner: Edgar; Richard A
Attorney, Agent or Firm: Andes; William S., Conte; Francis L.
Claims
1. A compressor airfoil for pressurizing air inside a surrounding casing, said
airfoil comprising:
laterally opposite pressure and suction sides joined together at chordally opposite
leading and trailing edges and extending in span from a root to a tip;
stagger increasing above said root, and decreasing above a midspan pitch section
of said airfoil; and
a dihedral angle relative to said casing increasing above said pitch section
to said tip.
2. An airfoil according to claim 1 further comprising a concave axial projection
along said leading edge, with said root and tip extending forward of said pitch
section along said leading edge.
3. An airfoil according to claim 2 wherein said stagger increases in magnitude
from said root to said pitch section, and decreases in magnitude above said pitch
section toward said root stagger magnitude.
4. An airfoil according to claim 3 wherein said dihedral angle above said pitch
section is opposite to said dihedral angle between said root and pitch section.
5. An airfoil according to claim 4 wherein dihedral angle along said leading
edge at said tip is greater than below said pitch section.
6. An airfoil according to claim 5 further comprising forward aerodynamic sweep
at both said leading and trailing edges of said tip.
7. An airfoil according to claim 6 further comprising aft aerodynamic sweep from
said root to said pitch section and to below said tip along said leading and trailing edges.
8. An airfoil according to claim 6 wherein said stagger varies along said leading
edge to bow said leading edge concave in span along said suction side.
9. An airfoil according to claim 6 wherein said dihedral angle is unidirectional
along said tip between said leading and trailing edges.
10. An airfoil according to claim 6 wherein said stagger has a maximum value
located in a range of about 60%-85% span from said root.
11. A compressor airfoil comprising:
laterally opposite pressure and suction sides joined together at chordally opposite
leading and trailing edges and extending in span from a root to a tip; and
stagger increasing above said root, and decreasing above a midspan pitch section
of said airfoil.
12. An airfoil according to claim 11 wherein said stagger increases in magnitude
from said root to said pitch section, and decreases in magnitude above said pitch
section toward said root stagger magnitude.
13. An airfoil according to claim 12 further comprising a dihedral angle relative
to a surrounding casing increasing above said pitch section to said tip.
14. An airfoil according to claim 13 wherein said dihedral angle above said pitch
section is opposite to said dihedral angle between said root and pitch section.
15. An airfoil according to claim 14 wherein dihedral angle along said leading
edge at said tip is greater than below said pitch section.
16. An airfoil according to claim 14 wherein said stagger varies along said leading
edge to bow said leading edge concave in span along said suction side.
17. An airfoil according to claim 14 wherein said dihedral angle is unidirectional
along said tip between said leading and trailing edges.
18. An airfoil according to claim 14 further comprising forward aerodynamic sweep
at both said leading and trailing edges of said tip.
19. An airfoil according to claim 18 further comprising aft aerodynamic sweep
from said root to said pitch section and to below said tip along said leading and
trailing edges.
20. An airfoil according to claim 14 further comprising a concave axial projection
along said leading edge, with said root and tip extending forward of said pitch
section along said leading edge.
Description
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and, more specifically,
to compressors therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel
in a combustor for generating hot combustion gases. The combustion gases are discharged
through turbine stages which extract energy therefrom for powering the compressor,
and producing output power for use in driving a fan in an exemplary turbofan aircraft
engine application.
A multistage axial compressor includes cooperating rows of stator vanes and rotor
blades which decrease in size to pressurize air in stages. The compressor vanes
and blades have corresponding airfoils which typically vary in configuration as
their size decreases from stage to stage for maximizing performance of the compressor.
Compressor performance includes, for example, efficiency of compression, flow capability,
and stall margin, which are all affected by the configuration of the vanes and blades.
More specifically, the flow or pressure distribution of the air as it is being
compressed through the stator vanes and rotor blades is a complex three dimensional
flow field varying circumferentially around the compressor, radially along the
span of the vane and blade airfoils, and axially along the circumferentially opposite
pressure and suction sides of the airfoils
The airfoil pressure side is a generally concave surface cooperating with the
opposite suction side, which is a generally convex surface, for efficiently pressurizing
the air as it flows between blades in the axial downstream direction between the
leading and trailing edges thereof. The pressure distribution of the air undergoing
compression varies from the radially inner root of the airfoil to the radially
outer tip of the airfoil which is spaced closely adjacent to a surrounding compressor
casing to provide a suitable radial gap or clearance therewith.
The airfoil, itself, may be supported from the compressor rotor in any suitable
manner such as being formed integrally therewith in a unitary blisk configuration,
or each rotor airfoil may have an integral platform and dovetail for mounting the
compressor blade in a corresponding dovetail slot formed in the perimeter of the
compressor rotor.
A significant feature affecting compressor performance is the radial clearance
provided between the airfoil tips and surrounding casing. The clearance should
be as small as possible to minimize undesirable flow losses therethrough, but must
be sufficiently large for accommodating transient operation of the compressor which
may occasionally lead to tip rubs. In a tip rub, material is removed from the airfoil
tip and may accumulate on the inner surface of the casing. The shortened tip increases
the clearance with the casing which decreases compressor performance, which is
further affected by any accumulation of rub material on the casing which disrupts
the smooth flow of air therealong.
Nevertheless, commercial experience of multistage axial compressors
in aircraft turbofan engines confirms long useful lives for the compressor rotor
blades and continued high performance of the compressor. However, the occasional
compressor blade tip rubs increase tip clearances and decrease compressor performance
over the useful blade lifetime. The loss in compressor performance due to tip rubs
further reduces performance of the engine since the pressurized air is used in
the combustion process, and energy is extracted from the combustion gases in the turbines.
Accordingly, it is desired to provide a compressor rotor airfoil having
improved aerodynamic efficiency notwithstanding increased blade tip clearances
due to tip rubs.
BRIEF DESCRIPTION OF THE INVENTION
A compressor airfoil includes opposite pressure and suction sides joined together
at leading and trailing edges and extending in span between a root and tip. The
airfoil includes stagger increasing above the root and decreasing above a midspan
pitch section thereof.
BRIEF DESCRIPTION OF THE DRAWINGS
The invention, in accordance with preferred and exemplary embodiments, together
with further objects and advantages thereof, is more particularly described in
the following detailed description taken in conjunction with the accompanying drawings
in which:
FIG. 1 is a partly sectional, axial projection side view of a row of compressor
rotor airfoils in a multistage axial compressor.
FIG. 2 is a isometric view of an exemplary one of the compressor rotor blades
illustrated in FIG. 1 in isolation.
FIG. 3 is a top radial view of the compressor airfoil illustrated in FIG. 2
and taken generally along line 3—3.
FIG. 4 is a graph plotting stagger in degrees over the radial span of the airfoil
illustrated in FIGS. 1-3 in an exemplary embodiment.
FIG. 5 is a graph plotting dihedral angle in degrees along the leading edge
over the radial span of the airfoil illustrated in FIGS. 1-3 in an exemplary embodiment.
FIG. 6 is a graph plotting aerodynamic sweep angle in degrees along the leading
and trailing edges over the radial span of the airfoil illustrated in FIGS. 1-3
in an exemplary embodiment.
DETAILED DESCRIPTION OF THE INVENTION
Illustrated in FIG. 1 is a row of compressor rotor blades
10 suitably
mounted to a compressor rotor
12 of a multistage axial compressor in a gas
turbine engine, shown in part. The compressor has several stages of stator vanes
(not shown) cooperating with corresponding compressor blades which decrease in
size in the downstream direction as air
14 is compressed during operation.
The rotor
12 is axisymmetrical around the axial centerline axis
16
of the engine and supports a full row of the blades
10 within an annular
outer casing
18.
Each compressor rotor blade
10 includes an airfoil
20 extending
in span along a radial axis Z between the perimeter of the rotor and the inner
surface of the casing
18. The airfoil may be integrally formed with the
rotor
12 in a blisk configuration (not shown), or may be removably joined
thereto in a conventional manner.
For example, each airfoil may include an integral platform
22 which defines
the inner boundary for the air being compressed. An integral dovetail
24
extends from the platform in a unitary configuration with the blade for being mounted
in a complementary dovetail slot in the perimeter of the rotor. In the exemplary
embodiment illustrated in FIG. 1, the dovetail
24 is a circumferential entry
dovetail suitably mounted in the perimeter of the rotor.
The compressor airfoil
20 is illustrated in a preferred embodiment in
FIGS. 1 and 2 and includes circumferentially or laterally opposite pressure and
suction sides
26,
28. The airfoil pressure side is generally concave
and precedes the generally convex suction side as the airfoil rotates in the circumferential
direction, represented by the Y axis, atop the rotor. The axial axis X is parallel
with the engine centerline axis and represents the generally downstream direction
of the air
14 as it undergoes compression through the multiple stages of
the compressor.
The corresponding surfaces of the pressure and suction sides are joined together
at axially or chordally opposite leading and trailing edges
30,
32
and extend in radial span from a radially inner root
34 at the junction
with the platform to a radially outer tip
36.
As shown in FIG. 1, the airfoil tip
36 is disposed closely adjacent to
the inner surface of the surrounding casing
18 and defines a substantially
constant radial clearance or gap therebetween extending between the leading and
trailing edges of the airfoil. The generally concave configuration of the airfoil
pressure side
26, and the generally convex configuration of the airfoil
suction side
28 are conventionally defined for pressurizing the air
14
as it flows downstream between the compressor rotor blades
10 in each stage
of the compressor.
The three-dimensional configuration of the airfoil may be defined in accordance
with conventional practice to maximize aerodynamic performance of the compressor
including efficiency, flow, and stall margin. And, the configuration of the airfoil
is also designed for minimizing centrifugal stresses created therein during rotary
operation of the blades in the compressor.
For example, conventional compressor rotor blades are designed with varying twist
or stagger from root to tip thereof. The various radial sections of the airfoil
have centers of gravity stacked along a suitable radial stacking axis which may
be straight or bowed for effecting reduced centrifugal stress during operation.
The surfaces of the airfoil are disposed relative to the incident air
14
being pressurized with suitable values of aerodynamic sweep which varies between
the leading and trailing edges and root to tip of the airfoil.
As indicated above, the occasional rubbing of the airfoil tip
36 with
the
casing
18 may increase the radial clearance therebetween and decrease compressor
performance for conventional compressor rotor blades. In order to reduce the sensitivity
of the compressor airfoil illustrated in FIGS. 1 and 2 to increased clearance due
to tip rubs, and for improving compressor performance, the airfoil
20 is
suitably modified as described hereinbelow.
For example, FIG. 3 illustrates a top view of the airfoil illustrated in FIG.
2 with a superimposed rectangular grid over the pressure and suction sides thereof.
Each radial section of the airfoil includes a straight chord
38 extending
from the leading edge to the trailing edge thereof which defines with the axial
axis X a twist or stagger angle A. The stagger angle A is plotted in FIG. 4 in
accordance with an exemplary embodiment varying in degrees from the root at zero
span to the normalized tip at unity (1.0).
A significant feature of the compressor airfoil illustrated in FIGS. 1-3 is the
introduction of bowed- or counter-stagger along the span thereof. Preferably, the
stagger increases above the root
34, and decreases above a midspan pitch
section
40.
In a conventional compressor rotor airfoil, the stagger angle typically increases
from root to tip of the blade. The desired stagger angle is primarily controlled
by the desired pressure distribution in the air being pressurized which varies
from root to tip of the airfoil.
In contrast, the stagger angle of the airfoil illustrated in FIGS. 2 and 3 increases
in magnitude from a minimum value at the root
34 to a larger value at the
pitch section
40, and decreases in magnitude above the pitch section toward
the root stagger magnitude.
In the exemplary graph illustrated in FIG. 4, the stagger angle has a minimum
value of 40 degrees at the airfoil root and increases to a maximum value of about
47 degrees above the pitch section. From its maximum value the stagger angle decreases
to the airfoil tip which has a stagger angle of 43 degrees which is slightly greater
than the stagger angle at the root. The maximum stagger angle is preferably located
above the midspan pitch section of the airfoil to promote the desired pressure
distribution over the airfoil span. In the exemplary embodiment illustrated in
FIG. 4, the maximum stagger value is located in the range of about 60%-85% span
from the airfoil root.
The introduction of the reverse or counterstagger in the compressor airfoil above
its pitch section results in the distinctive configuration of the airfoil illustrated
in FIGS. 2 and 3. The counterstagger in the outer span of the airfoil substantially
reduces the blade tip stagger over that found in conventional compressor airfoils
for significantly reducing clearance sensitivity due to tip rubs as confirmed by
three-dimensional computational fluid dynamic analyses. Rotor blade aerodynamic
efficiency is improved at nominal clearance levels, and is significantly improved
at deteriorated clearance levels following tip rubs. Analyses also confirm improved
flow pumping with the deteriorated tip clearances. And, improvement in stall margin
may also be possible.
The counterstagger is introduced or manifested in the airfoil illustrated in
FIGS. 2 and 3 primarily along the leading edge
30 relative to the trailing
edge
32. The trailing edge is generally straight from root to tip, whereas
the leading edge
30 includes a distinctive concave bow along the radial
span of the suction side
28, with the root and tip following in movement
the pitch section
40 of the airfoil as it rotates in the tangential or circumferential
Y direction illustrated in FIG.
3.
The introduction of reduced or lower blade tip stagger in the compressor airfoil
illustrated in FIGS. 2 and 3 facilitates the introduction of increased tip leading
edge dihedral. Tip dihedral is identified in FIG. 2 as the angle B between the
local surface of the airfoil and the surrounding casing
18 illustrated in
FIG.
1. Tip dihedral is a conventional parameter, with zero tip dihedral
resulting in a blade tip which is oriented normal or perpendicular to the casing.
Positive tip dihedral is achieved when the pressure or concave side, or both, of
the blade tip forms an acute angle with the surrounding casing.
FIG. 5 is a graph of an exemplary profile of the dihedral angle B along the
leading edge of the airfoil illustrated in FIGS. 1-3 relative to the surrounding
casing in which the dihedral angle increases above the pitch section to a maximum
value at the airfoil tip
36.
As shown in FIG. 5, the dihedral angle above the pitch section is positive and
opposite to the dihedral angle between the root and pitch sections which is primarily
negative. Preferably, the dihedral angle along the airfoil leading edge
30
at the tip
36 is greater in magnitude than below the pitch section
40.
In the exemplary embodiment illustrated in FIG. 5, the dihedral angle varies from
zero at the airfoil root to a maximum negative value of about -7 degrees at about
25% span returning to a zero value below the pitch section at about 40% span, and
then increasing in magnitude to a maximum positive value of about 25 degrees at
the airfoil tip at 100% span.
The dihedral angle B cooperates with the stagger angle A, both of which vary
along the airfoil leading edge
30 to bow the airfoil leading edge concave
in span along the suction side
28. The cooperation of the dihedral and stagger
permit desirable positive dihedral along the airfoil tip with a maximum value at
the airfoil leading edge, and relatively small but still positive magnitude of
dihedral at the airfoil trailing edge. Correspondingly, a negative value of dihedral
is provided immediately above the airfoil root, and along with the counterstagger
in the airfoil effects the distinctive counter bowed leading edge illustrated in
FIGS. 2 and 3.
FIGS. 3 and 5 illustrate a preferred configuration of the dihedral angle being
unidirectional with the same positive magnitude along the airfoil tip
36
from the leading edge
30 to the trailing edge
32. In this way, undesirable
negative dihedral is not found at the airfoil tip for improving aerodynamic performance
thereof, including performance following occasional tip rubs.
The compressor airfoil illustrated in FIG. 1 also includes aerodynamic sweep
C which is a conventional term of art. The counterstagger and tip dihedral permit
a new introduction of forward or negative aerodynamic sweep at both the leading
and trailing edges
30,
32 of the airfoil tip
36 for further
improving aerodynamic performance of the compressor airfoil.
FIG. 6 is an exemplary graph of the aerodynamic sweep angle C in degrees for
the leading and trailing edges
30,
32 of the airfoil illustrated in
FIG. 1 along the radial span thereof. Negative or forward aerodynamic sweep is
introduced at the airfoil tip along both the leading and trailing edges, with the
sweep having a larger magnitude at the trailing edge. And, aft or positive aerodynamic
sweep is introduced in the airfoil from the root
34 to the pitch section
40, and further radially outwardly to just below the airfoil tip along both
the leading and trailing edges. Along the leading edge
30, the sweep transitions
from positive to negative at about 90% span, and along the trailing edge, the sweep
transitions from positive to negative at about 80% span.
FIG. 1 illustrates an axial projection view, or meridional view of the airfoil.
The airfoil is shown with a concave axial projection along the leading edge
30,
with the root
34 and tip
36 extending forward of the pitch section
40 along the leading edge. In this axially bowed or concave leading edge
projection, the airfoil outer span is axially forward of the midspan region and
permits the introduction of the aerodynamically favorable forward blade tip sweep
at both the leading and trailing edges. Furthermore, the axially forward airfoil
tip section permits the trailing edge dihedral at the tip to maintain a favorable
positive value, and thusly avoiding undesirable negative dihedral along the airfoil tip.
As indicated above, compressor rotor airfoils are complex and sophisticated three-dimensional
elements typically designed with various compromises for the competing demands
of aerodynamic performance and mechanical strength. Stagger, dihedral, and aerodynamic
sweep are all conventional features used in designing modern compressor rotor blades
as indicated by the exemplary references of record, incorporated herein by reference.
However, the exemplary compressor rotor blade illustrated in FIGS. 1-3 includes
distinct configurations of stagger, dihedral, and aerodynamic sweep which are used
to advantage in a cooperation for enhancing compressor performance not only with
a nominal clearance with the compressor casing, but after increased clearance following
occasional tip rubs in extended use of the blade in a gas turbine engine.
The introduction of specific forms of stagger, dihedral, and aerodynamic sweep
at the compressor blade tip illustrated in the figures is blended with the stagger,
dihedral, and sweep in the inner span portion of the airfoil resulting in a distinctive
overall configuration and corresponding performance enhancement.
While there have been described herein what are considered to be preferred
and exemplary embodiments of the present invention, other modifications of the
invention shall be apparent to those skilled in the art from the teachings herein,
and it is, therefore, desired to be secured in the appended claims all such modifications
as fall within the true spirit and scope of the invention.
*
