DOUBLE BOWED COMPRESSOR AIRFOIL
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and,
more specifically, to compressors or fans therein.
In a turbofan aircraft gas turbine engine, air is pressurized in a fan and
compressor during operation. The fan air is used for propelling an aircraft in
flight. The air channeled through the compressor is mixed with fuel in a
combustor and ignited for generated hot combustion gases which flow
through turbine stages that extract energy therefrom for powering the fan
and compressor.
A typical turbofan engine includes a multistage axial flow compressor
which pressurizes the air sequentially to produce high pressure air for
combustion. The compressed air is diffused and decelerates as it is
compressed. Compressor airfoils must therefore be designed to reduce
undesirable flow separation which would adversely affect stall margin and
efficiency.
Conversely, combustion gases are accelerated through the turbine
stages, and the turbine blades have different aerodynamic designs for
maximizing efficiency of energy extraction.
Fundamental in compressor design is efficiency in compressing the air
with sufficient stall margin over the entire flight envelope of operation from
takeoff, cruise, and landing.
However, compressor efficiency and stall margin are normally
inversely related with increasing efficiency typically corresponding with
decrease in stall margin. The conflicting requirements of stall margin and
efficiency are particularly demanding in high performance military engine
applications which require high level of stall margin typically at the expense
of compressor efficiency, as opposed to less demanding commercial
applications.
Maximizing efficiency of compressor airfoils is primarily effected by
optimizing the velocity distributions over the pressure and suction sides of
the airfoil. However, efficiency is typically limited in conventional
compressor design by the requirement for a suitable stall margin. Any
further increase in efficiency typically results in a reduction in stall margin,
and, conversely, further increase in stall margin results in decrease in
efficiency.
High efficiency is typically obtained by minimizing the wetted surface
area of the airfoils for a given stage to correspondingly reduce airfoil drag.
This is typically achieved by reducing airfoil solidity or the density of airfoils
around the circumference of a rotor disk, or by increasing airfoil aspect ratio
of the span to chord lengths.
For a given rotor speed, this increase in efficiency reduces stall
margin. To achieve high levels of stall margin, a higher than optimum level
of solidity and/or lower than optimum aspect ratios may be used, along with
designing the airfoils at below optimum incidence angles. This reduces
axial flow compressor efficiency.
Increased stall margin may also be obtained by increasing rotor speed,
but this in turn reduces efficiency by increasing the airflow Mach numbers,
which increases airfoil drag.
And, compressor blades are subject to centrifugal stress which is
affected by aerodynamic design. Peak stress must be limited for obtaining
useful blade life, and this in turn limits the ability to optimize aerodynamic
performance.
Accordingly, typical compressor designs necessarily include a
compromise between efficiency and stall margin favoring one over the other,
which are further affected by allowable centrifugal stress.
It is, therefore, desired to further improve both compressor efficiency
and stall margin while limiting centrifugal stress for improving gas turbine
engine compressor performance.
BRIEF SUMMARY OF THE INVENTION
A compressor airfoil includes pressure and suction sides extending
from root to tip and between leading and trailing edges. Transverse sections
have respective chords and camber lines. Centers of gravity of the sections
are aligned along a double bowed stacking axis for improving performance.
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:
Figure 1 is an isometric view of a portion of a gas turbine engine
compressor rotor stage having bowed airfoils extending radially outwardly
from an integral rotor disk in accordance with an exemplary embodiment of
the present invention.
Figure 2 is an aft-looking-forward isometric view of one of the airfoils
illustrated in Figure 1 and taken generally along line 2-2 in a tangential and
radial plane.
Figure 3 is a side elevation view of one of the airfoils illustrated in
Figure 1 and taken generally along line 3-3 circumferentially projected in an
axial and radial plane.
Figure 4 is a top view of the airfoil illustrated in Figure 3 and taken
along line 4-4.
Figure 5 is a graph of an exemplary double bowed tangential stacking
axis for the airfoil illustrated in figures 1 -4.
DETAILED DESCRIPTION OF THE INVENTION
Illustrated in Figure 1 is a portion of an annular rotor blisk 10 defining
one stage of a multistage axial flow compressor for a gas turbine engine.
The blisk includes a plurality of circumferentially spaced apart rotor blades or
airfoils 12 extending radially outwardly from the perimeter of an integral
rotor disk 14 forming a one-piece unitary assembly. The blisk may be
manufactured using conventional milling and electrochemical machining.
Alternatively, the airfoils may be formed with integral dovetails for
being removably mounted in corresponding dovetail slots in the perimeter of
discrete rotor disk in another conventional configuration.
During operation, the blisk rotates in the exemplary clockwise
direction illustrated in Figure 1 for pressurizing air 16 as it flows between the
adjacent airfoils. The airfoils are aerodynamically configured in profile for
maximizing the efficiency of air compression while also providing a suitably
high stall margin for enhancing performance of the compressor. The blisk
10 illustrated in Figure 1 is only one of several stages of rotor airfoils which
may be configured in accordance with the present invention for enhancing
compressor performance by increasing together both efficiency and stall
margin, within allowable centrifugal stress limits.
Notwithstanding the conventional compromise made between
aerodynamic efficiency and stall margin, modern computer software is
conventionally available for solving three-dimensional (3D) viscous flow
equations for evaluating airfoil performance. The resulting airfoils generally
have distinctive 3D configurations which differ significantly over
conventional airfoils which vary less in radial section over the spans thereof.
Figure 1 illustrates a specifically doubled bowed airfoil 12 uncovered
from 3D analysis having improved performance for increasing both efficiency
and stall margin not previously possible due to stress limits.
The rotor disk 14 has three orthogonal axes including axial X,
tangential or circumferential Y, and radial Z. The axial axis X extends in the
downstream direction relative to the flow of air 16 through the compressor.
The tangential axis Y extends in the direction of rotation of the disk and
airfoils. And, the radial axis Z extends radially outwardly from the perimeter
of the disk for each of the airfoils thereon.
Each airfoil 12 includes a generally concave pressure side 18 and a
generally convex suction side 20 extending radially or longitudinally from a
root or hub 22 integrally joined with the perimeter of the disk to a radially
outer tip 24. The two sides extend chordally or axially between leading and
trailing edges 26, 28 from root to tip.
In accordance with one feature of the present invention, the airfoil
suction side 20 is laterally or tangentially bowed along the trailing edge 28
near or adjacent the root 22 at the intersection with the disk perimeter.
Flow separation of the air at this location may be substantially reduced or
eliminated for both increasing blade efficiency and improving stall margin.
The suction side trailing edge is bowed primarily only in the tangential
direction as illustrated in Figure 2. In the side projection of the axial and
radial plane X-Z illustrated in Figure 3, the suction side bow is imperceptible.
However, the airfoil may also be axially bowed as illustrated in Figure 3 for
further improvements in performance as later discussed hereinbelow.
The airfoil illustrated in Figures 1-3 is defined by a plurality of radially
or longitudinally stacked transverse sections from root to tip as illustrated in
Figure 4. Each section has an aerodynamic profile defined by respective
portions of the pressure and suction sides 18,20 extending between the
leading and trailing edges 26,28. Each profile is defined by a straight chord
30 extending axially between the leading and trailing edges, and an arcuate
camber line 32 which is a meanline spaced equidistantly between the
pressure and suction sides from leading to trailing edge.
The compressor airfoil 12 typically twists from root to tip for
maximizing compressor performance. The twist is defined by a stagger
angle A measured between the chord 30 and axial axis X at the leading
edge 26, for example, for each radial section. The stagger typically
increases from root to tip, and is larger at the tip than at the root.
Each airfoil section also has a center of gravity 34 which is aligned
radially along the longitudinal span of the airfoil in a stacking axis 36 as
illustrated in Figure 1 which is preferably double bowed in the tangential
direction in accordance with another feature of the present invention. The
stacking axis 36 in conjunction with the shapes of the corresponding airfoil
sections including their chords 30 and camber lines 32 permit 3D definition
of the airfoil for enhanced performance in accordance with the present
invention.
More specifically, the stacking axis 36 illustrated in Figure 1 has two
orthogonal components including a tangential stacking axis 36a illustrated in
Figures 2 and 5, and an axial stacking axis 36b illustrated in Figure 3. The
tangential stacking axis 36a is non-linear or bowed adjacent the airfoil root
22 to bow the suction side 20 of the airfoil near the trailing edge root or
hub.
As shown in Figures 1 and 5, the tangential stacking axis 36a
includes a first inversion or bow 38 having an initial lean onward or in the
forward direction of rotation of the airfoils and disk from the root 22 toward
the pressure side 18 of the airfoil. The first bow 38 then reverses lean
backward toward the radial axis Z.
The stacking axis 36a also includes a second inversion or bow 40
which leans hindward or backward past the radial axis Z from the first bow,
opposite to the direction of rotation of the airfoils and disk, toward the
suction side 20 adjacent the tip 24. The second bow then reverses lean
forward toward the radial axis Z. Correspondingly, stagger angle of the
airfoil transverse sections adjacent the root varies in turn to bow the suction
side along the trailing edge suction side.
The double bow of the tangential stacking axis 36a thusly has a
generally S-shape, and the corresponding shapes of the transverse sections
are selected for substantially reducing or eliminating flow separation of the
air along the suction side near the airfoil hub at the trailing edge, while also
reducing centrifugal stress. For example, the trailing edge 28 also has a
generally S-shape from root to tip.
The S-bowed stacking axis permits the trailing edge 28 as illustrated
in Figures 1 and 2 to be oriented substantially normal to the root of the
bowed suction side 20 and leans hindward thereabove. The trailing edge 28
intersects the perimeter or platform of the rotor disk at an intersection angle
B which would otherwise be significantly acute without the trailing edge
bow. Computer analysis indicates that acute trailing edge intersection
angles promote hub flow separation which decreases efficiency of the airfoil.
The suction side bow reduces the acuteness of the intersection angle B for
correspondingly reducing flow separation, with an attendant increase in
efficiency.
However, since the airfoil is a 3D design, its various sections are
aerodynamically and mechanically interrelated in a complex manner.
Accordingly, the shape and amount of tangential lean in the first bow 38 in
the direction of rotation are preferably controlled by aerodynamic analysis to
eliminate or reduce hub flow separation at the trailing edge. The first bow
correspondingly also moves the peak centrifugal stress away from the airfoil
root into the airfoil sections at the first bow.
In order to then reduce the centrifugal stress in the first bow region,
mechanical or stress analysis may then be used to control the remainder of
the tangential stacking axis profile in its transition outboard of the first bow
in the direction opposite to rotation. Centrifugal stress at the root and in the
first bow region may then be reduced by introducing the second bow 40
which leans the stacking axis once again in the direction of rotation for the
airfoil tip region.
The first and second bows 38,40 are disposed on opposite sides of
the radial axis Z extending through the center of gravity of the airfoil root to
limit peak centrifugal stress while maximizing aerodynamic performance at
the root. Both bows include inversion points at which the stacking axis
changes direction between onward and hindward. And, the second bow
may extend back across the radial axis if required to further reduce
centrifugal stress near the root.
The S-bowed stacking axis thusly permits centrifugal loads developed
during operation to slightly straighten the airfoil and introduce local
compressive bending stress which locally offsets centrifugal tensile stress.
Accordingly, the preferentially bowed airfoil reduces flow separation
at the hub, and is limited only by the degree of stacking axis bow which
may be introduced with acceptable bending stresses during operation. The
outboard second bow permits the inboard first bow to incline greater than it
otherwise could. Improved hub airflow increases airfoil efficiency without
compromising stall margin, both within acceptable stress limits.
Aerodynamic sweep is a conventional parameter for evaluating
performance of a compressor airfoil. Aft sweep may be limited by
configuring the airfoil leading edge 26 to have an axially coplanar radially
outer or outboard portion which includes the tip 24 as illustrated in Figure 3.
And, the remaining radially inner or inboard portion of the leading edge 26 is
inclined axially forwardly to the root 22 from the outboard portion.
Figure 3 illustrates an axial projection of the airfoil 12 from its suction
side 20 and shows a straight leading edge outboard portion which is
preferably positioned at a constant axial location. The inboard portion of the
leading edge 26 leans forward as the airfoil root is approached relative to the
radial line illustrated in phantom. Aerodynamic aft sweep of the airfoil is
thusly limited at the leading edge from the root to the tip of the airfoil.
Aft aerodynamic sweep may be further limited by preferentially
configuring the airfoil trailing edge 28 as illustrated in Figure 3. The axial
stacking axis 36b in conjunction with corresponding chord lengths may be
used to control trailing edge configuration. In a preferred embodiment, the
trailing edge 28 has an axially coplanar inboard portion including the root 22,
and an outboard portion inclined axially forwardly to the tip 24 from the
inboard portion.
Since the stacking axis includes both tangential and axial
components, the tangential component may be used to advantage to
introduce the bowed suction side 20 near the trailing edge at the root as
illustrated in Figures 1 and 2 for the advantages described above.
Correspondingly, the axial component of the stacking axis may be selected
for limiting the aft sweep along both the leading and trailing edges 26,28 as
illustrated in Figure 3. The stacking axis is configured in conjunction with
the shapes of the individual transverse sections of the airfoil including the
distribution in length of the chords 30 and the camber of the camber lines
32.
Accordingly, the two components of the stacking axis and the shape
of the airfoil transverse sections may be additionally configured based on 3D
viscous flow analysis to increase both airfoil efficiency and stall margin,
while controlling centrifugal stress, resulting in the distinctive 3D
configuration illustrated in the figures.
The degree of suction side bow and S-stack may be adjusted in
different combinations for different airfoil configurations to vary the benefits
of increased aerodynamic performance and reduced centrifugal stress. The
resulting airfoil 12 may thusly be designed for truly three dimensional
performance attributable to modern advances in computational analysis
which makes such improvements possible.
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.
Accordingly, what is desired to be secured Letters Patent of the
United States is the invention as defined and differentiated in the following
claims in which we claim:
WE CLAIM
1. A compressor airfoil 12 for a rotor disk 14 having axial, tangential,
and radial orthogonal axes, comprising:
pressure and suction sides 18, 20 extending radially from root 22 to
tip 24, and axially between leading and trailing edges 26,28;
transverse sections having respective chords and camber lines
extending between said leading and trailing edges, and centers of gravity 34
aligned in a double bowed stacking axis 36; and
said suction side 20 being bowed along said trailing edge 28 adjacent
said root 22 for reducing flow separation thereat.
2. An airfoil according to claim 1 wherein said stacking axis comprises
two orthogonal components including a tangential stacking axis 36a and an
axial stacking axis 36b, and said tangential stacking axis is bowed adjacent
said airfoil root 22 to bow said suction side 20 thereat.
3. An airfoil according to claim 2 wherein said tangential stacking axis
36a includes a first bow 38 having an initial lean onward from said root 22
toward said pressure side 18, and a second bow 40 joining said first bow
and leaning hindward toward said suction side 20 adjacent said tip 24, and
stagger of said sections adjacent said root varies to bow said suction side
thereat.
4. An airfoil according to claim 3 wherein said onward lean is in the
direction of rotation of said airfoil atop said disk 14, and said hindward lean
is opposite to said direction of rotation.
5. An airfoil according to claim 3 wherein said trailing edge 28 is
oriented substantially normal to said root at said bowed suction side 20, and
leans hindward thereabove.
6. An airfoil according to claim 3 wherein said first and second bows
38,40 are disposed on opposite sides of said radial axis extending through
said airfoil root 22.
7. An airfoil according to claim 3 wherein said stagger increases from
root to tip.
8. An airfoil according to claim 3 wherein said tangential stacking axis
has a generally S-shape from root to tip.
9. An airfoil according to claim 3 wherein said trailing edge 28 has a
generally S-shape from root to tip.
10. A compressor airfoil 12 for a rotor disk 14 having axial, tangential,
and radial orthogonal axes, comprising:
pressure and suction sides 18,20 extending radially from root 22 to
tip 24, and axially between leading and trailing edges 26,28;
transverse sections having respective chords and camber lines
extending between said leading and trailing edges, and centers of gravity 34
aligned in a bowed stacking axis 36;
said suction side 20 being bowed along said trailing edge 28 adjacent
said root 22 for reducing flow separation thereat; and
wherein said stacking axis has two orthogonal components including
a tangential stacking axis 36a and an axial stacking axis 36b, and said
tangential stacking axis is double bowed to bow said suction side along said
trailing edge at said root 22.
11. A compressor rotor airfoil 12 comprising a double bowed stacking
axis 36, and a suction side bowed along a trailing edge 28 adjacent a root
22 for reducing flow separation thereat.
12. An airfoil according to claim 11 wherein said stacking axis has a
generally S-shape, and said trailing edge has a generally S-shape.
13. An airfoil according to claim 12 wherein said trailing edge 28 is
oriented substantially normal to said root at said bowed suction side.
14. An airfoil according to claim 13 wherein said trailing edge 28 leans
from said bowed suction side to a tip of said airfoil.

A compressor airfoil (12) includes pressure and suction sides (18,20)
extending from root (22) to tip (24) and between leading and trailing edges
(26,28). Transverse sections have respective chords and camber lines.
Centers of gravity (34) of the sections are aligned along a double bowed
stacking axis for improving performance.