Hollow airfoil impact resistance improvement

Hollow airfoils, such as fan blades, nozzles and struts used in axial flow gas turbine engines, that have improved resistance to impact from foreign objects are disclosed. Relocation of airfoil material is used to preferentially strengthen the airfoil to respond to the stress from the impact of a foreign object. Internal spacers are redistributed toward the leading edge and the material from the skin of the concave side and the convex side is shifted from one side to the other and toward the leading edge of the airfoil where impact stress is greatest.

The present invention relates, in general, to hollow core airfoil 
structures and more particularly to hollow core airfoils including 
structures, such as fan blades, nozzles and struts for axial flow gas 
turbine engines, that are preferentially strengthened to resist the impact 
from a foreign object. 
BACKGROUND OF THE INVENTION 
Modern high bypass turbofan engines incorporate various types of airfoils 
such as fan blades, nozzles and struts. In particular, wide chord fan 
blades improve aerodynamic efficiency and improve tolerance to impact from 
foreign objects. To minimize system weight, airfoils and their root 
attachments (dovetails) are often made hollow by joining premachined 
sections together into a final assembly. Common internal core 
constructions rely on the use of radial ribs or trusses evenly spaced 
within the cavity. These features serve as the support structures for the 
outer skins, and act to carry centrifugal loading, support the skins 
against gas pressure and foreign object impact loading, such as from 
birds, and provide stiffness. 
Strength requirements in some components, particularly airfoils in the 
early stages of gas turbine engines, vary across the part. For instance, 
impact strength is a primary concern on the forward edge of the airfoil. 
Equally spaced rib or truss members provide uniform strength across the 
blade even though strength requirements are not uniform. Certain airfoil 
areas would benefit by more support. Adding more ribs or trusses or 
increasing skin thickness accomplishes this effect, but at a substantial 
weight penalty. 
In a gas turbine engine, added weight in blades necessitates added weight 
in other hardware, including the disk, stationary structures, and the 
containment system. This in turn, increases fuel burn and customer cost. 
More ribs or trusses also add to machining time, tooling costs, and 
inspection time, which further drives up production costs. 
Accordingly, the present invention provides a new arrangement of internal 
structural members that improves the impact strength of hollow airfoil 
components subjected to impact from foreign objects, such as birds, while 
maintaining existing aerodynamic function of the airfoil, minimizing 
airfoil weight and minimizing machining and inspection time. 
SUMMARY OF THE INVENTION 
In carrying out this invention, in one form thereof, a preferentially 
strengthened hollow airfoil that has a flow axis extending from a forward 
location to an aft location is provided for a gas turbine engine. The 
airfoil includes a root section located at the base of the airfoil; a tip 
section located distally from the root section, a leading edge connecting 
the root section and the tip section and facing forward along the flow 
axis; a trailing edge connecting the root section and the tip section and 
facing aft along the flow axis; an outer skin that extends between and 
connects the root section, the tip section, the leading edge and the 
trailing edge and forms an outer surface of the airfoil; and one or more 
spacers located between the leading edge and the trailing edge and 
enclosed by the skin and forming a plurality of non-uniformly sized 
cavities. 
Based on nonlinear transient dynamic analysis of an airfoil application, 
preferential strengthening is accomplished by rearranging one or more 
spacers between the leading edge and the trailing edge of the airfoil and 
by varying skin thickness on one side of the airfoil relative to the other 
side and along the length of each side to form a plurality of 
non-uniformly sized cavities. The location of each spacer is chosen to 
minimize the strains due to initial impact shock, and later to the 
transient response that occurs from a foreign object impact. The airfoil 
thickness or exterior shape is not changed. Preferential strengthening is 
further enhanced by preferential thickness control of the skin material 
without increasing weight. The skin material is rearranged in response to 
the predicted stress caused by the impact from a foreign object. For 
example, in one application for a fan blade of a gas turbine engine, the 
convex side skin thickness of the fan airfoil would be increased while the 
concave side skin thickness of the fan airfoil would be decreased. 
Additional skin thickness on the convex side is needed to prevent 
compressive buckling while a thinner concave side could survive the 
increased tension generated by a foreign object impact. Also, in this 
example, the skin thickness of both sides near the leading edge of the 
airfoil would be preferentially increased further to withstand the direct 
impact of a foreign object. Even though the convex side is generally 
thicker than the concave side, the skin thickness of the leading edge of 
both sides is preferentially increased at the expense of the skin 
thickness at the trailing edge of the blade thereby improving impact 
resistance without increasing weight. 
In one embodiment of this invention, the airfoil is in the form of a fan 
blade and has ten cavities of varying width as measured along the chord of 
the airfoil. Starting at the leading edge of the airfoil, the first cavity 
is the narrowest and the next four cavities are slightly wider in 
succession. The sixth cavity through the tenth cavity each have the same 
width which is the widest cavity dimension and designated as width L. The 
first cavity width is generally between 0.45 L and 0.55 L inclusive, the 
second cavity width is generally between 0.55 L and 0.65 inclusive, the 
third cavity width is generally between 0.65 L and 0.75 L inclusive, the 
fourth cavity is generally between 0.75 L and 0.85 L inclusive and the 
fifth cavity is generally between 0.85 L and 0.95 L inclusive. The 
decreased width of the cavities toward the leading edge of the airfoil is 
advantageous. The airfoil will be preferentially strengthened along the 
leading edge and will be able to resist impact from a foreign object, such 
as a bird ingested by a gas turbine engine. Also, by shifting the spacing 
between support members, the leading edge strength will be increased 
without increasing the overall weight of the airfoil. 
The support members, which serve as the spacers, separating the concave 
side and the convex side can be shaped several ways while still performing 
the function of supporting the outer skin in an airfoil shape. Most 
commonly, the support members are ribs, lands, stringers or trusses. Each 
structure will accommodate varied cavity size. There are several 
manufacturing processes and each is generally indicative of which 
structure to use. For example, a two piece diffusion bonding process would 
indicate ribs while a three piece diffusion bonding process would indicate 
a truss. 
In another embodiment of the present invention, either separately or in 
combination with the first embodiment, the skin thickness, T, of the 
airfoil is varied to improve the airfoil's strength distribution. The skin 
thickness is preferentially increased on the convex side and 
correspondingly decreased on the concave side which improves airfoil 
resistance to impact without adding weight. Additionally, the skin 
thickness is preferentially increased along the cavities near the leading 
edge of the airfoil to further improve airfoil strength in the leading 
edge area where an impact from a foreign object is most likely to occur. 
The skin thickness on the convex side is generally B inches thicker than 
the corresponding skin thickness on the concave side, where B is between 
0.005 inches and 0.025 inches. The skin thickness of the cavities nearest 
to the leading edge of the airfoil is preferentially increased between 40% 
and 60% for the first cavity closest to the leading edge, between 20% and 
40% for the second cavity from the leading edge and between 0% and 20% for 
the third cavity from the leading edge while typical increases in skin 
thickness are 50%, 30% and 10% for the first, second and third cavities 
from the leading edge respectively. 
In new airfoil designs, the positioning of internal support members for 
hollow airfoils and the choice of skin thickness distribution is done 
while minimizing airfoil weight and without altering aerodynamic or 
acoustic performance. 
When applied to an existing hollow airfoil design, the present invention 
applies to internal changes and can be understood as a redistribution of 
existing material to locations to preferentially increase the strength of 
the airfoil to match the calculated stress from the impact with a foreign 
object. Additional material is not added to the airfoil, therefore there 
is no increase in weight. 
In either an existing or a new hollow airfoil design, the positioning of 
internal support members and the choice of skin thickness distribution may 
also vary radially between the root section and the tip section and 
axially between the leading edge and the trailing edge. Each airfoil 
application has different strength requirements. The present invention 
contemplates redistributing the airfoil material in all dimensions of the 
airfoil to achieve the optimum resistance to the impact from a foreign 
object without compromising requirements for strength and life.

DESCRIPTION OF THE PREFERRED EMBODIMENTS 
Referring now to the Figures wherein like reference numerals have been used 
throughout to designate like parts. FIG. 1 illustrates a typical 
cross-sectional view of an axially symmetric gas turbine engine 10 having 
a nacelle 12 with an inlet 14 located forward along flow axis 16 of the 
gas turbine engine 10. The inlet 14 controls airflow 18 toward a plurality 
of hollow fan blades 20 that are encircled by nacelle 12 and are connected 
to gas turbine engine 10 at fan rotor 21. Fan blades 20 rotate about flow 
axis 16 and are vulnerable to being struck by any foreign object 22 
entering inlet 14. Generally, foreign object 22 travels parallel to flow 
axis 16, however foreign object 22 may enter inlet 14 at an angle 17 that 
is not parallel to flow axis 16. After entering inlet 14, foreign object 
22 will strike fan blade 20 along leading edge 26 at a relative velocity 
equivalent to the vector sum of fan blade velocity and foreign object 22 
velocity which can be as high as 500 mph for commercial applications and 
mach 3 or higher for military applications. 
Foreign object 22 in the normal operation of a gas turbine engine 10 ranges 
from small particles, for example sand, to large objects, for example, 
birds. The smaller foreign objects 22 generally do not cause major damage 
to the gas turbine engine or reduce its performance because they do not 
have much energy. However, a large bird has a much greater energy content 
and can cause significant damage. At a minimum, the ingestion of a large 
foreign object may damage fan blades 20 which may result in rotational 
imbalance. 
FIG. 2 illustrates a cutaway view of a preferred embodiment of the present 
invention configured as a hollow fan blade 20 with ribs 38 spaced a 
non-uniform distance apart. Fan blade 20 has a root section 28, located 
adjacent to base 30 of fan blade 20. Blade tip 24 is radially distal from 
root section 28 and is connected to root section 28 by leading edge 26 and 
trailing edge 32. Outer skin 34 connects root section 28, blade tip 24, 
leading edge 26 and trailing edge 32 and forms outer surface 36 of hollow 
fan blade 20. Leading edge 26 generally faces forward and trailing edge 32 
generally faces aft. Root section 28 supports and connects hollow fan 
blade 20 to gas turbine engine 10 (FIG. 1) at fan rotor 21. 
Outer skin 34 encloses a plurality of ribs 38 that serve as spacers to 
separate a plurality of cavities 40a through 40j. Cavities 40a through 40j 
have non-uniform widths 42a through 42j. Each of widths 42a through 42j of 
corresponding cavities 40a through 40j is, according to the present 
invention, narrower adjacent leading edge 26 than adjacent trailing edge 
32. Rib thickness 44 is generally uniform for each rib 38; however, the 
present invention contemplates that thickness 44 may vary as needed to 
achieve the desired strength distribution of fan blade 20. 
FIG. 3 illustrates a cross section of a hollow airfoil 20 of conventional 
design. Skin 34 has uniform thickness T on concave side 48 and on convex 
side 50. Cavities 40a through 40j have equal widths 42a through 42j and 
are uniformly spaced between leading edge 26 and trailing edge 32 along 
chord length 52. Likewise, rib thickness 44 is uniform for each rib 38 
that separates cavities 40a through 40j. 
The present invention improves the impact resistance of the conventional 
airfoil shown in FIG. 3 by rearranging the location, spacing and size of 
cavities 40a through 40j to preferentially stiffen the airfoil and by 
adjusting skin thickness, T, without adding weight or changing its 
aerodynamic or acoustic performance. 
The preferred embodiment of the present invention is again illustrated in 
FIG. 4. Fan blade 20 is illustrated; however any airfoil susceptible to 
impact, whether stationary or moving, would benefit from the present 
invention and is contemplated herein. 
Fan blade 20 illustrated in FIG. 4, has ten (10) cavities 40a through 40j 
of non-uniform width. In another embodiment of the present invention there 
may be a different number of cavities. Cavity 40a has the narrowest width 
42a and is located adjacent leading edge 26. Succeeding cavities 40b 
through 40j have widths 42b through 42j respectively that are greater than 
width 42a. Widths 42b through 42e are successively larger than width 42a. 
Widths 42f through 42j are equal for cavities 40f through 40j in the 
present embodiment; however, a non-uniform width is contemplated for these 
cavities in alternate embodiments of the present invention and would be 
determined by the specific airfoil application. Ribs 38 separate cavities 
40a through 40j and generally have a uniform width 44; however a 
non-uniform width is contemplated for ribs 38 in any alternate embodiment 
of the present invention. Cavity widths 42a through 42j are generally 
described by the following equation: 
EQU 0.4L.ltoreq.x&lt;L 
where, x, represents cavity width and L is the width of the widest cavity. 
Preferential strengthening of fan blade 20 is further enhanced by adjusting 
skin thickness, T, from FIG. 3 to accommodate the tension and compression 
stresses on concave side 48 and convex side 50 resulting from a foreign 
object 22 impact on leading portion 39. Care is taken to thicken convex 
skin 68 while thinning concave skin 70 of FIG. 4 Without increasing weight 
of fan blade 20. The concave skin thickness is generally represented as 
being AT inches and the convex side skin thickness is generally 
represented as being (AT+B) inches. B, as contemplated in the present 
invention, is between 0.005 inches and 0.025 inches and A is between 1 and 
1.6. In an alternate embodiment where the tension and compression stresses 
may be reversed, concave skin 70 may be thickened while thinning convex 
skin 68 in similar fashion. 
Preferential strengthening of fan blade 20 is further enhanced by further 
increasing the skin thickness for a selected number of cavities 40a, 40b 
and 40c, along leading portion 39 of fan blade 20. The additional 
thickening is incrementally increased for the selected cavities on both 
the concave and convex sides related to the uniform skin thickness for the 
cavities in trailing portion 72. Depending on the airfoil application, 
skin thickness may also vary radially. 
Concave skin 70 and convex skin 68 for leading portion 39, cavities 40a, 
40b and 40c, both have an increased skin thickness by a factor, A, 
relative to their uniform skin thickness, T and T+B, respectively. In the 
present invention A is between 1.0 and 1.6 where A for cavities 40a, 40b, 
and 40c is between 1.4 and 1.6 inclusive, between 1.2 and 1.4 inclusive 
and between 1.0 and 1.2 inclusive, respectively. A, for cavities 40d 
through 40j, is 1.0. Generally, concave skin thickness is AT and convex 
skin thickness is AT+B. In alternate embodiments, leading portion 39 of 
fan blade 20 may encompass a different number of cavities from the three 
shown in FIG. 3 and the value of A may vary radially and along chord 
length 52 depending on the application. 
At a given radial location along fan blade 20, cavity length and skin 
thickness, for the preferred embodiment illustrated in FIG. 4, can best be 
described by referring to Table 1. The size relationship among cavities 
and skin thicknesses may vary radially depending on airfoil application. A 
table similar to Table I may be generated for each radial cross section of 
a given airfoil for any given application. 
TABLE 1 
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Concave Side 
CAVITY WIDTH 
Convex Side Skin Thickness 
Skin Thickness 
.4 L .ltoreq. x .ltoreq. L 
T + .005 .ltoreq. y .ltoreq. 1.6 T + .025 
T .ltoreq. Z .ltoreq. 1.6 T 
CAVITY 
(X) (Y) (Z) 
__________________________________________________________________________ 
40a .4 L .ltoreq. x .ltoreq. .6 L 
1.4 T + .005 .ltoreq. y .ltoreq. 1.6 T + .025 
1.4 T .ltoreq. Z .ltoreq. 1.6 T 
40b .5 L .ltoreq. x .ltoreq. .7 L 
1.2 T + .005 .ltoreq. y .ltoreq. 1.4 T + .025 
1.2 T .ltoreq. Z .ltoreq. 1.4 T 
40c .6 L .ltoreq. x .ltoreq. .8 L 
T + .005 .ltoreq. y .ltoreq. 1.2 T + .025 
T &lt; Z .ltoreq. 1.2 T 
40d .7 .ltoreq. x .ltoreq. .9 L 
T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
40e .8 L .ltoreq. x .ltoreq. L 
T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
40f x = L T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
40g x = L T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
40h x = L T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
40i x = L T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
40j x = L T + .005 .ltoreq. y .ltoreq. T + .025 
Z = T 
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In Table 1 all dimensions are in inches, X is cavity width, Y is convex 
skin thickness, Z is concave skin thickness, L is the width of the widest 
cavity which usually is cavity 40j closest to trailing edge 32 and T is 
the thinnest skin thickness which usually is the concave side of cavity 
40j closest to trailing edge 32. L, T and the number of cavities are 
parameters that are application dependent and are selected by the designer 
to minimize airfoil weight while maintaining all functional requirements. 
In an alternate embodiment of the present invention, as illustrated in FIG. 
5, a truss 74 replaces ribs 38, illustrated in FIG. 4. Referring again to 
FIG. 5, the widths 42a through 42j of cavities 40a through 40j between 
portions of the truss that serve as spacers are measured along chord 
length 52 that bisects fan blade 20 and extends between leading edge 26 
and trailing edge 32. Cavity 40a is the shortest and is located adjacent 
leading edge 26 and cavities 40b through 40j have cavity width 42b through 
42j that increase in similar fashion to cavities 40b through 40j 
illustrated in FIG. 4. In FIG. 5, skin thickness, AT, of concave side 48 
is thinner than skin thickness, AT+B, of convex side 50. The skin 
thickness of leading portion 39 cavities 40a, 40b and 40c are further 
thickened on concave side 48 and on convex side 50 relative to skin 
thickness of trailing portion 72 cavities 40d through 40j. Concave side 
skin thickness AT and convex side skin thickness AT+B is likewise fully 
described in Table 1. 
The present invention has been described herein by way of example and is 
not intended to limit the scope of the invention claimed to the specific 
examples given. It is to be understood that the cavity length, skin 
thickness, spacer thickness such as ribs or trusses and the overall 
distribution of airfoil material can be varied beyond the specific limits 
given without exceeding the scope and intent of the present invention. 
Accordingly, the invention as anticipated by the inventors is limited only 
by the following; wherein,