Tip-shrouded blades and method of manufacture

Taught is an improved tip-shrouded blade and a method of manufacturing the same. In accordance with the present invention, the inventive blade comprises a blade including an airfoil having a residual compressive stress in areas thereof subject to high local stress.

The present invention relates generally to turbomachines such as gas 
turbine engines and specifically to an improved tip-shrouded blade and a 
method of manufacturing the same. 
BACKGROUND OF THE INVENTION 
The airfoils of a gas turbine engine, as well as other engine component 
parts, are subject to vibrational stress that can adversely affect the 
lifetime thereof as well as overall engine efficiency. The gas turbine 
industry, in consideration of this stress, has developed numerous ways of 
restraining or damping airfoil vibrations. One approach that has been 
taken involves frictionally damping particular modes of airfoil vibrations 
by interlocking the tips of tip-shrouded blades. To create the interlock, 
the blades are twisted during assembly of the engine. The applied twist is 
in a circumferential direction as viewed along the longitudinal axis of 
each blade. Such an assembly technique is taught in U.S. Pat. No. 
3,185,441 to Reuter. It is now, in the art to configure the interlock 
between adjacent tip shrouds substantially as a "Z" thereby providing 
frictional damping only along the diagonals of the "Z" rather than along 
the entire lengths of the adjacent tip shroud edges. 
Twisting the blades creates a load on the tip shrouds that is normal to the 
plane of contact between adjacent shrouds. This particular load, hereafter 
referred to as an elastic load, arises from the inherent tendency of the 
blade to rebound to its former shape. In other words, the blade tries to 
untwist. 
The elastic load causes the tip shrouds to rub against each other during 
engine operation and thereby dissipates the airfoil vibrations at least in 
part. The elastic load may be insufficient in and of itself, however, to 
provide the total desired blade damping. 
Another source of tip shroud load is available for rotating blades, 
however. During engine operation the centrifugal forces generated by the 
rotation will create an additional load on the tip shrouds. In other 
words, as the centrifugal forces act on the blade, the blade tries to 
lengthen and does so by trying to untwist. Thus, this rotationally 
generated, or centrifugal, load also lies normal to the plane of contact 
between the tip shrouds and adds to the load created by the inherently 
elastic blade material trying to return to its former, untwisted shape. 
The total tip shroud load, centrifugal and elastic, in many cases, may be 
sufficient to provide the desired level of blade damping. In such cases 
the centrifugal and elastic loads may be considered to be inversely 
proportional to each other. That is, for a given constant total load, as 
the centrifugal load is increased, the elastic load may correspondingly be 
decreased. The faster the blade rotates, in other words, the greater the 
centrifugal load is and the less the elastic load need be. Conversely, the 
slower the blade rotates, the smaller the centrifugal load is and the 
greater the elastic load must be to maintain the constant total load. 
Typically, then, as an airfoil rotates more and more slowly at normal 
operating conditions, it must be subjected to ever increasing degrees of 
assembly twist to provide the desired damping characteristics. 
With many features of gas turbine engine technology trade offs exist with 
many design features--i.e., solving one problem may create others with the 
solution--and so it is with twisting blades during assembly. Twisting the 
blades can introduce high stress levels in certain localized regions of 
the airfoil lying adjacent the tip shroud and the blade root. These areas 
of high stress are subject to the development of cracks during engine 
operation and thus create a potential failure mode for the blade. In other 
words, the need to twist the blades to create the load necessary for 
proper damping of the blade during engine operation stresses the blade 
material structure so that cracks can develop more quickly than they would 
without the twist. In turn, the blade cracking can lead to blade failure 
with the resultant loss of the blade itself as well as downstream blades 
which are struck by portions of the disintegrating blade. 
The degree of twist is directly related to the amount of resulting stress. 
Thus, the greater the twist is, the greater the resulting stress will be. 
Where the degree of twist is great due to the need to compensate for the 
loss of the centrifugal load due to a slowly rotating engine, the stress 
will correspondingly be greater as will the potential for failure of the 
blade itself and for other engine components. It would be desirable, 
therefore, to provide a blade having a reduced failure mode due to high 
local stresses introduced by twisting of the blade during engine assembly 
while still retaining the damping ability provided by interlocking tip 
shrouds. 
OBJECTS OF THE INVENTION 
It is a principle object of the present invention to provide new and 
improved apparatus and a new and improved method of manufacturing the same 
that is not subject to the foregoing disadvantages. 
It is another object of the present invention to increase the lifetime and 
integrity of gas turbine engine blades. 
It is yet another object of the present invention to reduce the high local 
stress created in certain areas of a gas turbine engine blade by the twist 
imposed on the blade during engine assembly. 
It is still another object of the present invention to provide a gas 
turbine engine blade, and a method of manufacturing the same, that has a 
reduced failure mode. 
SUMMARY OF THE PRESENT INVENTION 
The foregoing objects of the present invention are achieved by providing a 
tip-shrouded blade including a residual compressive stress in airfoil 
areas subject to large stresses during engine operation, and a method of 
manufacturing such a blade. The residual compressive stress may be created 
by twisting the airfoil beyond the elastic yield point of the airfoil 
material during manufacture of the blade. 
These and other objects of the present invention, as well as further 
features thereof, will become apparent from a reading of the following 
detailed description of the present invention in conjunction with the 
accompanying drawings, all of which are intended to be typical of, rather 
than in any way limiting on, the scope of the present invention.

DETAILED DESCRIPTION OF THE INVENTION 
While gas turbine engines are well known in the art, a brief description of 
the operation of such an engine will enhance appreciation of the 
interrelationship of the various components by way of background for the 
invention to be described below. Furthermore, while many different types 
of gas turbine engines exist, the present invention will be described in 
relation to its application to a particular type, it being recognized, of 
course, that the present invention could equally well be utilized in other 
types of gas turbine engines. 
Thus, there is depicted in FIG. 1 a gas turbine engine 10 of the type 
utilizing aft mounted counter rotating fans to provide thrust. Engine 10 
defines an annular flow path extending from an engine air inlet 12 to an 
engine exhaust nozzle 14 and comprises a gas generator 20 effective for 
producing a high energy gas stream, generally indicated by arrow 30, and a 
propulsor 40 for extracting energy from the gas stream to provide thrust. 
Gas generator 20 includes in an axial flow arrangement a compressor 22 for 
compressing air flowing into engine 10 through engine air inlet 12, a 
combustor 24 where fuel is mixed with the compressed air stream and 
ignited, thereby producing the high energy gas stream, and a turbine 26 
for extracting a portion of the energy of the gas stream to drive 
compressor 22. 
Further included within engine 10 is propulsor 40, which extracts 
additional energy from the gas stream produced by gas generator 20 by 
means of a relatively slowly rotating power turbine 50. The power turbine 
50 in turn drives forward and aft fans 42 and 46, respectively, that are 
circumferentially mounted on engine Forward fan 42 includes individual fan 
blades 44, while rear fan 46 includes individual fan blades 48. The 
forward and rear fans counter rotate, that is, one fan turns in a 
clockwise fashion and the other turns in a counter clockwise fashion. The 
fans are driven by a pair of turbine rotors that also counter rotate. 
Forward fan 42 is attached to a first rotor 52 of a first turbine. 
Likewise, rear fan 46 is attached to a second rotor 56 of a second 
turbine. First and second turbine rotors 52 and 56 each include a 
plurality of turbine blades, 54 and 58, respectively, circumferentially 
attached thereto. 
FIG. 2 illustrates a tip-shrouded turbine blade 70 of the type in which the 
present invention may find application. Blade 70 has a longitudinal axis 
90 and includes an airfoil 71 defined in part by an airfoil tip 72 and an 
airfoil root 73 disposed at opposite ends of airfoil 71. Airfoil 71 
extends between a tip shroud 74 and a blade root 75 and is attached to tip 
shroud 74 at airfoil tip 72 and to blade root 75 at airfoil root 73. Tip 
shroud 74 may include a seal 76. Blade root 75 typically includes means 
for attaching blade 70 to engine 10. As shown in FIG. 2, the means for 
attaching include hooks 79. Other means for attaching blades to a rotor 
are well known in the art and such means are not critical with respect to 
the present invention. Blade 70 is further defined by a leading or 
upstream edge 77 and a trailing or downstream edge 78 of airfoil 71. 
When a blade such as blade 70 is installed in a gas turbine engine, a 
circumferential twist is applied to the blade such that tip shroud 74 will 
interlock with the adjacent blade tip shroud. The circumferential 
direction is determined by the longitudinal axis 90 of blade 70 and can 
either be a clockwise direction or counterclockwise direction, as 
indicated by arrows 91 and 92, respectively, in FIG. 2. 
FIG. 3 shows in plan view a plurality of interlocked tip shrouds 74a 
through 74f, each tip shroud being viewed from the perspective of lines 
III--III in FIG. 2. The figure clearly indicates the substantially "Z" 
shaped configuration of the interlock. As indicated on shrouds 74a and 74b 
by arrows 85 and 86, respectively, the elastic and centrifugal loads act 
normal to the plane of contact between the shrouds, which is the diagonal 
of the "Z" interlock. Though a gap is indicated in the drawing at the 
location of the diagonal, it will be understood that such is for the 
purpose of clearly indicating individual tip shrouds only and that the 
shrouds actually are in contact along the surface of the "Z" interlock. 
The actual location of the high stress areas that develop as a result of 
the assembly twist will vary depending upon blade geometry and blade 
construction. In general, however, the stresses will be localized near the 
airfoil root and tip and near the leading and trailing edges or, in other 
words, at the corners of the airfoil. FIGS. 4a and 4b illustrate the 
localized stress that may be experienced by an all metal blade subjected 
to an assembly twist of three degrees. Thus, FIG. 4a depicts a prior art 
airfoil generally illustrating possible stress contour lines of a blade. 
FIG. 4b shows in greater detail the possible stress contour lines of the 
trailing edge of such a blade near the airfoil tip. The stress levels are 
indicated by the scale at the right hand side of the Figure. As 
illustrated, the stress contour lines indicate a high degree of localized 
stress is present. Such high stress levels are undesirable for the 
aforementioned reasons. 
As is well known, materials exhibit both elastic and plastic qualities 
depending upon the stress to which they are subjected. The point at which 
the material begins to exhibit plastic properties is known as the 
material's yield point. By yielding a material, that is, by stressing a 
material beyond its yield point and then releasing the stress, a residual 
compressive, or negative, stress due to yielding and torsion may remain in 
those areas subject to high stresses in normal operation. Such a yielding 
operation must be carefully controlled to avoid unacceptable damage to the 
material and, consequently, requires the yield point of the material to be 
known ahead of time. 
By yielding a blade such as blade 70, i.e., by twisting the blade past its 
known yield point and then letting it untwist, a residual compressive, or 
negative, stress will remain in the blade in the areas subject to high 
stresses resulting from the twist. When the positive stresses resulting 
from twisting the blade during assembly are added to the residual or 
negative stress resulting from yielding and torsion, a reduced total 
stress level results. Thus when a pre-yielded blade is installed in a gas 
turbine engine the total stresses resulting from the twist required to 
load the tip shrouds will be less than the twist stresses of a blade that 
has not been pre-yielded. By pre-yielded, it is meant that the blade has 
been twisted beyond its yield point in the same direction as it will be 
twisted when installed in the engine. 
Thus in accordance with the present invention a blade such as blade 70 will 
be pre-yielded during the manufacturing operation by twisting the blade 
past the blade's yield point. The reduced stress level which may be 
expected as a result of this novel manufacturing step is shown in FIG. 5a 
and FIG. 5b, each of which illustrates the reduced localized stresses of a 
blade that has been pre-yielded. While the stress contour lines remain 
substantially identical to those of FIGS. 4a and 4b, the stress levels are 
greatly reduced as can be seen comparing the scales of FIGS. 4a and 4b 
with those of FIGS. 5a and 5b. 
The desired amount of twist beyond the yield point will depend on the 
materials used to manufacture the blade and on the particular blade 
geometry. For example, for a blade manufactured at least in part of a 
nickel-based alloy, the amount of twist should be at least two-tenths of 
one percent (0.2%) beyond the yield point. The step of pre-yielding a 
blade can occur at several stages in the manufacturing process. Thus, a 
blade may be pre-yielded, for example, between the steps of casting and 
machining the blade or after the blade is machined and finished but not 
before the final heat treatment. 
The pre-yielding of the blades can be accomplished by different kinds of 
apparatus, none of which is critical to the present invention. For 
example, means comprising a first and a second clamp that securely holds a 
blade by the blade tip and the blade root, respectively, one clamp being 
relatively circumferentially rotatable with respect to the other clamp 
could be used to pre-yield a blade. Thus a blade would be securely clamped 
at its root and tip ends, and one end would be rotated circumferentially 
with respect to the other end wherein the circumferential direction is 
determined with respect to the longitudinal axis of the blade. 
While the present invention has been described in relation to a blade 
operating in a relatively low speed turbine, it should be apparent to 
those skilled in the art that the present invention would have application 
with regard to any tip shrouded blade that requires a loaded tip shroud. 
Furthermore, the present invention is not dependent upon the particular 
method by which the blade is attached to the engine rotor, nor on the 
particular basic manufacturing process of the blade itself. Furthermore, 
while the present invention has been described in relation to gas turbine 
engines in general, it is equally applicable to any alloyed or 
substantially single material structure that is subjected to high stress. 
Having thus described the present invention, additional numerous changes, 
substitutions, modifications and alterations will now suggest themselves 
to those skilled in the art, all of which fall within the spirit and scope 
of the present invention. Accordingly, it is intended that the invention 
be limited only by the scope of the appended claims.