Methods for fabricating a helicopter main rotor blade

A method for fabricating a helicopter blade subassembly includes the steps of providing a blade compaction apparatus having a lower assembly configured for mating with an upper assembly. The lower assembly includes a base having a contoured upper airfoil nest with at least two guide ramps and a plurality of leading-edge pusher cams disposed proximal thereto. The upper assembly includes a support structure supporting a flexible impervious membrane, wherein the flexible impervious membrane and the contoured upper airfoil nest define a molding cavity therebetween upon mating of the upper assembly to the lower assembly. The method further includes disposing the upper airfoil skin and the core in combination with the contoured upper airfoil nest, and then connecting a tip end support insert and a root end support insert in combination with the spar assembly, wherein each of the support inserts include guide members. The spar assembly is located in chordwise alignment in the contoured upper airfoil nest by driving the guide members into conics in the guide ramps and driving the spar assembly into a core conic with the leading-edge pusher cams. The lower airfoil skin is then disposed in combination with the core and the spar to form the helicopter blade subassembly, and then the upper assembly is mated to the lower assembly. The air within the molding cavity is then evacuated such that evacuating the air causes the flexible impervious membrane to apply compaction forces to the helicopter blade subassembly.

RELATED APPLICATION 
The present application is related to commonly-owned, co-pending U.S. 
patent application entitled APATUS FOR INSTALLING A LEADING-EDGE SHEATH 
ONTO A HELICOPTER MAIN ROTOR BLADE SUBASSEMBLY U.S. Ser. No. 09/838,214, 
and to commonly-owned, co-pending U.S. patent application entitled 
APATUS FOR ASSEMBLING A HELICOPTER MAIN ROTOR BLADE SUBASSEMBLY U.S. 
Ser. No. 08/843,527. 
TECHNICAL FIELD 
This invention relates generally to manufacturing methods, and more 
particularly, to methods for fabricating a helicopter main rotor blade. 
BACKGROUND ART 
There is a growing trend in the aerospace industry to expand the use of 
composite materials for a diverse array of structural and dynamic 
applications. One particular application for the use of composite 
materials lies in the fabrication of main rotor blades for helicopters. 
Sikorsky Aircraft has developed a parallel manufacturing protocol for 
fabricating helicopter main rotor blades wherein a blade subassembly and a 
leading-edge sheath are concurrently fabricated as individual components, 
and then the prefabricated blade subassembly and the prefabricated 
leading-edge sheath are integrated in combination to form an assembled 
main rotor blade. The assembled main rotor blade is placed in a clamshell 
and is subsequently cured in an autoclave to form a finished main rotor 
blade. 
In one prior art process for fabricating main rotor blades, the blade 
subassembly portion of the parallel manufacturing protocol begins with a 
composite skin and honeycomb core combination being placed in a plastic 
nest having a contour corresponding to the blade's upper airfoil, and 
manually positioned spanwise and chordwise using a plurality of locator 
pins. A titanium spar coated with adhesive is manually positioned in 
combination with the skin and honeycomb core combination and is seated 
within a conic in the honeycomb core. The honeycomb core is then coated 
with an adhesive and a second composite skin is placed over the honeycomb 
core and a portion of the spar, and is likewise positioned using the 
locator pins. A plastic lid having a contour corresponding to the blade's 
lower airfoil is then placed over the second composite skin and is mated 
in combination with the plastic nest using a plurality of clamps such that 
a compaction force is applied to the blade subassembly. 
The second part of this process for manufacturing main rotor blades 
involves the placement of the leading-edge sheath onto the exposed leading 
edge of the blade subassembly. The leading-edge sheath has a prefabricated 
configuration that does not allow the sheath to be inserted directly onto 
the blade subassembly. Rather, the aft edges of the leading-edge sheath 
must be spread apart to allow the leading-edge sheath to be inserted onto 
the blade subassembly. The prior art sheath spreader tool comprises 
segmented angular stainless steel sheet metal grabbers that are disposed 
spanwise in combination with the aft edges of the leading-edge sheath in 
contact with its inner mold line (IML) surfaces (which are formed of 
composite material). Each segment of the prior art grabber is individually 
actuated by means of a side cam lever, thereby spreading the aft edges of 
the leading-edge sheath. 
Adhesive is applied to the leading edge of the blade subassembly, followed 
by the blade subassembly tool being rotated 90.degree. such that the 
leading edge of the blade subassembly faces up. The sheath spreader tool 
is then hoisted up by a crane and lowered onto the blade subassembly tool 
such that the leading-edge sheath is placed over the leading edge of the 
blade subassembly. The sheath spreader tool is then drawn down toward the 
blade subassembly tool using a plurality of threaded rods until the sheath 
spreader tool engages predetermined tooling stops. This process of 
"drawing down" the sheath spreader tool generates significant stresses on 
the leading-edge sheath. The aft edges of the leading-edge sheath are then 
released by the sheet metal grabbers such that the leading-edge sheath is 
secured onto the leading edge of the blade subassembly. 
A significant drawback to the blade subassembly tool used in the process 
described above is the inadequate compaction provided to the blade 
subassembly by the plurality of clamps operating in combination with the 
plastic nest and lid. Specifically, the compaction forces imparted by the 
tool onto the blade subassembly are uneven over the span of the blade 
since each clamp provides its maximum compaction force on the blade 
subassembly only at its discrete spanwise location. Experience with this 
method and blade subassembly tool has shown that the compaction provided 
by the clamps is such that once the completed blade is removed from the 
blade subassembly tool, the blade has to be placed into the clamshell and 
autoclave within a half hour of compaction such that the entire blade 
assembly doesn't disbond. 
In addition to providing inadequate compaction over the span of the blade 
subassembly, the prior art blade subassembly tool described above also 
provides inadequate chordwise compaction coverage for the blade 
subassembly. Since the blade subassembly tool remains in place about the 
blade subassembly as the sheath spreading tool is being operated, it 
follows that the blade subassembly tool cannot fully extend to the leading 
edge of the blade subassembly. Therefore, the portions of the upper and 
lower airfoil skins located proximal to the leading edge of the blade 
subassembly are not subject to direct compaction, and therefore may not be 
adequately secured to the spar. A disadvantage of this incomplete 
compaction is that upon placement of the leading-edge sheath over the 
leading edge of the blade subassembly, if either of the composite skins 
separates from the spar, the result can be that one or both of the aft 
edges of the leading-edge sheath will slide under the composite skin, 
thereby providing an improper interface between the leading-edge sheath 
and the blade subassembly, resulting in rejection of the assembled blade. 
Another area of concern in the parallel manufacturing protocol described 
above is the sheath spreader tool used to integrate the leading-edge 
sheath with the blade subassembly. The prior art grabbers exert a shearing 
action against the IML surfaces of the leading-edge sheath when spreading 
the aft edges of the leading-edge sheath apart. This shearing action by 
the prior art grabbers can cause cracks and delaminations in the 
leading-edge sheath's composite materials, thereby resulting in component 
rejection and rework. In addition, the operation of the prior art grabbers 
is such that the grabbers can contaminate clean bond surfaces of the 
leading-edge sheath. Also, the segments of the grabber are individually 
actuated in a sequential manner such that multiple repetitive operations 
are necessary to spread apart the entire leading-edge sheath. Not only is 
such a procedure labor intensive, time consuming, and costly, such a 
procedure may induce unwanted stresses into the aft edges of the 
leading-edge sheath. 
Another approach for manufacturing main rotor blades is disclosed in U.S. 
Pat. Nos. 5,528,828, METHODS FOR FABRICATING A HELICOPTER MAIN ROTOR 
BLADE, and 5,570,631, APATUS FOR FABRICATING A HELICOPTER MAIN ROTOR 
BLADE, both assigned to the United Technologies Corporation (hereinafter 
"'828 patent" and "'631 patent" respectively). As depicted in FIGS. 1 and 
2, the apparatus disclosed in the '828 patent and the '631 patent 
comprises a compaction fixture 10 for assembling and compacting a blade 
subassembly 24, and a sheath spreading/insertion apparatus 50 for 
spreading and inserting a leading-edge sheath 22 onto the blade 
subassembly 24 during the compaction process, wherein the blade 
subassembly 24 comprises an upper airfoil skin 12, a lower airfoil skin 
18, a core 14, and a spar assembly 16. 
The compaction fixture 10 includes a lower assembly 26 having a contoured 
upper airfoil nest 28 mounted in combination with a support structure 30, 
and an upper assembly 32 having a pressure bag 34 affixed in sealed 
combination to a contoured backplate 36 affixed in combination to a 
structural support truss 38. The contoured upper airfoil nest 28 includes 
a plurality of tooling pins 31 for locating the upper airfoil skin 12 in 
aligned combination on the contoured upper airfoil nest 28 and a plurality 
of backwall pusher pins 39 for chordwise alignment of the spar assembly 16 
in the contoured upper airfoil nest 28. Spar stanchions 40 affixed to the 
support truss 38 provide spanwise alignment of the spar assembly 16 in the 
contoured upper airfoil nest 28. With the upper and lower assemblies 32, 
26 in locked combination, the pressure bag 34 is pressurized to compact 
the assembled blade subassembly components 12, 14, 16, 18. 
The sheath spreading/insertion apparatus 50 includes a movable stanchion 
52, upper and lower elongate carriage members 54, 56 mounted in 
synchronized movable combination with the stanchion 52, and rows of 
suction cups 58, 60 mounted in combination with the carriage members 54, 
56. Pneumatic cylinders 61, 62 are interposed between the stanchion 52 and 
each of the respective carriage members 54, 56. Pressurization of the 
pneumatic cylinders 61, 62 causes synchronized movement of the upper and 
lower carriage members 54, 56 between a disengaged position wherein the 
leading-edge sheath 22 can be inserted between the upper and lower rows of 
suction cups 58, 60, an engaged position wherein the suction cups 58, 60 
abuttingly engage respective outer mold line (OML) surfaces of the 
leading-edge sheath 22, and an operating position wherein the leading-edge 
sheath 22 is spread apart for insertion onto the blade subassembly 24 
during compaction thereof. A vacuum source 64 is pneumatically 
interconnected to the suction cups 58, 60 to generate suction forces 
therein in the engaged position to cause the suction cups 58, 60 to engage 
the respective OML surfaces of the leading-edge sheath 22 such that 
subsequent synchronized movement of the upper and lower carriage members 
54, 56 to the operating position causes the leading-edge sheath 22 to be 
spread apart. Movement of the movable stanchion 52 causes the spread-apart 
leading-edge sheath 22 to be placed onto the blade subassembly 24 during 
compaction. 
As with the blade subassembly tool described earlier, a drawback to the 
methods and apparatus disclosed for the '828 patent and '631 patent is 
that the compaction fixture 10 does not provide adequate compaction 
between the spar assembly 16 and the portions of the composite skins 12, 
18 located proximal to the leading edge 42 of the blade subassembly 24. As 
illustrated in FIG. 2, since the movable stanchion 52 places the 
leading-edge sheath 22 onto the blade subassembly 24 during compaction, 
there must be enough clearance about the leading edge of the blade 
subassembly 24 to physically allow the movable stanchion 52 to translate 
horizontally into proper position. Therefore, the pressure bag 34 does not 
fully cover the leading edge of the blade subassembly 24, and accordingly, 
the upper and lower airfoil composite skins 12, 18 cannot be properly 
compacted onto the spar assembly 16 in that area. As a result, it has been 
found that the leading edges of these composite skins 12, 18, also known 
as "joggles", can have a tendency to lift-up off the spar assembly 16 in 
this leading edge area. These joggles each define a "step" upon which the 
leading-edge sheath 22 must overlap the corresponding composite skin 12, 
18 in order to have a proper interface between the leading-edge sheath 22 
and the blade subassembly 24. In the event that the joggles lift-up off 
the spar assembly 16, the leading-edge sheath 22 could slide under the 
joggles during installation, thereby forming an improper and unacceptable 
interface between the leading-edge sheath 22 and the blade subassembly 24. 
In addition, drawbacks in the design of the sheath-spreading/insertion 
apparatus 50 disclosed in the '828 patent and '631 patent can further 
increase the possibility of an improper interface between the leading-edge 
sheath 22 and the blade subassembly 24 during installation. Specifically, 
when the leading-edge sheath 22 is positioned between the rows of suction 
cups 58, 60, and the upper and lower carriage members 54, 56 are in the 
operating position, the weight of the leading-edge sheath 22 generates a 
downward force upon the lower row of suction cups 60 and upon the lower 
carriage member 56. It has been found that this downward force on the 
lower carriage member 56 is not sufficiently counteracted by the pressure 
supplied to the lower pneumatic cylinders 62, and therefore, the lower 
carriage member 56 has a tendency to drift downward in response to the 
weight of the leading-edge sheath 22. In addition, since the suction cups 
58, 60 comprise rubber bellows, the weight of the leading-edge sheath 22 
can also collapse the bellows in the lower row of suction cups 60 and 
stretch the bellows in the upper row of suction cups 58. The downward 
forces on the carriage member 56, in combination with the tendency of the 
lower suction cups 60 to collapse, causes the aft edges of the 
leading-edge sheath 22 to skew downward such that a misalignment can occur 
between the leading-edge sheath 22 and the blade subassembly 24. The 
nature of this misalignment can be such that as the stanchion 52 is moved 
horizontally, and as the leading-edge sheath 22 approaches the leading 
edge of the blade subassembly 24, the skew of the leading-edge sheath 22 
can cause the upper aft edge of the leading-edge sheath 22 to dig under 
the joggle on the lower airfoil composite skin 18, thereby forming an 
improper interface. 
Compounding this problem is the fact that when the stanchion 52 is located 
proximal to the compaction fixture 10, the compaction fixture 10 and the 
sheath-spreading/insertion apparatus 50 occlude the aft edges of the 
leading-edge sheath 22 and make it very difficult for the operator to 
visually discern whether the aft edges of the leading-edge sheath 22 are 
in the proper position for installation. In an effort to compensate for 
the uncertainty in the alignment of the aft edges of the leading-edge 
sheath 22, the aft edges of the leading-edge sheath 22 can be spread wider 
apart (greater than a preferred 1.27 cm (0.5 in.) displacement on each 
side) such that it will be less likely for one of the aft edges to dig 
under one of the joggles. However, a drawback to spreading the aft edges 
in excess of 1.27 cm (0.5 in.) on each side is that the composite 
materials and the heater mat in the leading-edge sheath 22 may be more 
susceptible to failure and rework due to the stresses generated by such an 
overspread. 
DISCLOSURE OF THE INVENTION 
It is therefore an object of the present invention to provide a method for 
fabricating a helicopter main rotor blade that provides complete chordwise 
and spanwise compaction to the blade subassembly. 
Another object of the present invention is to provide a method for 
fabricating a helicopter main rotor blade that reduces the extent of 
manual labor needed to lift, transport, and accurately position main rotor 
blade components. 
Still another object of the present invention is to provide a method for 
fabricating a helicopter main rotor blade that provides uniform spreading 
and alignment of the aft edges of the leading-edge sheath. 
Yet another object of the present invention is to provide a method for 
fabricating a helicopter main rotor blade that provides a non-occluded 
view of the aft edges of the leading-edge sheath and the joggles of the 
blade subassembly during installation of the leading-edge sheath onto the 
blade subassembly. 
These objects and others are achieved in the present invention by a method 
for fabricating a helicopter blade subassembly comprising a lower airfoil 
skin, a core having a conic formed therein, an upper airfoil skin, and a 
spar assembly having a tip end and a root end. 
The method comprises the steps of providing a blade compaction apparatus 
having a lower assembly configured for mating with an upper assembly. The 
lower assembly comprises a base having a contoured upper airfoil nest, 
including a root end and a tip end, at least two guide ramps disposed 
proximal to the contoured upper airfoil nest, wherein the guide ramps 
define conics therein, and wherein at least one of the guide ramps is 
disposed proximal to the root end and at least one of the guide ramps is 
disposed proximal to the tip end, and a plurality of leading-edge pusher 
cams disposed proximal to the contoured upper airfoil nest. 
The upper assembly comprises a support structure supporting a flexible 
impervious membrane, wherein the flexible impervious membrane and the 
contoured upper airfoil nest defme a molding cavity therebetween upon 
mating of the upper assembly to the lower assembly. 
The method further comprises the steps of disposing the upper airfoil skin 
and the core in combination with the contoured upper airfoil nest, 
connecting a tip end support insert in combination with the tip end of the 
spar assembly, connecting a root end support insert in combination with 
the root end a of the spar assembly, wherein the tip end support insert 
and the root end support insert each has guide surfaces and guide members 
configured for use in combination with the guide ramps. The spar assembly 
is then located in spanwise alignment in the contoured upper airfoil nest 
by abutting at least one of the guide surfaces against at least one of the 
guide ramps. The spar assembly is located in chordwise alignment in the 
contoured upper airfoil nest by driving the guide members into the conics 
in the guide ramps and driving the spar assembly into the conic in the 
core with the leading-edge pusher cams. The lower airfoil skin is then 
disposed in combination with the core and the spar assembly to form the 
helicopter blade subassembly, and then the upper assembly is mated to the 
lower assembly. The air within the molding cavity is then evacuated, such 
that evacuating the air causes the flexible impervious membrane to apply 
compaction forces to the helicopter blade subassembly. 
Still other objects and advantages of the present invention will become 
readily apparent to those skilled in this art from the following detailed 
description, wherein the preferred embodiments of the invention are shown 
and described, simply by way of illustration of the best mode contemplated 
of carrying out the invention. As will be realized, the invention is 
capable of modifications in various respects, all without departing from 
the invention. Accordingly, the drawings and description are to be 
regarded as illustrative in nature, and not as restrictive.

BEST MODE FOR CARRYING OUT THE INVENTION 
The apparatus and methods described in further detail hereinbelow comprise 
part of the manufacturing protocol for fabricating main rotor blades for 
H-60 helicopters manufactured by the Sikorsky Aircraft Corporation. It 
will be appreciated, however, that the apparatus and methods described 
herein have applicability in fabricating main rotor blades in general. 
MAIN ROTOR BLADE 
An H-60 main rotor blade 100 is exemplarily illustrated in FIGS. 3A-3D, and 
includes a leading edge 102 and a trailing edge 104, which in combination 
define the chord of the rotor blade 100, and a root end 106 and a tip end 
108 which in combination define the span of the rotor blade 100 (a tip cap 
109 for the main rotor blade 100, is separately fabricated for connection 
to the tip end 108 of the main rotor blade 100). The main rotor blade 100 
comprises upper and lower airfoil skins 110, 112 that define the upper and 
lower aerodynamic surfaces of the blade 100, respectively, a core 114, a 
spar assembly 116, and a leading-edge sheath 120. The upper and lower 
airfoil skins 110, 112, the core 114, and the spar assembly 116, in 
combination defme a blade subassembly 132. 
In the described embodiment, the upper and lower airfoil skins 110, 112 are 
prefabricated components formed from several plies of prepreg composite 
material of a type known to those skilled in the art, e.g., woven 
fiberglass material embedded in a suitable resin matrix. A plurality of 
corresponding trailing-edge tooling tabs 130 extend from the trailing 
edges of the upper and lower airfoil skins 110, 112, wherein each of the 
trailing-edge tooling tabs 130 have apertures 131 formed therein. As will 
be more fully discussed below, the trailing-edge tooling tabs 130 
facilitate proper spanwise and chordwise positioning of the upper and 
lower airfoil skins 110, 112 during the blade assembly process. 
In the described embodiment, the core 114 is fabricated from honeycomb 
material of a type typically used in aerospace applications, e.g., 
NOMEX.RTM. (NOMEX.RTM. is a registered trademark of E. I. du Pont de 
Nemours & Co., Wilmington, Del. for aramid fibers or fabrics) and 
functions as a low weight, structural stiffening member between the upper 
and lower airfoil skins 110, 112. The leading edge of the core 114 defines 
a conic 121 configured for mating with the trailing edge of the spar 
assembly 116. The upper airfoil skin 110, lower airfoil skin 112, and core 
114 have a plurality of aligned locator apertures 134 formed therethrough 
to facilitate the location of the spar assembly 116 in a blade compaction 
apparatus 200 as described in further detail hereinbelow. After the main 
rotor blade 100 has been assembled, the locator apertures 134 are patched 
with composite material so that the upper airfoil skin 110 and lower 
airfoil skin 112 have aerodynamically smooth surfaces. 
The spar assembly 116 comprises a spar 117, one or more counterweights 118, 
and backwall blocks 119. The spar 117 functions as the primary structural 
member of the main rotor blade 100, reacting the torsional, bending, 
shear, and centrifugal dynamic loads developed in the rotor blade 100 
during operation of the helicopter. The spar 117 of the described 
embodiment is formed from titanium; however, in alternative embodiments, 
the spar 117 may be formed from other metals, may be formed from composite 
materials, or a combination thereof. 
The counterweights 118 are utilized to statically and dynamically balance 
the main rotor blade 100. In the described embodiment, the counterweights 
118 are fabricated from less dense to more dense materials, e.g., foam, 
tungsten, and lead, respectively, in the spanwise direction from the root 
end 106 to the tip end 108, and provide the necessary weight distribution 
for statically and dynamically balancing the main rotor blade 100. The 
counterweights 118 are fabricated to include hardpoints 136 that provide 
physical engagement between the counterweights 118 and the inner mold line 
(IML) surface of the leading-edge sheath 120. The counterweights 118 are 
adhesively bonded to the leading edge of the spar 117 such that the bonded 
counterweights 118 are interposed between the leading-edge sheath 120 and 
the leading edge of the spar 117. 
The backwall blocks 119 are adhesively bonded to the trailing edge of the 
spar 117 at discrete locations corresponding to the locator apertures 134 
in the upper airfoil skin 110 and the core 114. As with the locator 
apertures 134, the backwall blocks 119 facilitate the location of the spar 
assembly 116 in the blade compaction apparatus 200 as described in further 
detail hereinbelow. 
In the described embodiment, the leading-edge sheath 120, which is 
illustrated in greater detail in FIG. 3C, is a prefabricated hybrid 
component fabricated from composite materials and abrasion-resistive 
materials. The sheath 120 has a generally U-shaped configuration that 
defines the leading edge 102 of the main rotor blade 100. The sheath 120 
comprises one or more plies 122 of prepreg composite material, e.g., woven 
fiberglass material embedded in a suitable resin matrix, that define the 
inner mold line (IML) of the leading-edge sheath 120, a first abrasion 
strip 124, and a second abrasion strip 126. The leading-edge sheath 120 
provides abrasion protection for the leading edge 102 of the main rotor 
blade 100, controls airfoil tolerances of the main rotor blade 100, and 
houses the main rotor blade de-icer assembly (not shown). 
As will be discussed more fully below, the methods for manufacturing the 
main rotor blade 100 include the use of the blade compaction apparatus 
200, a support apparatus 300, and a leading-edge sheath installation 
apparatus 400. 
BLADE COMTION APATUS 
Referring to FIG. 4, the blade compaction apparatus 200 is provided having 
a lower assembly 202 configured for mating with an upper assembly 250. The 
lower assembly 202 comprises a base 206 having a contoured upper airfoil 
nest 208, wherein the contoured upper airfoil nest 208 is configured to 
match the contour of the upper airfoil of the blade subassembly 132, and 
has an inboard end 210 corresponding to the root end 106 of the blade 
subassembly 132 and an outboard end 212 corresponding to the tip end 108 
of the blade subassembly 132. Two guide ramps 214, 216 are disposed 
proximal to the contoured upper airfoil nest 208 such that the guide ramp 
214 is disposed proximal to the inboard end 210 of the contoured upper 
airfoil nest 208, and the guide ramp 216 is disposed proximal to the 
outboard end 212 of the contoured upper airfoil nest 208. Each of the 
guide ramps 214, 216 has a ramp surface 218 defined therein, terminating 
at a conic 222. In addition, a threaded bolt 224 is disposed in 
combination with each of the guide ramps 214, 216 such that rotation of 
the threaded bolt 224 causes the threaded bolt 224 to translate across the 
ramp surface 218 and toward the conic 222. 
Three backwall pusher pin recesses 226 are formed in the contoured upper 
airfoil nest 208 at spanwise and chordwise locations corresponding to the 
locator apertures 134 in the blade subassembly 132. As depicted in FIG. 5, 
three backwall pusher pins 228 are provided for use in combination with 
the locator apertures 134 and the backwall pusher pin recesses 226 such 
that as more fully described below, the backwall pusher pins 228 can be 
used for various tooling purposes during manufacture of the blade 
subassembly 132. In addition, three leading-edge pusher cams 230, each 
having a camming surface 232, are disposed proximal to the leading edge of 
the contoured upper airfoil nest 208 such that each of the leading-edge 
pusher cams 230 is disposed at a spanwise location corresponding to the 
backwall pusher pin recesses 226. The leading-edge pusher cams 230 are 
configured in a conventional manner as is known in the art for cams, such 
that activation of the leading-edge pusher cams 230 causes the camming 
surfaces 232 to advance in a chordwise direction toward the backwall 
pusher pins 228. 
Referring to FIGS. 4 and 6, a vacuum source 234 is disposed in combination 
with the base 206 and has a plurality of conduits 236 extending therefrom, 
wherein each of the conduits 236 terminates at a corresponding aperture 
238 in the base 206. In the described embodiment, the vacuum source 234 
comprises a conventional vacuum pump, and the apertures 238 are aligned 
into two rows, one located proximal to the leading edge of the contoured 
upper airfoil nest 208 and one located proximal to the trailing edge of 
the contoured upper airfoil nest 208. In addition, five spring-loaded 
tooling tab pins 240 are disposed in a row proximal to the trailing edge 
of the contoured upper airfoil nest 208 at chordwise and spanwise 
locations corresponding to the apertures 131 in the blade subassembly 132 
trailing-edge tooling tabs 130. 
Referring to FIGS. 4 and 6, the upper assembly 250 comprises a support 
structure 252 supporting a flexible impervious membrane 254. In the 
described embodiment, the support structure 252 is connected to the base 
206 with a plurality of hinges 256 such that the support structure 252 can 
be pivoted from a first position wherein the flexible impervious membrane 
254 does not cover the contoured upper airfoil nest 208 to a second 
position wherein the flexible impervious membrane 254 covers the contoured 
upper airfoil nest 208. The support structure 252, flexible impervious 
membrane 254, contoured upper airfoil nest 208, and base 206 are 
configured in combination such that when the support structure 252 is in 
the second position such that the flexible impervious membrane 254 covers 
the contoured upper airfoil nest 208, an airtight seal is formed between 
the support structure 252 and the base 206. In addition, when the support 
structure 252 is in the second position, the flexible impervious membrane 
254 and the contoured upper airfoil nest 208 define a molding cavity 258 
therebetween. In the described embodiment, the apertures 238 in the base 
206 are configured such that upon activation of the vacuum source 234, the 
vacuum source 234 evacuates the air from within the molding cavity 258, 
thereby drawing the flexible impervious membrane 254 toward the contoured 
upper airfoil nest 208. 
SUPPORT APATUS 
Referring to FIGS. 3A-D, 4 and 7, the support apparatus 300 is provided for 
use in combination with the blade compaction apparatus 200 during 
manufacturing of the blade subassembly 132. The support apparatus 300 
comprises a root end support insert 302 configured for connection to the 
root end 106 of the spar 117, a tip end support insert 304 configured for 
connection to the tip end 108 of the spar 117, a tip end gearbox 306 
connected to the tip end support insert 304, a root end gearbox 308 
connected to the root end support insert 302, and a crane apparatus 338 
for supporting the root end and tip end gearboxes 308, 306. For the 
described embodiment as depicted in FIG. 8, the root end support insert 
302 comprises a base 310 having a pair of members 312, 314 extending 
therefrom that are configured for insertion into the root end 106 of the 
spar 117. One of the members 312 has two spring-loaded prongs 316 
extending therefrom for use in combination with a corresponding pair of 
apertures 117A in the spar 117, and functions to secure the root end 
support insert 302 in combination with the spar 117. The tip end support 
insert 304 comprises angular flanges 320 inserted into the tip end 108 of 
the spar 117 and having a plurality of apertures 322 formed therein 
corresponding to apertures in the tip end 108 of the spar 117 provided for 
securing the tip cap 109 to the blade subassembly 132. A plurality of 
bolts 324 are provided for use in combination with the apertures 322, 
thereby locating and securing the tip end support insert 304 to the spar 
117. 
Referring to FIGS. 6 and 8, each of the support inserts 302, 304 further 
comprises two guide members 326, 328 extending in a spanwise direction 
therefrom and terminating at a gearbox attachment plate 329 configured to 
facilitate attachment of the support inserts 302, 304 to their respective 
gearboxes 308, 306. In the described embodiment, the guide members 328 
proximal to the trailing edge of the spar assembly 116 are cylindrical 
rollers configured for engaging the ramp surfaces 218 in the guide ramps 
214, 216. In addition, each of the gearbox attachment plates 329 includes 
a guide surface 330 located on the side of the attachment plate 329 
proximal to the guide members 326, 328. As will be more fully discussed 
below, the guide members 326, 328 and the guide surfaces 330 are 
configured for use in combination with the guide ramps 214, 216 to 
properly locate the spar assembly 116 during manufacture of the blade 
subassembly 132. 
Referring to FIGS. 7 and 8, the tip end gearbox 306 and the root end 
gearbox 308 each are configured with conventional gearing arrangements 
such that each of the gearboxes 306, 308 are capable of translating the 
spar assembly 116 in a substantially vertical direction and also capable 
of rotating the spar assembly 116 about its longitudinal axis 123. In the 
described embodiment, the gearboxes 306, 308 are provided with rotatable 
cranks 334, 336 for manually urging the spar assembly 116 in the vertical 
and/or rotational directions. In alternative embodiments, electrical 
motors may be used in combination with the gearboxes 306, 308 to impart 
vertical and/or rotational movement to the spar assembly 116. 
The crane apparatus 338 comprises a conventional overhead crane as is known 
in the art for hoisting materials in a manufacturing environment. In the 
described embodiment, the crane apparatus 338 includes hoist cables 340 
connected to each of the gearboxes 306, 308, wherein the crane apparatus 
338 is capable of translating the gearboxes 306, 308 in the vertical 
and/or horizontal directions. 
LEADING-EDGE SHEATH INSTALLATION APATUS 
Referring to FIGS. 9, 10, and 11, the leading-edge sheath installation 
apparatus 400 comprises a lower assembly 402 connected to an upper 
assembly 450. 
The lower assembly 402 comprises a base 404 having a plurality of pneumatic 
cylinders 406 connected thereto and aligned in opposed rows. Each of the 
pneumatic cylinders 406 includes a translating member 408 which is capable 
of translational movement in response to pressurized air provided to the 
pneumatic cylinders 406 by a pressurized air source 410. The translating 
members 408 are connected to opposed carriage members 412, wherein the 
carriage members 412 are configured to support a plurality of suction cups 
414. The pneumatic cylinders 406, carriage members 412, and suction cups 
414 are configured such that the opposed rows of suction cups 414 are 
capable of synchronized translational movement in response to the 
pressurized air provided by the pressurized air source 410. 
In the described embodiment, the lower assembly 402 further comprises four 
vacuum pumps 416 connected to four vacuum accumulators 418, wherein the 
vacuum accumulators 418 are connected to the plurality of suction cups 414 
using a plurality of corresponding conduits 422. The vacuum pumps 416 and 
vacuum accumulators 418 function in combination to provide suction forces 
to the plurality of suction cups 414. In the described embodiment, the 
plurality of suction cups 414, and their corresponding conduits 422, are 
divided into four circuits wherein each circuit is connected to one of the 
vacuum accumulators 418. A plurality of valves 424 are disposed in 
combination with the conduits 422, and function to regulate the vacuum 
pressure provided to the plurality of suction cups 414. The vacuum 
accumulators 418 are of a conventional design and have a predetermined 
storage capacity designed for quickly providing each suction cup 414 with 
vacuum pressure between approximately 67.73 kPa and 84.66 kPa (20 in. Hg 
and 25 in. Hg). A leading-edge sheath contour nest 420 is provided for 
interposition between the opposed rows of suction cups 414 and is 
configured to support the leading-edge sheath 120. In alternative 
embodiments, the number of vacuum pumps 416, vacuum accumulators 418, and 
circuits can vary from those in the described embodiment in order to meet 
operational requirements of those embodiments. 
The upper assembly 450 includes five stanchions 452 extending substantially 
vertically from the base 404 and positioned spanwise at locations 
corresponding to the locations of the trailing-edge tooling tabs 130 of 
the blade subassembly 132. Each of the stanchions 452 has a contour clamp 
454 connected thereto, wherein each of the contour clamps 454 includes a 
contour surface 456, a hinged securing member 458, and a tooling tab pin 
460. The contour surfaces 456 and hinged securing members 458 are 
configured for clamping the blade subassembly 132 therebetween, wherein 
the tooling tab pins 460 function in combination with the blade 
subassembly 132 tooling tabs 130 to properly position the blade 
subassembly 132. In the described embodiment, each of the contour clamps 
454 is capable of assuming an open configuration wherein the hinged 
securing member 458 is positioned distal from the contour surface 456, and 
a closed configuration wherein the hinged securing member 458 is 
positioned proximal to the contour surface 456 and functions to secure the 
blade subassembly 132. In addition, the upper assembly 450 also comprises 
three backwall pusher pin recesses 451 formed in the contour surfaces 456 
and/or bridging members (not shown) extending between adjacent contour 
clamps 454 and positioned at spanwise and chordwise locations 
corresponding to the locator apertures 134 in the blade subassembly 132. 
The contour clamps 454 are connected to the stanchions 452 such that the 
contour clamps 454 are capable of translational movement relative to the 
leading-edge sheath contour nest 420. In the described embodiment, a 
rotational wheel 462 and differential 464 are provided in a conventional 
gearing arrangement and mechanically coupled to the contour clamps 454 for 
manually urging the contour clamps 454 into translational movement 
relative to the contour nest 420. In alternative embodiments, the 
translational movement can be provided utilizing electrical motors and/or 
hydraulic systems. 
METHOD FOR FABRICATING THE BLADE SUBASSEMBLY 
Referring to the blade subassembly 132, blade compaction apparatus 200, and 
support apparatus 300 illustrated in FIGS. 3A-D and 4-8, and referring to 
the flowchart of FIG. 12, a method MF for fabricating the blade 
subassembly 132 will now be discussed in greater detail. 
In step 500, the upper airfoil skin 110 and core 114 are disposed upon the 
upper airfoil nest 208. In the described embodiment, the upper airfoil 
skin 110 and core 114 are adhesively bonded together prior to disposal 
upon the upper airfoil nest 208 in order to reduce fabrication time. In 
step 502, the upper airfoil skin 110 and core 114 combination is located 
spanwise and chordwise upon the upper airfoil nest 208 by disposing the 
apertures 131 in the trailing-edge tooling tabs 130 about the 
spring-loaded tooling tab pins 240. In addition, the backwall pusher-pins 
228 are inserted through the locator apertures 134 in the core 114 and 
upper airfoil skin 110, and into the backwall pusher pin recesses 226. 
In the described embodiment, adhesives (not shown) of a type known in the 
art for adhesively bonding blade spars to skin/core combinations are 
applied to the spar assembly 116 prior to proceeding to step 504. In step 
504, the root end support insert 302 is connected to the root end 106 of 
the spar 117, and the tip end support insert 304 is connected to the tip 
end 108 of the spar 117. In steps 506 and 508, the root end support insert 
302 is connected to the root end gearbox 308, the tip end support insert 
304 is connected to the tip end gearbox 306, and then both gearboxes 306, 
308 are, in turn, connected to the hoist cables 340 such that the spar 
assembly 116 is supported by the crane apparatus 338. 
In step 510, the spar assembly 116 is hoisted by the crane apparatus 338 
and positioned such that the root end support insert 302 is proximal to 
the guide ramp 214 on the inboard end 210 of the contoured upper airfoil 
nest 208, and such that the tip end support insert 304 is proximal to the 
guide ramp 216 on the outboard end 212 of the contoured upper airfoil nest 
208. In step 512, the guide members 326, 328 of both the root end support 
insert 302 and the tip end support insert 304 are engaged with the ramp 
surfaces 218 in the guide ramps 214, 216. The spar assembly 116 is then 
positioned spanwise by abutting the guide surface 330 on the tip end 
support insert 304 against the guide ramp 216 on the outboard end 212 of 
the contoured upper airfoil nest 208. In step 514, the root end gearbox 
308 is disconnected from the root end support insert 302, and the tip end 
gearbox 306 is disconnected from the tip end support insert 304. 
In step 516, the root end 106 and tip end 108 of the spar assembly 116 are 
properly positioned chordwise by driving the guide members 326, 328 into 
the ramp surfaces 218 with the threaded bolts 224 until the guide members 
328 proximal to the trailing edge of the spar assembly 116 abut the conic 
222 in each of the ramp surfaces 218. Proper positioning of the root end 
106 and tip end 108 of the spar assembly 116 does not necessarily ensure 
that the rest of the spar assembly 116 will have the proper chordwise 
alignment. To properly locate the entire spar assembly 116 chordwise, in 
step 518, the leading-edge pusher cams 230 are actuated such that the 
camming surfaces 232 engage the hardpoints 136 on the spar assembly 116 
and drive the backwall blocks 119 of the spar assembly 116 against the 
backwall pusher pins 228. It will be appreciated that the backwall pusher 
pins 228 are located chordwise such that contact between the backwall 
blocks 119 and the backwall pusher pins 228 ensures proper seating of the 
trailing edge of the spar assembly 116 within the conic 121 defined by the 
core 114. After the spar assembly 116 has been properly located, adhesives 
(not shown) are applied to both the core 114 and spar assembly 116 in 
preparation for receiving the lower airfoil skin 112. 
In step 520 the lower airfoil skin 112 is disposed onto the core 114 and 
spar assembly 116 and is located spanwise and chordwise using the lower 
airfoil skin's 112 trailing-edge tooling tabs 130 in combination with the 
spring-loaded tooling tab pins 240. In step 522, a caulplate 242 is 
disposed over the blade subassembly 132 (see FIG. 6), and functions to 
distribute the compaction forces applied to the blade subassembly 132 by 
the flexible impervious membrane 254 during compaction. In step 524, a 
fiberglass breather bag 244 is disposed over the caulplate 242 (see FIG. 
6) and functions to assist in the evacuation of air trapped between the 
caulplate 242 and the flexible impervious membrane 254 during compaction, 
and to reduce friction between the blade subassembly 132 and the flexible 
impervious membrane 254. 
In step 526, the support structure 252 is pivoted to the second position 
and secured in place such that the flexible impervious membrane 254 covers 
the blade subassembly 132 and the contoured upper airfoil nest 208, 
thereby forming the airtight molding cavity 258 therebetween. In step 528, 
the vacuum source 234 is activated, thereby evacuating air from within the 
molding cavity 258. In the described embodiment, the vacuum pressure 
supplied by the vacuum source 234 is approximately equal to 44.02 kPa (13 
in. Hg). Evacuation of the air from within the molding cavity 258 causes 
the flexible impervious membrane 254 to be drawn toward the contoured 
upper airfoil nest 208, thereby providing compaction forces to the 
components of the blade subassembly 132. In the described embodiment, for 
a vacuum pressure of approximately 44.02 kPa (13 in. Hg), a compaction 
duration of approximately 30 minutes provides appropriate compaction to 
blade subassembly 132. 
In step 530, the vacuum source 234 is deactivated, and the support 
structure 252 is pivoted back to the first position, wherein the flexible 
impervious membrane 254 is no longer covering the blade subassembly 132. 
To prepare the compacted blade subassembly 132 for removal from the blade 
compaction apparatus 200, the breather bag 244, caulplate 242, 
leading-edge pusher cams 230, and backwall pusher pins 228 are all 
removed. In steps 532 and 534, the root end support insert 302 is 
reconnected to the root end gearbox 308, the tip end support insert 304 is 
reconnected to the tip end gearbox 306, and then both gearboxes 306, 308 
are, in turn, reconnected to the hoist cables 340 such that the blade 
subassembly 132 is supported by the crane apparatus 338. 
In step 536, the threaded bolts 224 are loosened and the guide members 326, 
328 are slid out of the ramp surfaces 218 such that the root end support 
insert 302 and the tip end support insert 304 are no longer constrained by 
the blade compaction apparatus 200. The blade subassembly 132 is then 
transpositioned by the support apparatus 300 from the blade compaction 
apparatus 200 to the leading-edge sheath installation apparatus 400. 
METHOD FOR INSTALLING THE LEADING-EDGE SHEATH 
Referring to the blade subassembly 132, support apparatus 300, and 
leading-edge sheath installation apparatus 400 illustrated in FIGS. 3A-D 
and 7-11, and referring to the flowchart of FIG. 13, a method MI for 
installing the leading-edge sheath 120 in combination with the compacted 
blade subassembly 132 will now be discussed in greater detail. 
In preparation for placement of the leading-edge sheath 120 within the 
leading-edge sheath installation apparatus 400, in step 600, the opposed 
rows of suction cups 414 are checked to ensure that the suction cups 414 
are in a disengaged position wherein the opposed rows of suction cups 414 
are spaced-apart a distance greater that a width w.sub.LES of the 
leading-edge sheath 120 (see FIG. 3C). In step 602, the leading-edge 
sheath 120 is placed within the leading-edge sheath contour nest 420, and 
positioned spanwise by abutting the tip end of the leading-edge sheath 120 
against a tooling stop (not shown). In step 604, the opposed rows of 
suction sups 414 are moved by the pneumatic cylinders 406 into an engaged 
position such that each of the plurality of suction cups 414 is in 
abutting engagement with the respective outer mold line (OML) surfaces of 
the aft edges of the leading-edge sheath 120. 
In step 606, the vacuum pumps 416 are activated, thereby accumulating 
vacuum pressures in each of the corresponding vacuum accumulators 418. In 
step 608, the valves 424 are opened such that the vacuum pressures 
accumulated in the vacuum accumulators 418 are provided to the plurality 
of suction cups 414, thereby generating suction forces between each of the 
plurality of suction cups 414 and the OML surfaces of the leading-edge 
sheath 120. In the described embodiment, each of the suction cups 414 
comprises a bellow-shaped design such that upon opening of the valves 424, 
the bellow-shaped portion of each suction cup 414 collapses, thereby 
causing the OML surfaces of the aft edges of the leading-edge sheath 120 
to be pulled towards the opposed carriage members 412. In the described 
embodiment, the suction cups 414 are configured such that upon application 
of a vacuum pressure between approximately 67.73 kPa and 84.66 kPa (20 in. 
Hg and 25 in. Hg), the bellow-shaped portion of each suction cup 414 
collapses approximately 1.27 cm (0.5 in.) such that the width W.sub.LES of 
the leading-edge sheath 120 increases by approximately 2.54 cm (1 in.). In 
alternative embodiments, the design of the suction cups 414 and the 
magnitude of the suction forces applied to the suction cups 414 can be 
modified from the described embodiment to increase or decrease the spread 
of the leading-edge sheath 120. In addition, the pneumatic cylinders 406 
can be used to translate the opposed rows of suction cups 414 away from 
each other such that additional spread of the leading-edge sheath 120 is 
achieved. 
Prior to placement of the blade subassembly 132 in combination with the 
leading-edge sheath installation apparatus 400, adhesives (not shown) can 
be applied to the leading-edge 133 of the blade subassembly 132 (see FIG. 
9) for use in adhesively bonding the leading-edge sheath 120 onto the 
blade subassembly 132. In step 610, the root end and tip end gearboxes 
306, 308 of the support apparatus 300 are adjusted to ensure that the 
blade subassembly 132 is in a substantially vertical orientation with the 
leading-edge 133 of the blade subassembly 132 facing down. In step 612, 
the backwall pusher pins 228 are inserted into the backwall pusher pin 
recesses 451 in the upper assembly 450 of the leading-edge sheath 
installation apparatus 400, and the contour clamps 454 are placed in the 
open configuration such that the hinged securing members 458 are 
positioned distal from the contour surfaces 456. In step 614, the blade 
subassembly 132 is transported by the support apparatus 300 to the open 
contour clamps 454 and is placed therein such that the backwall pusher 
pins 228 are inserted through the locator apertures 134 in the upper 
airfoil skin 110 and core 114, and such that the tooling tab pins 460 are 
inserted through the corresponding tooling tab apertures 131, thereby 
ensuring proper alignment of the blade subassembly 132 in the leading-edge 
sheath installation apparatus 400. In step 616, the contour clamps 454 are 
placed in the closed configuration such that the hinged securing members 
458 are positioned in abutting engagement with the lower airfoil skin 112 
of the blade subassembly 132, thereby securing the blade subassembly 132 
within the contour clamps 454. 
In step 618, the root end gearbox 308 is disconnected from the root end 
support insert 302, and the tip end gearbox 306 is disconnected from the 
tip end support insert 304. In step 620, the contour clamps 454 are 
translated downward using the rotational wheel 462 such that the leading 
edge 133 of the blade subassembly 132 is inserted into the spread 
leading-edge sheath 120. In step 622, the leading edge 133 of the blade 
subassembly 132 and the leading-edge sheath 120 are visually inspected to 
ensure that the leading edge 133 of the blade subassembly 132 is properly 
inserted into the leading-edge sheath 120 such that the aft edges of the 
leading-edge sheath 120 overlap both the upper airfoil skin 110 and the 
lower airfoil skin 112. In particular, step 622 is a check to ensure that 
the aft edges of the leading-edge sheath 120 do not cause separation 
between either the upper airfoil skin 110 or the lower airfoil skin 112 
and the spar assembly 116. 
In step 624, the valves 424 disposed in combination with the conduits 422 
are closed such that the supply of vacuum pressure to the plurality of 
suction cups 414 is discontinued, thereby causing the leading-edge sheath 
120 to be fitted in combination with the leading edge 133 of the blade 
subassembly 132. In step 626, the contour clamps 454 are translated upward 
using the rotational wheel 462 such that the leading edge 102 of the 
assembled main rotor blade 100 clears the opposed rows of suction cups 
414. In steps 628 and 630, the root end support insert 302 is reconnected 
to the root end gearbox 308, the tip end support insert 304 is reconnected 
to the tip end gearbox 306, and then both gearboxes 306, 308 are, in turn, 
reconnected to the hoist cables 340 such that the main rotor blade 100 is 
supported by the crane apparatus 338. 
In steps 632 and 634, the contour clamps 454 are opened and the main rotor 
blade 100 is separated from the tooling tab pins 460 and backwall pusher 
pins 228 such that the main rotor blade 100 can be transported by the 
support apparatus 300 away from the leading-edge sheath installation 
apparatus 400. 
It will be readily seen by one of ordinary skill in the art that the 
present invention fulfills all the objects set forth above. After reading 
the foregoing specification, one of ordinary skill will be able to effect 
various changes, substitutions of equivalents and various other aspects of 
the invention as broadly disclosed herein. It is therefore intended that 
the protection granted hereon be limited only by the definition contained 
in the appended claims and equivalents thereof.