Note: Descriptions are shown in the official language in which they were submitted.
CA 02521125 2005-10-19
. _=
APPARATUS AND METHODS FOR FABRICATING
A HELICOPTER MAIN ROTOR BLADE
This application is a division of Application Serial No. 2,195,078 filed June
21, 1995.
Related Application
The instant application is related to commonly-owned, U.S. Patent No.
5,598,760 issued
February 4, 1997, and entitled AN EOP SCRIBE DEVICE.
Technical Field .
The present invention is directed to manufacturing apparatus. and methods, and
more specifically, to apparatus and methods for fabricating a helicopter main
rotor blade.
Background of the Invention
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 materiais lies in the fabrication of
'main rotor blades
for helicopters.
With increased usage of composite materials to fabricate main rotor blades,
the
helicopter industry is continually seeking to improve the tooling and/or
niethods used to
fabricate main-rotor blades so as to reduce the per unit fabrication costs
associated with
the main rotor blades. Typically, the per blade fabrication costs are higher
than need be
due to part rejections or rework that occurs during the main rotor blade
fabrication
process. Part rejections typically occur where the composite material has been
so
substantially damaged during the fabrication process that rework is not cost
effective or
where a finished fabricated part exceeds the tolerance limits established for
the part.
Rework occurs where the composite material has been damaged during the
fabrication
process, and the damage may be repaired in a relatively cost effective manner.
Sikorsky Aircraft has developed a parallel manufacturing protocol for
fabricating
helicopter main rotor blades wherein a blade subassembly and a leading-edge
sheath are
concurrentl~ 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 subs-
equently
cured to form a finished main rotor blade. This protocol was adopted in large
measure
because experience has shown that the leading edges of main rotor blades are
subjected to'
varying degrees of abrasion during helicopter operations. As a result of such
abrasion
effects, the leading edge of a helicopter main rotor blade at some point
becomes
= 35 aerodynamically unsuitable for further use. Rather than replacing the
entire main rotor
CA 02521125 1995-06-21
blade, it was determined that a replaceable leading-edge sheath would allow
abrasion-
degraded main rotor blades to be efficaciously and economically repaired.
The prior art process for fabricating blade subassemblies involved the use of
a
"clamshell" tooling fixture and a "wet" lay-up process for the composite
materials. It was
determined that the rejection rate for blade subassemblies fabricated using
the clamshell
tooling fixture and the wet lay-up process was unacceptable in light of the
today's
competitive market. The dependability and accuracy of the clamshell tooling
fixture
depended upon the stability of the laid up tooling contours, the proper
securing and
pinning of all fasteners and locators, and the variability in applying blade
outer mold line
pressures. The clamshell tooling fixture and the wet lay-up process were
subjected to
shrinkage and lose of tolerances, which led to component rejection. The
clamshell
configuration result in asymmetrical pressure distributions across the layed-
up blade
subassembly.
Another area of concern in the parallel manufacturing protocol was the sheath
spreader tool used to integrate the leading-edge sheath in combination with
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 mounted
spanwise on the
aft edges of the leading-edge sheath in contact with the inner mold Iine
(IMI.) surfaces
(which are formed of composite material) of the leading-edge sheath. Each
segment of the
prior art grabber is individually actuated by means of a side cam lever. The
prior grabbers
exert a shearing action against the MfL surfaces of the leading-edge sheath in
spreading
the aft edges of the sheath apart. The shearing action caused by the prior art
grabbers
caused cracks and delaminations in the composite material subjected to the
shearing action
thereof, resulting in component rejections or rework. In addition to the
foregoing
deficiency of the prior art leading-edge sheath spreader tool, the segments of
the grabber
are individually actuated in a sequential manner such that to spread apart the
entire
leading-edge sheath involves multiple, repetitive operations. Not only is such
a procedure
labor intensive and time consuming, and hence costly, such a procedure may
induce
unwanted stresses into the aft edges of the leading-edge sheath.
A need exists to provide an apparatus for spreading a leading-edge sheath for
insertion onto a blade subassembly without inducing cracks and/or
delanrinations in the
composite material of the leading-edge sheath. Preferably, the apparatus
should spread the
leading-edge sheath apart in a single operation to reduce the time required to
spread the
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CA 02521125 1995-06-21
leading-edge sheath apart. A need also exists to provide a fixture for
assemblage and
compacting of a blade subassembly that provides a uniform pressure
distribution during the
compaction of the blade subassembly, that facilitates the use of prepreg
composite
materials, and that ensures proper chordwise and spanwise alignment of the
components of
the blade subassembly layed-up in the fixture. A need also exists to provide a
sheath
spreading apparatus and compaction fixture which in combination simplify the
insertion of
a spread-apart leading-edge sheath onto the blade subassembly.
Summary of the Invention
One object of the present invention is to provide a sheath spreadingTnsertion
apparatus that spreads apart a leading-edge sheath without inducing cracks
and/or
delaminations in the composite material thereof.
Another object of the present invention is to provide a sheath
spreading/'insertion
apparatus that spreads apart a leading-edge sheath in a single operation.
A further object of the present invention is to provide a compaction fixture
for
assemblage and compaction of blade subassembly components that provides a
uniform
pressure distribution during compaction of the blade subassembly.
Still another object of the present invention is to provide a compaction
fixture that
ensures proper chordwise and spanwise alignment of the components comprising
the blade
subassembly as assembled in the compaction fixture.
One more object of the present invention is to provide a sheath
spreading/insertion
apparatus and a compaction fixture which, in combination, greatly simplify the
insertion of
a spread-apart leading-edge sheath assembly onto the blade subassembly.
These and other objects of the present invention are achieved by a sheath
spreadingrnsertion apparatus according to present invention for spreading a
leading-edge
sheath and inserting the spread-apart leading-edge sheath in combination with
a blade
subassembly. The sheath spreading/insertion apparatus comprises a movable
stanchion, an
upper elongate carriage member mounted in movable combination with the movable
stanchion and a lower elongate carriage member mounted in movable combination
with the
movable stanchion. A plurality of suction cups are mounted in combination with
each of
the upper and lower elongate carriage members. A means is provided by
imparting
synchronized movement to the upper and lower carriage members between a
disengaged
position wherein the leading-edge sheath may be inserted between the
pluralities of suction
cups mounted in combination with the upper and lower carriages without contact
therewith, an engaged position wherein the pluralities of suction cups
abuttingly engage
respective outer mold line (OML) surfaces of the leading-edge sheath, and an
operating
3
CA 02521125 2005-10-19
position wherein the leading-edge sheath is spread apart for insertion onto
the blade
subassembly. A means is provided for generating suction forces in the
pluralities of
suction cups in the engaged position to cause the suction cups to hold the
respective OML
surfaces of the leading-edge sheath such that subsequent synchronized movement
of the
upper and lower carriage members to the operating position causes the leading-
edge
sheath to be spread apart. A means is provided for moving the movable
stanchion to insert
the spread-apart leading-edge sheath onto the blade subassembly.
There is provided, in accordance with an aspect of the present invention, a
method
for assembling and compacting a blade subassembly suitable for use in a
helicopter main
rotor blade that includes an upper composite skin, a lower composite skin, a
honeycomb
core, and a spar assembly, comprising the steps of: providing a composite
fixture that
includes a lower assembly having a contoured upper airfoil nest mounted on a
support
structure and an upper assembly having a structural support truss, a contoured
backplate
affixed to the structural support truss, and a pressure bag fastened in sealed
combination
with the contoured backplate; laying up the upper composite skin and the
honeycomb
core in combination in the contoured upper airfoil nest; locating the spar
assembly in
chordwise and spanwise alignment in the contoured upper airfoil nest; laying
up said
lower composite skin in combination with the spar assembly and the honeycomb
core;
locking the upper assembly and the lower assembly in combination; and
pressurizing the
pressure bag to compact the blade subassembly.
There is further provided, in accordance with another aspect of the present
invention, a method for spreading and positioning a leading-edge sheath in
relation to a
blade subassembly, comprising the steps of: mounting the leading-edge sheath
between
first and second rows of suction cups; imparting synchronized movement to the
first and
second rows of suction cups to an engaged position wherein the suction cups
are in
abutting engagement with respective outer mold line surfaces of the leading-
edge sheath;
generating suction forces in the first and second rows of suction cups to
cause the suction
cups to hold the respective outer mold line surfaces of the leading-edge
sheath; imparting
synchronized movement to the first and second rows of suction cups to an
operating
position to cause the leading-edge sheath to be spread apart; and positioning
the spread-
apart leading-edge sheath in relation to the blade subassembly.
There is further provided, in accordance with yet another aspect of the
present
invention, a compaction fixture for assembling and compacting a blade
subassembly
suitable for use in a helicopter main rotor blade that includes an upper
composite skin, a
lower composite skin, a honeycomb core, and a spar assembly, comprising: a
lower
assembly including a support structure, a contoured upper airfoil nest mounted
in
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CA 02521125 2005-10-19
combination with said support structure, said contoured upper airfoil nest
having an outer
mold line surface that defines the airfoil surface of the upper composite
skin, a plurality
of tooling pins for locating the upper composite skin in aligned combination
with said
contoured upper airfoil nest, and a plurality of pusher pins for locating the
spar assembly
in chordwise alignment in said contoured upper airfoil nest, and a spar
stanchion mounted
in combination with said support structure at the inboard and outboard ends of
said
contoured upper airfoil nest, respectively, for locating the spar assembly in
spanwise
alignment in said contoured upper airfoil nest; an upper assembly including a
structural
support truss, a contoured backplate affixed to said structural support truss,
and a pressure
bag fastened in sealed combination with said contoured backplate, said
pressure bag
having spanwise and chordwise dimensions corresponding to the blade
subassembly;
means for locking said upper assembly in combination with said lower assembly
wherein
compaction of the blade subassembly assembled in said lower assembly may be
effectuated; and means for pressurizing said pressure bag with said upper and
lower
assemblies in locked combination to compact the blade subassembly assembled in
said
lower assembly.
The sheath spreading/insertion apparatus further includes a means for
indicating
that the spread-apart leading-edge sheath has been fully inserted onto the
blade
subassembly. The synchronized movement imparting means comprises a plurality
of
pneumatic cylinders mounted in combination with the upper elongate carriage
member
and the movable stanchion, a plurality of pneumatic cylinders mounted in
combination
with the lower elongate carriage members and the movable stanchion, and a
pressure
source pneumatically interconnected to the pluralities of pneumatic cylinders.
Actuation
of the pressure source provides pressurized air to the pluralities of
pneumatic cylinders to
cause synchronized movement of the upper and lower elongate carriage members
between the disengaged position wherein the leading-edge sheath may be
inserted
between the pluralities of suction cups mounted in combination with the upper
and lower
elongate carriage members without contact therewith, the engaged position
wherein the
pluralities of suction cups abuttingly engage respective OML surfaces of the
leading-edge
sheath, and the operating position wherein the leading-edge sheath is spread
apart for
insertion onto the blade subassembly. For the described embodiment, ninety
suction cups
are mounted in combination with the upper elongate carrier member and ninety
suction
cups are mounted in combination with the lower elongate carrier member.
To spread and insert the leading-edge sheath in combination with a blade
subassembly, the leading-edge sheath is mounted between upper and lower rows
of
suction cups, the upper and lower rows of suction cups are displaced in
synchronized
4a
CA 02521125 2005-10-19
movement to an engaged position wherein the suction cups abuttingly engage
respective
OML surfaces of the leading-edge sheath, suction forces are generated in the
upper and
lower suction cups to cause the suction cups to hold the respective OML
surfaces of the
leading edge sheath, the upper and lower rows of suction cups are displaced in
synchronized movement to an operating position to cause the leading-edge
sheath to be
spread apart, and the spread-apart leading-edge sheath is inserted onto the
blade
subassembly.
A compaction fixture according to the present invention is provided for
assembling and compacting a blade subassembly that includes upper and lower
composite
skins, a honeycomb core, and a spar assembly (a spar with at least one
counterweight
bonded
4b
CA 02521125 1995-06-21
thereto). The compaction fixture comprises a lower assembly that includes a
support
structure and a contoured upper airfoil nest mounted in combination with the
support
structure. The contoured upper airfoil nest has an OML surface that defines
the airfoil
surface of the upper composite skin, a plurality of tooling pins for locating
the upper
composite skin, honeycomb core combination in the contoured upper airfoil
nest, and a
plurality of pusher pins for locating the spar assembly in chordwise alignment
in the
contoured upper airfoil nest. A spar stanchion is mounted in combination with
the inboard
and outboard ends of the contoured upper airfoil nest, respectively, for
locating the spar
assembly in spanwise alignment in the contoured upper airfoil nest. The
compaction
fixture further comprises an upper assembly that includes a structural support
truss, a
contoured backplate affixed to the structural support truss, and a pressure
bag having
chordwise and spanwise dimensions corresponding to the blade subassembly
fastened in
sealed combination with the contoured backplate. A means is provided for
locking the
upper and lower assemblies in combination so that compaction of the blade
subassembly
assembled in the lower assembly may be effectuated. A means is provided for
pressurizing
the pressure bag to compact the blade subassembly disposed in the locked upper
and lower
assemblies. The compaction fixture may further include a caul plate interposed
between
the upper and lower assemblies to provide uniform pressure distribution over
the layed-up
blade subassembly during compaction thereof.
To assemble and compact the blade subassembly, a composite fixture as
described
in the preceding paragraph is provided. The upper composite skin and the
honeycomb
core are layed-up in combination in the contoured upper airfoil nest. The spar
assembly is
located in chordwise and spanwise alignment in the contoured upper airfoil
nest, and the
lower composite skin is layed-up in combination with the spar assembly and the
honeycomb core. The upper and lower assemblies are locked in combination and
the
pressure bag is pressurized to compact the assembled blade subassembly.
Brief Description of the Drawings
A more complete understanding of the present invention and the attendant
features
and advantages thereof may be had by reference to the following detailed
description when
considered in conjunction with the following drawings wherein:
Figure lA is a top plan view of an exemplary main rotor blade for an H-60
helicopter.
Figure 1 B is a cross-sectional view of the main rotor blade of Figure 1 A
taken
along line 1B-1B thereof.
5
CA 02521125 1995-06-21
Figure 1C is an enlarged partial perspective view of the leading edge sheath
illustrated in Figure 1B.
Figure 1D is an enlarged partial perspective view of a counterweight for the
exemplary main rotor blade of Figure 1 A.
Figure 2 is a perspective view of a compaction fixture and a sheath
spreadingrnsertion apparatus according to the present invention.
Figure 3 is a partial plan view of the apparatus of Figure 2.
Figure 3A is a partial perspective view of the sheath spreadinglinsertion
apparatus
of Figure 3.
Figure 4A is a flow chart illustrating the assemblage and compaction process
according to the present invention.
Figure 4B is a flow chart illustrating the sheath spreading and sheath
insertion
process according to the present invention.
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 Division of United Technologies
Corporation. In
particular, the apparatus and methods described herein have particular utility
for
fabricating the H-60 growth main rotor blade developed by Sikorsky Aircraft.
It will be
appreciated, however, that the apparatus and methods described herein have
applicability
in fabricating main rotor blades in general.
An H-60 growth main rotor blade 100 is exemplarily illustrated in Figures lA-
1D,
and includes a leading edge 102 and a trailing edge 104, which in combination
define the
chord of the rotor blade 100, and an inboard end 106 and an outboard (tip) end
108 (an
anhedral tip portion of the main rotor blade 100, which is the portion of the
blade outboard
of the dashed line 109 in Figure 1 A, is separately fabricated as a
replaceable component for
the main rotor blade 100), which in combination define the span of the rotor
blade 100.
The main rotor blade 100 comprises upper and lower composite skins 110, 112
that define
the upper and lower aerodynamic surfaces of the blade 100, respectively, a
honeycomb
core 114, a spar 116, one or more counterweights 118, and a leading-edge
sheath 120.
Adjustable trim tabs 130 (two for the illustrated embodiment) extend
rearwardly from the
trailing edge 104. The upper and lower composite skins 110, 112, the honeycomb
core 114, the spar 116, and the counterweights 118 in combination define a
blade
subassembly 132.
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CA 02521125 2005-10-19
The composite 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., for the
described embodiment woven fiberglass material embedded in a suitable resin
matrix.
The upper composite skin 110 has a plurality of locator apertures 134 (see
Figure 1 A)
formed therethrough to facilitate the location of the spar assembly 116/118 in
a
compaction fixture 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 composite skin 110 has an aerodynamically smooth surface. The
honeycomb core 114 is fabricated of material type typically used in aerospace
applications, e.g., for the described embodiment NOMEXS (NOMEX is a registered
trademark of E.I. du Pont de Nemours & Co., Wilmington, DE for aramid fibers
or
fabrics) and functions as a low weight, structural stiffening member between
the upper
and lower composite skins 110, 112.
The spar 116 is a prefabricated component and 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 116 of the described embodiment is a composite spar of
the type
disclosed and claimed in commonly-owned U.S. Patent No. 5,755,558 issued May
26,
1998 and entitled FIBER REINFORCED COMPOSITE SPAR FOR A ROTARY WING
AIRCRAFT AND METHOD OF MANUFACTURE THEREOF. The composite spar 116
comprises upper and lower side walls corresponding to upper and lower airfoil
surfaces,
respectively, and forward and aft conic closures corresponding to leading and
trailing
edges, respectively, of the main rotor blade 100. The upper and lowerside
walls comprise
a plurality of pre-ply layers, each pre-ply layer including unipack plies and
cross plies of
prepreg composite material, i.e., fibrous material embedded in a resin matrix.
The
unipack plies, which have longitudinally (axial) orientated fibers, are of
equal width and
staggered to provide a tapered edge with the upper and lower side walls. The
cross plies,
which have f40 orientated fibers, have varying widths that form a staggered
butt joint
about the periphery of the forward and aft conic closures. While the described
embodiment of the main rotor blade 100 incorporates a composite spar 116, one
skilled in
the art will appreciate that the apparatus and methods of the present
invention may also be
utilized in manufacturing main rotor blades wherein the spar is fabricated as
a metallic
structural member, e.g., a titanium spar.
One or more counterweights 118, one of which is illustrated in further detail
in
Figure 1D, are utilized to statically and dynamically balance the main rotor
blade 100.
The counterweights 118 are fabricated from less dense to more dense materials,
e.g.,
foam, tungsten, and lead, respectively, for the described embodiment, in the
spanwise
direction from the inboard end 106 to the outboard end 108 to provide the
necessary
weight
7
CA 02521125 2005-10-19
distribution for statically and dynamically balancing the main rotor blade
100. The
counterweights 118 are fabricated to include hardpoints 136 that provide the
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
spar 116 to form a spar assembly 116/118 wherein the bonded counterweights 118
are
in an interposed position between the leading edge sheath 120 and the leading
edge of
the spar 116.
The leading edge sheath 120, which is illustrated in greater detail in Figure
1 C, is a prefabricated hybrid component fabricated from composite materials
and
abrasion-resistive materials. The sheath 120 has a generally V-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 for the described embodiment,
that
define the inner mold line (IML) of the leading edge sheath 120, a first
abrasion strip
124, and a second abrasion strip 126. For the described embodiment of the
leading
edge sheath 120, the first abrasion strip 124 is fabricated from titanium and
the second
abrasion strip 126 is fabricated from nickel. The tip end 108, i.e., outboard
end, of the
leading edge sheath 120 has the nickel strip 126 bonded to the titanium strip
124 as
illustrated in Figure 1 C. The titanium strip 124 with the nickel strip 126
overlay is
adhesively bonded to the prepreg composite plies 122 to form the leading edge
sheath
120. Exposed segments 128 of the prepreg composite plies 122 facilitate
adhesive
bonding of the leading edge sheath 120 in combination with the blade
subassembly
132. The exposed segments 128 include finished edges 128A (a method and
apparatus
for defiuiing the finished edges 128A of the leading-edge sheath 120 is
described in
commonly-owned, U.S. Patent No. 5,598,760 issued February 4, 1997 and entitled
AN EOP SCRIBE DEVICE that have been formed to define the proper integration of
the leading-edge sheath 120 in combination with blade subassembly 132. The
leading
edge sheath 120 is removable to facilitate replacement thereof. The leading
edge
sheath 120, and in particular the titanium strip 124 and the nickel strip 126
overlay,
provides abrasion protection for the leading edge 102 of the main rotor blade
100. The
leading edge sheath 120 also provides control of airfoil tolerances of the
main rotor
blade 100.
With reference to Figures 2-3, 3A, the apparatus according to the present
invention includes a compaction fixture 10 and a sheath spreading/insertion
apparatus
50. The compaction fixture 10 includes a lower assembly 12 and an upper
assembly
30, which, in secured combination, define the compaction fixure 10 which is
operative to compact the assembled blade subassembly 132. The lower assembly
12
comprises a contoured upper airfoil nest 14 mounted on a support structure 16.
The
contoured upper airfoil nest 14 has
8
CA 02521125 1995-06-21
an outer mold line (OML) surface 18 that defines the OML of the upper airfoil
surface of
the rotor blade 100, i.e., the upper composite skin 110.
Affixed in combination with the ONiL surface 18 of the contoured upper airfoil
nest 14 are a plurality of tooling pins 20 (five for the described
embodiment). The tooling
pins 20 function as location markers for locating the upper composite skin 110
in aligned
combination on the contoured upper airfoil nest 14. Also affixed in
combination with the
OML surface 18 of the contoured upper airfoil nest 14 are a plurality of
pusher pins 22
(three for the described embodiment). The pusher pins 22 are operative to
define the
chordwise alignment of the spar assembly 116/118 in combination with the upper
composite skin l 10, honeycomb core 114 combination as layed-up in the
contoured upper
airfoil nest M. The pusher pins 22 are sized to allow insertion of the pins 22
through the
locator apertures 134 formed in the upper composite skin 110.
Affixed in combination with the support structure 12 at the inboard and
outboard
ends of the contoured upper airfoil nest 14 are spar stanchions 24. The spar
stanchions 24
in operative combination define the spanwise alignment of the spar assembly
116/118 in
combination with the upper composite skin I 10, honeycomb core 114 combination
as
layed-up in the contoured upper airfoil nest 14. Also affixed in combination
with the
support structure 16 at each end thereof are a pair of locking members 26. One
or more
hard stops 28 also form part of the support structure 16 (see Figure 3).
The upper assembly 30 comprises a pressure bladder or bag 32, a contoured
backplate 34, and a structural support truss 36. The pressure bag 32 is sized
to the
spanwise and chordwise dimensions of the blade subassembly 132. For the
described
embodiment of the growth main rotor blade 100, the pressure bag 32 has
dimensions of
about 3 feet in the chordwise direction and about 24 feet in the spanwise
direction. The
pressure bag 32 is sealingly fastened to the contoured backplate 32, and is
pressurized
during the compaction process to develop the pressure forces required. to
compact the
blade subassembly 132. The contoured backplate 34 defines the OIv1L of the
lower airfoil
surface, i.e., the lower composite skin 112, and is operative to function as a
reaction
surface against the back pressures developed in the pressure bag 32 during the
compaction
process.
Due to the overall size of the pressure bag 32, large pressure forces are
developed
during pressurization thereof during the compaction process (within the range
of about
52,000 pounds to about 103,000 pounds total). The structural support truss 36
is
operative to counteract such large pressure forces to prevent damage to the
upper
assembly 30 during the compaction procedure. Each end of the support truss 36
includes
a pair of complementary locking members 38. When the upper assembly 30 is
lowered
9
CA 02521125 1995-06-21
(e.g., by means of a crane) into combination with the lower assembly 12, the
complementary locking members 38 interact with the locking members 26 to allow
the
upper and lower assemblies 30, 12 to be temporarily locked in combination,
e.g., by pinned
connections, to effectuate the compaction process. A pressure source 40 is
pneumatically
interconnected with the pressure bag 32 and operative to provide the
pressurizing gas to
pressurize the pressure bag 32 during the compaction process.
As is evident from an examination of Figures 2-3, the sheath
spreadinglinsertion
apparatus 50 is precisely co-located adjacent the lower assembly 12 of the
compaction
fixture 10 inasmuch as the compaction fixture 10 and the sheath
spreadingrnsertion
apparatus 50 have an interactive functional relationship during the
fabrication protocol of
the main rotor blade 100. The sheath spreading/insertion apparatus 50 is
operative, during
the compaction process effectuated by the compaction fixture 10, to insert the
prefabricated leading edge sheath 120 in combination with the blade
subassembly 132
layed-up in the compaction fixture 10. The sheath spreadingrnsertion apparatus
50
comprises an elongate stanchion 52 having a length corresponding to the span
of the
leading-edge sheath 120 that is movably supported by rolling members 54. The
rolling
members 54 interact with rails 56 secured in combination with an elongate
support
table 58 so that the stanchion 52 is movable with respect to the compaction
fixture 10.
The support table 58 is precisely positioned with respect to the lower
assembly 12 to
facilitate insertion of the leading edge sheath 120 in combination with the
blade
subassembly 132 during the compaction procedure.
The sheath spreading/'msertion apparatus 50 further includes upper and lower
elongate carriage members 60U, 60L that are mounted in movable combination
with the
elongate stanchion 52. A plurality of pneumatic cylinders 62U, 62L operatively
interconnect the respective carriage members 60U, 60L to the stanchion 52. A
pressure
source 64 is pneumatically interconnected to the pneumatic cylinders 62U, 62L
and
operative to provide pressurized air thereto for synchronized movement of the
carriage
members 60U, 60L with respect to the stanchion 52 between a disengaged
position, an
engaged position, and an operating position. While the embodiments of the
compaction
fixture 10 and the sheath spreading/'insertion apparatus 50 described herein
utilize separate
pressure sources 40, 64, it will be appreciated that a common pressure source
may be
utilized for the compaction fixture 10 and the sheath spreading/insertion
apparatus 50
according to the present invention in lieu of the separate pressure sources
40, 64 described
herein.
Pluralities of suction cups 66U, 66L are disposed in aligned combination,
i.e.,
rows, with the respective carriage members 60U, 60L along the spanwise length
thereof.
CA 02521125 1995-06-21
For the described embodiment, one hundred and eighty suction cups 66U, 66L are
mounted in combination with the respective carriage members 60U, 60L (ninety
suction
cups per carriage member). Each individual suction cup 66U, 66L has a bellowed
configuration (to facilitate engagement of the cups with the contours of the
respective
OML surfaces of the leading-edge sheath 120), and an outer diameter of about
two and
one-half inches. Each suction cup 66U, 66L is capable of exerting a suction
force of about
100 to about 175 pounds. Suction cups of the type manufactured by PIAB AB,
Akersberga, Sweden, may be used in practicing the present invention. The
individual
suction cups 66U, 66L are fluidically interconnected to a vacuum source 68
which
provides suction pressure therefor. The suction cups 66U, 66L are operative to
engage
and hold the respective OML surfaces of the leading edge sheath 120 with the
vacuum
source 68 actuated. Subsequent synchronized movement of the respective
carriage
members 60U, 60L away from one another to the operating position causes
spreading of
the sheath 120 to facilitate insertion thereof in combination with the blade
subassembly 132.
A tip end locator 70 is secured in combination with one end of the elongate
stanchion 52 and a plurality of leading edge stops 72 are secured in
combination with the
stanchion 52 (see particularly Figure 3A) along the length thereof. The tip
end locator 70
is operative to provide spanwise alignment of the leading-edge sheath 120
between the
upper and lower suction cups 66U, 66L to ensure proper insertion thereof in
combination
with the blade assembly 132. The leading-edge stops 70 are operative to ensure
that the
leading-edge sheath 120 is properly inserted between the upper and lower
suction
cups 66U, 66L so that the suction cups 66U, 66L can engage and hold the OML,
surfaces
of the leading-edge sheath 120.
A means 74 is provided for moving the elongate stanchion 52 along the rails 56
to
insert the leading-edge sheath 120 onto the blade subassembly 132 and for
moving the
stanchion 52 away from the compaction fixture 10 once the leading-edge sheath
120 is
inserted onto the blade subassembly 132 . For the described embodiment, the
means 74
comprises one or more screw jacks. Mounted on the stanchion 52 are one or more
complementary locator rods 76 that interact with the respective hard stops 28
of the
support structure 16 during movement of the stanchion 52 towards the
compaction
fixture 10. For the described embodiment, interaction between the locator rods
76 and the
hard stops 28 causes a displacement of the locator rods 76. Continued movement
of the
stanchion 52 towards the compaction fixture 10 causes a corresponding
displacement of
the locator rods 76 until a red band thereon becomes visible, indicating to
the operator of
the sheath spreading/'insertion apparatus 50 that the leading-edge sheath 120
has been
11
CA 02521125 1995-06-21
properly inserted onto the blade subassembly 132. One skilled in the art will
appreciate
that other means may be utilized to indicate that movement of the stanchion 52
should be
terminated inasmuch as the leading-edge sheath 120 has been properly inserted
onto the
blade subassembly 132. For example, the locator rods 76 and the respective
hard stops 28
could be functionally configured and positioned so that contact therebetween
automatically
terminates the operation of the moving means 74 such that the stanchion 52
ceases
moving.
The steps of an assemblage and compaction process 200 according to the present
invention are schematically illustrated in Figure 4A. The purpose of the
assemblage and
compaction process according to the present invention is to assemble the
components of
the main rotor blade 100 described hereinabove into a cure configuration. The
cure
configuration of the main rotor blade 100 is inserted into an autoclave (not
shown) for
final cure to form the finished main rotor blade assembly 100 exemplarily
illustrated in
Figure lA. The upper composite skin 110 and the honeycomb core 114 are
provided as a
prefabricated combination 110/114 for the initial step 202 of the described
embodiment of
the assemblage and compaction protocol. This is achieved by applying a
suitable film
adhesive to the honeycomb core 114 which is then mounted on the upper
composite
skin 110 and the combination 110/114 is then cured. In step 202, the
prefabricated
combination 110/114 is layed-up in the contoured upper airfoil nest 14 by
aligning the
trailing edge 104 of the upper composite skin 110 with the tooling pins 20 and
inserting
the locator apertures 134 of the upper composite skin 110 onto the pusher pins
22. While
providing the upper composite skin 110 and the honeycomb core 114 as a
prefabricated
combination 1l0/l14 simplifies the assemblage and compaction process according
to the
present invention, one skilled in the art will appreciate that the upper
composite skin 110
and the honeycomb core 114 may alternatively be individually layed-up in
combination
with the contoured upper airfoil nest 14 in separate sequential steps.
A suitable film adhesive is then applied in step 204 prior to lay-up of the
spar
assembly 116/118. The film adhesive may be applied directly to the spar
assembly 116/118, or alternatively, directly to the layed-up upper composite
skin 110. The
spar assembly 116/118 is then layed-up in combination with the layed-up upper
composite
skin, honeycomb core combination 110/114 in step 206. Proper chordwise
alignment of
the spar assembly 116/118 is achieved by abutting the spar assembly 116/118
against the
pusher pins 22 protruding through the upper composite skin 102. Proper
spanwise
alignment of the spar assembly 116/118 is achieved by ensuring that the ends
of the spar
assembly 116/118 abut the spar stanchions 24. A suitable adhesive is applied
to the
exposed surfaces of the honeycomb core 114 and the spar assembly 1161118 in
step 208,
12
CA 02521125 1995-06-21
and then the lower composite skin 112 is layed-up on adhesive-coated honeycomb
core 114 and spar assembly 116/118 in step 210. The assembled combination of
the upper
composite skin 110 and honeycomb core 114 combination, the spar assembly
116/118, and
the lower composite skin 112 define the blade subassembly 132 components
assembled in
the lower assembly 12 of the compaction fixture 10.
Prior to lowering and locking the upper assembly 30 in combination with the
lower
assembly 12, a caul plate 42 is preferably inserted between the spar
subassembly 132 and
the upper assembly 30 in step 212. The caul plate 42 is formed from a
plurality of
composite plies, e.g. for the described embodiment five to nine prepreg
fiberglass plies.
The caul plate 42 is configured to conform to the lower composite skin 112 of
the blade
subassembly 132 and is operative to provide an even pressure distribution over
the blade
subassembly 132 components during compaction thereof. After insertion of the
caul
plate 42 in optional step 212, the upper assembly 30 is lowered and locked in
combination
with the lower assembly 12 by means of pinned connections between the
respective
locking members 26, 38 in step 214. Close-out blocks (not illustrated)
preferably have
been inserted at the inboard and outboard ends 106, 108 of the blade
subassembly 132
prior to locking the upper and lower assemblies 12, 30 in combination. The
close-out
blocks are operative to prevent blade tip and root end round out during
compactiori of the
blade subassembly 132.
With the upper assembly 30 locked in combination with the lower assembly 12,
the
leading-edge segment (see reference numeral 138 in Figure 3) of the assembled
blade
subassembly 132 protrudes outwardly from the compaction fixture 10, i.e., is
not enclosed
by the compaction fixture 10. The pressure source 40 is actuated to pressurize
the
pressure bag 32 in step 216, which exerts pressure forces via the caul plate
42 (or directly
if the caul plate 42 is not utilized) to compact the assembled blade
subassembly 132. For
the described embodiment, the pressure bag 32 is pressurized so that pressure
forces
within the range of about 5 psi to about 10 psi are exerted against the
assembled blade
subassembly 132. The assembled blade subassembly 132 is subjected to
compaction
pressure for a predetermined compaction period in step 218. For the described
embodiment of the main rotor blade 100, the described assemblage and
compaction
process has a compaction period within the range of about ten minutes to about
fifteen
minutes.
During the compaction period the leading-edge sheath 120 is spread and
inserted
onto the exposed leading-edge segment 138 of the assembled blade subassembly
132 by
means of a sheath spreading and insertion process 300 according to the present
invention
as illustrated in Figure 4B. Prior to insertion of the leading-edge sheath
120, a suitable
13
CA 02521125 1995-06-21
adhesive is applied to the exposed leading-edge segment 138 in step P1
(alternatively the
adhesive may be applied to the IIviL surfaces of the prepreg composite plies
122 of the
leading-edge sheath 120).
With the upper and lower carriage members 60U, 60L in a disengaged position,
the
leading-edge sheath 120 is inserted between the upper and lower suction cups
66U, 66L in
step 302. The disengaged position of the carriage members 60U, 60L facilitates
such
insertion without any physical contact between the leading-edge sheath 120 and
the
suction cups 66U, 66L. Proper insertion of the leading-edge sheath 120 is
ensured by the
abutting engagement of the sheath 120 against the tip end locator 70 and the
leading edge
stops 72. The pressure source 64 is actuated to allow synchronized movement of
the
carriage members 60U, 60L to an engaged position in step 304 wherein the
suction
cups 66U, 66L abuttingly engage the leading-edge sheath 120.
The vacuum source 68 is then actuated, causing the suction cups 66U, 66L to
exert
suction forces against the respective OML surfaces of the leading-edge sheath
120 in
step 306. The suction forces exerted by the suction cups 66U, 66L are of
sufficient
strength that the leading-edge sheath 120 remains in engaged combination,
i.e., held, by
the suction cups 66U, 66L. The pressure source 64 is actuated to allow
synchronized
movement of the carriage members 60U, 60L back towards the open position to
the
operating position. Since the suction cups 66U, 66L are exerting suction
forces to hold
the leading-edge sheath 120, the synchronized movement of the carriage members
60U,
60L to the operating position causes the exposed segments 128 of the sheath
120 to
spread apart. For the described embodiment, the finished edges 128A (see
Figure 1 C) of
the exposed segments 128 of the sheath 120 are typically spaced apart by about
one and
one-half inches in the normal spaced-apart state. Due to the action of the
sheath
spreading/'insertion apparatus 50, the fcnished edges 128A of the exposed
segments 128 of
the sheath 120 are spread apart to a spread-apart condition defined by a
separation
distance of about two and one-half to three inches between the opposed
segments 128
when the carriage members 60U, 60L are moved to the operating position. The
spread-
apart configuration of the leading-edge sheath 120 facilitates insertion
thereof onto the
exposed segment 138 of the blade subassembly 132.
To insert the spread-apart leading-edge sheath 120, the stanchion moving means
74
is actuated to move the stanchion 52 towards the compaction fixture 10 for
insertion of the
spread-apart leading-edge sheath 120 in step 310. Proper insertion of the
leading-edge
sheath 120 is indicated by the visibility of the red band of the locator rods
76. With the
leading-edge sheath 120 fully inserted onto the exposed segment 138 of the
blade
subassembly 132, the hardpoints 136 of the counterweights 118 abuttingly
engage the INIl.
14
CA 02521125 1995-06-21
of the leading-edge sheath 120, the exposed segments 128 of the prepreg
composite
plies 122 underlie the upper and lower composite skins 110, 112, and the edges
of the
upper and lower composite skins 110, 112 abuttingly engage the respective
edges of the
titanium strips 124.
Once the leading-edge sheath 120 is fully inserted onto the exposed segment
138
of the blade subassembly 132, the 'pressure source 64 is actuated to cause
synchronized
movement of the upper and lower carriage members 60U, 60L to the engaged
position in
step 312. Such synchronized movement allows the spread-part leading-edge
sheath 120 to
return to its normal spread-apart state, i.e., for the described embodiment,
from a spread-
apart condition of about two and one-half to three inches to about one and one-
half inches.
In the normal spaced-apart state, the leading-edge sheath 120 exerts a
compaction force
against the blade subassembly 132 to facilitate adhesive bonding of the
leading-edge
sheath 120 in combination with the blade subassembly 132. The vacuum source 68
is shut
down in step 314, which terminates the suction forces exerted by the upper and
lower
suction cups 66U, 66L against the leading-edge sheath 120. The pressure source
is
actuated to cause synchronized movement of the carriage members 60U, 60L back
to the
disengaged position in step 316. Finally, in step 318 the stanchion 52 is
returned to the
starting position wherein the spreading and insertion process may be repeated
as required..
Once the compaction period has elapsed, the locking members 26, 38 are
unlocked, and a crane is utilized to remove the upper assembly 30 from the
lower
assembly 12 in step 222. The assembled main rotor blade 100 is then removed
from the
contoured upper airfoil nest 14. The assembled main rotor blade 100 is
subsequently
cured in an autoclave to form a finished main rotor blade assembly 100.
A variety of modifications and variations of the above-described apparatus and
methods for fabricating a main rotor blade are possible in light of the above
teachings. It is
therefore to be understood that, within the scope of the appended claims, the
present
invention may be practiced otherwise than as specifically described
hereinabove.
What is claimed is:
