Adhesive bonding and miscellaneous chemical manufacture – Methods – Surface bonding and/or assembly therefor
Reexamination Certificate
1998-02-04
2001-01-30
Heitbrink, Jill L. (Department: 1732)
Adhesive bonding and miscellaneous chemical manufacture
Methods
Surface bonding and/or assembly therefor
C156S212000, C156S245000, C264S162000, C264S229000, C264S250000, C264S258000, C181S292000
Reexamination Certificate
active
06179943
ABSTRACT:
FIELD OF THE INVENTION
This invention is directed to a thrust reverser assembly for a high bypass jet engine, and more specifically, a method for forming an acoustic panel for a thrust reverser assembly.
BACKGROUND OF THE INVENTION
Airplane manufacturers are under increasing pressure to produce lightweight, strong, and durable aircraft at the lowest cost for manufacture and lifecycle maintenance. An airplane must have sufficient structural strength to withstand stresses during flight, while being as light as possible to maximize the performance of the airplane. To address these concerns, aircraft manufacturers have increasingly used fiber-reinforced resin matrix composites.
These composites provide improved strength, fatigue resistance, stiffness, and strength-to-weight ratio by incorporating strong, stiff, carbon fibers into a softer, more ductile resin matrix. The resin matrix material transmits forces to the fibers and provides ductility and toughness, while the fibers carry most of the applied force. Unidirectional continuous fibers can produce anisotropic properties, while woven fabrics produce quasi-isotropic properties. Honeycomb core is often sandwiched between composite sheets to provide stiff panels having the highest specific strength.
Because of the noise regulations governing commercial transport aircraft, high bypass engines incorporate acoustic panels within the nacelles. Conventionally, these elements are made with an inner perforated skin, a surrounding buried septum honeycomb core, and a non-perforated outer skin, such as described in U.S. Pat. Nos. 4,600,619; 4,421,201; 4,235,303; 4,257,998; and 4,265,955, which we incorporate by reference. The inner and outer skins are metal, usually aluminum, or composite, and the honeycomb core is either aluminum or composite. Manufacturing these acoustic panels is a challenge because of their size, their complex curvature and the close tolerances necessary for them to function properly.
As shown in
FIG. 1
, a nacelle
10
for a commercial high bypass jet engine includes a thrust reverser assembly having a fore-and-aft translating sleeve
11
to cover or expose thrust reverser cascades
12
when deploying thrust reverser blocker doors
15
carried on the translating sleeve. The thrust reverser assembly is positioned just aft of a jet engine, not shown, as is used on an airplane. The thrust reverser assembly is fitted within the nacelle
10
. The thrust reverser cascades
12
are circumferentially spaced around the interior of the nacelle.
During normal flying operations the translating sleeve
11
is in a closed, or forward, position to cover the thrust reverser cascades
12
. For landing an airplane, the translating sleeve
11
is moved from the closed position to the rearwardly extended, or deployed, position by means of actuator rods
18
. This positioning routes exhaust gas to flow through the thrust reverser cascades
12
so as to slow down the aircraft on the ground. Exhaust is rerouted through the thrust reverser cascades
12
by closing the circumferentially positioned blocker doors
15
.
The translating sleeve
11
is usually formed from a pair of semi-cylindrical outer cowl panels
13
(only one shown in
FIG. 2
) and a pair of semi-cylindrical inner acoustic panels
14
(only one shown in
FIG. 2
) bonded together to form the aft portion of the cylindrical nacelle
10
. The outer cowl and acoustic panels
13
,
14
are bonded at their aft ends and branch or diverge to provide a chamber for containing and concealing the thrust reverser cascades
12
and the associated support structures.
When the translating sleeve
11
is in the stowed position (FIG.
2
), the leading ends of the acoustic panel
14
and the outer cowl panel
13
extend on opposite sides of the thrust reverser cascades
12
. When the thrust reverser is deployed, the translating sleeve
11
is moved aft to expose the cascades
12
(FIG.
3
). The fan duct blocker doors
15
at the forward end of the acoustic panel
14
are deployed to divert fan flow through the cascades
12
. The blocker door assembly is described in U.S. Pat. No. 4,852,805.
To form an acoustic composite sandwich panel, prior art methods used a male upper mandrel. The perforated composite inner skin was laid against lay-up surface of the mandrel and buried septum honeycomb core was laid over the inner perforated skin. A composite non-perforated skin was then laid over the honeycomb core, and the three layers were cured or co-cured so as to form a single part.
This method did not provide index control for the inner or outer surface of the perforated sheets. Inexact tolerances on the inner and outer surfaces made locating and attaching details on the inside or outside of the acoustic panel difficult.
Thermal residual stresses produced during the curing process caused acoustic panels to warp. Although the warpage was predictable to some extent, it was usually not uniform over the entire surface, leaving the part less than the desired design shape. Joining the acoustic panel and the outer cowl panel at a continuous aft joint with a smooth connection was difficult. Significant rework and shimming were required to correctly position the outer cowl panel and attach fittings against the outer side of the warped acoustic panel to complete the connection. Resin flowed into the perforations of the honeycomb core during the curing process, requiring rework of the perforated surface.
The acoustic panel must include recesses for receiving the fan duct blocker doors to provide a streamlined continuation during normal operation of the engine. In addition, deployment of the fan duct blocker doors imposes large bending moments at the leading end of the acoustic panel. As can be seen in
FIG. 2
, to receive the fan duct blocker doors
15
and to oppose the load of the fan duct blocker doors
15
when the thrust reverser is deployed, prior art acoustic panels
14
typically included a separate diaphragm
16
that was fastened at the leading end of the acoustic panel
14
, and reinforced by gussets (not shown). An aft ring
17
extended between the acoustic panel
14
and the diaphragm
16
. The blocker doors
15
were hinged from the leading end of the diaphragm
16
. A forward ring
18
, sometimes called a “bullnose ring”, extended from the leading edge of the diaphragm
16
toward the thrust reverser cascades
12
.
Because the gussets, the forward ring
18
, the aft ring
17
, and the diaphragm
16
were separate pieces, assembly of the thrust reverser and acoustic panel was laborious. The associated fasteners and connecting parts added significant weight to the acoustic panel
14
. The steep angles formed by forward and aft rings deterred anyone from trying to form them in a single piece. Reducing the number of parts will reduce assembly time and will improve performance because an integral part made to close tolerance with the method of the present invention will be more aerodynamically efficient.
The fan duct blocker doors
15
fold downward and fit within recesses oil the inner side of the acoustic panel. The blocker doors are trapezoidal so when stored, they create a triangular gap that needed to be filled for proper efficient airflow. The triangular gaps were filled by separate “wedge fairings” that were difficult to install with precision. The wedge fairings did not provide significant sound absorption. Attempts to form the wedge fairings integrally with the acoustic panel have not been successful.
The acoustic panels usually require reinforcement in the areas of attachment so that stresses applied by fittings attached to the acoustic panel will not damage the skins or core during sustained ultimate loads. To provide support at areas of fastened detail, prior art added plies to the skins to make the areas for fastening thicker than surrounding areas of the panel. The plies decreased in width so that the edges of the added plies formed “steps” or “ramps”. The ramps help to dissipate forces applied through the fasteners into the skin. The presence of extra composite material at the ramps m
Baca Leonard L.
Popp Thomas D.
Welch John M.
Christensen O'Connor Johnson & Kindness PLLC
Heitbrink Jill L.
Staicovici Stefan
The Boeing Company
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