RCC fabrication begins with
a rayon cloth graphitized and impregnated with a phenolic resin.
This impregnated cloth is layed up as a laminate and cured in
an autoclave. After being cured, the laminate is pyrolized to
convert the resin to carbon. This is then impregnated with furfural
alcohol in a vacuum chamber, then cured and pyrolized again to
convert the furfural alcohol to carbon. This process is repeated
three times until the desired carbon-carbon properties are achieved.
To provide oxidation resistance for reuse capability, the outer
layers of the RCC are converted to silicon carbide. The RCC is
packed in a retort with a dry pack material made up of a mixture
of alumina, silicon and silicon carbide. The retort is placed
in a furnace, and the coating conversion process takes place in
argon with a stepped-time-temperature cycle up to 3,200º F. A diffusion
reaction occurs between the dry pack and carbon-carbon in which
the outer layers of the carbon-carbon are converted to silicon
carbide (whitish-gray color) with no thickness increase. It is
this silicon-carbide coating that protects the carbon-carbon from
oxidation. The silicon-carbide coating develops surface cracks
caused by differential thermal expansion mismatch, requiring further
oxidation resistance. That is provided by impregnation of a coated
RCC part with tetraethyl orthosilicate. The part is then sealed
with a glossy overcoat. The RCC laminate is superior to a sandwich
design because it is light in weight and rugged; and it promotes
internal cross-radiation from the hot stagnation region to cooler
areas, thus reducing stagnation temperatures and thermal gradients
around the leading edge. The operating range of RCC is from minus
250º F to about 3,000º F. The RCC is highly resistant to fatigue
loading that is experienced during ascent and entry.
The RCC panels are mechanically attached to the wing with a series
of floating joints to reduce loading on the panels caused by wing
deflections. The seal between each wing leading edge panel is
referred to as a T-seal. The T-seals allow for lateral motion
and thermal expansion differences between the RCC and the orbiter
wing. In addition, they prevent the direct flow of hot boundary
layer gases into the wing leading edge cavity during entry. The
T-seals are constructed of RCC.
Since carbon is a good thermal conductor, the adjacent aluminum
and the metallic attachments must be protected from exceeding
temperature limits by internal insulation. Inconel 718 and A-286
fittings are bolted to flanges on the RCC components and are attached
to the aluminum wing spars and nose bulkhead. Inconel-covered
cerachrome insulation protects the metallic attach fittings and
spar from the heat radiated from the inside surface of the RCC
The nose cap thermal insulation ues a blanket made from ceramic
fibers and filled with silica fibers. HRSI or FRCI tiles are used
to protect the forward fuselage from the heat radiated from the
hot inside surface of the RCC.
During flight operations, damage has occurred in the area between
the RCC nose cap and the nose landing gear doors from impact during
ascent and excess heat during entry. The HRSI tiles in this area
are to be replaced with RCC.
In the immediate area surrounding the forward orbiter/ET attach
point, an AB312 ceramic cloth blanket is placed on the forward
fuselage. RCC is placed over the blanket and is attached by metal
standoffs for additional protection from the forward orbiter/ET
attach point pyrotechnics.