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NASA, Technical Note - Experimental Investigation of Rocket Engine Ablative Material Performance After Postrun Cooling at Altitude Pressures

Within the past few years, interest in using ablative materials for the
combustion chamber and the nozzle walls of rocket engines has intensified. ew
of the advantages of using ablative materials to maintain the physical integrity
of a rocket engine are increased versatility in varying the propellant flows by
elimination of regenerative cooling requirements, simplicity of the propellant-
flow systems, and rapidity of engine development.l

K The use of ablative materials in production engines has thus far been con—
fine to those portions of rocket engines that are difficult or impossible to
cool regeneratively, for example, engine throats or solid-propellant rocket noz—
zles. The materials used have been carbon, graphite, or refractory metals. Re-
cently, however, new ablative materials, which consist of mixtures of plastic
resin binder and silica or graphit fiber reinforcement, have been developed for
the fabrication of entire chambers (refs. 1 to 5).

One area of concern, which has only been lightly explored, is the effect of
low pressures, such as would be found in space, on ablative chambers, particular—
ly when the ablation material is still hot, following a run. The resins in these
proposed ablative materials start to decompose when their temperatures rise above
3000 to 4000 F at a pressure of 1 atmosphere. After a run, the wall temperature
of an ablative rocket engine can theoretically be greater than 15000 F. These
organic resins could possibly volatilize at an extreme rate until they cooled
below their decomposition temperatures. Such volatilization would theoretically
be even greater when in the vacuum conditions of space. Madorsky'and Straus
(ref. 4) have reported that a typical phenolic resin, when heated in a vacuum
furnace for 5 minutes, was volatilized 29 percent at 9500 F and was volatilized
47 percent at 14700 F. More material did not volatilize only because a carbona—
ceous char layer formed over the surface.

Previous efforts to investigate this problem by use of ablative materials
have consisted of heating small pieces electrically in a vacuum apparatus to tem—
peratures up to 3000 F (ref. 5) or of heating samples with a torch and then put—
ting them in a vacuum atmosphere (ref. 6). These test methods reported no im—
portant effects due to cooling in a vacuum as compared with cooling at atmos-
pheric pressure (less than 1 percent difference in weight loss). Such methods,
however, do not reproduce the actual conditions that may be encountered in a
full—size rocket engine operating in deep space.