Heat-Sink Structures

Prior to the X-15 flight-test program, there were several theories pre­dicting the amount of friction heat that would transfer to the surface of a winged aircraft, with substantial differences in the theories. Wind tunnels, ballistic ranges, and high-temperature facilities such as arc-jets were unable to adequately duplicate the flight environment necessitating

full-scale flight test to determine which theory was correct. The X-15 was that test aircraft. The design needed to be robust in order to survive the worst-case heating predictions if the theories proved to be correct.

A heat-sink structure was selected as the safest and simplest method for providing thermal protection. Inconel X, a nickel alloy normally used for jet engine exhaust pipes, was selected as the primary structural mate­rial. It maintained adequate structural strength to about 1,200 °F. The design proceeded by first defining the size of each structural member based on the air loads anticipated during entry, then increasing the size of each member to absorb the expected heat load that would occur dur­ing the short exposure time of an X-15 flight.

As with most of the first missile and aircraft explorations, early hyper­sonic flights in the X-15 showed that none of the prediction methods was completely accurate, although each method showed some validity in a cer­tain Mach range. In general, the measured heat transfer was less than pre­dicted. Thus, one of the most significant flight-test results from the X-15 program was development of more accurate prediction methods based upon real-world data for the thermal protection of future hypersonic and entry vehicles.[750] The majority of aerodynamic heating issues that required attention during the X-15 flight-test program were associated with local­ized heating: that is, unexpected hot spots that required modification. Some typical examples included loss of cockpit pressurization because of a burned canopy seal (resolved by installing a protective shield in front of the canopy gap), cockpit glass cracked because of deformation of the glass retainer ring (resolved by increasing the clearance around the glass), wing skin buckling behind the slot in the leading edge expansion joint (resolved by installing a thin cover over the expansion joint), thermal expansion of the fuselage triggering nose gear door deployment with resulting damage to internal instrumentation (resolved by increasing the slack in the deploy­ment cable), and buckling of skin on side tunnel fairings because of large temperature difference between outer skin and liquid oxygen (LOX) tank (resolved by adding expansion joints along the side tunnels).

Most of these issues were discovered and resolved fairly easily since the flight envelope was expanded gradually on successive flights with small increases in Mach number on each flight. Had the airplane been exposed to the design entry environment on its very first flight, the

combined results of these local heating problems would probably have been catastrophic.