The X-15 offered investigators a unique opportunity to measure heat transfer and skin friction under quasi-steady flight conditions at high Mach numbers and low wall-to-recovery temperature ratios. This allowed them to make a direct comparison between measured flight data and calculated values. A considerable amount of heat-transfer data and some skin-friction data were obtained during the flight program, and these data indicated that the level and rate of change of turbulent skin friction and heat transfer were lower than predicted by the most widely used theories, such as those of Van Driest and Eckert. However, comparisons of the X-15 data and the theory were inconclusive due to uncertainties about the boundary layer conditions because of
non-uniform flow and conduction losses. To evaluate the problem, researchers wanted to use a highly instrumented panel in a location with known flow characteristics. They also wanted the panel to be shielded from aerodynamic heating until the airplane was in a steady-state cruise condition.
Researchers selected the X-15-3 with the sharp-leading-edge modification on the dorsal rudder to carry the experiment. The test panel was located just behind the right-side leading-edge boundary-layer trips 15.1 inches below the top of the rudder, and was constructed of 0.0605- inch-thick Inconel X. Researchers installed a removable panel on the left side of the rudder to provide access to the instrumentation used for the test panel. To obtain the desired wall-to – recovery temperature ratios and ensure an isothermal test surface when the airplane reached the desired speed and altitude, it was necessary to insulate the test panel during the initial phase of the flight. Explosive charges jettisoned the insulating cover from the test panel in approximately 50 milliseconds, resulting in an instantaneous heating of the test panel (the so-called "cold wall" effect). Researchers instrumented the test panel with thermocouples, static-pressure orifices, and a skin-friction gage with the data recorded on tape by a PCM data acquisition system at a rate of 50 samples per second. A Millikan camera operating at 400 frames per second was in the upper bug-eye camera bay to record the events. The measurements obtained were in general agreement with previous X-15 data.12101
Researchers used the same general location for another test panel, but without the cold wall. This panel, which was flush with the normal surface of the rudder, had a microphone and static – pressure orifice mounted flush, and an "L"-shaped total pressure probe sticking out and forward. The microphone was located 28.8 inches from the original rudder leading edge (not the sharp extension) and 20.3 inches from the top of the rudder. The data was recorded onboard the airplane and evaluated after the flight. The intent of the experiment was to determine when the boundary layer transitioned to turbulent flow. The highest noise levels occurred during reentry as the Reynolds numbers reached their peak value. The data gathered provided a qualitative indication of the end of the transition that agreed reasonably well with wind-tunnel data. Interestingly, researchers also recorded some data while the NB-52 carried the X-15, and described the noise level as "very high" due to aerodynamic interference with the carrier aircraft. This confirmed predictions made before the first glide flight.12111