The X-15 was the first manned aircraft in which aerodynamic heating was the dominant problem of structural design. It was, of course, out of the question in 1954 to hope to be able to identify an optimum structural concept representative of distant future developments and no attempt was made to do so. The approach taken was to utilize the concept best suited to the short duration X-15 mission itself, with its particular requirements for heat-transfer research for drastically different trajectories ranging from equilibrium glides to steep high-angle-of-attack space reentries. A thick-skinned heat-sink approach was adopted. This has proved satisfactory both as a structure and as a calorimeter for heat-transfer measurements. Aside from its relatively thick. heat-sink skin, the X-15 structure contains many features representative of current advanced concepts for Mach 6-8 cruise aircraft, such as corrugated shear webs, combined use of superalloys and materials of a lower expansion coefficient such as Titanium, and the use of segmented leading edges. In retrospect, the major advances in the area of primary structure occurred during design and development. Great reliance was placed on the results of structural tests in which heat was applied electrically and loads mechanically according to schedules representing actual flight environmental histories. Measurements of the behavior of the primary structure in actual flight have verified these ground simulations. Thus the X-15 confirmed that ccMplex high-temperature structures can be reliably developed with ground-based "partial simulation" techniques using only rather simple and inexpensive equipment. Here again, early notions that a true-temperature aerodynamic type of structures test facility would be required were dispelled by the X-15 findings.

Although no surprises were found in flight for the primary structure, many unanticipated problems came to light in the secondary structures. Early in the program the pilot reported a rumbling noise at high dynamic pressures. This turned out to be panel flutter of large areas of the skin on the side fairings and tails. It was found to be related to certain design features incorporated to reduce thermal stress. Solution of this problem produced important advances in flutter prediction criteria (ref. 20).

Figure 7 - wing skin buckle during flight and corrective modification made
Figure 7. Wing skin buckle on flight to Mach 5.3 and corrective modification.

Figure 8 - Temperature distribution of the leading-edge expansion slot
Figure 8. Temperature distribution aft of leading-edge expansion slot.

Unanticipated leading-edge distortions and buckling of the adjacent wing skin developed during the first flight to Mach 5.3 (fig. 7). The slots, perhaps through vortex action and partly through triggering transi tion, cause intense local heating. Covering the slot and providing an internal shear tie between segments solved this problem. Evidence of the local heating is seen in figure 8 (ref. 5).

Figure 9 - cracked windshield
Figure 9. Cracked windshield glass in flight to Mmax= 6.04.

Figure 10 - X-15 windshield retainer modifications
Figure 10. X-15 windshield retainer modifications.

Although the soda-lime glass windshields of the X-15 were designed conservatively from the standpoint of glass temperatures, a failure occurred, fortunately involving only the right outer pane (fig. 9). Close examination of the failure revealed that it was initiated by buckling of the retainer frame. This problem was solved by replacing the Inconel X frame with Titanium, for which buckling does not occur because of the lower expansion coefficients of Titanium (fig. 10). The aft portion of the frame was removed because it was apparently causing a shock-induced hot spot on the aft part of the window (ref. 7). These fixes have proved successful.

One could argue that these local problems could have been solved by more comprehensive ground testing and analysis of the local flow field and heating mechanisms. True, if someone had had the foresight to identify these new problems in advance of their disclosure by the flight tests. The really important lesson here is that what are minor and unimportant features of a subsonic or supersonic aircraft must be dealt with as prime design problems in a hypersonic airplane. This lesson was applied effectively in the precise design of'a host of important details on the manned space vehicles.

Figure 11 - wing covered with spray-on ablator prior to sanding
Figure 11. Wing covered with spray-on ablator prior to sanding.

During the past year one of the airplanes has been flown with a radical change in the structural concept (ref. 10). The airplane was covered with ablative insulation designed to permit flights to Mach 7.4 (fig. 11). A silicone elastomeric ablator was sprayed on in variable thickness appropriate to the local heat loads. Leading edges were protected by a related erosion-resistant material applied in preformed sections. This program was of interest to lifting reentry vehicle development, because the spray-on technique has often been advanced as a possible means of refurbishing the incremental protection required for metal reentry structures. In the initial phases of a lifting reentry where the heating rates are beyond the limits of the metallic radiation cooled structure, the material would ablate. Then, as the speed reduces the bare metal would be exposed and radiation cooling would take over.

The X-15 experience rather clearly shows that this approach to a refurbishable structure is impractical. Some 5 weeks time and over 2000 hours were required for the total job including special treatment of removable hatches, covers, etc., hardly a practical operation for a refurbishable logistics vehicle.

Performance of the material was generally satisfactory on flight which reached Mach 6.7. Charring occurred only in regions of highest heating, and a few local failures occurred where the thickness of the material was inadequate or where bond failures occurred due to back surface heating.

Figure 12 - Telemetry antenna showing effect of bow shock heating on fuselage
Figure 12. Telemetry antenna showing effect of bow
shock heating on fuselage, Mmax= 6.7.

A most interesting by-product of the ablative insulation flights was the revelation of a number of heating phenomena not detectable on previous flights with the metal heat-sink structure. The ablator tended to show up areas of intense heating by charring. Many of these areas did not become very hot in the bare-metal flights because of heat-sink cooling, but the ablation material blocked the cooling effect and showed up the hot spots. An example is the high-pressure heating zone behind the detached bow shock of the telemetry antenna (fig. 12).

Figure 13 - cavity heating revealed by charring of ablator
Figure 13. Cavity heating revealed by charring of ablator, Mmax= 6.7.

Figure 14 - tail-fuselage juncture heating
Figure 14. Tail-fuselage juncture heating, Mmax= 6.7.

Figure 15 - local heating on wing
Figure 15. Local heating on wing due to
impingement of fuselage bow shock and side-fairing shock.

Figure 16 - dummy ramjet engine installation
Figure 16. Dummy ramjet engine installation.

Cavity heating is evident in figure 13. On this flight the reaction controls were not used and the nozzles were filled with recessed plugs. Tail-body juncture heating (fig. 14) and shock-intersection heating (fig. 15) have been the subject of many wind tunnel studies but have not been seen previously in flight. A far more serious example of impingement heating occurred under the airplane where a dummy ramjet nacelle was mounted (fig. 16).

Figure 17 - failure of pylon due to interference heating from dummy ramjet
Figure 17. Failure of pylon due to interference
heating from dummy ramjet.

The supporting pylon was subject to strong shock fields from the spike and from the closed cowl inlet. Heat protection material was applied to the leading edge of the pylon in roughly the same thicknesses as for the wing leading edge but the complex local heating phenomena on the pylon were not accounted for in the design. A serious failure of the system occurred with actual burn-through of the pylon skin (fig. 17). Pylon leading-edge heating rates of the order of seven times those without interference were estimated. This is the closest the X-15 ever came to a major structural failure in flight due to heating. Again these results underscore the need for maximum attention to aerothermodynamic detail in design and preflight testing.

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