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THE effects of elevated temperatures up to 600 F on the mechanical properties of common structural materials are treated briefly. The shortcomings of both short- and long-time exposure data are set forth, as are the inadequacies of present methods of analyzing certain members of aircraft structures. The changes in material properties at the higher temperatures and the manufacturing considerations relevant to material selection, such as cost, formabiliry, and availability, are mentioned as a part of the design problem in general as regards aircraft structures.

TWO principal methods for providing safety against catastrophic aircraft structural fatigue — safe-life and fail-safe — are treated. The author concludes that the safe-life method is generally inadequate, while the fail-safe method is practical and sound. Some test results of the fail-safe attributes of certain materials, stress levels, and design details are summarized, and some observations on fail-safe testing techniques are made. From a fail-safe standpoint, it is concluded that 24ST aluminum alloy is superior to 75ST, that certain stress levels and design features will fail safe, and that underwater pressure testing is probably unnecessary because dynamic over-stresses in fail-safe tests are probably small and fatigue cracks can be quickly simulated in the laboratory.

PROBLEMS resulting from increased temperature associated with flight at supersonci speeds, the serious effects of aerodynamic heating on material properties, and the ultimate strength of assembled structures are discussed in this paper. Ways are outlined which will alleviate the intrinsic weight penalties accompanying the use of high-density alloys at elevated temperatures. Primarily, this can be accomplished by the proper selection of materials and the choice of more suitable structural configurations. Suggestions are made how to lessen the peaks of thermal stresses induced by temperature gradients during the interval of transient heating. The need for efficient stabilization of compression panels subjected to thermal environment is pointed out in a general way; and recent innovations which might accomplish this purpose are then discussed with special regard to thin wing and fuselage design.

THE PROPERTIES of presently available and advanced development alloys are reviewed in each of the following groups: (1) magnesium alloys, (2) aluminum alloys, (3) titanium alloys, (4) low-alloy steels, (5) corrosion resistant steels, and (6) nickel, cobalt, and mixed-base alloys. The strongest materials in each group are selected to make comparisons of the strength of typical airframe components at various temperatures. From these comparisons, material-efficiency curves are made to show the strength and thermal limitations of each material group.

KNOWLEDGE of loads on major structural parts of an airplane while in flight is needed for complete utilization of the airplane's strength and capabilities. Some in-flight measuring techniques and equipment for reducing the data thus obtained are discussed in this paper. Instrumentation and calibration of strain gages for measuring in-flight air loads are described in detail. Wing bending-moment measurements obtained on a Navy PV-2 airplane are used to illustrate data reduction procedure.

STRUCTURAL MATERIALS for Mach 3 jet transports pose difficult problems for the design engineer. Reasons for this problem are the incomplete information available on the many possible metals and the diversity of critical properties that are added by supersonic requirements. The material properties discussed in this paper include tensile strength, resistance to crack propagation, ease of fabrication, weldability, and thermal expansion. Cost factors are also considered. The structural configuration of the wing and fuselage is an example of the complexity of the material selection problem. The wing may be rigidity-critical, and the fuselage strength-critical; each requires diferent material properties to solve the problem.*