Journal of Aerospace Engineering

A structural analysis is performed for the rim support of a pressurized off‐axis paraboloidal membrane that serves as a space‐based solar concentrator. The function of the elliptical rim support is to take up the tensile forces created by the stretched membrane. This paper extends earlier load and deformation analyses to include the effect of a tension element added along the major axis of the ellipse. Such an element, in the form of a string or cable, restrains the tendency for increased ovality of the rim support under load without seriously interfering with the packaging and deployment of the device, at negligible added mass. The internal forces and moments and the deformations resulting from the membrane‐applied loads have been calculated. The analysis has shown that a properly designed tension element can reduce the maximum bending moment by more than a factor of 2. The deformations are reduced by more than a factor of 20.

An optimization procedure has been developed to address the complex and conflicting requirements associated with the design of high‐speed proprotor aircraft. Since the key technical issues are maintenance of propulsive efficiency and aeroelastic stability in high‐speed cruise without deteriorating figure of merit in hover, rotor aerodynamic performance and aeroelastic analysis are coupled, inside a closed loop, to the optimizer. The discipline couplings provide actual blade air‐loads, during hover and cruise, and also provide realistic blade designs. The propulsive efficiency in high‐speed cruise is used as the objective function. Constraints are also imposed on the aeroelastic stability in axial flight and the rotor figure of merit in hover. Both structural and planform design variables are used. The optimization procedure yields significant improvements in the aerodynamic characteristics of the rotor. Off‐design performance studies, conducted with the optimum blade, show overall design improvements.

Various multitasking approaches are investigated for optimization of large space structures. Judicious combination of microtasking, macrotasking, and autotasking is explored with the goal of achieving a vectorized and multitasked algorithm for optimization of a large structure with maximum speedup performance. Speedup results are presented and compared for three space truss structures, with 526, 1,046, and 3,126 members. The processing time required for optimization of large structures increases exponentially with the size of the structure (number of design variables). Example 3 of this paper has 3,126 members and 2,226 displacement degrees of freedom. Development of efficient concurrent algorithms utilizing the unique architecture and capabilities of high‐performance computers results in substantial reduction in the overall execution time.

Soil plays an important role in the construction of foundations of roads and buildings. The utilization of soil traditionally involves the compaction of the in‐situ or previously loosened material to achieve a desired strength. The paper reports an initial investigation on the effect of the percentage of fines on the densification and strength of lunar‐soil simulant. The results of vibratory and static compaction tests in the laboratory suggest that the amount of fines should be reduced from the existing 50% that is typical for lunar regolith. The measurement of density and cone resistance of soil mix with only 10% showed large differences when compared to soils with 30% or 50% fines. No attempt has been made to find optimal soil mixes for compaction.

The development of design criteria for lunar structures must begin soon in order to establish adequate criteria. Some of the items that need consideration in such criteria are discussed. The categorization of the structures will provide designers with information on the purpose and level of complexity of the structure. Various construction materials and structure types that will be critical for the design of lunar structures, are considered. The environment of the moon and its possible effects on structures are presented and lead to the development of a few load cases that need to be considered in design. A probabilistic format for the criteria and design lifetimes are also discussed.

This project's goal was to evaluate the feasibility of pneumatic transfer for the movement of regolith (lunar soil) at a lunar base. Operation of pneumatic conveying systems at partial (lunar and Mars) gravity conditions on NASA's KC‐135 aircraft allowed the determination of some key parameters necessary for the design of an operable system. Both horizontal and vertical transfer were studied. In the vertical experiment, the choking velocity for 150 μm glass spheres at lunar gravity was determined to be 1/2–1/3 the velocity required at 1 g. Pressure drops were reduced by roughly the same amount. Choking velocity and pressure‐drop data for runs at Mars gravity were problematic due to the intermittent nature of the experiment. Determination of the saltation velocity in the horizontal run was also problematic, but qualitatively similar results were obtained. Comparison of the partial g results to 1 g behavior and theoretical analysis is made.

A preliminary design concept for a weldable joint for on‐orbit assembly of large space structures is described. The joint was designed for ease of assembly, for structural efficiency, and to allow passage of fluid (for active cooling or other purposes) along the member through the joint. The members were assumed to consist of graphite/epoxy tubes to which were bonded 2219‐T87 aluminum alloy end fittings for welding on‐orbit to nodes of the same alloy. A modified form of gas tungsten arc welding was assumed to be the welding process. The joint was designed for the thermal and structural loading associated with a 37 m diameter tetrahedral truss intended as an aerobrake for a mission to Mars. It was concluded that the assembly process could lock large loads into the truss members and that the assembly robot could be required to exert large forces while aligning pairs of nodes during assembly. It was also concluded that the connections between the composite struts and the aluminum fittings will be subjected to very high service stresses due to the effects of differential thermal expansion.

The use of inflatable structures has often been proposed for aerospace applications. The advantages of such structures include low launch weight and easy assembly. To consider using inflated fabric beams in aerospace applications, an accurate and efficient method of structural analysis must be developed. This paper develops a method of analysis for inflated fabric beams that is analogous to the shear‐moment method. The bending model is developed from basic assumptions about the state of stress in the fabric of the inflated beam. The resulting model shows that the bending behavior of inflatable fabric beams is identical to conventional solid elastic beams as long as the beam fabric remains unwrinkled. Experimental data were obtained and verify the static bending model for beams with length/diameter ratios greater than 10.

Damage in composite laminates caused by low‐velocity impact may produce significant reductions in compressive strength. Lockheed conducted an experimental program to investigate the damage‐containment capability of simple composite laminates of two graphite‐epoxy systems: T300/5208 and AS/3501‐6. Four different lay‐ups, including two hybrid laminates, were investigated for each material system. Results are presented for comparison. Laminates with Kevlar layers, especially where the layer is on the surface of the laminate, demonstrate better impact resistance.