Superplastically formed diffusion bonded metallic structure

A metallic sandwich structure particularly suited for use in aerospace industries comprising a base plate, a cover plate, and an orthogonally corrugated core is described. A pair of core plates formed of a superplastic alloy are interposed between the base plate and the cover plate and bonded. Each of the core plates is characterized by a plurality of protrusions comprising square-based, truncated pyramids uniformly aligned along orthogonally related axes perpendicularly bisecting the legs of the bases of the pyramids and alternately inverted along orthogonally related planes diagonally bisecting the pyramids, whereby an orthogonally corrugated core is provided

Accuracies of southwell and force/stiffness methods in the prediction of buckling strength of hypersonic aircraft wing tubular panels

Accuracies of the Southwell method and the force/stiffness (F/S) method are examined when the methods were used in the prediction of buckling loads of hypersonic aircraft wing tubular panels, based on nondestructive buckling test data. Various factors affecting the accuracies of the two methods were discussed. Effects of load cutoff point in the nondestructive buckling tests on the accuracies of the two methods were discussed in great detail. For the tubular panels under pure compression, the F/S method was found to give more accurate buckling load predictions than the Southwell method, which excessively overpredicts the buckling load. It was found that the Southwell method required a higher load cutoff point, as compared with the F/S method. In using the F/S method for predicting the buckling load of tubular panels under pure compression, the load cutoff point of approximately 50 percent of the critical load could give reasonably accurate predictions

Prediction of service life of aircraft structural components using the half-cycle method

The service life of aircraft structural components undergoing random stress cycling was analyzed by the application of fracture mechanics. The initial crack sizes at the critical stress points for the fatigue-crack growth analysis were established through proof load tests. The fatigue-crack growth rates for random stress cycles were calculated using the half-cycle method. A new equation was developed for calculating the number of remaining flights for the structural components. The number of remaining flights predicted by the new equation is much lower than that predicted by the conventional equation

Stress concentration around a small circular hole in the HiMAT composite plate

Anisotropic plate theory is used to calculate the anisotropic stress concentration factors for a composite plate (AS/3501-5 graphite/epoxy composite, single ply or laminated) containing a circular hole. This composite material is used on the highly maneuverable aircraft technology (HiMAT) vehicle. It is found that the anisotropic stress concentration factor could be greater or less than 3 (the stress concentration factor for isotropic materials), and that the locations of the maximum tangential stress points could shift with the change of fiber orientation with respect to the loading axis. The effect of hole size on the stress concentration factor is examined using the Point Stress Criterion and the Averaged Stress Criterion. The predicted stress concentration factors based on the two theories compared fairly well with the measured values for the hole size 0.3175 cm (1/8 in). It is also found that through the lamination process, the stress concentration factor could be reduced drastically, indicating an improvement in structural performance

Stress analyses of B-52 pylon hooks

The NASTRAN finite element computer program was used in the two dimensional stress analysis of B-52 carrier aircraft pylon hooks: (1) old rear hook (which failed), (2) new rear hook (improved geometry), (3) new DAST rear hook (derated geometry), and (4) front hook. NASTRAN model meshes were generated by the aid of PATRAN-G computer program. Brittle limit loads for all the four hooks were established. The critical stress level calculated from NASTRAN agrees reasonably well with the values predicted from the fracture mechanics for the failed old rear hook

Application of Finite Elastic Theory to the Deformation of Rubbery Materials

The purpose of this discussion, then, is to show how the nature of the strain energy function can be deduced from experiments on rubbery materials

Preflight transient dynamic analyses of B-52 aircraft carrying Space Shuttle solid rocket booster drop-test vehicle

This paper concerns the transient dynamic analysis of the B-52 aircraft carrying the Space Shuttle solid rocket booster drop test vehicle (SRB/DTV). The NASA structural analysis (NASTRAN) finite element computer program was used in the analysis. The B-52 operating conditions considered for analysis were (1) landing and (2) braking on aborted takeoff runs. The transient loads for the B-52 pylon front and rear hooks were calculated. The results can be used to establish the safe maneuver envelopes for the B-52 carrying the SRB/DTV in landings and brakings

Thermal response of Space Shuttle wing during reentry heating

A structural performance and resizing (SPAR) finite element thermal analysis computer program was used in the heat transfer analysis of the space shuttle orbiter that was subjected to reentry aerodynamic heatings. One wing segment of the right wing (WS 240) and the whole left wing were selected for the thermal analysis. Results showed that the predicted thermal protection system (TPS) temperatures were in good agreement with the space transportation system, trajectory 5 (STS-5) flight-measured temperatures. In addition, calculated aluminum structural temperatures were in fairly good agreement with the flight data up to the point of touchdown. Results also showed that the internal free convection had a considerable effect on the change of structural temperatures after touchdown

Thermal behavior of a titanium honeycomb-core sandwich panel

Finite element thermal stress analysis was performed on a rectangular titanium honecomb-core sandwich panel which is subjected to thermal load with a temperature gradient across its depth. The distributions of normal stresses in the face sheets and the face-sheet/sandwich-core interfacial shear stresses are presented. The thermal buckling of the heated face sheet was analyzed by assuming the face sheet to be resting on an elastic foundation representing the sandwich core. Thermal buckling curves and thermal buckling load surface are presented for setting the limit for temperature gradient across the panel depth

Reentry heat transfer analysis of the space shuttle orbiter

A structural performance and resizing finite element thermal analysis computer program was used in the reentry heat transfer analysis of the space shuttle. Two typical wing cross sections and a midfuselage cross section were selected for the analysis. The surface heat inputs to the thermal models were obtained from aerodynamic heating analyses, which assumed a purely turbulent boundary layer, a purely laminar boundary layer, separated flow, and transition from laminar to turbulent flow. The effect of internal radiation was found to be quite significant. With the effect of the internal radiation considered, the wing lower skin temperature became about 39 C (70 F) lower. The results were compared with fight data for space transportation system, trajectory 1. The calculated and measured temperatures compared well for the wing if laminar flow was assumed for the lower surface and bay one upper surface and if separated flow was assumed for the upper surfaces of bays other than bay one. For the fuselage, good agreement between the calculated and measured data was obtained if laminar flow was assumed for the bottom surface. The structural temperatures were found to reach their peak values shortly before touchdown. In addition, the finite element solutions were compared with those obtained from the conventional finite difference solutions