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An externally deployable honeycomb structure is investigated with respect to crash energy management for light aircraft. The new concept utilizes an expandable honeycomb-like structure to absorb impact energy by crushing. Distinguished by flexible hinges between cell wall junctions that enable effortless deployment, the new energy absorber offers most of the desirable features of an external airbag system without the limitations of poor shear stability, system complexity, and timing sensitivity. Like conventional honeycomb, once expanded, the energy absorber is transformed into a crush efficient and stable cellular structure. Other advantages, afforded by the flexible hinge feature, include a variety of deployment options such as linear, radial, and/or hybrid deployment methods. Radial deployment is utilized when omnidirectional cushioning is required. Linear deployment offers better efficiency, which is preferred when the impact orientation is known in advance. Several energy absorbers utilizing different deployment modes could also be combined to optimize overall performance and/or improve system reliability as outlined in the paper. Results from a series of component and full scale demonstration tests are presented as well as typical deployment techniques and mechanisms. LS-DYNA analytical simulations of selected tests are also presented.

The feasibility of using scale model testing for predicting the full-scale behavior of flat composite coupons loaded in tension and beam-columns loaded in flexure is examined. Classical laws of similitude are applied to fabricate and test replica model specimens to identify scaling effects in the load response, strength, and mode of failure. Experiments were performed on graphite-epoxy composite specimens having different laminate stacking sequences and a range of scaled sizes. From the experiments it was deduced that the elastic response of scaled composite specimens was independent of size. However, a significant scale effect in strength was observed. In addition, a transition in failure mode was observed among scaled specimens of certain laminate stacking se quences. A Weibull statistical model and a fracture mechanics based model were applied to predict the strength scale effect since standard failure criteria cannot account for the in fluence of absolute specimen size on strength.

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The effect of specimen size upon the response and strength of ± 45° angle-ply laminates has been investigated for two graphite fiber-reinforced plastic systems and several stacking sequences. The first material system was an epoxy-based system, AS4 fibers in 3502 epoxy, and the second was a thermoplastic-based system, AS4 fibers in PEEK matrix. For the epoxy-based system, two generic ±45° layups were studied; ( + 45°n/-45°n)2s (blocked plies), and ( + 45°/-45°)2nS (distributed plies), where n = 1, 2, 3, and 4. In the case of the thermoplastic system, only the layup with distributed plies was investigated, ( + 45°/-45°)2nS, for n = 1 and 2. The in-plane dimensions of the specimens were varied such that the width/length relationship was 12.7 × n/127 × n mm, for n = 1, 2, 3, or 4. It is shown that the stress/strain response and the ultimate strength of these angle-ply laminates depends on the laminate thickness and the type of generic layup used. The ultimate strength of the epoxy matrix material was found to be much more sensitive to specimen size when compared to the thermoplastic matrix system. Scaling effects defined with respect to the first ply failure, strain at ultimate failure, and ultimate strength are isolated and discussed. Furthermore, it is shown that first ply failure occurs in the surface plies as a result of normal rather than shear stresses. The implications of the experimental findings upon the validity of the ±45° tension test, which is used to determine the in-plane shear response of unidirectional composites, are discussed.

A research program has been initiated to study and isolate the factors responsible for scale effects in the tensile strength of graphite/epoxy composite laminates.Four lay-ups, (±30° n/9002n)s, (±45° n/O° n/90° n)s, (90° nlO° n/90° n/O° n)s, and (±45° n/±45 ° n)s, have been chosen with appropriate stacking sequences so as to highlight individual and interacting failure modes. Four scale sizes have been selected for investigation including full scale size, 3/4, 2/4, and 1/4, with n equal to 4, 3,2, and 1, respectively. The full scale specimen size was 32 plies thick as compared to 24, 16, and 8 plies for the 3/4, 2/4, and 1/4 specimen sizes respectively.Results were obtained in the form of tensile strength, Stress/strain curves and damage development. Problems associated with strength degradation with increasing specimen size have been isolated and discussed. Inconsistencies associated with strain measurements have also been identified. Enhanced X-ray radiography was employed for damage evaluation, following step loading.It has been shown that fiber dominated lay-ups were less sensitive to scaling effects compared to the matrix dominated lay-ups. Further, it has been shown that fabrication induced damage was partly responsible for the observed behavior.Extrapolation to the full scale strength was attempted by means of three basic methods: a Weibull statistics based model, a fracture mechanics based model, and a combination model involving the previous two models in conjunction with a failure criterion. The predictive performance of each one of these models has been assessed and their applicability to the present problem has been discussed.

A combination of aerodynamic analysis and testing, aerothermodynamic analysis, structural analysis and testing, impact analysis and testing, thermal analysis, ground characterization tests, configuration packaging, and trajectory simulation are employed to determine the feasibility of an entirely passive Earth entry capsule for the Mars Sample Return mission. The design circumvents the potential failure modes of a parachute terminal descent system by replacing that system with passive energy absorbing material to cushion the Mars samples during ground impact. The suggested design utilizes a spherically blunted 45-degree half-angle cone forebody with an ablative heat shield. The primary structure is a hemispherical, composite sandwich enclosing carbon foam energy absorbing material. Though no demonstration test of the entire system is included, results of the tests and analysis presented indicate that the design is a viable option for the Mars Sample Return Mission.

A composite fuselage concept for light aircraft has been developed to provide improved crashworthiness. The fuselage consists of a relatively rigid upper section, or passenger cabin, including a stiff structural oor and a frangible lower section that encloses the crash energy management structure. The crashworthy performance of the fuselage concept was evaluated through impact testing of a one-fth-scale model fuselage section. The impact design requirement for the scale model fuselage is to achieve a 125-g average oor-level acceleration for a 31-ft/s vertical impact onto a rigid surface. The energy absorption behavior of two different sub oor con gurations was determined through quasi-static crushing tests. For the dynamic evaluation, each sub oor con guration was incorporated into a one-fth-scale model fuselage section, which was dropped from a height of 15 ft to achieve a 31-ft/s vertical velocity at impact. The experimental data demonstrate that the fuselage section with a foam-block sub oor con guration satis ed the impact design requirement. A second drop test was performed to evaluate the energy absorption performance of the fuselage concept for an off-axis impact condition. The experimental data are correlated with analytical predictions from a nite element model developed using the nonlinear, explicit transient dynamic code MSC/DYTRAN.