Journal of Aerospace Engineering

A key element of the President’s Vision for Space Exploration is the development of a new space transportation system to replace Shuttle that will enable manned exploration of the moon, Mars, and beyond. The National Aeronautics and Space Administration has created the Constellation Program to develop this architecture, which includes the Ares launch vehicle and Orion manned spacecraft. The Orion spacecraft must carry six astronauts and its primary structure should be reusable, if practical. These requirements led the Constellation Program to consider a baseline land landing on return to earth. To assess the landing system options for Orion, a review of current operational parachute landing systems such as those used for the F-111 escape module and the Soyuz is performed. In particular, landing systems with airbags and retrorockets that would enable reusability of the Orion capsule are investigated. In addition, Apollo tests and analyses conducted in the 1960s for both water and land landings are reviewed. Finally, test data and dynamic finite-element simulations are presented to understand land landings for the Orion spacecraft.

A review of astronaut whole body impact tolerance is discussed for land or water landings of the next generation manned space capsule named Orion. LS-DYNA simulations of Orion capsule landings are performed to produce a low, moderate, and high probability of injury. The paper evaluates finite-element (FE) seat and occupant simulations for assessing injury risk for the Orion crew and compares these simulations to whole body injury models commonly referred to as the Brinkley criteria. The FE seat and crash dummy models allow for varying the occupant restraint systems, cushion materials, side constraints, flailing of limbs, and detailed seat/occupant interactions to minimize landing injuries to the crew. The FE crash test dummies used in conjunction with the Brinkley criteria provides a useful set of tools for predicting potential crew injuries during vehicle landings.

Prior to the Columbia accident, quantitative impact assessment tools were not available to analyze debris impacts onto the Shuttle’s thermal protection system. Following the accident, the Columbia Accident Investigation Board recommended changes to increase the safety of future shuttle flights; one component was the development of physics-based analytical capabilities to evaluate damage due to debris impacts. This paper will present an overview of real time debris assessment impact analysis conducted by the Boeing Philadelphia Advanced Structural Analysis Impact Analysis Team in support of Space Shuttle missions since Return to Flight, the first mission after the Columbia accident. Specifically, analyses performed in support of missions STS-114, 121 and 117 will be presented. For each of these cases, an overview of the structural and material model development will be provided, and results of each analysis will be presented followed by a discussion of how the results lead into real time mission decisions. This work illustrates the importance of maintaining a physics-based real-time analysis capability as a vital instrument in supporting the safety of future spaceflight missions.

This paper provides an overview of rotorcraft crashworthiness research being conducted at the National Aeronautics and Space Administration Langley Research Center under sponsorship of the Subsonic Rotary Wing aeronautics program. The research is focused in two areas: development of an externally deployable energy attenuating concept and improved prediction of rotorcraft crashworthiness. The deployable energy absorber (DEA) is a composite honeycomb structure, with a unique flexible hinge design that allows the honeycomb to be packaged and remain flat until needed for deployment. The capabilities of the DEA have been demonstrated through component crush tests and vertical drop tests of a retrofitted fuselage section onto different surfaces or terrain. The research on improved prediction of rotorcraft crashworthiness is focused in several areas including simulating occupant responses and injury risk assessment, predicting multiterrain impact, and utilizing probabilistic analysis methods. A final task is to perform a system-integrated simulation of a full-scale helicopter crash test onto a rigid surface. A brief description of each research task is provided along with a summary of recent accomplishments.

A ballistic impact test program was conducted to provide validation data for the development of numerical models of blade out events in fabric containment systems. The impact response of two different fiber materials—Kevlar 49 (E.I. DuPont Nemours and Company) and Zylon AS (Toyobo Co., Ltd.) was studied by firing metal projectiles into dry woven fabric specimens using a gas gun. The shape, mass, orientation, and velocity of the projectile were varied and recorded. In most cases the tests were designed such that the projectile would perforate the specimen, allowing measurement of the energy absorbed by the fabric. The results for both Zylon and Kevlar presented here represent a useful set of data for the purposes of establishing and validating numerical models for predicting the response of fabrics under conditions simulating those of a jet engine blade release situation. In addition some useful empirical observations were made regarding the effects of projectile orientation and the relative performance of the different materials.

The development of a robust and reliable material model for fabrics used to prevent fan blade-out events in propulsion engines has significant importance in the design of fan-containment systems. Currently, Kevlar is the only fabric approved by the Federal Aviation Administration to be used in fan-containment systems. However, very little work has been done in building a mechanistic-based material behavior model, especially one that can be used to quantify the behavior of Kevlar when subjected to high-velocity projectiles. Experimental static and high strain rate tensile tests have been conducted at Arizona State University to obtain the material properties of Kevlar fabric. In this paper we discuss the development and verification of a constitutive model for dry fabrics for use in an explicit finite-element program. Results from laboratory tests such as tension tests including high-strain rate tests, picture frame shear tests, and friction tests yield most of the material properties needed to define a constitutive model. The material model is incorporated in the LS-DYNA commercial program as a user-defined subroutine. The validation of the model is carried out by numerically simulating actual ballistic tests conducted at NASA-GRC and fan blade out tests conducted at Honeywell Aerospace (Propulsion Engines).

In accordance with the FAA certification requirements, all modern commercial turbofan engines must successfully demonstrate its ability to withstand a fan blade-out (FBO) event through actual test. Possibility of losing a rotating fan blade from a running engine is a flight safety consideration, which must be addressed during the design phase of the engine. A typical fan blade-out event involves very complex nonlinear transient dynamics with large deflection of the release blade and rigid body rotation of the trailing blade as well as progressive failure and fragmentation of various components. Due to the nature of the impact type loading, the solution to the problem should also address dependence of the material behavior such as yield strength as a function of strain rates. In short, the transient dynamic analysis of a fan blade-out event highlights the complexity of the numerical technique, which includes all the nonlinearities of structural dynamics: plastic behavior of the materials, large displacements, and contact interaction between structural elements. In this paper, we present the results of a LS-DYNA simulation of a FBO event on a full-engine analytical model, which covers both the primary as well as secondary damages.

Carbon fiber composites that utilize flattened, large tow yarns in woven or braided forms are being used in many aerospace applications. The complex fiber architecture and large unit cell size in these materials present challenges for both understanding the deformation process and measuring reliable material properties. In this paper composites made using flattened 12k and 24k (referring to the number of fibers in the fiber tow) standard modulus carbon fiber yarns in a $0°∕+60°∕-60°$ triaxial braided architecture are examined. Standard straight-sided tensile coupons were tested with the 0° axial braid fibers either parallel to (axial tensile test) or perpendicular to (transverse tensile test) the applied tensile load. The nonuniform surface strain resulting from the triaxial braided architecture was examined using photogrammetry. Local regions of high strain concentration were examined to identify where failure initiates and to determine the local strain at the time of failure initiation. Splitting within fiber bundles was the first failure mode observed at low to intermediate strains. For axial tensile tests the splitting was primarily in the $\pm 60°$ bias fibers, which were oriented 60° to the applied load. At higher strains in the axial tensile test, out-of-plane deformation associated with localized delamination between fiber bundles or damage within fiber bundles was observed. For transverse tensile tests, the splitting was primarily in the 0° axial fibers, which were oriented transverse to the applied load. The initiation and accumulation of local damage caused the global transverse stress-strain curves to become nonlinear and caused failure to occur at a reduced ultimate strain for both the axial and transverse tensile tests. Extensive delamination at the specimen edges was also observed. Modifications to the standard straight-sided coupon geometry are needed to minimize these edge effects when testing the large unit cell type of material examined in this work.

Metal failure in many applications, such as ballistic impact, containment, shielding, metal forming, and crashworthiness, occurs while the material is in a three-dimensional state of stress. Many previous definitions of triaxiality use two invariants to define the relative stress state in a virtual element, leading to a characterization that can be better thought of as biaxial. In this paper, an additional parameter based upon the third stress invariant is defined, which extends the characterization of the state of stress to three dimensions and to true triaxiality. The relation of the two parameters is explored and limits are found in the failure surface, which is used in defining the critical failure regions. Standard tests are examined to determine if they can provide enough data to construct these regions of interest and new tests are proposed, which envelope the limits and thus define this failure surface.

Uncontained aircraft engine failure can cause catastrophic damaging effects to aircraft systems if not addressed in the aircraft design. Mitigating the damaging effects of uncontained engine failure and improving the numerical modeling capability of these uncontained engine events are crucial. In this paper, high strain rate material behavior of one of the most extensively used materials in the aircraft industry is simulated and the results are compared against ballistic impact tests. Ballistic limits are evaluated by utilizing explicit finite-element (FE) simulations based on the corresponding ballistic impact experiments performed at different material thicknesses. LS-DYNA is used as a nonlinear explicit dynamics FE code for the simulations. A Johnson–Cook material model with different sets of parameters is employed as a thermo-viscoplastic material model coupled with a nonlinear equation of state and an accumulated damage evaluation algorithm for the numerical simulations. Predictive performance of the numerical models is discussed in terms of material characterization efforts, material model parameters, mesh sensitivities, and effects of stress triaxiality. It is shown that mesh refinement does not necessarily provide better results for ballistic limit simulations without considering and calibrating these interrelated factors. Moreover, it is shown that current models that can only fit a specific function for damage evaluation as a function of stress triaxiality are not always successful in predicting failure, especially if the state of stress changes significantly.

A framework is presented for analyzing the inelastic behavior and fracture of polymer matrix composites. Physics-based viscoplastic constitutive equations are used and supplemented with a matrix cracking model and an energy-based debonding model. The capabilities of the framework are illustrated by finite-element solutions of boundary-value problems under plane strain conditions using the unit-cell concept. In the application, focus is centered on the effect of manufacturing induced voids on local modes of fracture under both tension and compression.

For polymer matrix composites subjected to large strain rates, it is important to correctly characterize the nonlinear and strain-rate dependent response of polymers. Viscoplastic constitutive material models have been developed to account for the effects of hydrostatic effects and inelastic strains in polymer materials. The effective implementation of such viscoplastic models is important for development of composite models geared toward practical applications. Goldberg’s polymer model numerical implementation into a commercial finite-element code constitutes the main objective of this paper. Special attention is given to the use of effective algorithms for solving the model nonlinear rate dependent viscoplastic equations. Existent experimental data are used to verify the accuracy and robustness of the computational polymer model. A phenomenological fiber model and a simplified iso-strain mixture theory used to obtain the resultant stresses in the composite by averaging the stresses of the individual constituents are also defined. The validation of the simplified mixture theory for the composite model will be presented later on.

In this paper, a two-dimensional triaxial braided composite model has been studied using the nonlinear explicit finite-element code LSDYNA. The unit cell consists of six subcells and material properties associated with shell element integration point simulate braiding architecture. The local material properties were selected by correlation of the global behavior of a coupon model with static specimen tests. By changing subcell size and orientation angle at integration points, different braids architectures were obtained. Panel ballistic models were performed with benefits of computation efficiency of shell elements. Mechanical properties, panel impact threshold velocities, and failure initiations for braids with bias angles of 75, 60, 45, and 30° were studied. Boundary effects were also investigated.

Based on a structural model for wave-controlled impact, a modified Hertzian contact law was used to investigate the impact responses of composite laminates. The original nonlinear governing equation was transformed into two linear equations using asymptotic expansion. Closed-form solution can be derived for the first linear homogeneous equation, which is the equation of motion for single degree of freedom system with viscous damping. The second linear nonhomogeneous equation was solved numerically. The overall impact responses for wave-controlled impacts can be obtained semianalytically and agree well with the numerical solutions of nonlinear governing equations. The proposed methodology is useful for providing guidance to numerical simulation of impact on complex composite structures with contact laws fitting from experimental data.

Over the past several decades, two intriguing classes of problems, having a wide range of applications in engineering, have been of interest to many researchers: (1) coupled dynamics of a distributed parameter system traversed by one or more moving oscillators; and (2) transient dynamic analysis of axially moving media (and associated phenomena of parametric resonances). Bridge vehicle interaction falls into the first class of problems, and the analysis of flexible appendages deployed from a satellite or a spacecraft is typical of the second class. Mathematically, these two problems are dual to each other, and they often are highly nonlinear in nature and typically involve large overall motion in space with complex effects of convective inertia terms in their governing equations of dynamic equilibrium. The “nonstandard” analytical nature of these problems stems from the fact that we are dealing with one or more of the following peculiarities: (1) variable problem domain; (2) varying spatial distribution of forces over the time duration of the analysis; and/or (3) changing location and type of constraints. Many researchers are trying to formulate the response of these problems, each having a different approach, but applicable only to certain specific details. Moreover, few researchers have concluded that these problems are beyond the scope of the present commercial finite-element (FE) software packages. However, we believe that if the nature and details of these problems are studied properly and carefully, it is immediately possible to simulate these problems in available commercial FE programs. An added advantage would also be the avoidance of many unrealistic simplifying assumptions that are often introduced to reduce the mathematical complexity; e.g., neglecting possible separation (after periods of prior contacts) in beam-moving vehicle problems, assuming linear behavior of suspension systems, and restriction to beam configuration only, among many others. For demonstration, we use ABAQUS in a large number of test cases to be presented. The results are compared with those presented in literatures.