Ivatury S. Raju

Simulating the Structural Response of a Preloaded Bolted Joint

The present paper describes the structural analyses performed on a preloaded bolted-joint configuration. The joint modeled was comprised of two L-shaped structures connected together using a single bolt. Each L-shaped structure involved a vertical flat segment (or shell wall) welded to a horizontal segment (or flange). Parametric studies were performed using elasto-plastic, large-deformation nonlinear finite element analyses to determine the influence of several factors on the bolted-joint response. The factors considered included bolt preload, washer-surface-bearing size, edge boundary conditions, joint segment length, and loading history. Joint response is reported in terms of displacements, gap opening, and surface strains. Most of the factors studied were determined to have minimal effect on the bolted-joint response; however, the washer-bearing-surface size affected the response significantly.

Recent Advances in Durability and Damage Tolerance Methodology at NASA Langley Research Center

Durability and damage tolerance (D&DT) issues are critical to the development of lighter, safer and more efficient aerospace vehicles. 1 Durability is largely an economic life-cycle design consideration whereas damage tolerance directly addresses the structural airworthiness (safety) of the vehicle. Both D&DT methodologies must address the deleterious effects of changes in material properties and the initiation and growth of damage that may occur during the vehicle’s service lifetime. The result of unanticipated D&DT response is often manifested in the form of catastrophic and potentially fatal accidents. As such, durability and damage tolerance requirements must be rigorously addressed for commercial transport aircraft and NASA spacecraft systems. This paper presents an overview of the recent and planned future research in durability and damage tolerance analytical and experimental methods for both metallic and composite aerospace structures at NASA Langley Research Center (LaRC).

NASA Structural Analysis Report on the American Airlines Flight 587 Accident- Local Analysis of the Right Rear Lug

A detailed finite element analysis of the right rear lug of the American Airlines Flight 587 - Airbus A300-600R was performed as part of the National Transportation Safety Board’s failure investigation of the accident that occurred on November 12, 2001. The loads experienced by the right rear lug are evaluated using global models of the vertical tail, local models near the right rear lug, and a global -local analysis procedure. The right rear lug was analyzed using two modeling approaches. In the first approach, solid-shell type modeling is used, and in the second approach, layered-shell type modeling is used. The solid-shell and the layered-shell modeling approaches were used in progressive failure analyses (PFA) to determine the load, mode, and location of failure in the right rear lug under loading representative of an Airbus certification test conducted in 1985 (the 1985-certification test). Both analyses were in excellent agreement with each other on the predicted failure loads, failure mode, and location of failure. The solid-shell type modeling was then used to analyze both a subcomponent test conducted by Airbus in 2003 (the 2003-subcomponent test) and the accident condition. Excellent agreement was observed between the analyses and the observed failures in both cases. From the analyses conducted and presented in this paper, the following conclusions were drawn. The moment, Mx (moment about the fuselage longitudinal axis), has significant effect on the failure load of the lugs. Higher absolute values of Mx give lower failure loads. The predicted load, mode, and location of the failure of the 1985-certification test, 2003-subcomponent test, and the accident condition are in very good agreement. This agreement suggests that the 1985-certification and 2003subcomponent tests represent the accident condition accurately. The failure mode of the right rear lug for the 1985-certification test, 2003-subcomponent test, and the accident load case is identified as a cleavage-type failure. For the accident case, the predicted failure load for the right rear lug from the PFA is greater than 1.98 times the limit load of the lugs.

Meshless Local Petrov-Galerkin Euler-Bernoulli Beam Problems: A Radial Basis Function Approach

A radial basis function implementation of the meshless local Petrov-Galerkin (MLPG) method is presented to study Euler-Bernoulli beam problems. Radial basis functions, rather than generalized moving least squares (GMLS) interpolations, are used to develop the trial functions. This choice yields a computationally simpler method as fewer matrix inversions and multiplications are required than when GMLS interpolations are used. Test functions are chosen as simple weight functions as in the conventional MLPG method. Compactly and noncompactly supported radial basis functions are considered. The non-compactly supported cubic radial basis function is found to perform very well. Results obtained from the radial basis MLPG method are comparable to those obtained using the conventional MLPG method for mixed boundary value problems and problems with discontinuous loading conditions.

Best Practices for Reliable and Robust Spacecraft Structures

,A study was undertaken to capture the best practices for the development of reliable and robust spacecraft structures for NASA’s next generation cargo and crewed launch vehicles. In this study, the NASA heritage programs such as Mercury, Gemini, Apollo, and the Space Shuttle program were examined. A series of lessons learned during the NASA and DoD heritage programs are captured. The processes that “make the right structural system” are examined along with the processes to “make the structural system right”. The impact of technology advancements in materials and analysis and testing methods on reliability and robustness of spacecraft structures is studied. The best practices and lessons learned are extracted from these studies. Since the first human space flight, the best practices for reliable and robust spacecraft structures appear to be well established, understood, and articulated by each generation of designers and engineers. However, these best practices apparently have not always been followed. When the best practices are ignored or short cuts are taken, risks accumulate, and reliability suffers. Thus program managers need to be vigilant of circumstances and situations that tend to violate best practices. Adherence to the best practices may help develop spacecraft systems with high reliability and robustness against certain anomalies and unforeseen events. I. Introduction ASA is currently in the process of developing the next generation crewed and cargo launch vehicles and spacecraft to return to the moon and beyond. With the experience and knowledge base available from past similar programs, a document that captures salient aspects of successful programs is being developed. This document serves as an important guide in evaluating next generation and future spacecraft concepts and proposals. As a part of this guide, guides for individual technical disciplines are being developed. Reliable and robust structural systems design is one of these technical disciplines. The structures document describes pertinent issues, best practices, errors, miss-steps, lessons learned, and summarizes the previously used design processes (tools and standards) for the structures discipline. Structural systems provide the basic framework to distribute external and internal loads resulting from all flight loads, ground loads, and associated operational and environmental loads. The primary objective of a structural system is to remain intact and experience minimal deformation when exposed to various environments, including ground processing, testing, launch, on-orbit, and re-entry operations. Structural systems also provide containment for pressures as in pressure vessels, pressure components, and pressurized structures. Structures tend to be a dependent subsystem in the sense that many requirements flow to structures from other subsystems. Space systems are very complex products of multiple disciplines, and therefore require a multidisciplinary analysis and optimization approach to capture various system interactions and sensitivities in order to obtain optimum system

Elastic-Plastic Nonlinear Response of a Space Shuttle External Tank Stringer. Part 2; Thermal and Mechanical Loadings

Elastic-plastic, large-deflection nonlinear thermo-mechanical stress analyses are performed for the Space Shuttle external tank’s intertank stringers. Detailed threedimensional finite element models are developed and used to investigate the stringer’s elastic-plastic response for different thermal and mechanical loading events from assembly through flight. Assembly strains caused by initial installation on an intertank panel are accounted for in the analyses. Thermal loading due to tanking was determined to be the bounding loading event. The cryogenic shrinkage caused by tanking resulted in a rotation of the intertank chord flange towards the center of the intertank, which in turn loaded the intertank stringer feet. The analyses suggest that the strain levels near the first three fasteners remain sufficiently high that a failure may occur. The analyses also confirmed that the installation of radius blocks on the stringer feet ends results in an increase in the stringer capability.

Elastic-Plastic Nonlinear Response of a Space Shuttle External Tank Stringer. Part 1; Stringer-Feet Imperfections and Assembly

Elastic-plastic, large-deflection nonlinear stress analyses are performed for the external hat-shaped stringers (or stiffeners) on the intertank portion of the Space Shuttle s external tank. These stringers are subjected to assembly strains when the stringers are initially installed on an intertank panel. Four different stringer-feet configurations including the baseline flat-feet, the heels-up, the diving-board, and the toes-up configurations are considered. The assembly procedure is analytically simulated for each of these stringer configurations. The location, size, and amplitude of the strain field associated with the stringer assembly are sensitive to the assumed geometry and assembly procedure. The von Mises stress distributions from these simulations indicate that localized plasticity will develop around the first eight fasteners for each stringer-feet configuration examined. However, only the toes-up configuration resulted in high assembly hoop strains.

Fracture mechanics analyses of the slip-side joggle regions of wing-leading-edge panels

ach Space Shuttle Orbiter wing is comprised of 22 leading edge panels. These panels are part of the thermal protection system that protects the Orbiter wings from extreme heating that take place on the reentry in to the earth atmosphere. On some panels that experience extreme heating, liberation of silicon carbon (SiC) coating was observed on the slip-side regions of the panels. Global structural and local fracture mechanics analyses were performed on these panels as a part of the root cause investigation of this coating liberation anomaly. The wing-leadingedge reinforced carbon-carbon (RCC) panels, Panel 9, T-seal 10, and Panel 10, are shown in Fig. 1, and the progression of the stress analysis models is presented in Fig. 2. The global structural analyses showed minimal interaction between adjacent panels and the T-seal that bridges the gap between the panels. A bounding uniform temperature is applied to a representative panel and the resulting stress distribution is examined. For this loading condition, the interlaminar normal stresses showed negligible variation in the chord direction and increased values in the vicinity of the slip-side joggle shoulder. As such, a representative span–wise slice on the panel can be taken, and the cross section can be analyzed using plane strain analysis.

Lessons Learned from Recent Failure and Incident Investigations of Composite Structures

During the past few decades, NASA Langley Research Center (LaRC) has supported several large-scale failure and incident investigations and numerous requests for engineering consultations. Although various extenuating circumstances contributed to each of these incidents, in all cases, the failure resulted from accumulation and/or propagation of damage that reduced the load carrying capability of the structure to a level below that which was needed to sustain structural loads. A brief overview of various failure and incident investigations supported by LaRC, including some of the computational and experimental methodologies that have been applied, is presented. An important outcome of many of these failure and incident investigations is the development of an improved understanding of not only the state-of-the-art in experimental and analytical methods but also the state-of-the-art in the design and manufacturing processes that may contribute to such failures. In order to provide insight into such large-scale investigations, a series of lessons learned were captured. Awareness of these lessons learned is highly beneficial to engineers involved in similar investigations. Therefore, it is prudent that the lessons learned are disseminated such that they can be built upon in other investigations and in ensuing research and development activities.

Moving Aerospace Structural Design Practice to a Load and Resistance Factor Approach

Aerospace structures are traditionally designed using the factor of safety (FOS) approach. The limit load on the structure is determined and the structure is then designed for FOS times the limit load - the ultimate load. Probabilistic approaches utilize distributions for loads and strengths. Failures are predicted to occur in the region of intersection of the two distributions. The load and resistance factor design (LRFD) approach judiciously combines these two approaches by intensive calibration studies on loads and strength to result in structures that are efficient and reliable. This paper discusses these three approaches.