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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Global and Local Stress Analyses of McDonnell Douglas Stitched/RFI Composite Wing Stub Box

This report contains results of structural analyses performed in support of the NASA structural testing of an all-composite stitched/RFI (resin film infusion) wing stub box. McDonnell Douglas Aerospace Company designed and fabricated the wing stub box. The analyses used a global/local approach. The global model contains the entire test article. It includes the all-composite stub box, a metallic load-transition box and a metallic wing-tip extension box. The two metallic boxes are connected to the inboard and outboard ends of the composite wing stub box, respectively. The load-transition box was attached to a steel and concrete vertical reaction structure and a load was applied at the tip of the extension box to bend the wing stub box upward. The local model contains an upper cover region surrounding three stringer runouts. In that region, a large nonlinear deformation was identified by the global analyses. A more detailed mesh was used for the local model to obtain more accurate analysis results near stringer runouts. Numerous analysis results such as deformed shapes, displacements at selected locations, and strains at critical locations are included in this report.

Wang, John T.↗
NASA NTRS1980

Dual methods and approximation concepts in structural synthesis

Approximation concepts and dual method algorithms are combined to create a method for minimum weight design of structural systems. Approximation concepts convert the basic mathematical programming statement of the structural synthesis problem into a sequence of explicit primal problems of separable form. These problems are solved by constructing explicit dual functions, which are maximized subject to nonnegativity constraints on the dual variables. It is shown that the joining together of approximation concepts and dual methods can be viewed as a generalized optimality criteria approach. The dual method is successfully extended to deal with pure discrete and mixed continuous-discrete design variable problems. The power of the method presented is illustrated with numerical results for example problems, including a metallic swept wing and a thin delta wing with fiber composite skins.

Fleury, C.↗
NASA NTRS1980

Discrete-continuous variable structural synthesis using dual methods

Approximation concepts and dual methods are extended to solve structural synthesis problems involving a mix of discrete and continuous sizing type of design variables. Pure discrete and pure continuous variable problems can be handled as special cases. The basic mathematical programming statement of the structural synthesis problem is converted into a sequence of explicit approximate primal problems of separable form. These problems are solved by constructing continuous explicit dual functions, which are maximized subject to simple nonnegativity constraints on the dual variables. A newly devised gradient projection type of algorithm called DUAL 1, which includes special features for handling dual function gradient discontinuities that arise from the discrete primal variables, is used to find the solution of each dual problem. Computational implementation is accomplished by incorporating the DUAL 1 algorithm into the ACCESS 3 program as a new optimizer option. The power of the method set forth is demonstrated by presenting numerical results for several example problems, including a pure discrete variable treatment of a metallic swept wing and a mixed discrete-continuous variable solution for a thin delta wing with fiber composite skins.

Schmit, L. A.↗
NASA NTRS1926

Albatros Commercial Airplane L 73

The Albatros was a two engine commercial biplane carrying 2 pilots, eight passengers, and 160 KG of baggage. The framework is metal, the wings having plywood and fabric over the steel tubing. The L 73 was the first 2 engine biplane to be made in Germany.

Ruhl, Karl↗
NASA NTRS1927

Junkers Commercial Airplane G 31

The three engine Junkers commercial Airplane, type G 31 is a further development of the earlier G 24. It is an all metal low wing monoplane with a total engine output of 1100-1200 HP.

Source record↗
NASA NTRS2010

The Reconstruction and Failure Analysis of The Space Shuttle Columbia

This viewgraph presentation describes a very detailed reconstruction plan and failure analysis of The Space Shuttle Columbia accident. The contents include: 1) STS-107 Timeline; 2) Foam Impact; 3) Recovery; 4) Reconstruction; 5) Reconstruction Plan; 6) Reconstruction Hanger; 7) Pathfinders; 8) Aluminum Pathfinder; 9) Early Analysis - Left MLG Door Area; 10) Emphasis Switched to Left Hand Wing Leading Edge; 11) Wing Leading Edge Subsystem (LESS); 12) 3D Reconstruction of Left WLE; 13) Left Wing Tile Table; 14) LESS Observations; 15) Left Hand Wing Debris Points to RCC 8/9 - Slumped Tile; 16) Reconstructed View of LC/P 9 tile with I/B Tile; 17) Reconstructed View of Lower C/P 9 Tile; 18) Carrier Panel 8 - Upper; 19) Left Hand Wing Debris Points to RCC 8/9 - Erosion and RCC with attach hole intact; 20) Erosion on Panel 8 Upper Outboard Rib; 21) RCC Panels 8 & 9 Erosion Features; 22) Slumping Source for Carrier Panel 9 Tile was Revealed; 23) Debris Indicated Highest Probability Initiation Site; 24) Left Hand Wing Debris Points to RCC 8/9- Metallic Deposits; 25) Relative Metallic Deposition on L/H Wing Materials; 26) Metallic Deposit Example, LH RCC 8; 27) High Level Questions; 28) Analysis Plan Challenges; 29) Analysis Techniques; 30) Analysis Approach; 31) RCC Panel 8 Erosion Features; 32) Radiographic Features; 33) Radiography WLE LH Panel 8; 34) LH RCC 8 Upper Apex; 35) LH RCC 8 - Deposit Feature: Thick Tear Shaped; 36) LH RCC 8 - Deposit Feature: Thick Globules; 37) LH RCC 8 - Deposit Feature: Spheroids; 38) LH RCC 8 - Deposit Feature: Uniform Deposit; 39) Significant Findings - Sampling All Other panels; 40) Proposed Breach Location and Plasma Flow; 41) Corroborating Information - RCC Panel Debris Locations; 42) Corroborating Information - LH OMS Pod Analysis; 43) Corroborating Information - Impact Testing; and 44) Overall Forensic Conclusions.

Russell, Richard W.↗
NASA NTRSDate not supplied

TPSAS-NF1676L-18502-DND

1. Aeroelastic tailoring with tow-steered composites 2. Aeroelastic tailoring with functionally-graded metals 3. Wing box topology optimization - Curvilinear rib/spar/stiffener placement - Cellular-based stringer topologies - Optimal topology of lightening holes within ribs/spars - 3D level set methods

Stanley R Cole↗
NASA NTRSDate not supplied

Manufacturing Trials of Integrally Stiffened Composite Panels Using Automated Fiber Placement

Commercial aircraft structures are frequently manufactured from carbon-epoxy materials because of their weight and stiffness advantages compared to metallic materials. Wing cover panels are regularly manufactured using an automated fiber placement (AFP) process, but current design and manufacturing methodologies do not fully take advantage of the opportunities afforded by AFP. Design and manufacturing studies were undertaken at the NASA Langley Research Center at the Integrated Structural Assembly of Advanced Composites (ISAAC) facility to quantify manufacturing benefits and limitations associated with AFP to create structurally efficient integral stiffeners as an alternative to bonded or mechanically fastened stiffeners. This methodology could save weight and remove failure mechanisms by reducing the need for rivets and bonding materials since the stiffener plies are interleaved within the skin plies. The use of AFP with integral stiffeners can open the design space, but a fundamental, systematic evaluation of manufacturing limitations is necessary. Manufacturing trials are described herein, where considered manufacturing variables included stiffener location, stiffener course staggering, stiffener widths, stiffener intersections, and material thicknesses for both the skin and stiffener plies. The manufacturing process and lessons learned from each trial are described, including the most successful current design which contains staggered stiffeners, non-traditional laminate angles, and a combination of multiple material thicknesses within the same laminate.

Composite manufacturing↗
NASA NTRSDate not supplied

Manufacturing Trials of Integrally Stiffened Composite Panels Using Automated Fiber Placement

Commercial aircraft structures are frequently manufactured from carbon-epoxy materials because of their weight and stiffness advantages compared to metallic materials. Wing cover panels are regularly manufactured using an automated fiber placement (AFP) process, but current design and manufacturing methodologies do not fully take advantage of the opportunities afforded by AFP. Design and manufacturing studies were undertaken at the NASA Langley Research Center at the Integrated Structural Assembly of Advanced Composites (ISAAC) facility to quantify manufacturing benefits and limitations associated with AFP to create structurally efficient integral stiffeners as an alternative to bonded or mechanically fastened stiffeners. This methodology could save weight and remove failure mechanisms by reducing the need for rivets and bonding materials since the stiffener plies are interleaved within the skin plies. The use of AFP with integral stiffeners can open the design space, but a fundamental, systematic evaluation of manufacturing limitations is necessary. Manufacturing trials are described herein, where considered manufacturing variables included stiffener location, stiffener course staggering, stiffener widths, stiffener intersections, and material thicknesses for both the skin and stiffener plies. The manufacturing process and lessons learned from each trial are described, including the most successful current design which contains staggered stiffeners, non-traditional laminate angles, and a combination of multiple material thicknesses within the same laminate.

Automated Fiber Placement↗
NASA NTRS2009

Dynamic Structural Fault Detection and Identification

Aircraft structures are designed to guarantee safety of flight in some required operational envelope. When the aircraft becomes structurally impaired, safety of flight may not be guaranteed within that previously safe operational envelope. In this case the safe operational envelope must be redefined in-flight and a means to prevent excursion from this new envelope must be implemented. A specific structural failure mode that may result in a reduced safe operating envelope, the exceedance of which could lead to catastrophic structural failure of the aircraft, will be addressed. The goal of the DFEAP program is the detection of this failure mode coupled with flight controls adaptation to limit critical loads in the damaged aircraft structure. The DFEAP program is working with an F/A-18 aircraft model. The composite wing skins are bonded to metallic spars in the wing substructure. Over time, it is possible that this bonding can deteriorate due to fatigue. In this case, the ability of the wing spar to transfer loading between the wing skins is reduced. This failure mode can translate to a reduced allowable compressive strain on the wing skin and could lead to catastrophic wing buckling if load limiting of the wing structure is not applied. The DFEAP program will make use of a simplified wing strain model for the healthy aircraft. The outputs of this model will be compared in real-time to onboard strain measurements at several locations on the aircraft wing. A damage condition is declared at a given location when the strain measurements differ sufficiently from the strain model. Parameter identification of the damaged structure wing strain parameters will be employed to provide load limiting control adaptation for the aircraft. This paper will discuss the simplified strain models used in the implementation and their interaction with the strain sensor measurements. Also discussed will be the damage detection and identification schemes employed and the means by which the damaged aircraft parameters will be used to provide load limiting that keeps the aircraft within the safe operational envelope.

Smith, Timothy↗
NASA NTRS1977

Study of metallic structural design concepts for an arrow wing supersonic cruise configuration

A structural design study was made, to assess the relative merits of various metallic structural concepts and materials for an advanced supersonic aircraft cruising at Mach 2.7. Preliminary studies were made to ensure compliance of the configuration with general design criteria, integrate the propulsion system with the airframe, select structural concepts and materials, and define an efficient structural arrangement. An advanced computerized structural design system was used, in conjunction with a relatively large, complex finite element model, for detailed analysis and sizing of structural members to satisfy strength and flutter criteria. A baseline aircraft design was developed for assessment of current technology. Criteria, analysis methods, and results are presented. The effect on design methods of using the computerized structural design system was appraised, and recommendations are presented concerning further development of design tools, development of materials and structural concepts, and research on basic technology.

Turner, M. J.↗