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Dawn C Jegley

Publications and source records attributed to Dawn C Jegley.

18 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2023

A Tool for Defining Tow-Steered Laminates for Finite Element Grids

Tailoring of composite laminates is traditionally performed by changing the orientation of straight fibers in one or more plies. Modern automated fiber placement machines facilitate placement of bundles of curved fibers (tows) in a process called tow-steering, but additional variables must be used to define the shapes of tow-steered fiber paths. In this paper, a Python-based tool, called Automated Tool for Steered COmposite Optimizable Laminates (ATSCOOL), is presented for modeling steered tows as chains of circular arcs. The geometry for defining tow path shapes using four input variables is described. The formulae for determining gaps and overlaps between clusters of consecutive tows, called courses, and for determining the number of steered courses needed to cover a rectangular panel are presented. Graphical representation of courses and thickness distribution in an example panel using the ATSCOOL software are presented. Finally, an example of a finite element analysis performed using the property information output from ATSCOOL is shown.

finite element analysis↗
NASA NTRS2023

Lessons Learned from Large-Scale Aerospace Structural Testing

Large-scale testing of aerospace structures is frequently the final step in a development project to validate the structural performance, and that step typically involves a large cost and time investment. To ensure that the testing provides the required data, avoiding errors that can result in an unsuccessful test and failure to meet objectives is critical. Five lessons learned are presented herein to provide insight to those conducting tests in order to help them avoid known pitfalls that may result in an unsuccessful test. Five large-scale tests are described, and include two composite wing tests, a composite hybrid-wing body center section test, a full-scale 27.5-ft diameter metallic barrel test, and an 8-ft diameter metallic barrel test. Problems identified during the testing and mitigation approaches to solve the problems are presented, then the lessons learned are identified and discussed.

Lessons Learned↗
NASA NTRS2022

Technology Maturation Report for Damage Arresting Composites under the Environmentally Responsible Aviation Project

The goal of the NASA Environmentally Responsible Aviation (ERA) Project was to develop technologies that will lead to commercial aircraft that will burn less fuel and produce fewer emissions than the 2010 state of the art. The purpose of this Technology Maturation Report (TMR) is to summarize both the development and lessons learned for the Damage Arresting Composites (DAC) Demonstration and to summarize how the test results were used to predict the impact the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) technology would have on future transport aircraft weights. The TMR includes the technical work plans for the DAC Demonstration for the six-year effort of the NASA ERA Project. System studies indicated that DAC would reduce structural weight by 20% for both hybrid wing body and tube and wing aircraft. A lighter vehicle requires less energy to fly so weight savings becomes reduced fuel burn. Additionally, the hybrid wing body configuration has improved lift-to-drag ratios compared to the tube and wing aircraft, further reducing fuel burn.

carbon-epoxy↗
NASA NTRS2021

ISAAC

This document describes the ISAAC (Integrated Structural Assembly of Advanced Composites) robotic system. ISAAC is a highly accurate, automated robotic platform used to support research on the design, analysis, manufacturing and evaluation of advanced composite materials and structures. ISAAC can build composite parts using automated fiber placement (AFP) and stitching through the thickness of dry fabric. The system is frequently upgraded to include new capabilities. ISAAC is located at the NASA Langley Research Center

carbon-epoxy↗
NASA NTRS2021

Testing of Composite Stiffened Panels with Mid-Length Lateral Notches

Two hat-stiffened carbon-epoxy panels were loaded to failure to determine their failure load and mode under the Advanced Composites Project (ACP) Rapid Design Tools activity. Each specimen contained a lateral notch located at the mid-length location which severed the central stiffener. A description of the specimens, instrumentation, test fixtures and procedures is presented, followed by the presentation of the experimental results.

Dawn C Jegley↗
NASA NTRS2020

Testing of Composite Laminate Specimens with Mid-Length Lateral Notches

Eight fiber-reinforced, carbon-epoxy composite laminate specimens without external stiffeners (unstiffened) were loaded to failure to determine their failure load and mode under the Advanced Composites Project (ACP) Rapid Design Tools activity. Each specimen contained a lateral notch located at the mid-length location. A description of the specimens, instrumentation, test fixtures and procedures is presented, followed by the presentation of the experimental results.

Dawn C Jegley↗
NASA NTRS2012

TPSAS-NF1676L-13678-DND

An area that shows promise in enhancing structural integrity of aircraft and aerospace structures is integral stitched composite technology. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The goal of the PRSEUS concept is to provide damage-containment capability for composite structures while reducing overall structural weight. The National Aeronautics and Space Administration, the Boeing Company, and the Federal Aviation Administration (FAA) have partnered in an effort to assess the damage-containment features of a full-scale curved PRSEUS panel using the FAA Full-Scale Aircraft Structural Test Evaluation and Research facility. The background, test plan, fixture modifications, pre-test analysis, and planned experimental procedure were presented at the 2011 Aircraft Airworthiness and Sustainment Conference. This follow-on paper and presentation will provide details of the experimental procedure, test results, nondestructive inspection results, and preliminary test analysis comparison. The test program included three phases of loading and inspections: Phase I, as-built; Phase II, with barely visible impact damage (BVID); and Phase III, with a two-bay notch severing the central stiffener. Axial tension, internal pressure, and combined axial tension and internal pressure load conditions were applied during each phase using load levels that demonstrate compliance with the strength, deformation, and damage-tolerance requirements of Title 14 Code of Federal Regulations Part 25. Pressure loads were based on an operating pressure of 9.2 psi, designated as 1P, and the axial loads were based on a design limit load (DLL) of 227 kip. The Phase I test results provided a baseline. First, 50 percent limit-load levels were applied for each of the three load cases to verify proper load introduction and repeatability by examining strain and displacement results. Next, three limit-load cases were applied: limit pressure of 12.2 psi (1.33 P), axial DLL of 227 kip, and combined 1P pressure and axial DLL. Linear strain and displacement results were observed in all load cases. For Phase-II testing, the panel was impacted to introduce BVID with an energy of 40 ft-lb, with the impact located between the central stringer stitch line and flange edge. The impact created non-penetrating visual damage with fiber breaks and delamination between the skin and stringer flange. Limit-load conditions were applied followed by pressure overload of 18.4 psi (2P) and ultimate load conditions, defined as 1.5 times DLL. Inspections were performed after each load case; no damage growth was observed. For Phase-III testing, a two-bay notch was machined severing the central stiffener. The panel was then subjected to limit-load conditions, followed by combined 1P pressure while increasing axial tension load to catastrophic failure. As limit combined loads were approached, damage initiation was observed in the form of 45? splitting at 1P + 79% DLL. Axial loading was increased to 1P + 150 percent DLL causing damage to propagate to the adjacent stringer flanges. Damage was contained within the two-bay region by the stitching rows up to 1P + 176% DLL. Axial load was further increased until catastrophic failure, when a stringer rod failed at 1P + 192% DLL. These test results further verify the damage containment features of the PRSEUS concept and suggest its appropriateness for future flight vehicles.

Andrew Bergan↗
NASA NTRS2012

TPSAS-NF1676L-13538-DND

Highly loaded composite struts from a proposed truss-based Altair lunar lander descent stage concept were selected for development under NASA's Advanced Composites Technology program. Predicted compressive member forces during launch and ascent of over -100,000 lbs were much greater than the tensile loads. The development of a capability for experimental evaluation of the structural performance of these struts is presented. Strut lengths range from 60 to over 120 inches, and compressive launch and ascent loads can exceed -100,000 lbs, or approximately two times the corresponding tensile loads. Allowing all possible compressive structural responses, including elastic buckling, were primary considerations for designing the test hardware. NASA's industry partner designed and built highly loaded struts that were delivered to NASA for testing. Their design, fabricated on a washout mandrel, had a uniform-diameter composite tube with composite tapered ends. Each tapered end contained a titanium end fitting with facing conical ramps that are overlaid and overwrapped with composite materials. The highly loaded struts were loaded in both tension and compression, with ultimate failure produced in compression. Test results for these struts are presented and discussed, along with measured deflections, strains and observed failure mechanisms.

K Chauncey Wu↗
NASA NTRSDate not supplied

STITCHED COMPOSITES WITH THREE-DIMENSIONAL STITCH PATHS

Stitched composites have been shown to exhibit damage tolerance and to reduce weight compared to traditional layered composites through unitization of the structure by elimination of fasteners. Stitching capabilities have been incorporated into the Integrated Structural Assembly of Advanced Composites (ISAAC) system at NASA Langley Research Center with the introduction of two stitching heads. Stitching path control was initially implemented as straight lines in space, as was done for previous stitching development. However, more complex stitched structures such as a wind tunnel blade or around cutouts within a fuselage or wing skin, require that the stitching paths be implemented as three-dimensional (3D) stitching paths in space. Unfortunately, control programming output by an existing preprocessor program cannot stitch these curved paths due to problems that arise in stitch formation and the introduction of side forces on the needles using the conventional programming approach whereby the head is simultaneously controlled through translations and rotations. This lack of capability is most significant for the single-sided stitching head, where two needles are in the preform at the same time for the majority of the stitching process. A means to program 3D stitching paths in space was developed whereby the translation and rotation of each stitch were decoupled, thereby eliminating the problems associated with current control programming approach. Using this newly developed stitching path definition and control programming, complex stitching paths have successfully been stitched at the ISAAC facility. The ability to stitch general 3D stitching paths in space enables the use of stitching on more complex parts.

carbon-epoxy↗
NASA NTRSDate not supplied

Exploration and Manufacturing with Automated Fiber Placement

Highly repeatable and nearly defect-free fabrication of composite parts is critical to the success and widespread acceptance of composite materials. Through optimization using thin-ply materials, composite parts can be manufactured to be lighter and tailored more specifically to anticipated design loads than with standard prepreg materials alone. However, defects arising from the thin-ply manufacturing process are not always similar to defects found with standard tows. These new defects warrant evaluation. At NASA Langley Research Center, the manufacturing process parameters associated with automated fiber placement (AFP), a slit tape-based composite manufacturing process, were optimized for the use of a thin-ply prepreg carbon-epoxy material. Carbon-epoxy tows with areal weights of 30 g/m2 and 70 g/m2 were used in these manufacturing trials. The AFP process parameters of heater output, compaction force, tow feed rate, and tow tension were adjusted and optimized for successful manufacturing. This article documents an exploration of thin-ply fabrication on both flat and complex-shaped surfaces. Ultimately, aerospace-quality laminates were made from the 70-g/m2 material, but imperfections in the 30-g/m2 material itself and the fact that the AFP machine was not designed for such a thin material meant that more research and trials are required to obtain flight-quality 30-g/m2 laminates.

carbon-epoxy↗
NASA NTRSDate not supplied

Stitched Composites with Three-Dimensional Stitch Paths

Stitched composites have been shown to exhibit damage tolerance and to reduce weight compared to traditional layered composites through unitization of the structure by elimination of fasteners. Stitching capabilities have been incorporated into the Integrated Structural Assembly of Advanced Composites (ISAAC) system at NASA Langley Research Center with the introduction of two stitching heads. Stitching path control was initially implemented as straight lines in space, as was done for previous stitching development. However, more complex stitched structures such as a wind tunnel blade or around cutouts within a fuselage or wing skin, require that the stitching paths be implemented as three-dimensional (3D) stitching paths in space. Unfortunately, control programming output by an existing preprocessor program cannot stitch these curved paths due to problems that arise in stitch formation and the introduction of side forces on the needles using the conventional programming approach whereby the head is simultaneously controlled through translations and rotations. This lack of capability is most significant for the single-sided stitching head, where two needles are in the preform at the same time for the majority of the stitching process. A means to program 3D stitching paths in space was developed whereby the translation and rotation of each stitch were decoupled, thereby eliminating the problems associated with current control programming approach. Using this newly developed stitching path definition and control programming, complex stitching paths have successfully been stitched at the ISAAC facility. The ability to stitch general 3D stitching paths in space enables the use of stitching on more complex parts.

carbon-epoxy↗
NASA NTRSDate not supplied

Lessons Learned from Large-Scale Aerospace Structural Testing

Large-scale testing of aerospace structures is frequently the final step in a development project to validate the structural performance, and that typically involves a large cost and time investment. In order to ensure that the testing provides the required data, avoiding errors that can result in an unsuccessful test and failure to meet objectives is critical. Presented herein are five lessons learned to provide insight to those conducting tests in order to help them avoid known pitfalls that may result in an unsuccessful test. Five subject large-scale tests are described, and include two composite wing tests, a composite hybrid-wing body center section test, a full-scale 27.5-ft diameter metallic barrel test, and an 8-ft diameter metallic barrel test. Problems identified during the testing and mitigation approaches to solve the problems are presented, then the lessons learned are summarized.

Lessons Learned↗
NASA NTRSDate not supplied

Lessons Learned from Large-Scale Aerospace Structural Testing

Large-scale testing of aerospace structures is frequently the final step in a development project to validate the structural performance, and that typically involves a large cost and time investment. In order to ensure that the testing provides the required data, avoiding errors that can result in an unsuccessful test and failure to meet objectives is critical. Presented herein are five lessons learned to provide insight to those conducting tests in order to help them avoid known pitfalls that may result in an unsuccessful test. Five subject large-scale tests are described, and include two composite wing tests, a composite hybrid-wing body center section test, a full-scale 27.5-ft diameter metallic barrel test, and an 8-ft diameter metallic barrel test. Problems identified during the testing and mitigation approaches to solve the problems are presented, then the lessons learned are summarized.

Lessons Learned↗
NASA NTRSDate not supplied

Out-of-Plane Restraint under Tension In-Plane Loading

The development and description of a test fixture designed to restrict out-of-plane motion in the center region of tension-loaded carbon-epoxy panels are presented herein. The test fixture was used to impose displacement conditions consistent with those in an analytical tool being developed to predict panel behavior in the vicinity of a central notch, which included the assumption that there would be no out-of-plane displacement and no buckling. However, pretest analysis using finite element models for a panel loaded without any out-of-plane restraint in the region of the notch indicated that with an imperfection magnitude equal to 20 percent of the thickness of the thin-skin panel, an unacceptable amount of out-of-plane deformation would occur as Poisson effects induced compression loads in the region of the notch. Therefore, to validate this tool in a test program, a restraint fixture which would suppress out-of-plane motion was required. The panel could not be encased in restraining plates because instrumentation and visibility were required in the vicinity of the notch. Therefore, a fixture that would restrict out-of-plane motion while still allowing access to the surface of the panel at the notch edges for instrumentation and line-of-sight access for cameras was required. To satisfy these requirements, a fixture was designed to restrict only out-of-plane motion near the center of the notch. Two 1.78-m-long test panels were loaded in tension to failure using this fixture. Out-of-plane deformations were not directly measured during testing, so back-to-back strain gages were used to obtain an indication of buckling. Strain results indicated that the restraint fixture performed as designed and buckling did not occur.

Fixtures↗
NASA NTRSDate not supplied

Manufacturing Trials of Integrally Stiffened Panels for Flight Applications

Reduced weight is an ever-increasing demand within the aerospace industry, and non-traditional means of achieving this demand are increasingly being considered. There is also a desire to reduce the manual labor involved in attaching thousands of parts together with rivets and other fasteners. Using advanced stiffener concepts can avoid the weight of fasteners and streamline fabrication by reducing the part count, while still providing the mandatory strength requirements for flight certification. Described herein is a set of manufacturing trials that were completed at NASA Langley Research Center to investigate the feasibility of various integrated stiffener concepts. Fully composite stiffeners that can be integrated or attached by other means, either through bonding or interleaving them within the skin, can remove the fasteners and still maintain a rapid manufacturing rate. This study contains these advanced stiffener concepts, and the lessons learned and design iterations that resulted from the series of manufacturing trials. Each set of trials was completed with an end goal of using tow-steered integral stiffeners for flight structure in composite aircraft wings.

Composite↗
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↗