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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

NDE Methods and Capabilities Handbook - Resources for the Composites NDE Community

Composite structures are often difficult to inspect due to the anisotropic nature of the material systems and the range of flaws and defects that can arise in the manufacturing process. NASA’s Advanced Composites Project, a public-private partnership between government, universities and industry, sought to address improved methods, tools, and protocols to reduce the development and certification timeline for composite materials and structures. The rapid inspection of composites was one of the three focused research areas of the project. One deliverable of the project was to develop an NDE Methods and Capabilities Handbook. This Handbook provides a publicly available guidance document that facilitates the selection of appropriate NDE techniques and provides recommended protocols for detecting and characterizing common flaw types in solid laminate carbon fiber reinforced polymer composite (CFRP) structures. The Handbook seeks to reduce the time required to develop qualified inspection processes for composite aircraft structures during the development, certification, and manufacturing phases by providing a reference that helps minimize trial and error and provides guidance on best practices, techniques, and settings, for specific flaw types and geometries. Techniques include, but are not limited to ultrasound, laser based ultrasound, infrared thermography, and x-ray CT. Flaw types include, but are not limited to porosity, delaminations, voids, Automated Fiber Placement (AFP) defects and impact damage. The Handbook contains details such as inspection parameters, data analysis techniques, inspection efficiency and limitations. This paper describes key Handbook activities with examples of the techniques and flaw types examined by the Advanced Composite Consortium members. Availability to the NDE community of the fabricated NDE standards as well as the raw data collected on the standards by the Consortium members is discussed.

nondestructive evaluation↗
NASA NTRSDate not supplied

Design of a Manufacturing Demonstration Unit for a Composite Lunar Landing Gear Strut

The goal of the current work was to develop an analytical framework for design of composite struts using various levels of model fidelity. Rapid optimization trade studies were performed using low fidelity two-dimensional (2D) axisymmetric models with smeared composite properties. The optimum 2D model was compared with three-dimensional (3D) models with higher levels of fidelity in material property representation. Good agreement was found with all models. The buckling performance of the highest fidelity 3D model was found to be satisfactory for the intended loading conditions. An additional goal of this work was to perform an initial assessment of using automated fiber placement (AFP) and other advanced manufacturing methods to explore their feasibility for the fabrication of struts for lunar landers, strut-braced wings, and other aerospace components. The designs produced during this study are intended to be used to develop manufacturing demonstration units (MDU) that can be fabricated at the Integrated Structural Assembly of Advanced Composites (ISAAC) facility at Langley Research Center (LaRC) and tested in lab facilities at LaRC.

carbon-epoxy↗
NASA NTRSDate not supplied

Design of a Manufacturing Demonstration Unit for a Composite Lunar Landing Gear Strut

The goal of the current work was to develop an analytical framework for design of composite struts using various levels of model fidelity. Rapid optimization trade studies were performed using low fidelity two-dimensional (2D) axisymmetric models with smeared composite properties. The optimum 2D model was compared with three-dimensional (3D) models with higher levels of fidelity in material property representation. Good agreement was found with all models. The buckling performance of the highest fidelity 3D model was found to be satisfactory for the intended loading conditions. An additional goal of this work was to perform an initial assessment of using automated fiber placement (AFP) and other advanced manufacturing methods to explore their feasibility for the fabrication of struts for lunar landers, strut-braced wings, and other aerospace components. The designs produced during this study are intended to be used to develop manufacturing demonstration units (MDU) that can be fabricated at the Integrated Structural Assembly of Advanced Composites (ISAAC) facility at Langley Research Center (LaRC) and tested in lab facilities at LaRC.

carbon-epoxy↗
NASA NTRSDate not supplied

Demonstration of Linear Covariance Analysis Techniques to Evaluate Entry Descent and Landing Guidance Algorithms, Vehicle Configurations, Analysis Techniques, and Trajectory Profiles

Linear covariance analysis techniques have been previously developed to analyze closed-loop entry, descent, and landing (EDL) scenarios and the initial validation efforts are under-way confirming the generated GN&C system performance results. Given both the theoretical foundation and previous conceptual demonstration, this work begins to flex the potential of linear covariance analysis for atmospheric flight and highlight its versatility and reliability by evaluating multiple entry guidance algorithms, vehicle configurations, trajectory profiles, environment conditions, and analysis techniques for a variety of trade studies. To demonstrate the benefit linear covariance analysis can provide in producing rapid yet accurate performance data, two entry profiles are adopted including the NASA Mars Science Laboratory (MSL) and Exploration Flight Test-1 (EFT-1) while utilizing two different guidance algorithms, the Apollo Final Phase (AFP) and the Fully Numeric Predictor-Corrector Entry Guidance (FNPEG) with different navigation sensor suites in a 6 degree-of-freedom (6-DOF) simulation environment. Results are shown using both linear covariance and Monte Carlo analysis techniques to high-light the consistency between the two methodologies and continue the validation maturation of linear covariance analysis for entry, descent, and landing.

EDL↗
NASA NTRSDate not supplied

Composites From in-Situ Consolidation Automated Fiber Placement of Thermoplastics for High-Rate Aircraft Manufacturing

The National Aeronautics and Space Administration (NASA) project Hi-Rate Composites Aircraft Manufacturing (HiCAM) aims to significantly increase commercial aircraft composite structures manufacturing rate. Thermoplastic composites offer attractive solutions to rapid manufacturing due to their ability to be formed and consolidated quickly. NASA has a particular interest in assessing composite structure manufacturing utilizing an in-situ consolidation automated fiber placement (AFP) of thermoplastics (ICAT) process employing current state-of-the-art laser heating systems. Three semi-crystalline polyaryletherketone (PAEK) thermoplastic tape materials were characterized to ascertain the ICAT process parameters. The required laser power settings were determined at Electroimpact, Inc., measuring material temperatures utilizing a forward looking infrared (FLIR) thermal imaging camera and thermocouples. The material temperature, tool temperature, and placement speed were varied for resulting consolidation quality assessment. The resulting temperature data were also utilized to calibrate thermal analysis models under development at NASA. The experimental temperature data confirmed analytical results. The quality of the resulting test panels was evaluated by both non-destructive evaluation as well as destructively by photo-microscopy. The effect on interlaminar strength was determined by short beam strength testing. Test results of carbon fiber laminates fabricated by ICAT using polyetheretherketone (PEEK), polyetherketoneketone (PEKK), and low-melt polyaryletherketone (LM-PAEK) at various placement temperatures and placement speeds are presented.

thermoplastic composites↗
NASA NTRS2024

The Effects of Peel Ply on the Damage Tolerance Characteristics of the Space Launch System (SLS) Payload Adapter Fitting (PAF)

A payload adaptor fitting (PAF) has been manufactured at NASA’s Marshall Space Flight Center (MSFC) in support of the Space Launch System (SLS) program. This structure is essentially a large cone made up of eight curved “lobes” that are joined together to form the cone. The lobes were manufactured via automated fiber placement (AFP). The PAF will carry predominantly compressive loads thus compression strength will be the focus of this study. The lobes are sandwich structure made with aluminum honeycomb core and carbon/epoxy face sheets. The lobes will be inserted into metallic rings at the top and bottom of the cone to form the full cone. Both front and back exterior surfaces of each lobe have a plain weave fabric placed at ±45⁰ to the vertical direction to help prevent fiber “breakout” during drilling. The face sheet of each lobe is thickest at the top and plies are dropped as the cone reaches the main acreage in the bottom section of the cone. The entire structure used resin impregnated peel ply that, once removed, supplies a good bonding surface for secondary bonding steps. The peel ply can be removed either directly after cure, or just before any bonding operation. Since foreign object impact can occur to the structure at any time during its life, damage tolerance of the PAF structure both with and without peel ply needed to be addressed because a previous study, albeit with limited data, showed some significant differences in damage tolerance whether the peel ply was present or not [1]. A literature search showed no results for comparing impact damage on composite laminates with and without peel ply. Load versus deflection of impact curves, visual damage, dent depth, damage as ascertained by thermography and cross-sectional microscopy were evaluated as part of the damage resistance of the structure. Compression After Impact (CAI) was performed to assess damage tolerance. The impact tests were conducted on representative specimens both with and without the peel ply present.

Sandwich Structure↗
NASA NTRSDate not supplied

Effect of Barely Visible Impact Damage on Thermoplastic Welded Structure

Spirit Aerosystems recently conducted screening tests on thermoplastic skin/stringer panels manufactured with Automated Fiber Placement (AFP) and stamp forming which were joined using Spirit’s Co-Fusion process. Skin and stringer gauge were both set to 0.168 cm. The resulting panel design is 48 cm tall, 49 cm wide, and has 3 stringers welded at a 21.5 cm pitch. Barely Visible Impact Damage (BVID) and Residual Compressive Strength were chosen to conduct a damage tolerance study. Zee stringers, made with IM8/PEKK-FC material, were Co-Fusion welded to a T1100/TC1225 skin. Out of these 5 panels, 1 was chosen to be a BVID survey panel, 1 for pristine compressive strength determination, and 3 to be impacted to create BVID followed by testing for residual compressive strength. On the survey panel, impacts ranged in energy from 40J – 65J and dent depth measurements were conducted at 24 hours, 48 hours, and two weeks after impacting. The residual compressive strength panel failed near the predicted load of 214 MPa. For the residual compressive strength panels, BVID was created at two different mid-bay locations: 1) striking the OML coincident with the stringer and 2) striking the skin IML 2.5 cm away from the stringer flange. Pulse Echo Ultrasonic Testing was used to inspect the panels before and after creating BVID.

HiCAM↗
NASA NTRS2026

Smart Process Planning for Automated Fiber Placement

Many industries, including aerospace, automotive, wind energy, maritime, and sporting goods, rely on strong, lightweight materials called composites. These materials are made by layering fibers, which can come in the form of narrow strips or wider sheets, and setting them in a polymer matrix. One of the most advanced ways to make these parts is through automated fiber placement, where a machine lays down the fibers in precise patterns. This method can create very efficient and strong designs, but it is complex, expensive, and often depends heavily on the experience of skilled engineers. Today, the design, manufacturing, and inspection stages of composite production are usually handled separately. This separation means that important information, such as how a part will be built or what defects might occur, is not always shared between stages. As a result, parts may not be as lightweight, strong, or defect-free as possible, and the process can take longer and cost more. This research develops a smart process planning system that connects design, manufacturing, and inspection into one continuous process. Built as software that works with existing tools, the system can automatically plan how the fibers are placed, predicting and reducing defects while improving both manufacturability and strength. The system optimizes not only individual layers but also how defects are distributed across all layers, preventing them from stacking up in ways that weaken the final part. It also uses inspection results from completed parts to improve future designs, creating a feedback loop where each stage informs the others. The system was tested by designing a composite panel using this new approach and comparing it to a panel made with state-of-the-art manual planning methods. The results showed that the system could intentionally control where defects appeared and increase the efficiency of the planning process. By unifying design, manufacturing, and inspection, this research shows a way to make advanced composite manufacturing more efficient, consistent, and cost-effective. This approach lowers the barrier to using automated fiber placement and opens the door for its wider adoption not only in aerospace but also in industries such as automotive, wind energy, maritime, and sporting goods, where strong and lightweight structures are essential.

Computer-Aided Process Planning↗