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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Closed-Loop Control of Laser Surface Treatment via Laser Induced Breakdown Spectroscopy

Current composite manufacturing rates need to be improved to meet the projected demand of on-demand mobility (ODM) and commercial aircraft transports. Some processes in the composite manufacturing chain require touch labor such as manual sanding of surfaces prior to bonding and installing thousands of mechanical fasteners in bonded primary structures to achieve Federal Aviation Administration (FAA) certification. The consequences are insufficient repeatability in bond quality resulting in increased manufacturing time. The FAA has indicated that robust process control is one potential means of certifying bondlines in primary structure. This work describes an integrated closed-loop control system between a laser machining system and laser induced breakdown spectroscopy (LIBS) instrumentation as a means of assuring bond performance and achieving certified composite bondlines. Preliminary results from in-situ treatment and inspection are presented. Results indicate that LIBS inspection and laser treatment can be conducted simultaneously to achieve high quality, reliable adhesive bonds.

Rodolfo Ledesma↗
NASA NTRSDate not supplied

Closed-Loop Control of Laser Surface Treatment via Laser Induced Breakdown Spectroscopy

Current composite manufacturing rates need to be improved to meet the projected demand of on-demand mobility (ODM) and commercial aircraft transports. Some processes in the composite manufacturing chain require touch labor such as manual sanding of surfaces prior to bonding and installing thousands of mechanical fasteners in bonded primary structures to achieve Federal Aviation Administration (FAA) certification. The consequences are insufficient repeatability in bond quality resulting in increased manufacturing time. The FAA has indicated that robust process control is one potential means of certifying bondlines in primary structure. This work describes an integrated closed-loop control system between a laser machining system and laser induced breakdown spectroscopy (LIBS) instrumentation as a means of assuring bond performance and achieving certified composite bondlines. Preliminary results from in-situ treatment and inspection are presented. Results indicate that LIBS inspection and laser treatment can be conducted simultaneously to achieve high quality, reliable adhesive bonds.

Rodolfo Ledesma↗
NASA NTRSDate not supplied

Improved Method for Increased-Rate Stitched Composites Manufacturing

Stitched composites, as defined herein, are created by stitching a dry preform, infusing the preform with resin and curing the resin. Stitched composites have been shown to have benefits over unstitched composites for stiffened structures, including improved damage tolerance, reduced weight, and fewer fasteners. However, conventional stitched composite structure production is very time and manual-labor intensive, and therefore is not conducive for high-rate production of commercial aircraft main structure. The National Aeronautics and Space Administration (NASA) Hi-Rate Composite Aircraft Manufacturing (HiCAM) Project has the objective to increase the manufacturing rate for future composite aircraft. Stitched resin infused (SRI) composites are one of the technologies being considered under the HiCAM Project, but to be viable, production rates must be increased (i.e., production time reduced). Previous work has shown that it is possible to reduce the time required to stitch a dry composite preform, such as a skin with integral stiffeners, but the stitching process is a small portion of the total time required to produce a stitched preform. To significantly reduce overall stitched preform production time, a study was undertaken to examine a new stitching method that would yield time reduction in the pre- and post-stitching activities that include all portions of a stitched preform production with the exception of the actual stitching process. The new method resulted in significant reduction in production time, from 27% to 40%, while at the same time reducing the costs associated with fabricating a stitched preform by eliminating stations within the production line, simplifying tooling, reducing labor, and reducing consumables.

Stitching↗
NASA NTRSDate not supplied

Improved Method for Increased-Rate Stitched Composites Manufacturing

Stitched composites, as defined herein, are created by stitching a dry preform, infusing the preform with resin and curing the resin. Stitched composites have been shown to have benefits over unstitched composites for stiffened structures, including improved damage tolerance, reduced weight, and fewer fasteners. However, conventional stitched composite structure production is very time and manual-labor intensive, and therefore is not conducive for high-rate production of commercial aircraft main structure. The National Aeronautics and Space Administration (NASA) Hi-Rate Composite Aircraft Manufacturing (HiCAM) Project has the objective to increase the manufacturing rate for future composite aircraft. Stitched resin infused (SRI) composites are one of the technologies being considered under the HiCAM Project, but to be viable, production rates must be increased (i.e., production time reduced). Previous work has shown that it is possible to reduce the time required to stitch a dry composite preform, such as a skin with integral stiffeners, but the stitching process is a small portion of the total time required to produce a stitched preform. To significantly reduce overall stitched preform production time, a study was undertaken to examine a new stitching method that would yield time reduction in the pre- and post-stitching activities that include all portions of a stitched preform production with the exception of the actual stitching process. The new method resulted in significant reduction in production time, from 27% to 40%, while at the same time reducing the costs associated with fabricating a stitched preform by eliminating stations within the production line, simplifying tooling, reducing labor, and reducing consumables.

Stitching↗
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 NTRS2024

Ultra-Light, Strong, and Self-Reprogrammable Mechanical Metamaterials

Versatile programmable materials have long been envisioned that can reconfigure themselves to adapt to changing use cases in adaptive infrastructure, space exploration, disaster response, and more. We introduce a robotic structural system as an implementation of programmable matter, with mechanical performance and scale on par with conventional high-performance materials and truss systems. Fiber reinforced composite truss-like building-blocks form strong, stiff, and lightweight lattice structures as mechanical metamaterials. Two types of mobile robots operate over the exterior surface and through the interior of the system, performing transport, placement, and reversible fastening using the intrinsic lattice periodicity for indexing and metrology. Leveraging programmable matter algorithms to achieve scalability in size and complexity, this system design enables robust collective automated assembly and reconfiguration of large structures with simple robots. We describe the system design and experimental results from a 256-unit cell assembly demonstration and lattice mechanical testing, as well as demonstration of disassembly and reconfiguration. The assembled structural lattice material exhibits ultra-light mass density (0.0103 g/cc) with high strength and stiffness for its weight (0.01117 MPa and 1.1129 MPa, respectively), a material performance realm appropriate for applications like space structures. With simple robots and structure, high mass-specific structural performance, and competitive throughput, this system demonstrates potential for self-reconfiguring autonomous metamaterials for diverse applications.

Robotics↗
NASA NTRS2022

Nickel Accounting for the Psyche Spacecraft

(16) Psyche is a metallic asteroid located in the asteroid belt between Mars and Jupiter. The composition of Psyche is unknown, but the scientific community believes it is composed primarily of iron, nickel, and silicates. The Psyche mission will be the first space mission to study this asteroid in close proximity, and as such, determining the elemental composition of the asteroid is one of the primary objectives of this mission. The spacecraft measures this elemental composition using the Gamma Ray Spectrometer (GRS), which detects cosmic-ray induced gamma rays from Psyche’s surface. In addition to gamma rays from Psyche, the GRS will also detect background gamma rays from spacecraft material nearby the GRS. Because Ni is such a key element for achieving the mission science objectives, a special effort has been made to minimize background gamma rays from Ni, thus maximizing sensitivity to the Ni gamma rays from Psyche. The Psyche mission approached this potential Ni background issue with two methods, locating the GRS away from the spacecraft/source of Ni and setting a threshold requirement against the total Ni mass of the spacecraft. Verifying that the total amount of Ni on the spacecraft is within this threshold is thus a key part of ensuring the GRS can meet its science objectives. Ni alloys are commonly used on spacecraft, but the mass and location of elemental Ni is not typically accounted for. Ni can be found in many materials all over the spacecraft – in fasteners, optical assemblies, magnetic shields, electromagnetic interference (EMI) tape, under gold plating, in connectors, and so on. The task proved further challenging because of the distributed nature of the build of this spacecraft and the use of vendors external to JPL. While the analysis was atypical and source information difficult to cull, the approach described in this paper used tools and data available to the project in some form to achieve this goal. For Psyche, current analysis shows there is margin in meeting the Ni mass requirement. This paper explains the systematic approach used to accurately and precisely record Ni content on the Psyche spacecraft. While this approach was defined for Ni on the Psyche mission, it can be used for any future project with the need to accurately estimate the total mass of a single element or material on a spacecraft.

Cullinan, Joe↗
NASA NTRS2022

Nickel Accounting for the Psyche Spacecraft

(16) Psyche is a metallic asteroid located in the asteroid belt between Mars and Jupiter. The composition of Psyche is unknown, but the scientific community believes it is composed primarily of iron, nickel, and silicates. The Psyche mission will be the first space mission to study this asteroid in close proximity, and as such, determining the elemental composition of the asteroid is one of the primary objectives of this mission. The spacecraft measures this elemental composition using the Gamma Ray Spectrometer (GRS), which detects cosmic-ray induced gamma rays from Psyche’s surface. In addition to gamma rays from Psyche, the GRS will also detect background gamma rays from spacecraft material nearby the GRS. Because Ni is such a key element for achieving the mission science objectives, a special effort has been made to minimize background gamma rays from Ni, thus maximizing sensitivity to the Ni gamma rays from Psyche. The Psyche mission approached this potential Ni background issue with two methods, locating the GRS away from the spacecraft/source of Ni and setting a threshold requirement against the total Ni mass of the spacecraft. Verifying that the total amount of Ni on the spacecraft is within this threshold is thus a key part of ensuring the GRS can meet its science objectives. Ni alloys are commonly used on spacecraft, but the mass and location of elemental Ni is not typically accounted for. Ni can be found in many materials all over the spacecraft – in fasteners, optical assemblies, magnetic shields, electromagnetic interference (EMI) tape, under gold plating, in connectors, and so on. The task proved further challenging because of the distributed nature of the build of this spacecraft and the use of vendors external to JPL. While the analysis was atypical and source information difficult to cull, the approach described in this paper used tools and data available to the project in some form to achieve this goal. For Psyche, current analysis shows there is margin in meeting the Ni mass requirement. This paper explains the systematic approach used to accurately and precisely record Ni content on the Psyche spacecraft. While this approach was defined for Ni on the Psyche mission, it can be used for any future project with the need to accurately estimate the total mass of a single element or material on a spacecraft.

Cullinan, Joe↗
NASA NTRSDate not supplied

Space Environmental Effects on Additively Manufactured (AM) Parts on MISSE-9/10

The Materials on International Space Station Experiment (MISSE) experiments have made a continuous presence on the ISS and allowed for testing of performance and durability of materials exposed to the Low Earth Orbit (LEO) space environment. Additive Manufacturing (AM) has been included on several MISSE flights, and many companies (Made in Space and Stratasys) and NASA centers wanted to test various AM material samples. The goal of studying Space Environmental Effects on AM Parts with MISSE 9/10 was to fly promising metal and polymer AM material samples (dogbones) that could have future manufacturing capabilities in space and conduct post-flight analysis of these materials via mechanical and optical properties. On these flights, MISSE-9 contained three Inconel and 12 polymer samples and MISSE-10 included three Inconel and 16 polymer samples. Before flying to the ISS, all the samples underwent thermal vacuum bakeout (at 60°C for 24 hours and 10-6 torr). After bakeout, control and flight experiments were weighed, and optical properties were taken. Optical properties such as solar absorptance and infrared emittance were taken on the control samples and the flight samples (preflight and postflight). To understand more about the mechanical properties of the AM samples, tensile testing was conducted on the samples to determine the yield and ultimate tensile strength of the dogbone samples against the control. Lastly, the torque on the fasteners securing the dogbone samples were loosened and measured using a torque wrench, and they were compared to their preflight torque specification required. Optical properties and masses were recorded before flight, and the postflight results were recorded and collected in June 2022 to study the effects. With the postflight analysis, NASA engineers can now better characterize how UV radiation and thermal cycling affect the material properties.

Materials on International Space Station Experimen↗
NASA NTRSDate not supplied

Windform® XT 2.0 Use as 3U CubeSat Primary Structure

CubeSats provide a platform for small-scale space research and technology demonstration at reduced complexity, cost, and development time. These advantages drove the NASA Langley Research Center (LaRC) to develop and launch the GPX2 3U CubeSat to explore the viability of using Commercial Off-The-Shelf (COTS) differential GPS (dGPS) in low earth orbit. To reduce manufacturing costs and increase design flexibility, the project chose additive manufactured Windform® XT 2.0 as the primary bus material rather than traditional subtractive manufactured (milled) metal. The bus is a two-part, Selective Laser Sintered, 3D-print structure consisting of a single-piece, five-walled chassis and single-walled cover. The bus was specially designed to allow the project to accommodate the payload electronics stack as well as antennas, receivers, and deployable mechanisms. By using an additive manufactured solution, LaRC was able to design in features unrealizable through traditional milling, with a lead-time of roughly two weeks. In comparison, traditional subtractive manufacturing limits geometry options due to toolpath reach and bus construction would have required multiple components for each wall. This would have resulted in a more costly, longer lead-time article with more joints, fasteners, and complexity with a commensurate increase in overall mass. A number of lessons-learned were captured during the design, analysis, and testing of the GPX2 CubeSat covering thermal and structural analysis, vibration modeling, and geometric tolerancing. Additionally, a variety of material testing and verification were performed before and during spacecraft design and integration to assure the suitability of Windform® XT 2.0 for the launch and mission environments. This article provides the highlights of designing and testing the GPX2 bus.

CubeSat↗
NASA NTRSDate not supplied

NASA Hypersonic Technology Project (HTP) Sic/Sic Research for Hypersonic Airframe Hot Structures

NASA, DoD, and industry have a vision for reusable hypersonic vehicles (Mach > 5). The realization of this vision will require high-specific-strength, high-temperature, reusable structures. In the sub-sonic and super-sonic flight regimes, high-specific-strength structures can be obtained via the use of aluminum, titanium, and composite materials. In the high-temperature environments associated with hypersonic flight, high specific strength is obtained via refractory-composite materials. NASA Langley Research Center has multiple efforts focused on evaluating refractory-composites for use as airframe hot structures for reusable hypersonic vehicles. Current efforts span the entire building-block approach, including coupon, element, subcomponent, and component manufacturing and testing. The coupon level work includes progressive damage testing and analysis, restrained thermal growth testing, stressed oxidation testing, and evaluation of laminate properties for several layup architectures. Element testing includes evaluation of over-temperature capability and co-processed bonding, fastener development, and analysis and testing of beaded (corrugated) panels. Subcomponent level work includes a solid rocket motor booster fin and a portion of a control surface, and component level work includes a control surface static test article. The current presentation will provide an update of work being performed at NASA, which is expected to transition refractory-composite materials to reusable hypersonic vehicles.

Hot structures↗
NASA NTRSDate not supplied

The Repair Maintenance and Fabrication Facility in the Common Habitat Architecture

The Common Habitat Architecture seeks to increase the habitability of long-duration human spaceflight systems. A key aspect of this is vehicle survivability. Missions beyond low Earth orbit need onboard capabilities for Repair, Maintenance, And Fabrication (RMAF) to overcome potential contingency scenarios. Strategies employed in historic human spaceflight such as redundancy management, reliability, sparing, orbital replacement units, and aborts may be insufficient by themselves. Based on subject matter input, a list of 53 critical failures defining a set of incidents that can render a key spacecraft subsystem inoperable were generated. A subsequent analysis found that a robust in-space RMAF system capable of performing 14 key functions can potentially repair a subsystem plagued by any of these failures. An ancillary benefit is this capability may provide psychological benefits to the crew, by enabling greater self-sufficiency in earth-independent problem solving. A basic RMAF facility has been defined for the Common Habitat, consisting of five workstations. A work bench and computer workstation provide a multipurpose horizontal work surface, computing interface, and tools storage. A CNC machining center provides a subtractive manufacturing capability for metals and plastics. A multi-material 3D printing facility provides additive manufacturing capabilities for plastic, metals, and printed electronics. A welding facility is used for joining metal components where a higher strength is needed than can be achieved with fasteners or adhesives. A glovebox facility is used to perform work that is too hazardous for any of the other workstations. This may include hardware brought in from outside the spacecraft that could potentially contaminate the cabin environment. Forward work includes considering the accommodation of additional manufacturing processes not modeled in the current system, assessing the ability of systems to operate in partial gravity and microgravity environments, incorporation of the system into the Common Habitat Computer Aided Design (CAD) model, bottoms-up mass estimating, and a crew time analysis.

Habitat↗
NASA NTRSDate not supplied

Thermoplastic Aircraft Manufacturing Model

A manufacturing process to produce 80 single aisle sized thermoplastic composite aircraft a month is presented. Within the Hi-Rate Composite Aircraft Manufacturing (HiCAM) program a baseline aircraft fuselage was designed and modeled based on the Boeing 787 one-piece barrel build, but scaled down to a single aisle aircraft. Key Performance Parameters (KPPs) of recurring cost, non-recurring cost, and weight are calculated from the model. The outcome of this simulation indicated a prohibitively high non-recurring cost, so the thermoplastic team modeled the manufacturing flow for a panelized fuselage concept. The minimum success criteria is a 30% cost reduction and less than two percent increase in weight compared to the baseline aircraft. Automated Fiber Placement (AFP) is used to layup skins, stamp forming is used to make body frames, stringers, and passenger door surround parts. In the model, Co-Fusion was used to consolidate the skin while simultaneously welding on stringers to make stiffened panels. Additional welding technologies were considered to minimize fastener use and a vertical build concept was used for final assembly to reduce factory size. This manufacturing concept for a thermoplastic aircraft indicated a 39% cost and 12% weight savings compared to the baseline. Additional trade studies were conducted using Northrop Grumman’s Automated Stiffener Forming technology to make stringers and stamp forming instead of AFP to make skins.

thermoplastic composites↗
NASA NTRS2024

Advances in Thermoplastic Composites Over Three Decades – A Literature Review

Recently, there has been a renewed interest in thermoplastic composites driven mainly by advances in automation which can lead to significant cost reductions by increasing manufacturing rates while simultaneously reducing the part count and energy consumption relative to the manufacturing of thermoset composites. At the same time, new material systems have been developed and the thermoplastic composites prepreg material quality has improved over time. Additionally, thermoplastic composites have nearly infinite room-temperature shelf life and production scraps can be reused and retired parts can be recycled, providing opportunities for more sustainable operations and downstream markets. These factors have contributed to a strong interest in advancing thermoplastic composites for aerospace, automotive, and other industrial applications. The objective of the present study is to assess the state of the art regarding the maturity and performance of thermoplastic composite materials and to document the advances made in manufacturing and assembly of thermoplastic composite structures. Therefore, more than 200 NASA reports, conference proceedings and journal papers spanning three and a half decades (1986-2022) were reviewed and the findings summarized. The current study, however, is limited in scope with a focus on the application of thermoplastic composites to aircraft structures. Further, the study does not provide a comprehensive assessment of the available literature but rather offers an overview of past and present research being conducted in the field of thermoplastic composites. The information provided may be used to identify gaps and help guide future research and development. In this report, the opportunities offered by the use of thermoplastic composites in general and challenges encountered in particular are presented first. Second, an overview of thermoplastic materials is presented, manufacturing methods are discussed, and new methods for fastener-less assembly such as welding are introduced. Third, tests that have been performed on different levels of the building block, ranging from coupon to structural level, are presented. In particular, fracture toughness results for different thermoplastic composite materials are used to demonstrate their performance compared to state-of-the-art thermoset composites. Fourth, several examples of analyses are discussed including process modelling and progressive damage analysis (PDA). An extensive list of references and appendices with tables support the narrative. Lastly, a brief summary with outlook and recommendations for future research is provided.

composites↗
NASA NTRSDate not supplied

Thermoplastic Composite Aircraft Manufacturing Model

A manufacturing process to produce 80 single aisle sized thermoplastic composite aircraft a month is presented. Within the Hi-Rate Composite Aircraft Manufacturing (HiCAM) program a baseline aircraft fuselage was designed and modeled based on the Boeing 787 one-piece barrel build, but scaled down to a single aisle aircraft. Key Performance Parameters (KPPs) of recurring cost, non-recurring cost, and weight are calculated from the model. The outcome of this simulation indicated a prohibitively high non-recurring cost, so the thermoplastic team modeled the manufacturing flow for a panelized fuselage concept. The minimum success criteria is a 30% cost reduction and less than two percent increase in weight compared to the baseline aircraft. Automated Fiber Placement (AFP) is used to layup skins, stamp forming is used to make body frames, stringers, and passenger door surround parts. In the model, Co-Fusion was used to consolidate the skin while simultaneously welding on stringers to make stiffened panels. Additional welding technologies were considered to minimize fastener use and a vertical build concept was used for final assembly to reduce factory size. This manufacturing concept for a thermoplastic aircraft indicated a 39% cost and 12% weight savings compared to the baseline. Additional trade studies were conducted using Northrop Grumman’s Automated Stiffener Forming technology to make stringers and stamp forming instead of AFP to make skins.

thermoplastic composites↗
NASA NTRS2024

Offset-Stoichiometric Reflowable Composite Bonding Method with Adhesive for Mitigating Strict Faying Surface Tolerances

Inherent susceptibility of adhesive bonds to miniscule quantities of contamination can cause undetectable weakened bonds. For this reason, the Federal Aviation Administration (FAA) places strict regulations on adhesively bonded joints in primary aircraft structures. To meet certification requirements aircraft manufactures resort to redundant load paths in the form of fasteners which inherently add weight to the structure and increase manufacturing time. In prior work, a secondary bonding technique called AERoBOND was developed, which utilized off-stoichiometric epoxymatrix resins to facilitate reflow and diffusion of the resin within the joint interface during a secondary bonding/cure process, thus achieving a bond similar to a co-cured joint. However, the AERoBOND process required tight spatial tolerances between the two parts being joined. This study examined the utilization of conventional adhesive with the AERoBOND method to act as a filler in the joint line, effectively reducing the need for tight tolerances on the joining parts and serving as a flexible alternative for existing manufacturing processes. Ultrasonic inspection, optical microscopy, and ASTM International standard tests were performed to analyze the joints for defects and to quantify the mode-I and mode-II interlaminar fracture toughness and short beam strength of the proposed methodology with varying manufacturing parameters. The comprehensive results indicate that the AERoBOND+ method with and without surface preparation performs comparably to co-cured and conventional, adhesively bonded joints when secondary cured in an autoclave with 791 kPa of pressure. As an example, the AERoBOND+ panel without surface preparation bonded with 791 kPa of pressure (referred to as AB+3 throughout paper) had a mode-I fracture toughness (G Ic ) of 0.643 kJ/m 2 and a mode-II fracture toughness (G IIc ) of 4.000 kJ/m 2 in the non-precracked condition and 4.218 kJ/m 2 in the precracked condition at the adhesive-to-prepreg interface. These results were 96%, 142%, and 217%, respectively, of a co-cured baseline panel (referred to as C1 throughout paper).

Composites↗
NASA NTRS2024

Transitional Flow in the Wake of a Pseudorandom Roughness on a Supersonic Flat Plate

Prediction of roughness-induced transition is a practical requirement for the optimized design of high-speed vehicles, which rely on a reduction of thermal stresses and drag for higher performance. This reduction is affected by boundary layer transition, which on real vehicles may be caused by unavoidable roughness such as inherent surface material roughness and deterioration or roughness in the form of sensing elements, fasteners, etc. A better understanding of transition resulting from roughness can help inform acceptable manufacturing tolerances or maintenance requirements on real vehicles.

Amanda Chou↗