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All sourcesNASA NTRSDOE OSTI
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↗
NASA NTRSDate not supplied

Sources of Uncertainty in Ultrasonic Phase-Based Assessment of Interfacial Stiffness in Adhesive Joints

Quantitative nondestructive evaluation (NDE) of adhesive joint strength is critical for certifying fastener-free bonded aerospace structures. Unfortunately, available adhesive NDE approaches are only suitable for detecting gross defects and delamination and cannot reliably detect variations in strength in intact bonds. Intact (i.e., nominally defect free) bonds are challenging to inspect because they, by definition, contain no macroscopic defects that would produce a strong NDE signature. Even in the absence of obvious defects, however, bond strength can vary substantially due to factors such as contamination, variations in the adhesive cure processes, and environmental exposure. For ultrasonic bond inspections, the effect of interactions from an imperfect adhesive interface are subtle and may include changes in the frequency content or phase of the ultrasonic signal. The bond condition can be deduced from these subtle changes in the acoustic response using an accompanying physics-based model of adhesive joint interaction. For example, researchers have used the spring network adhesion model of Baik and Thompson [1] to determine the interfacial stiffness of the joint from its ultrasonic response (e.g., [2]). This interfacial stiffness metric describes the bonds resistance to separation between the adhesive and adherent and has shown a strong correlation with joint strength. In this study, interfacial stiffness is assessed using the ultrasonic phase change associated with bond interaction in varying bond conditions. The phase change is measured precisely using a phase-locked-loop instrument over a swept range of frequencies. The phase transition at resonance, where the bond thickness is half the acoustic wavelength, produces a unique signature with which to fit a physics-based model. Interfacial stiffness is determined by minimizing the difference between the theoretical and experimental phase response. Experimental results are presented for glass specimens bonded with aerospace adhesives and a sensitivity analysis is performed to examine the effects of parameters that may be uncertain in the industrial inspection environment – namely bond thickness, adhesive and adherent properties, and porosity. The results are a first step toward an uncertainty quantified version of the measurement technique, show correlation between interfacial stiffness and bond strength, and are another key step toward maturing this approach for use in complex materials and joint geometries. [1] Baik, J.M. and Thompson, R.B., 1984. “Ultrasonic scattering from imperfect interfaces: a quasi-static model”. Journal of Nondestructive Evaluation, 4. [2] Haldren, H.A., Perey, D.F., Yost, W.T., Cramer, K.E., and Gupta, M.C., 2019. “Swept-frequency ultrasonic phase evaluation of adhesive bonding in tri-layer structures”. Journal of the Acoustical Society of America, 145.

Matt Webster↗
NASA NTRSDate not supplied

Structural Analysis, Modeling, and Testing of Multi-nut Joint for TriTruss Structure

The Multi-nut joint, developed at NASA Langley Research Center (LaRC), was designed to assemble TriTruss modules for a doubly curved 65.6-feet telescope support structure. Axial and bending tests for a multi-nut joint in the TriTruss structure were conducted to characterize the structural performance of the multi-nut joint and to obtain the installation torque value of the attachment fastener. This paper describes the summary of the axial and bending tests of the multi-nut joint, along with the test results. The equivalent axial stiffness (EA) and equivalent bending stiffness (EI) of the multi-nut joint were characterized based on these test results and are presented in this report. Additionally, finite element models (FEMs) of the axial and bending test specimens were developed, and the predicted axial and bending responses of the multi-nut joint were compared with the test results. Test-analysis comparison results show that the predicted axial load-displacement response of the test specimen was within 5 percent of the test results, while the predicted load-bending response of the test specimen exhibited a consistent trend with the test data.

TriTruss↗
NASA NTRSDate not supplied

Improvements to Stitching Controls for Manufacturing Advanced Stitched Composites

Stitched composites have been extensively studied and have been shown to have benefits over unstitched composites for stiffened structures, including improved damage tolerance, reduced weight, and fewer fasteners. These benefits make stitched composites attractive from a performance perspective, however, the nature of stitching the preforms prior to infusion includes additional manufacturing steps and equipment compared to conventional resin infusion. Stitching of composite material structures at NASA Langley Research Center (LaRC) is performed at the Integrated Structural Assembly of Advanced Composites (ISAAC) facility using control algorithms developed by the lead author. Two deficiencies have been identified in the current control algorithms, and mitigations for these deficiencies were developed within the Hi-Rate Composite Aircraft Manufacturing (HiCAM) project and are presented herein. First, decoupling stitch translation and rotation takes more time for stitching since two robot operation control steps are executed for each stitch. Therefore, a method was developed to determine when to decouple the translation and rotation for stitching, which reduces stitch time by decoupling the operation control steps only when necessary to execute the stitch properly. Second, a correction to the sideslip stitching approach that eliminates a significant portion of the pre- and post-stitching activities was developed. The developed method adjusts the sideslip stitch angle based on the rotations of previous stitches to ensure that the width of the stitching seam remains constant along its length. Constant seam width eliminates observed overlap of the insertion and catcher thread portions for highly curved paths, which maintains the desired uniformity of through-the-thickness architecture along the seam and eliminates the local degradation of properties where the insertion needle is too close to, or crosses, the catcher needle portion of the seam.

stitching↗
NASA NTRSDate not supplied

Structural Analysis, Modeling, and Testing of Multi-nut Joint for TriTruss Structure

The Multi-nut joint, developed at NASA Langley Research Center (LaRC), was designed to assemble TriTruss modules for a doubly curved 65.6-feet telescope support structure. Axial and bending tests for a multi-nut joint in the TriTruss structure were conducted to characterize the structural performance of the multi-nut joint and to obtain the installation torque value of the attachment fastener. This paper describes the summary of the axial and bending tests of the multi-nut joint, along with the test results. The equivalent axial stiffness (EA) and equivalent bending stiffness (EI) of the multi-nut joint were characterized based on these test results and are presented in this report. Additionally, finite element models (FEMs) of the axial and bending test specimens were developed, and the predicted axial and bending responses of the multi-nut joint were compared with the test results. Test-analysis comparison results show that the predicted axial load-displacement response of the test specimen was within 5 percent of the test results, while the predicted load-bending response of the test specimen exhibited a consistent trend with the test data.

Structural analysis↗
NASA NTRS2025

Mechanics of Preloaded Bolt Tensile Loading With Focus on Load Introduction Factor

The bolt tensile and joint separation loads are directly influenced by the locations at which the external loads enter the clamped members of a preloaded bolted joint (PBJ) and the associated load-paths through the joint. This physical load introduction mechanism affecting the bolt tensile loading is typically represented in the bolt tensile load equation, in part, by a load introduction factor (LIF), which was shown by H.M. Lee of Marshall Spaceflight Center to be a natural product of the bolt tensile load equation using a linear spring stiffness model. This LIF, being a function of load-path stiffness, has subsequently been denoted as the stiffness-based LIF (SBLIF), providing a framework to calculate the LIF using whatever load-path stiffness approximations are appropriate. Expanding upon the work of Lee, it is shown that the SBLIF and the joint stiffness factor are functions of the stiffnesses of the same load-paths and regions within a PBJ, and thus they should not be treated as independent variables. Mathematical expressions for the SBLIF are presented. Comparisons are shown between the analytically calculated SBLIF, the analytically calculated geometric LIF (GLIF), which is a simple clamped-member thickness ratio, the experimentally derived LIF, and the LIF determined by finite element analysis (FEA). Using experiment and FEA as a benchmark, the SBLIF, using traditional load-path stiffness approximations, enables a more accurate prediction of bolt tensile loading than the GLIF, although it can be unconservative near joint separation. The GLIF generally attributes more of the externally applied tensile load to the bolt than does the SBLIF, potentially resulting in heavier and/or more costly bolted joints. Mathematical relationships between the SBLIF and the GLIF are developed. Supplemental material is provided in the appendixes where the historical practice of using the joint compressive stiffness in place of the joint tensile stiffness is evaluated. The appendixes include step-by-step examples demonstrating the calculation of the SBLIF using traditional stiffness approximations and conclude with the development of the joint diagram in terms of the SBLIF, culminating into formulas for the key features of a joint diagram, which is useful for programming.

Load Path↗
NASA NTRS2015

Evaluation of Long-Term Calibrations of the AVHRR Visible Radiances

Two systematic calibrations have been compiled for the visible radiances measured by the series of AVHRR instruments flown on the NOAA operational polar weather satellites: one by the International Satellite Cloud Climatology Project (ISCCP), anchored on NASA ER-2 underflights in the 1980s and early 1990s and covering the period 1981-2009, and one by the PATMOS-x project, anchored on comparisons to the MODIS instruments on the Aqua and Terra satellites in the 2000s and covering the period 1979-2010 (this result also includes calibration for the near-IR channels). Both methods have had to extend their anchor calibrations over a long series of instruments using different vicarious approaches, so a comparison provides an opportunity to evaluate how well this extension works by cross-checking the results at the anchor points. The basic result of this comparison is that for the ''afternoon'' series of AVHRRs, the calibrations agree to within their mutual uncertainties. However, this retrospective evaluation also shows that the representation of the time variations can be simplified. The ISCCP procedure had much more difficulty extending the calibration to the ''morning'' series of AVHRRs with the calibrations for NOAA-15 and NOAA-17 exceeding the estimated uncertainties. Given the general agreement, a new calibration for all AVHRR visible radiances (except TIROS-N, NOAA-6, NOAA-19, and MetOp-A) is proposed that is based on the average of the best linear fits to the two time records. The estimated uncertainty of these calibrations is 63% absolute (scaled radiance units).

Advanced Very High Resolution Radiometer↗
NASA NTRS2022

Calculating Factors of Safety and Margins of Safety from Interaction Equations for Preloaded Bolts

This report presents the derivation of the factored interaction equations for preloaded bolts in two- or three-dimensional space. This is an extension of the previous work of the first author which covered factored interaction equations consisting of load or stress-ratios which were either solely constant or solely varying quantities. The factored interaction equations for preloaded bolts consist of an axial load-ratio that is a combination of a constant quantity, preload and thermal load at preload, and a varying quantity, tensile and bending loads, and was not covered under the previous work. Two forms of the factored interaction equations for preloaded bolts, arising from two perspectives, are presented and compared. Closed form solutions for the tensile, shear, and bending loads as functions of the other concurrent loads that produces zero margin of safety are presented.

interaction equations↗