SEARCH · Search NASA

Results for “HONEYCOMB CORE”

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

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 271 records · Page 15

All sources NASA NTRS DOE OSTI

NASA NTRS2018

Applications of Parallel-Element, Embedded Mesh-Cap Acoustic Liner Concepts

This study explores progress achieved with 2DOF, 3DOF, and MDOF acoustic liners constructed with mesh caps embedded within a honeycomb core. These liner configurations offer potential for broadband noise reduction, and are suitable for conventional aircraft implementation. Samples for each configuration are tested in the NASA normal incidence tube and grazing flow impedance tube, with and without a wire mesh facesheet. Impedances based on these measured data compare favorably with those predicted using a transmission line impedance prediction model. Predicted impedances are then used as input for an aeroacoustic propagation code to compute axial acoustic pressure distributions in the grazing flow tube. These predicted distributions compare favorably with the corresponding measured distributions at frequencies away from the frequency of peak attenuation, but suffer slight degradation for frequencies very near the peak attenuation frequency, where the predicted results are sensitive to input impedance changes. As expected, the noise reduction frequency range increases as more degrees of freedom are included. Although the specific results achieved herein may differ from those that would be achieved with other 2DOF, 3DOF, and MDOF liners, this comparison highlights some of the key features that can be exploited in the design of parallel-element, embedded mesh-cap liners.

Jones, M. G.↗

NASA NTRS2019

Structural Analysis of Test Flight Vehicles for Application of Multifunctional Energy Storage System

Under the NASA Aeronautics Research Mission Directorate (ARMD) Convergent Aeronautical Solutions (CAS) project, NASA Glenn Research Center has been leading Multifunctional Structures for High Energy Lightweight Load-bearing Storage (M-SHELLS) research efforts. The technology of integrating load-carrying structures with electrical energy storage capacity has the potential to reduce the overall weight of future electric aircraft. The proposed project goals were to develop M-SHELLS in the form of honeycomb coupons and subcomponents, integrate them into the structure, and conduct low-risk flight-tests onboard a remotely piloted small aircraft. Experimental M-SHELLS energy-storing coupons were fabricated and tested in the laboratory for their electrical and mechanical properties. In this report, finite element model development and structural analyses of two small test aircraft candidates are presented. The finite element analysis of the initial two-spar wing is described for strain, deflection, and weight estimation. After a test aircraft Tempest was acquired, a load-deflection test of the wing was conducted. A finite element model of the Tempest was then developed based on the test aircraft dimensions and construction detail. The component weight analyses from the finite element model and test measurements were correlated. Structural analysis results with multifunctional energy storage panels in the fuselage of the test vehicle are presented. Although the flight test was cancelled because of programmatic reasons and time constraints, the structural analysis results indicate that the mid-fuselage floor composite panel could provide structural integrity with minimal weight penalty while supplying electrical energy. To explore potential future applications of the multifunctional structure, analyses of the NASA X-57 Maxwell electric aircraft and a NASA N+3 Technology Conventional Configuration (N3CC) fuselage are presented. Secondary aluminum structures in the fuselage sub-floor and cargo area were partially replaced with reinforced five-layer composite panels with M-SHELLS honeycomb core. The N3CC fuselage weight reduction associated with each design without risking structural integrity are described. The structural analysis and weight estimation with the application of composite M-SHELLS panels to the N3CC fuselage indicate a 3.2% reduction in the fuselage structural weight, prior to accounting for the additional weight of core material required to complete the energy storage functionality.

Mukhopadhyay, Vivek↗

NASA NTRS2019

Structural Analysis of Test Flight Vehicles with Multifunctional Energy Storage

Under the NASA Aeronautics Research Mission Directorate (ARMD) Convergent Aeronautical Solutions (CAS) project, NASA Glenn Research Center has been leading Multifunctional Structures for High Energy Lightweight Load-bearing Storage (M-SHELLS) research efforts. The technology of integrating load-carrying structures with electrical energy storage capacity has the potential to reduce the overall weight of future electric aircraft. The proposed project goals were to develop M-SHELLS in the form of honeycomb coupons and subcomponents, integrate them into the structure, and conduct low-risk flight tests onboard a remotely piloted small aircraft. Experimental M-SHELLS energy-storing coupons were fabricated and tested in the laboratory for their electrical and mechanical properties. In this paper, finite element model development and structural analyses of two small test aircraft candidates are presented. The finite element analysis of the initial two-spar wing is described for strain, deflection, and weight estimation. After a test aircraft Tempest was acquired, a load- deflection test of the wing was conducted. A finite element model of the Tempest was then developed based on the test aircraft dimensions and construction detail. The component weight analysis from the finite element model and test measurements were correlated. Structural analysis results with multifunctional energy storage panels in the fuselage of the test vehicle are presented. Although the flight test was cancelled because of programmatic reasons and time constraints, the structural analysis results indicate that the mid-fuselage floor composite panel could provide structural integrity with minimal weight penalty while supplying electrical energy. To explore potential future applications of the multifunctional structure, analyses of the NASA X-57 Maxwell electric aircraft and a NASA N+3 Technology Conventional Configuration (N3CC) fuselage are presented. Secondary aluminum structure in the fuselage sub-floor and cargo area were partially replaced with reinforced five-layer composite panels with M-SHELLS honeycomb core. The N3CC fuselage weight reduction associated with each design without risking structural integrity are described. The structural analysis and weight estimation with the application of composite M-SHELLS panels to the N3CC fuselage indicate a 3.2% reduction in the fuselage structural weight, prior to accounting for the additional weight of core material required to complete the energy storage functionality.

Mukhopadhyay, Vivek↗

NASA NTRS2020

Design and Analysis of Buckling-Critical Large-Scale Sandwich Composite Cylindrical Test Articles

It has long been established in the literature that the buckling response of thin-shell structures can be very sensitive to the presence of small geometric and loading imperfections. The Shell Buckling Knockdown Factor Project (SBKF) was established by the NASA Engineering and Safety Center (NESC) to develop analysis-based shell buckling design recommendations for stiffened-metallic and composite launch-vehicle shell structures. Large-scale buckling tests were used to validate the modeling and analysis methods applied in developing these analysis-based recommendations. Herein, the test article design methodology for 8-ft-diameter, honeycomb-core sandwich composite cylinder validation tests is discussed and cylinder designs are presented. In this methodology, first, the sandwich composite design space was defined using several nondimensional parameters, and the desired test article design space was determined by examining the designs of launch-vehicle cylinder structures. Essentially all test article designs within certain design parameters were generated and then downselected based on simple closed-form failure calculations and the nondimensional design-space parameters. Four of these designs that spanned a significant portion of the design space of interest and had global buckling as the first predicted failure mode were selected and subjected to higher-fidelity finite element analyses (FEAs): shell-element-based analyses, axisymmetric-element-based analyses, and global-local analyses. The analysis flow discussed in this report supported the design objective. As the analysis flow progressed, designs were downselected so the fidelity of the analysis methods, and consequently their computational cost and accuracy, was increased. The selection of the FEA types created an analysis framework where particular methods complemented each other and reduced the uncertainty of the predicted test article responses. The analysis results are illustrated using several designs when the computationally expeditious closed-form analysis stage is discussed. Once this stage is complete, the higher-fidelity FEA types are illustrated using one selected detailed test article design. Both perfect and imperfect test article geometries were considered.

Buckling↗

NASA NTRS2024

Test and Analysis of the 8-foot Diameter Cylindrical Sandwich Composite Test Article CTA8.2B: As part of the NASA Engineering and Safety Center Shell Buckling Knockdown Factor Project

The buckling response of thin-walled cylindrical structures have long been shown to be sensitive to various imperfections. That is, imperfections related to load introduction, boundary conditions, material property variation, and geometric deviations can all contribute to experimentally obtained buckling loads being lower than the buckling loads predicted for perfect representations of the same thin-walled cylindrical structure. To account for the influence of geometry-based imperfections of a thin-walled cylindrical structure on buckling response, a design factor (also known as a buckling knockdown factor) is used during the design phase. Guidelines for buckling knockdown factors are readily available in NASA SP-8007, Buckling of Thin Walled Circular Cylinders, which was updated in 2020. Prior to this recent update, NASA SP-8007 had not been updated since 1969, and since that time, computational analysis, experimental testing, and manufacturing methods have significantly improved. The NASA Engineering and Safety Center (NESC) Shell Buckling Knockdown Factor Project (SBKF, NESC Assessment 07-010-E) had the goal of developing buckling design recommendations for select classes of metallic and composite launch-vehicle structures [3]. Specifically, the test and analysis results from the fourth SBKF composite test article SBKF-P3-CYL-CTA8.2B, which is referred to as CTA8.2B, are described in this paper. This test was the fourth in a series of four tests on sandwich composite cylinders that will be used to experimentally validate analysis methods, which in turn can be used to develop new analysis-based shell buckling design guidelines for typical sandwich composite launch vehicle cylindrical structures. Publications describing previous test validation with analysis are available in Refs. CTA8.2B was designed to occupy one corner of the design space with approximately equal bending stiffnesses in the axial and transverse directions and have a larger ratio of radius to effective shell thickness. Each test article had a unique design and was designed to occupy a specific region of the design space. CTA8.2B was an 8-foot diameter honeycomb-core sandwich cylinder that was fabricated and tested at the NASA Marshall Space Flight Center (MSFC). The primary objectives of this test were to interrogate the structural capability of a composite test article, and to verify the test article design and analysis approach for cylinders subjected to axial compression and combined axial and bending loads. First, descriptions of the test article design and test are given in Section 2.0, and modeling and analysis methods used in support of the test article design and testing activities are described briefly in Section 3.0. Then, select test results are presented and compared to predicted results in Section 4.0, and concluding remarks are presented in Section 5.0. Finally, drawings of CTA8.2B and references used to support the test and analysis are provided in the Appendix.

Nonlinear analysis↗

NASA NTRS2022

Buckling Test and Analysis of the 8-Foot-Diameter Sandwich Composite Cylinder Test Article CTA8.2 as Part of the Shell Buckling Knockdown Factor Project: Test Dates 5-7 December 2017

It is well known that the buckling response of thin shell structures can be sensitive to small imperfections in the geometry and loading. The NASA Engineering and Safety Center (NESC) Shell Buckling Knockdown Factor Project (SBKF) has the goal of developing buckling design recommendations for select classes of metallic and composite shells. Part of completed SBKF work is described in this report. In particular, the test and analysis results from the second SBKF composite test article, CTA8.2, are described. This test was the second in a series of tests on sandwich composite cylinders that can be used to experimentally validate analysis methods, which in turn can be used to develop analysis-based shell buckling design guidelines for sandwich composite launch-vehicle cylindrical structures. CTA8.2 was an 8-foot diameter honeycomb-core sandwich cylinder that was fabricated and tested at the Marshall Space Flight Center (MSFC). The primary objectives of this test were to interrogate the structural capability of the composite test article, and to verify the test-article design and analysis approach for cylinders subjected to axial compression loads. This report contains the descriptions of the test-article design, fabrication, and test. The pre-test modeling and analysis methods, and corresponding results used in support of the test-article design and test planning, are also described. Additional post-test modeling and analysis efforts and results follow. Selected test results are compared to pre-test predictions and post-test analyses.

Adam Przekop↗

NASA NTRS2022

The Effects of Co-cured Versus Precured Carbon/Epoxy Face Sheets on the Modulus and Compression After Impact Strength of Honeycomb Sandwich Structure

This study presents experimental results of compression after impact (CAI) and stiffness testing of aluminum honeycomb core sandwich structure with quasi-isotropic face sheets that have either been co-cured or precured (and then bonded to the core in a secondary operation). The co-cured face sheets contain some ply waviness whereas the precured face sheets did not. The effect of this waviness was ascertained for two important design parameters for the Space Launch System (SLS) Payload Adapter Fitting (PAF); compressive modulus and CAI strength. It was found that there was no effect on the modulus between co-cured and precured face sheets using the material and processes in this study. In addition, there was no difference in CAI strength behavior between the co-cured and precured face sheet.

sandwich structure↗

NASA NTRS2023

Effects of Layer Spacing for a Multilayered Facesheet Acoustic Liner

A single degree-of-freedom acoustic liner, a perforate facesheet over honeycomb core, is modified to have a facesheet comprised of three layers. The top and bottom perforate layers have the same dimensions with aligned perforations while the middle layer perforate can be longitudinally shifted up to 0.075 inches to change the effective porosity of the facesheet. These configurations range in porosity from 100% open down to 10% open in addition to two closed configurations (0% porosity). Gaps are introduced between the facesheet layers ranging from no gap to 0.016 inches. These configurations are tested in the NASA Langley Grazing Flow Impedance Tube to determine their acoustic impedance spectra. Introduction of a gap affects the acoustic impedance of the liner where the resonant frequency of the liners decreases with increased gap thickness for liners with an open area ratio above 25% but increases with gap thickness for liners with an open area ratio below 25%. Additionally, it is observed that increasing the gap height results in a reduction of the effect of porosity on acoustic impedance. The results show that the presence of such gaps can have a significant effect on liner impedance especially as porosity decreases.

acoustic liners↗

NASA NTRS2024

Characterizing Face Sheet/Core Disbonding Using the Single Cantilever Beam Test: Results from an International Round Robin

Test results from an international round robin are presented. The single cantilever beam (SCB) test was used to characterize face sheet/core disbonding in sandwich components in an effort to help standardization. Each of the seven participating laboratories performed a set of five baseline tests using the same protocol, however, using their own specific equipment. In addition, each laboratory performed two sets of tests with altered tests conditions which included using a test fixture with a translating carriage and performing the tests with different loading and unloading speeds. The orientation of the disbond front with respect to the honeycomb core cells was varied and the effect on the fracture toughness was also studied. Additional factors that could influence test results were investigated, such as the use of a saw cut starter disbond as an alternative to Teflon® release film, and the effects of using a face sheet doubler to increase the bending stiffness. For each set of tests, summary results such as load/displacements plots, observed disbond growth location and calculated energy release rates are reported and a comparison of results between labs is presented. Critical strain energy release rate measurements from SCB tests are also compared with measurements made from an alternative test, namely the double cantilever beam with uneven bending moments (DCB-UBM) test, which has been studied extensively at the Technical University of Denmark and proposed as a mixed-mode test standard. A set of recommendations are made with respect to improving the SCB test and a path to standardization is presented.

Ronald Krueger↗

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 NTRS2025

ISS Radiator Face Sheet Anomaly Investigation and Return to Function

Maintaining sufficient heat rejection on the International Space Station (ISS) is critical to the function of the Low-Earth Orbit station. The External Active Thermal Control System (EATCS) rejects the excess heat generated by the US On-Orbit Segment (USOS) modules. The system uses single-phase liquid ammonia to collect the heat and reject it to six radiators (three on the Starboard side, three on the Port side). Each radiator is made up of eight panels. In September 2008, imagery of the Starboard radiators was conducted and showed one of the panels’ face sheets had peeled up from the internal honeycomb core structure. Out of an abundance of caution, the radiator was isolated from the rest of the EATCS and the ammonia was vented to space. Since the radiator was vented, periodic imagery of all radiator panels was taken to monitor any changes in the face sheet and inspect for other radiator anomalies. In January 2025, after years of trending and inspections, the radiator was reintegrated into the system to increase heat rejection capabilities. This paper will document the multi-year investigation that took place to determine root cause and mitigations implemented to reduce risk to the system. A summary of the data since the investigation will show rationale for reintegrating the radiator even with the damaged panel. The goal of the paper is to document the investigation for historical purposes and provide future programs with findings discovered during the investigation.

Aaron Rodriguez↗

NASA NTRS2026

ISS Radiator Face Sheet Anomaly Investigation and Return to Function

Aaron Rodriguez↗

NASA NTRS2017

Constitutive Modeling of the Facesheet to Core Interface in Honeycomb Sandwich Panels Subject to Mode I Delamination

A new cohesive zone traction-separation law, which includes the effects of fiber bridging, has been developed, implemented with a finite element (FE) model, and applied to simulate the delamination between the facesheet and core of a composite honeycomb sandwich panel. The proposed traction-separation law includes a standard initial cohesive component, which accounts for the initial interfacial stiffness and energy release rate, along with a new component to account for the fiber bridging contribution to the delamination process. Single cantilever beam tests on aluminum honeycomb sandwich panels with carbon fiber reinforced polymer facesheets were used to characterize and evaluate the new formulation and its finite element implementation. These tests, designed to evaluate the mode I toughness of the facesheet to core interface, exhibited significant fiber bridging and large crack process zones, giving rise to a concave downward concave upward pre-peak shape in the load-displacement curve. Unlike standard cohesive formulations, the proposed formulation captures this observed shape, and its results have been shown to be in excellent quantitative agreement with experimental load-displacement and apparent critical energy release rate results, representative of a payload fairing structure, as well as local strain fields measured with digital image correlation.

composites↗

NASA NTRS2012

Structural Analysis and Optimization of a Composite Fan Blade for Future Aircraft Engine

This report addresses the structural analysis and optimization of a composite fan blade sized for a large aircraft engine. An existing baseline solid metallic fan blade was used as a starting point to develop a hybrid honeycomb sandwich construction with a polymer matrix composite face sheet and honeycomb aluminum core replacing the original baseline solid metallic fan model made of titanium. The focus of this work is to design the sandwich composite blade with the optimum number of plies for the face sheet that will withstand the combined pressure and centrifugal loads while the constraints are satisfied and the baseline aerodynamic and geometric parameters are maintained. To satisfy the requirements, a sandwich construction for the blade is proposed with composite face sheets and a weak core made of honeycomb aluminum material. For aerodynamic considerations, the thickness of the core is optimized whereas the overall blade thickness is held fixed so as to not alter the original airfoil geometry. Weight is taken as the objective function to be minimized by varying the core thickness of the blade within specified upper and lower bounds. Constraints are imposed on radial displacement limitations and ply failure strength. From the optimum design, the minimum number of plies, which will not fail, is back-calculated. The ply lay-up of the blade is adjusted from the calculated number of plies and final structural analysis is performed. Analyses were carried out by utilizing the OpenMDAO Framework, developed at NASA Glenn Research Center combining optimization with structural assessment.

Coroneos, Rula M.↗

NASA NTRS2015

A Numerical Study on the Effect of Facesheet-Core Disbonds on the Buckling Load of Curved Honeycomb Sandwich Panels

A numerical study on the effect of facesheet-core disbonds on the post-buckling response of curved honeycomb sandwich panels is presented herein. This work was conducted as part of the development of a damage tolerance approach for the next-generation Space Launch System heavy lift vehicle payload fairing. As such, the study utilized full-scale fairing barrel segments as the structure of interest. The panels were composed of carbon fiber reinforced polymer facesheets and aluminum honeycomb core. The panels were analyzed numerically using the finite element method. Facesheet and core nodes in a predetermined circular region were detached to simulate a disbond induced via low-speed impact between the outer mold line facesheet and honeycomb core. Surface-to-surface contact in the disbonded region was invoked to prevent interpenetration of the facesheet and core elements. The diameter of this disbonded region was varied and the effect of the size of the disbond on the post-buckling response was observed. A significant change in the slope of the edge load-deflection response was used to determine the onset of global buckling and corresponding buckling load.

structural failure↗

NASA NTRS1990

A comparison of mechanical properties of some foams and honeycombs

A comparative study is conducted of the mechanical properties of foam-core and honeycomb-core sandwich panels, using a normalizing procedure based on common properties of cellular solids and related properties of dense solids. Seven different honeycombs and closed-foam cells are discussed; of these, three are commercial Al alloy honeycombs, one is an Al-alloy foam, and two are polymeric foams. It is concluded that ideal, closed-cell foams may furnish compressive strengths which while isotropic can be fully comparable to the compressive strengths of honeycombs in the thickness direction. The shear strength of ideal closed-cell foams may be superior to the shear strength of honeycombs.

Bhat, Balakrishna T.↗

DOE OSTI2020

Catalytic converters having non-linear flow channels

Disclosed is a honeycomb catalyst substrate core having geometrically non-linear flow channels. In an embodiment, the honeycomb catalyst substrate core includes helical flow channels. In another embodiment, the honeycomb catalyst substrate core includes sinusoidal flow channels. In yet another embodiment, the honeycomb catalyst substrate core includes helical plus sinusoidal flow channels. The honeycomb catalyst substrate core comprises a plurality of parallel non-linear flow channels formed along a longitudinal axis of symmetry of the catalyst substrate core, each non-linear flow channel configured such that eddies occurs during engine exhaust gas flow. Also disclosed is a method for manufacturing a ceramic honeycomb having non-linear flow channels, comprising the steps extrusion soft ceramic material through a die whilst the die moves through six degrees of freedom along its axis of symmetry. Disclosure includes a method for manufacturing a ceramic honeycomb having non-linear flow channels using three-dimensional printing.

Masoudi, Mansour↗

DOE OSTI2020

Catalytic converters having non-linear flow channels

Disclosed is a honeycomb catalyst substrate core having geometrically non-linear flow channels. In an embodiment, the honeycomb catalyst substrate core includes helical flow channels. In another embodiment, the honeycomb catalyst substrate core includes sinusoidal flow channels. In yet another embodiment, the honeycomb catalyst substrate core includes helical plus sinusoidal flow channels. The honeycomb catalyst substrate core comprises a plurality of parallel non-linear flow channels formed along a longitudinal axis of symmetry of the catalyst substrate core, each non-linear flow channel configured such that a turbulent vortical flow occurs during engine exhaust gas flow. Also disclosed is a method for manufacturing a ceramic honeycomb having non-linear flow channels, comprising the steps extrusion soft ceramic material through a die whilst the die moves through six degrees of freedom along its axis of symmetry. Disclosure includes a method for manufacturing a ceramic honeycomb having non-linear flow channels using three-dimensional printing.

33 ADVANCED PROPULSION SYSTEMS↗