Daniel N. Kosareo

Vantage Partners LLC, Brook Park, Ohio

Stanley T. Oliver 

Marshall Space Flight Center, Huntsville, Alabama

Brett A. Bednarcyk

Glenn Research Center, Cleveland, Ohio

Buckling Design and Analysis of a Payload 

Fairing 1/6th Cylindrical Arc-Segment Panel

NASA/TM—2013-216574

September 2013
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Daniel N. Kosareo

Vantage Partners LLC, Brook Park, Ohio

Stanley T. Oliver 

Marshall Space Flight Center, Huntsville, Alabama

Brett A. Bednarcyk

Glenn Research Center, Cleveland, Ohio

Buckling Design and Analysis of a Payload 

Fairing 1/6th Cylindrical Arc-Segment Panel

NASA/TM—2013-216574

September 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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NASA/TM—2013-216574 1

Buckling Design and Analysis of a Payload Fairing 1/6th 

Cylindrical Arc-Segment Panel

Daniel N. Kosareo

Vantage Partners LLC

Brook Park, Ohio 44142

Stanley T. Oliver

National Aeronautics and Space Administration

Marshall Space Flight Center

Huntsville, Alabama 35812

Brett A. Bednarcyk

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Design and analysis results are reported for a panel that is a 1/6th arc-segment of a full 33-ft diameter 

cylindrical barrel section of a payload fairing structure. Six such panels could be used to construct the 

fairing barrel, and, as such, compression buckling testing of a 1/6th arc-segment panel would serve as a 

validation test of the buckling analyses used to design the fairing panels. In this report, linear and nonlinear 

buckling analyses have been performed using finite element software for 1/6th arc-segment panels composed 

of aluminum honeycomb core with graphite/epoxy composite facesheets and an alternative fiber reinforced 

foam (FRF) composite sandwich design. The cross sections of both concepts were sized to represent 

realistic Space Launch Systems (SLS) Payload Fairing panels. Based on shell-based linear buckling 

analyses, smaller, more manageable buckling test panel dimensions were determined such that the panel 

would still be expected to buckle with a circumferential (as opposed to column-like) mode with significant 

separation between the first and second buckling modes. More detailed nonlinear buckling analyses were 

then conducted for honeycomb panels of various sizes using both Abaqus and ANSYS finite element codes, 

and for the smaller size panel, a solid-based finite element analysis was conducted. Finally, for the smaller 

size FRF panel, nonlinear buckling analysis was performed wherein geometric imperfections measured from 

an actual manufactured FRF were included. It was found that the measured imperfection did not 

significantly affect the panel's predicted buckling response.

1.0 Introduction

The Advanced Exploration Systems (AES), Composite for Exploration (CoEx) payload fairing 

project was tasked to design a composite honeycomb sandwich construction for use in the Space Launch 

Systems (SLS) Payload Fairing. The CoEx payload fairing project is a follow-on of the Lightweight 

Spacecraft Structures and Materials (LSSM) project, and its predecessor, the Structural Concepts Element 

of the Advanced Composites Technology (ACT) project. The goal of these three projects was to advance 

the Technology Readiness Level (TRL) of large scale composite structures for heavy lift launch vehicles.

The CoEx Payload Fairing project was focused on a honeycomb sandwich construction for its maturity,

affordability, and performance capability to support the SLS Payload Fairing design. This report also 

discusses the design and analysis of an alternate 1/6th Fiber Reinforced Foam (FRF) (Bednarcyk et al., 

2010, 2011; Zalewski and Bednarcyk, 2010) panel design, whose main purpose was to serve as a 

pathfinder for the honeycomb sandwich panel manufacturing and testing. This was referred to as the Tool 

Try Panel (TTP) for the first manufacturing demonstration of the new 1/6th tool. 
NASA/TM—2013-216574 2

Figure 1.—CoEx payload fairing configuration, PF2-COEX-28.

Figure 2.—Fiber placement at Spirit Aerosystems.

This paper documents the design of the 1/6th arc segment panel, which represents a segment of the 

barrel (or cylindrical) section of the payload fairing as shown in Figure 1. Although the SLS Payload 

Fairing Block 1A baseline design has an 8.4-m (27.6-ft) diameter, a 10-m (33-ft) diameter payload fairing 

design, the SLS Block 2 configuration was chosen to maintain continuity with, and to build upon, 

previous work performed in the LSSM and ACT projects. The goal was to design and build an arcsegment buckling test panel that mimicked the structural behavior of a full cylindrical section in 

compression. This buckling test panel would be cut from a 27- by 18-ft (1/6th) composite sandwich panel 

based on a 33-ft diameter fairing built by Spirit Aerospace in Wichita, Kansas.

The 1/6th honeycomb sandwich panel was constructed of a 1-in. thick aluminum honeycomb core 

with 8-ply graphite/epoxy facesheets. As shown in Figure 2, Spirit Aerospace used automated fiber 

placement technology to produce the composite panel on NASA's panel tool and demonstrate the viability 

of this manufacturing process for large scale fairing structures. Figure 3 shows the overhead gantry and 

NASA's arc segment panel tool built by Janicki Industries on the floor.
NASA/TM—2013-216574 3

Figure 3.—Gantry and 1/6th panel tool at Spirit Aerosystems.

Figure 4.—Fiber reinforced foam (FRF) sandwich panel concept, (a) schematic, and (b) photograph.

The 1/6th FRF sandwich panel was constructed with a nominal total thickness of 1-in., with non-crimp 

fabric graphite/epoxy facesheets equivalent to 8 plies. As discussed in Bednarcyk et al., (2010, 2011) and 

Zalewski and Bednarcyk (2010) , the FRF sandwich core consists of long rectangular cross-section sticks 

of foam that are wound with dry fibers (see Figure 4). These wound sticks are then assembled side-byside, and a light polymeric scrim fabric is applied on either side to hold the sticks together and produce a 

continuous dry core. The core is then stacked with the hand-layed non-crimp fabric facesheets on a tool, 

and the entire assembly is vacuum bagged, infiltrated with epoxy resin, and oven cured. This out-ofautoclave process requires a suitable, infiltratable, epoxy resin. As shown in Figure 4, the final panel 

configuration includes composite webs within the core, which are formed by the infiltrated winding 

layers. These webs provide FRF with some additional axial extensional and bending stiffness and strength 

compared to foam core sandwich panels, plus a great deal of additional axial through-thickness shear 

stiffness and strength.

The remainder of this report describes the two composite panel types considered in the buckling 

analyses and then provides a detailed presentation of the buckling analyses, performed using the 

commercial finite element codes Abaqus and ANSYS. First, linear eigenvalue-based analyses were 

performed considering the full 1/6th arc segment panel dimensions, along with reduced-size panels, to 

determine an expedient test panel size in terms of the expected buckling mode and ease of testing. 

Nonlinear buckling analyses were also performed for the honeycomb sandwich configuration using both 

shell- and solid-based finite element models. Finally, nonlinear buckling analyses of an as-manufactured 

FRF panel were conducted, wherein measured imperfections were included in the analyses.

(a) (b)
NASA/TM—2013-216574 4

Figure 5.—CoEx payload fairing design dimensions (in inches).

2.0 Panel Layout and Material Properties

As stated earlier, the 1/6th sandwich panel designs were based on the cylindrical barrel section of the 

CoEx Payload Fairing configuration, with a 33-ft diameter. Referring to Figure 5, the actual dimensions 

of a 1/6 arc-segment of the barrel are 26.25- by 17.28-ft. The CoEx project chose 27- by 17.25-ft as the 

baseline panel size for the tool design, and this panel size became the initial design for analysis. This 

initial full 1/6th arc segment sandwich panel (27- by 17.25-ft) was investigated (along with other smaller 

panel sizes) using two finite element analysis (FEA) software programs, ANSYS and Abaqus.

2.1 Honeycomb Sandwich Panels

An example ANSYS 2-D layered structural shell element model of the honeycomb sandwich 

configuration is shown in Figure 6. An example Abaqus shell model for the honeycomb sandwich panel is 

shown in Figure 7. In both cases, the facesheet plies, along with the core, are modeled as layers in the 

layered shell elements.

As shown in Figure 6 and Figure 7, the stacking sequence of the 8-ply carbon fiber/epoxy facesheets 

was [45°/90°/–45°/0°]S with 0.0053 in. thick plies. The facesheet material was IM7/977-3 tape. The 

IM7/977-3 elastic properties allowables were obtained from the Orion Materials Database (Lockheed

Martin, 2010), and they are not listed due to ITAR restrictions. Referring to Table 1, the aluminum (5052-

T6 alloy) honeycomb properties were obtained from commercially available literature (Hexcel, 2013).

Note that the honeycomb in-plane longitudinal and transverse moduli (E1 and E2) were kept at 21.28 psi, 

and the in-plane shear modulus (G12) was kept at 5.32 psi. It is also common for these in-plane properties 

to be assigned an arbitrary extremely low value (c.f. Myers et al., 2013).
NASA/TM—2013-216574 5

Figure 6.—ANSYS FEM of the 1/6th honeycomb sandwich panel.

Figure 7.—Abaqus FEM of the 1/6th honeycomb sandwich panel.

TABLE 1.—ALUMINUM HONEYCOMB MATERIAL PROPERTIES,

3.1 pcf, 1/8 in.-5052-0.0007 in. THICKNESS

Property/units Value Property/units Value

E1, psi ...........................................21.28

E2, psi ...........................................21.28

E3, ksi ................................................75

ν12 .................................................0.333

ν23 ............................................ 1.0×10–5

ν13 ............................................ 1.0×10–5

G12, psi ...........................................5.32

G13, ksi ..............................................45

G23, ksi ..............................................22

γ, lb/in3 .....................................0.00179

ρ, pcf................................................ 3.1

Ft1, psi............................................. 215

Fc1, psi ............................................ 215

Ft2, psi............................................. 215

Fc2, psi ............................................ 215

Ft3, psi............................................. 130

Fc3, psi ............................................ 300

Fs12, psi ............................................. 90

Fs23, psi ............................................. 90

Fs13, psi ............................................. 90
NASA/TM—2013-216574 6

2.2 FRF Sandwich Panels

The FRF sandwich panel cross-section design considered is shown in Figure 8. This panel design was 

taken from the barrel section of a full FRF fairing design that considered both strength and stability 

failure modes. The facesheets each were designed to consist of two layers of T700 VectorPly C-QX 1800 

non-crimp fabric (VectorPly, 2013), which has an internal ply stacking of [–45°/90°/45°/0°] and a per ply 

thickness of 0.006 in. As shown in Figure 9, each sheet of the non-crimp fabric consists of unidirectional

plies that are not woven together, but rather stitched with polyester fiber. Two sheets of this fabric, layed 

up in reverse order, form each facesheet, with an effective stacking sequence of [–45°/90°/45°/0°]s. The 

FRF panel was designed using E862 epoxy resin. Because ply properties for this facesheet material were 

unavailable, they were predicted from the constituent properties (fiber/matrix) using NASA's MAC/GMC 

micromechanics software (Bednarcyk and Arnold, 2002), with an estimated fiber volume fraction of 60

percent. While the stiffness properties predicted by micromechanics are known to be very reliable, 

prediction of composite allowables using micromechanics is known to be problematic (Aboudi et al., 

2013). Therefore, ply level B-Basis stress allowables for the T700/E862 facesheets were estimated based 

on the B-Basis strain allowables of the IM7/977-3 material used in the honeycomb sandwich panel 

design, with an assumed linear relation between the estimated stress allowables and assumed strain 

allowables. The ply level material properties for the 60 percent volume fraction T700/E862 facesheet 

material are given in Table 2.

Figure 8.—FRF sandwich panel cross-section design.

Figure 9.—Schematic of VectorPly C-QX 1800

non-crimp fabric (VectorPly, 2013).
NASA/TM—2013-216574 7

The winding layers (see Figure 8) cannot be represented with these same material properties, even 

though they have the same T700/E862 constituent materials, because the fiber volume fraction that 

results from the winding and infiltration process is much lower. Based on an estimated fiber volume 

fraction of 32 percent, the winding layer ply properties were approximated using the identical process 

described above for the facesheet plies. These properties are given in Table 3. Note that the thickness 

of each winding ply was estimated to be 0.01 in., and the panel design called for two winding plies with 

[+45°/–45°] layup. Thus, each web, which consists of the windings from two adjacent foam sticks, has a 

symmetric [+45°/–45°]s layup. The foam material included in the design was Rohacell 31HF 

polymethacrylimide foam (Evonik Industries, 2013), with material properties given in Table 4.

TABLE 2.—ESTIMATED 60 PERCENT T700/E862

PLY PROPERTIES USED IN THE FRF 

PANEL FACESHEET DESIGN

Property/units Value

E1, Msi......................................... 20.17

E2, Msi......................................... 1.147

ν12................................................. 0.257

G12, Msi ....................................... 0.465

G13, Msi ....................................... 0.465

G23, Msi ....................................... 0.330

γ, lb/in3

......................................... 0.057

Thickness, in................................ 0.006

Ft1, ksi............................................. 251

Fc1, ksi ............................................ 142

Ft2, ksi............................................ 12.5

Fc2, ksi ........................................... 10.2

Fs12, psi .......................................... 5.89

TABLE 3.—ESTIMATED 32 PERCENT T700/E862

PLY PROPERTIESUSED IN THE FRF

PANEL WINDING LAYER DESIGN

Property/units Value

E1, Msi......................................... 10.94

E2, Msi......................................... 0.705

ν12................................................. 0.300

G12, Msi ....................................... 0.237

G13, Msi ....................................... 0.237

G23, Msi ....................................... 0.207

γ, lb/in3

....................................... 0.0496

Thickness, in.................................. 0.01

Ft1, ksi............................................. 136

Fc1, ksi ........................................... 77.0

Ft2, ksi............................................ 7.73

Fc2, ksi ........................................... 6.27

Fs12, psi .......................................... 3.00

TABLE 4.—ROHACELL 31HF 

POLYMETHACRYLIMIDE 

FOAM MATERIAL PROPERTIES

Property/units Value

E, ksi.............................................. 5.22

G, ksi............................................ 1.885

Ft, psi .............................................. 145

Fc, psi................................................ 58

Fs, psi................................................ 58

γ, lb/in3

....................................0.001157
NASA/TM—2013-216574 8

The complexity of the FRF sandwich panel cross-section (see Figure 4) renders it more difficult to 

represent using simple shell properties. As such, the commercial HyperSizer structural sizing software 

(Collier Research Corp., 2013) was used to generate effective extensional, coupling, and bending stiffness 

matrices with the software’s “Reinforced Core Sandwich (RCS)” uniaxially stiffened panel concept.

These matrices, referred to a reference plane at the midpoint of the top facesheet, are,

10 lb/in

0 0 385

368 931 0

943 368 0

3 ×





















A =

 

10 in lb/in

0 0 183

175 443 0

449 175 0

3 × −





















B = 10 in lb/in

0.5 0.5 171

163 418 0.5

421 163 0.5

3 2 × −





















D =

The FRF panel buckling response was evaluated using only the Abaqus finite element software.

3.0 Linear Eigenvalue Buckling Analysis

The finite element model of the panel is shown in Figure 10, along with the loading and axial 

boundary conditions. The lateral panel edges were left free. These identical conditions were employed, 

along with very similar mesh sized, for the ANSYS and Abaqus analyses of the honeycomb sandwich 

panel and for the Abaqus analysis of the FRF sandwich panel. Linear eigenvalue analysis was performed 

to predict the first buckling eigenvalue and mode shape for these models of the full size 27- by 17.25-ft 

panels. The results are shown in Figure 11. Clearly, the predicted first buckling mode is the same for all 

three cases, involving two outward radial waves at the panel edges. The ANSYS and Abaqus models of 

the honeycomb sandwich panel predicted eigenvalues that were very close (1.968 versus 1.972, 

respectively), while the Abaqus FRF sandwich panel has predicted a slightly lower eigenvalue of 1.935.

This slightly lower eigenvalue can be attributed to a number of differences between the FRF and 

honeycomb sandwich panels, despite their identical shell dimensions. First, the FRF composite facesheet 

material has a lower stiffness than does the honeycomb facesheet material. Further, the FRF panel total 

thickness (1 in.) is lower than the honeycomb panel (1 in. core plus 16 facesheet plies that are each 

0.0053 in. thick = 1.0848 in.). However, the FRF panel facesheets are thicker (0.048 in. each versus

0.0424 in.), and the FRF panel has the additional stiffness contribution of the composite winding plies.

The combination of these effects resulted in the honeycomb panel having slightly higher predicted 

buckling resistance.

To simplify, lower cost, and provide for additional potential testing locations of the 1/6th panel, it was 

desirable to determine a reduced size panel that could be cut from the full size panel. This would also 

provide additional test specimens from the single panel. As such, the analyzed panel size was reduced, 

while keeping the 27 to 17.25 = 1.565 panel aspect ratio constant. It was found that, not only the predicted 

minimum eigenvalue of the panel changed, but also the predicted first mode changed. The Abaqus results 

for this panel size study for the honeycomb sandwich panel are shown in Figure 12. In this figure, the 

eigenvalues associated with the first four buckling modes of the full size honeycomb panel are plotted 

versus the panel height. For the full size panel (height = 27 ft), the first four eigenvalues are fairly close.

The first mode (A), as discussed above, has an eigenvalue of 1.971 and the mode involves an axial halfwave in the same radial direction on the panel sides (edges). The second mode (B) (eigenvalue = 2.059) 

involves a full-wave axially in the same radial direction on the panel sides. The third mode (C) 

(eigenvalue = 2.140) involves a half-wave axially, but with alternating radial directions on the panel 

sides, and finally, the fourth mode (D) (eigenvalue = 2.311) involves a full-wave axially with alternating 

radial directions on the panel sides. As the panel size is reduced (while maintaining the aspect ratio), these 

four predicted buckling modes remain as the first four, but their order and spread changes. At a panel 

height of 12- to 13-ft, mode C has the lowest eigenvalue, plus a large (approximately 50 percent) 

difference from the next lowest eigenvalue. A large spread between eigenvalues is desirable in terms of 
NASA/TM—2013-216574 9

test versus prediction correlation because potential imperfections in the panel construction and the test set 

up are then less likely to cause a different actual buckling mode. Furthermore, the mode shapes with 

alternating radial direction on the sides are more representative of a classical periodic full cylinder 

buckling mode, where waves with alternating radial direction progress circumferentially around the 

cylinder as shown in Figure 13 (Sleight et al., 2011). For these regions, the panels with heights from 

approximately 10- to 14-ft were identified as a “sweet spot” for a potential test article that could be cut 

from the full size 1/6th honeycomb sandwich panel. This identical panel size study was repeated for the 

honeycomb sandwich panel using ANSYS, with nearly identical results, as shown in Figure 14. The panel 

size study for the FRF sandwich panel was conducted using Abaqus, with the results shown in Figure 15.

The results are very similar to those of the honeycomb sandwich panel, with a clear candidate panel 

height of approximately 12 to 13 ft emerging as the most desirable.

Figure 10.—Applied static load and boundary 

constraints on the panel finite element 

models.
NASA/TM—2013-216574 10

Figure 11.—First buckling mode shapes for the 27- by 17.25-ft panel. (a) Honeycomb sandwich 

ANSYS prediction - eigenvalue = 1.968, (b) Honeycomb sandwich Abaqus prediction -

eigenvalue = 1.971, and (c) FRF sandwich Abaqus prediction - eigenvalue = 1.935.

(a) (b)

(c)
NASA/TM—2013-216574 11

Figure 12.—First four Abaqus-predicted eigenvalues and mode shapes as a function of honeycomb 

sandwich panel size while maintaining the full panel aspect ratio.

Figure 13.—Buckling mode in a cylindrical, composite honeycomb 

sandwich shell in axial compression.
NASA/TM—2013-216574 12

Figure 14.—First four ANSYS-predicted eigenvalues and mode shapes as a function of honeycomb sandwich panel 

size while maintaining the full panel aspect ratio.

Figure 15.—First four Abaqus-predicted eigenvalues and mode shapes as a function of FRF sandwich 

panel size while maintaining the full panel aspect ratio.
NASA/TM—2013-216574 13

4.0 Honeycomb Sandwich Panel Nonlinear Buckling Analysis

4.1 Shell Element Based Buckling Analyses

Shell element based, geometrically nonlinear progressive collapse (buckling) analyses were 

performed for the honeycomb sandwich 1/6th arc-segment panel using ANSYS and Abaqus. In all cases, 

the first 10 linear eigenvalue buckling modes were used to seed the model with a small degree of 

geometric imperfection. Then, the actual loading was applied to the geometrically imperfect model in an 

incremental fashion, with the effects of large displacements activated. The buckling load is then evident 

as a nonlinearity in the resulting compressive load versus displacement plot for the panel.

It should be noted that, in the case of ANSYS, when performing nonlinear analysis of a structure like 

the 1/6th panel shown in Figure 10, some sort of imperfection or asymmetry is needed to trigger buckling.

That is, if the geometry of the panel is perfect, the applied loading is uniform compression, and the shell 

properties are such that there is no extensional-bending coupling, then the panel will just compress 

axially, with no out of plane deformation, even in the presence of nonlinear geometric effects. This is not 

true in the case of Abaqus (unless the panel is flat), although nonlinear buckling simulations on the 

perfect geometry can lead to numerical issues. Thus, even for Abaqus, seeding of the model with 

imperfections is highly desirable. 

An alternative method for introducing asymmetry to the model is to apply small perturbations to the 

applied loads or enforced displacements. This method is not ideal, however, because it is not clear how to 

realistically introduce such perturbations. For example, varying the load across the top of the panel too 

drastically could change the problem completely. Another alternative method to introducing geometric 

imperfections is to slightly modify the coordinates of the nodes with random amplitudes. A disadvantage 

of using such random imperfections is that they cannot be repeated, and the results would differ for each 

realization of the same panel. For these reasons, the more straightforward and standard method for

imposing geometric imperfections on the finite element model, through use of the linear buckling mode 

shapes, was employed herein.

Towards this end, the displacements associated with the linear buckling mode shapes is multiplied by 

a scale factor and added to the initial panel geometry. The scale factor should be on the order of the 

manufacturing tolerances and thus the initial bowed shape. A factor of 0.020 in. was chosen for the 1/6th 

panel based on previous work on 3- by 5-ft panels of similar construction (Myers et al., 2013). Either a 

positive or negative scale factor was used to correctly represent the direction of the initial bow shape of 

the panels in that study. Without preliminary photogrammetry data of the 1/6th panel, initial 

imperfections were unknown. Thus, a positive scale factor was used in the Abaqus assessment. Figure 16

shows the first 10 linear buckling mode shapes for the honeycomb sandwich 1/6th arc-segment panel.

Figure 17 shows the resulting initial geometry with an exaggerated scale factor for the honeycomb 

sandwich panel.

Six honeycomb sandwich panel sizes (Figure 18) were analyzed using the nonlinear buckling 

procedure in both ANSYS and Abaqus. Note that Panel 1 is the full sized 1/6th panel, whereas, Panel 6 is 

one quarter of the full sized panel. Unlike the linear eigenvalue analysis above, the aspect ratio of all of 

these panels has not been maintained. The results of the study are shown in Figure 19. In general, the 

ANSYS and Abaqus result match well. Some differences are evident in the knee region of the curves 

associated with the initial onset of nonlinear buckling. An example is the 20.25 by 12.95 panel result, 

where ANSYS appears to have overshot the initial buckling nonlinearity, but after this knee, the results 

once again match Abaqus. The nonlinear prediction in this knee region appears to be somewhat dependent 

on the load increment size, which, in these cases, is variable and automatically chosen by ANSYS and 

Abaqus. Obviously, these codes have different algorithms for determining the increment sizes, which can 

affect the details of the predicted load—displacement curves.
NASA/TM—2013-216574 14

Figure 16.—First 10 linear buckling mode shapes of the honeycomb sandwich 1/6 panel (FRF panel was virtually 

identical).

Figure 17.—Initial deformed panel shape with exaggerated imperfections.
NASA/TM—2013-216574 15

Figure 18.—Sizes of six honeycomb sandwich panels analyzed for nonlinear buckling in ANSYS and Abaqus.

4.2 Solid Element Based Buckling Analyses

The 12- by 7.67-ft arc segment sandwich panel was investigated using two different ANSYS finite 

element models. The first model was a 2-D layered structural shell element from which the results have 

been presented earlier. The second model was a 3-D layered structural solid-shell element model in which 

the test fixturing was also included. In ANSYS, the 2-D layered shell element is called a SHELL281, and 

the 3-D solid-shell is called a SOLSH190. These finite element types are shown in Figure 20. The 

SOLSH190 element can be used to model shell structures with a wide range of thickness from thin to 

moderately thick. Furthermore, the SOLSH190 element formulation permits small thickness to length 

ratios without producing errors due to large aspect ratios, as can occur in standard solid elements. The 

thickness between the nodes must equal the sum of the individual layer thicknesses. Otherwise, the 

software will scale the individual layer thicknesses to fit the actual element thickness. Since the results of 

the 2-D shell model were included with the Abaqus and ANSYS shell FEM results in previous sections, 

these results will only be used as a basis of comparison with the 3-D finite element model.

Figure 21 shows the solid-shell model of the test panel geometry with the end plates. For clarity, the 

finite element edges are suppressed. The panel was modeled as 149-in. tall including the 2.5-in. aluminum 

end plates and 92-in. wide along the chord. The top and bottom 2.5 in. portions of the panel were 

supported in the potting material and end plates. Figure 22 shows the slot in the end plate shaped like the 

arc segment test specimen model. The end of the specimen model is centered in the slot, and the 0.5 in.

space around the specimen is filled with solid elements modeling the isotropic UNISORB V-100 potting 

compound with the elastic properties Young’s Modulus (E) of 436 ksi and Poison’s Ratio (ν) of 0.35.

Each color represents a different material assignment: cyan for the IM7/977-3 facesheets, purple for the 

honeycomb core, red for the potting compound, and orange for the aluminum end plates.
NASA/TM—2013-216574 16

Figure 19.—Nonlinear buckling honeycomb sandwich panel size study. Force versus compression comparisons 

between Abaqus and ANSYS.

Figure 20.—ANSYS 2-D shell versus 3-D solid-shell (brick) finite elements.
NASA/TM—2013-216574 17

Figure 21.—The 12- by 7.67-ft arc segment test panel configuration.

Figure 22.—The 12- by 7.67-ft arc segment test panel and test fixture end view.
NASA/TM—2013-216574 18

Figure 23.—Facesheet 8-ply stacking sequence.

The solid model was generated using a script file containing ANSYS preprocessing, finite element 

solution, and post-processing commands in the ANSYS Parametric Design Language (APDL) (ANSYS, 

2011). The basic geometry, ply stacking sequence, material properties, and geometric imperfections were 

defined using parameters. The script file can generate the finite element model, perform each required 

analysis in sequence, and process the results in tabular or graphic format.

As shown in Figure 23, as in the previous shell finite element models, the stacking sequence of the 

facesheets was [45°/90°/–45°/0°]S with 0.0053 in. thick plies. Again, the IM7/977-3 elastic properties and 

B-basis allowables were obtained from Orion Materials Database (Lockheed Martin, 2010), and they are 

not listed due to ITAR restrictions. Referring to Table 1, the aluminum (5052-T6 alloy) honeycomb 

properties were obtained from commercially available literature. Unlike the HyperSizer properties, the 

honeycomb in-plane longitudinal and transverse moduli (E1 and E2) were kept at the actual value of

21.28 psi, and the in-plane shear modulus (G12) was kept at 5.32 psi in the ANSYS solid models. The 

honeycomb normal (out-of-plane) modulus (E3) was also kept at 75 ksi.

At the time of writing this report, no tests on the 12- by 7.67-ft honeycomb sandwich panel had been 

conducted. The actual test panel would be secured between two loading platens, with the bottom end plate 

fixed on the loading platen. The top end plate would move with the top platen in the axial direction to 

place the panel in uniaxial compression. Referring to Figure 24, all three displacements in the solid-shell 

finite element model were fixed along the bottom edge of the panel and end plate. The same boundary 

condition was applied to the top panel edge and top end plate, except a uniform, monotonic displacement 

was applied in the negative z-axis direction.

It has been pointed out earlier in this report that the initial geometric imperfections have an impact on 

the buckling load and direction, especially for an axially compressed, arc segment composite shell 

structure. The 12- by 7.67-ft curved panel is expected to buckle like a classical full cylinder, with halfwave nodes along the edges. The expected shape is similar to the first linear buckling mode shape shown 

in Figure 25 and Figure 26. Therefore, the initial imperfect geometry is important to capture in a 

nonlinear buckling analysis. If the 12- by 7.67-ft panel finite element model is perfectly symmetric (i.e., 

in geometry and loading), nonsymmetrical buckling does not occur numerically in ANSYS. Rather, 
NASA/TM—2013-216574 19

nonlinear buckling analysis fails because nonsymmetrical buckling responses cannot be triggered in such 

a geometrically perfect finite element model.

As in the nonlinear buckling analysis of the shell-based model, geometric imperfections are obtained 

by running a preliminary linear buckling analysis, then updating the geometry of the finite element model 

to the deformed configuration. This technique was implemented by adding the displacements of the first 

ten mode shapes reduced by a scaling factor. As in the case of the shell-based simulations, a scale factor 

of 0.020 in. was used for the solid FEM of the 12- by 7.67-ft panel. The shell FEM produced similar 

linear buckling mode shapes. However, the first mode shape of the shell FEM (Figure 25) was a mirror 

reflection of the solid FEM's first mode shape (Figure 26). The eighth and ninth modes are also mirror 

reflections, but the first mode is the desired dominant mode shape for the panel. In order to compare strain 

results, a negative scale factor (-0.020 in.) was applied to all the modes of the shell FEM. This ensured 

consistent out-of-plane displacement directions between the two models. Figure 27 shows the solid FEM 

perfect and exaggerated imperfect geometries.

After the geometric imperfections were added to the finite element model, the nonlinear buckling 

analysis was performed. In ANSYS, a nonlinear buckling analysis is a static analysis with large 

deflections active. The magnitude of the applied axial compression is extended beyond the first linear 

(Eigen) buckling mode. In this analysis the compression was increased gradually using 50 small time 

increments to predict the critical buckling load. Figure 28 shows the reaction load versus end compression 

for the linear and nonlinear shell and solid ANSYS models of the 12- by 7.67-ft panel. The two linear 

buckling analyses appear to over-predict slightly the buckling loads predicted by the two nonlinear 

analyses. In both nonlinear ANSYS analyses, the panel initial stiffness was nearly the same.

Figure 29 shows the panel radial displacement plot (radial component of the total displacement field) 

when -0.221 in. was applied to the top edge, which is after the buckling knee apparent in Figure 28. This 

figure shows that the geometric imperfections applied to the model are dominated by the first linear 

buckling mode shape (see Figure 26). As expected, the 12- by 7.67-ft arc-segment honeycomb sandwich 

panel buckles with circumferential waves rather than in a column-like mode.

Figure 24.—ANSYS Solid Model boundary conditions.
NASA/TM—2013-216574 20

Figure 25.—First 10 linear buckling mode shapes for the ANSYS Shell FEM

Figure 26.—First 10 linear buckling mode shapes for the ANSYS Solid FEM.
NASA/TM—2013-216574 21

Figure 27.—Panel model with the applied geometric imperfections (exaggerated).

Figure 28.—Total reaction load versus end shortening for ANSYS Solid FEM.
NASA/TM—2013-216574 22

Figure 29.—ANSYS Solid FEM, post-buckling radial displacement contour plot at 0.221-in. end 

shortening (90,094 lb).

Potential strain gage locations for the 12- by 7.67-ft arc-segment honeycomb sandwich panel are 

shown in Figure 30. These gage locations were used to extract local strain data from the ANSYS solid 

and shell FEMs.

Figure 31 shows load versus minimum principal strain plots for both the solid FEM and shell FEM at 

gages 15 and 16 near the left edge at the vertical centerline. The plots confirm that applying a negative 

scale factor to the imperfections (linear buckling modes) of the shell FEM produced similar results to the 

solid FEM.

An additional analysis was performed on the solid FEM of 12- by 7.67-ft panel to refine the ANSYS 

results near the predicted buckling load. In this analysis the compression was increased gradually using 50 

small time increments up to 0.325 in., which was just beyond the compression of 0.201 in. predicted by 

the linear buckling analysis. Note that this increment size is approximately one-half of that used 

previously (e.g., in Figure 31). A plot of the load versus compression in Figure 32 shows the additional 

refinement in the results. Figure 33 shows how the load versus compression results from the refined 

analysis compare with the original analysis results. The main purpose in the refinement was to capture 

more strain data near the predicted buckling load. 

ANSYS 14.0

MAR 5 2013

11:45:18

PLOT NO. 2

NODAL SOLUTION

STEP=1

SUB =34

TIME=.68

UX

MIDDLE

LAYR=8

RSYS=1

DMX =1.6743

SMN =-1.19402

SMX =1.6607

1

MN

MX

Global Radial Displacement Contours

Initial Panel Bow Amplitude = 0"

Geometric Imperfections Scale Factor = .02

X Y

Z

-1.19402

-.876832

-.55964

-.242449

.074743

.391935

.709126

1.02632

1.34351

1.6607

12x7.667 Panel, 8-Ply, IM7/977-3 Face Sheets, -0.221" Disp
NASA/TM—2013-216574 23

Figure 30.—Proposed strain gage locations for the 12- by 7.67-ft 

honeycomb sandwich composite panel, (odd (OML), even (IML)).
NASA/TM—2013-216574 24

Figure 31.—Load versus minimum principal strain plots for solid versus shell FEM at gages 15 and 16.

Figure 32.—Total reaction load versus end shortening for the ANSYS Solid FEM (refined analysis).
NASA/TM—2013-216574 25

To determine the panel buckling loads, a method from Singer et al. (1998), shown in Figure 34, was 

used. This method utilizes global load versus local axial strain gage data to determine the onset of 

buckling. In Figure 34, a vertical tangent line intersects the load-strain curve at a local strain where the 

local strain increment reverses. The load corresponding to that local strain is designated the local buckling 

load, with the lowest local buckling load indicating the panel buckling initiation load. It should be noted 

that the buckling strain, and hence buckling load, can only be determined at monitored locations and 

therefore a value lower than the lowest detected value could be present. Thus, the postulated buckling 

initiation load is somewhat subjective and based upon the location where the strains are being monitored 

for reversal

The strains in the curved panel are, in general, three dimensional. In a panel compression test, strains 

are typically measured on a free surface using strain gages or digital image correlation techniques. Max 

and min principal strains are frequently used to assess structural integrity in conjunction with the finite 

element method because they provide a convenient scalar quantity representing an extreme that can be 

plotted over the structure. Principal strains, however, are influenced by the entire strain field. Thus, where 

significant shear strains are present, such as in the panel corners and near the buckling wave inflection 

points, the principal strains can deviate from the global axial and tangential strains. Far from the panel 

ends, and away from inflection points, shear strains are minimal, and the principal strains should 

correspond to the axial and tangential strains. To examine this potential effect, the minimum principal 

strains, axial strains, and z-strains were plotted on each surface in Figure 35 and Figure 36 at a 0.2145-in. 

end compression of the panel. 

Figure 33.—Refined versus original coarse load step, buckling analysis.

.
NASA/TM—2013-216574 26

Figure 34.—Method for determining 

buckling load from Singer et al. (1998).

Figure 35.—Minimum principal and component strain comparisons on the OML surface at 0.2145-in. compression.

Figure 36.—Minimum principal and component strain comparisons on the IML surface at 0.2145-in. compression.
NASA/TM—2013-216574 27

Figure 37.—Element coordinate systems for the OML facesheet elements.

Both Figure 35 and Figure 36 show little variations in the strain contours. The ANSYS time history 

post-processing module warns the user of potential problems extracting data from the structural solidshell finite elements with layered composite capability. Averaging performed in the module assumes 

consistent coordinate systems for all the elements used in the averaging. This is not a condition required 

by the general post-processing module which was used to produce the contour plots. The model generated 

for this paper aligned each element coordinate system with the axial direction in order to define ply angles 

along the fiber directions. Furthermore, one axis of each element coordinate system is also aligned with 

the surface normal as shown in Figure 37 in which only the OML facesheet elements have been selected.

This led to incorrectly averaged nodal results for component strains (x, y, or z) because the strains in each 

layer were in different reference frames. However, principal strains are independent of the element 

coordinate systems, and the averaged nodal results for the principal strains should be representative, 

although different for each ply. Thus, the min. principal strains in the 12- by 7.67-ft panel were used 

instead of the axial strains in Singer's method for determining the onset of buckling.

Buckling loads as indicated by the various strain gage pairs using the method of Singer et al. (1998) 

are tabulated in Table 5. All strain gage pairs are listed, but some strain gage pairs did not exhibit a strain 

reversal. The first detected minimum principal strain reversal occurs at the location of gages 1 and 2 

located near the top left corner of the panel and the location of gages 29 and 30 located near the lower left

corner of the panel, at a load of 75,598 lb. Figure 38 to Figure 49 show plots of the load versus minimum 

principal strain for all strain gage pairs. Figure 38 and Figure 48 show that, in the left corners, the IML

panel strains went into compression up to the onset buckling, after which the strain increment reverses 

and the axial strain is alleviated as the load increases during post-buckling. However, Figure 40 and 

Figure 49 show that, in the right corners, the OML panel strains went into compression up to the onset 

buckling, after which the strain increment reverses and the strain is alleviated as the load increases.

Furthermore, Figure 42 and Figure 43 show that, near the left edge in the panel vertical center, the OML

panel strains went into compression up to the onset buckling, after which the strain increment reverses. 
NASA/TM—2013-216574 28

Near the right edge at the panel vertical center, Figure 45 and Figure 46 show that axial IML panel strains 

went into compression up to the onset buckling, after which the strain increment reverses. Figure 44

shows that there is no strain reversal at the panel center, and Figure 39 shows no strain reversal at the 

panel top center. The panel vertical center lies at an inflection point, as shown by the radial displacement 

contour plot in Figure 29. These panel strain results are consistent with the solid FEM buckling mode

shown in Figure 29. The strains at gages 13, 14 (Figure 41) and gages 27, 28 (Figure 47) show no strain 

reversal.

TABLE 5—PREDICTED BUCKLING LOADS 

FROM THE SOLID FEM ANALYSIS

Gage Analysis results, 

lb

Microstrain

1,2 75,598 1006

5,6 --------- ---------

9,10 85,617 1431

13,14 --------- ---------

15,16 81,008 1277

17,18 81,008 1149

19,20 --------- ---------

21,22 106,533 1525

23,24 78,333 1239

27,28 --------- ---------

29,30 75,598 858

37,38 85,617 1359

Figure 38.—Load versus minimum principal strain plots for gages 1 and 2.
NASA/TM—2013-216574 29

Figure 39.—Load versus minimum principal strain plots for gages 5 and 6.

Figure 40.—Load versus minimum principal strain plots for gages 9 and 10.
NASA/TM—2013-216574 30

Figure 41.—Load versus minimum principal strain plots for gages 13 and 14.

Figure 42.—Load versus minimum principal strain plots for gages 15 and 16.
NASA/TM—2013-216574 31

Figure 43.—Load versus minimum principal strain plots for gages 17 and 18.

Figure 44.—Load versus minimum principal strain plots for gages 19 and 20.
NASA/TM—2013-216574 32

Figure 45.—Load versus minimum principal strain plots for gages 21 and 22.

Figure 46.—Load versus minimum principal strain plots for gages 23 and 24.
NASA/TM—2013-216574 33

Figure 47.—Load versus minimum principal strain plots for gages 27 and 28.

Figure 48.—Load versus minimum principal strain plots for gages 29 and 30.
NASA/TM—2013-216574 34

Figure 49.—Load versus minimum principal strain plots for gages 37 and 38.

Figure 50 illustrates another method to predict quantitatively the onset of panel buckling. This method 

uses a sudden change in panel stiffness to predict the buckling load. The panel compressive secant 

stiffness, P/U, is determined from the reaction load versus end-shortening data, as plotted in Figure 32.

The tangent stiffness, dP/dU, is derived from the load-displacement curve. Initially, while the panel 

deforms linearly, the secant and tangent stiffnesses coincide. The vertical line is drawn at the point where 

the tangent stiffness, dP/dU, suddenly decreases and deviates from the secant stiffness curve. For the 12-

by 7.67-ft honeycomb composite panel, this sudden change in the stiffness occurs at an axial compression 

(end-shortening) of 0.182 in. From the reaction load versus end-shortening data, this axial compression 

corresponds to a load of 78,332 lb. Thus, the predicted panel buckling load of 78,332 lb is a global panel 

prediction as opposed to local strain reversal prediction method of Singer et al. (1998). This is also in 

agreement with the middle left side strain gage (Figure 45, Gages 23 and 24) which shows a buckling 

onset load of 78,333 lb. Again, the postulated buckling load is somewhat subjective and based upon the 

judgment of the location where a sudden decrease in the stiffness occurs. In Figure 50, the vertical dotted

line is drawn where the decrease in stiffness changes from 1.85 to 2.72 percent. Beyond the vertical 

dashed line, the decrease in stiffness jumps successively to 5, 11, 25, and 40 percent.

In the ANSYS solid element model, the radial (out-of-plane) displacements do not vary noticeably 

between the OML and IML surfaces as seen in Figure 51 and Figure 52. Examining these model 

displacement fields, Figure 51 and Figure 52 show the evolution of the out-of-plane or radial 

displacement, obtained from the numerical analysis, as the applied compressive load is increased. Both 

figures show that the out-of-plane (radial) displacement was negligible up to the load of 75,598 lb, which 

is the buckling load predicted by the axial strains at gage locations 1 and 2 and locations 29 and 30, as 

listed in Table 5. However, Figure 53 and Figure 54 show that the axial displacements vary significantly 

between the OML and IML surfaces, especially in the regions near the top and bottom test fixtures.

Figure 53 and Figure 54 show the evolution of the in-plane or longitudinal displacement, obtained from 

the numerical analysis, as the applied compressive load is increased. The contour plot variations between 
NASA/TM—2013-216574 35

the OML axial displacements in Figure 53 and the IML axial displacements in Figure 54 explain why the 

slopes of the load versus strain curves for the OML and IML surfaces are so different even before the 

panel buckles at the top (Figure 38, Figure 39, and Figure 40) and the bottom of the panel (Figure 47, 

Figure 48, and Figure 49). Yet, the load versus strain curve along the midplane (Figure 42 to Figure 46) 

for the OML and IML are the same until the buckling load, because the midplane axial displacements 

(Figure 53 and Figure 54) remain the same until the point of buckling at 75,598 lb.

For the OML surface, Figure 55 shows that the minimum principal strain (compressive) was minimal

up until the point of buckling at 75,598 lb as seen in the out-of-plane (radial) displacements in Figure 51. 

Similarly, Figure 56 shows that the minimum principal strain (compressive) on the IML surface was 

minimal up until the same point of buckling which is also seen in the out-of-plane (radial) displacements 

in Figure 52. Both strain contour plots display asymmetry which was due to the scaled asymmetric 

geometric imperfections superimposed from the linear buckling modes.

Figure 50.—Stiffness versus axial compression for the ANSYS Solid FEM of the 12- by 7.67-ft

honeycomb sandwich panel
NASA/TM—2013-216574 36

Figure 51.—ANSYS Solid FEM, OML radial (x-axis or out-of-plane) displacement contour plots versus load.

Figure 52.—ANSYS Solid FEM, IML radial (x-axis, or out-of-plane) displacement contour plots versus load.
NASA/TM—2013-216574 37

Figure 53.—ANSYS Solid FEM, OML axial (z-axis or longitudinal) displacement contour plots versus load.

Figure 54.—ANSYS Solid FEM, IML axial (z-axis or longitudinal) displacement contour plots versus load.
NASA/TM—2013-216574 38

Figure 55.—ANSYS Solid FEM, OML minimum principal strain (z-axis or longitudinal).

Figure 56.—ANSYS Solid FEM, IML minimum principal strain (z-axis or longitudinal).
NASA/TM—2013-216574 39

5.0 FRF Sandwich Panel Nonlinear Buckling Analysis

A 1/6th FRF sandwich panel was constructed according to the design shown in Figure 8. Based on the 

linear eigenvalue buckling analyses presented in Section 3.0, a buckling test article size of 12- by 7.67-ft 

was selected. This panel size maintains the full 1/6th panel aspect ratio and is also within the “sweet spot”

identified in Figure 15. Measurements of the actual 1/6th FRF sandwich panel revealed some variance of 

the panel cross-section dimensions from the design shown in Figure 8. In particular, the average height of 

the panel was 0.94 in. (versus 1-in. in the design), the internal webs were 0.042 in. thick and spaced at 

2.54-in. (versus 0.04- and 2.5-in. in the design), and the facesheet plus the winding layer thickness was

0.056-in. (versus 0.068 in. in the design). New properties for the FRF panel were generated based on 

these actual dimensions, with exception of the facesheet plus winding layer thickness. This thickness 

variation is due to the presence of a different amount of matrix material than expected during the 

infiltration, bagging, and curing process, whereas the amount of fiber material is known and fixed. It 

would thus be incorrect to increase the thickness of the facesheet without also decreasing its fiber volume 

fraction, which would necessitate changing all of the facesheet mechanical properties. Instead, the 

original design facesheet thickness of 0.048 in. was maintained, as this will provide close to the proper 

contribution of the facesheet to the panel stiffness.

The updated panel effective stiffness matrices, again referred to a reference plane at the midpoint of 

the top facesheet, are,

10 lb/in

0 0 391

373 937 0

948 373 0

3 ×





















A = 10 in lb/in

0 0 174

166 418 0

423 166 0

3 × −





















B = 10 in lb/in

0.5 0.5 152

145 369 0.5

372 145 0.5

3 2 × −





















D =

While the extensional (A) and coupling (B) stiffness matrices have changed slightly, a significant change 

in the bending stiffness (D) has occurred, mainly because of the 6 percent reduction in the overall panel 

thickness compared to the design.

The Abaqus nonlinear buckling analysis results for the 12- by 7.67-ft FRF sandwich panel model are 

shown in Figure 57. Note that a fixed increment size was used for these analyses. Cases in which the first 

ten linear eigenvalue buckling modes were used to seed the model were analyzed with scale factors of 

both 0.015 in. and 0.02 in. In addition, a case with a scale factor of 0.02 in., but only seeded with the first 

linear eigenvalue buckling mode was analyzed. All three of these cases gave nearly identical loaddisplacement results, characterized by the onset of nonlinearity at a load of 92,400 lb. This compares well 

with the first linear buckling eigenvalue for the panel of 95,900 lb (3.8 percent difference). Also shown in 

Figure 57 is the buckling mode associated with the three nearly coincident seeded nonlinear cases and the 

first mode associated with the linear eigenvalue buckling analysis. These modes are the same, and they 

correspond to the circumferential buckling mode C from Figure 15.

Figure 57 also includes Abaqus nonlinear buckling results for the same FRF panel with no 

imperfection, using the same increment size that was sufficient for the cases seeded with imperfection.

Clearly, it appears that the case with no imperfection is giving a completely different buckling prediction.

The panel response remains linear up to a load of 133,700 lb (a difference of 45 percent from the seeded 

cases) and takes a different buckling mode, similar to the axial or column buckling mode A in Figure 15.

However, close examination of the Abaqus output for this case reveals that, once the load associated with 

the linear eigenvalue of the panel was exceeded, Abaqus was reporting negative eigenvalues. The 

increments associated with these negative eigenvalues are indicated in Figure 57. Negative eigenvalues, 

in a linear regime such as this, are an indication of numerical problems in the finite element code, and the 

associated results are thus unreliable. Note that, although not shown herein, decreasing the load increment 

size did not affect these results.
NASA/TM—2013-216574 40

Figure 57.—Abaqus nonlinear buckling analyses of a 12- by 7.67-ft FRF sandwich panel.

The actual 1/6th FRF panel was cut to produce a 12- by 7.67-ft buckling test article. This test article 

was scanned to measure its geometric imperfection. Measurements of the deviation of the panel from 

cylindrical, normal to the panel surface, were taken along the 12-ft length of the panel close to the left 

edge, close to the right edge, and along the center. The results are shown in Figure 58. The panel takes a 

slight saddle shape, with the most severe imperfection measured along the left edge with an approximated 

maximum magnitude of 0.02 in.

To assess the impact of the imperfection on the FRF panel predicted buckling response, the measured

left edge imperfection was considered. First, the measured curve, which ends at 133 in., was extended to 

the full panel length of 144 in. by scaling all axial point locations by 144/133 = 1.083. Then a three point 

central moving average was taken to filter the data, and a few outlier points were removed, resulting in 

the approximation of the imperfection shown in Figure 59. This approximation was then used as spline 

points in Abaqus to generate the geometry of the long edge of the panel and subsequently swept along the 

panel θ-direction to generate the full panel shell geometry. This procedure does not capture any variation 

of the imperfection in the θ-direction, but has been shown by Myers et al. (2013) to provide a good 

approximation of cylindrical panel imperfection for buckling analyses.

Nonlinear buckling analysis results for the 12- by 7.67-ft FRF sandwich panel, including the left 

imperfection (bow), are shown in Figure 60. The prediction that includes the geometric imperfection 

matches very closely with the nonlinear analysis with no imperfection, but is also similarly plagued by the 

negative eigenvalues, which indicate an unreliable result. The predicted column buckling mode also 

matches the no imperfection case.

0

20000

40000

60000

80000

100000

120000

140000

160000

0 0.1 0.2 0.3 0.4 0.5 0.6

Compressive Load (lb)

Applied Axial Compressive Displacement (in.)

Nonlinear Analysis, Seeded with 10 Modes @ 0.015 in.

Nonlinear Analysis, Seeded with 10 Modes @ 0.02 in.

Nonlinear Analysis, Seeded with 1st Mode Only @ 0.02 in.

Linear Eignevalue Buckling Load

Nonlinear Analysis, No Imperfection

Negative Eigenvalues

Linear Eigenvalue

Nonlinear - No Imperfection Nonlinear - Seeded with Modes
NASA/TM—2013-216574 41

Figure 58.—Measured geometric imperfections in the FRF 12- by 7.67-ft buckling test article.

Figure 59.—Measured and approximated left edge geometric imperfection in 

the FRF 12- by 7.67-ft buckling test article.

0

20

40

60

80

100

120

140

160

-0.025 -0.02 -0.015 -0.01 -0.005 0 0.005

Long Axis (in.)

Out-of-plane Axis (in.)

Scaled Data

Approximation
NASA/TM—2013-216574 42

To overcome the unreliable predictions, the Riks arc-length method, available within Abaqus, was 

used, in combination with a very small loading increment. It should be noted that, using the Riks method 

without limiting Abaqus to a very small loading increment resulted in prediction similar to those shown in 

Figure 60 with negative eigenvalues, which indicate a numerical problem in the solution. Forcing Abaqus 

to use a very small loading increment without using the Riks method resulted in divergence near the load 

associated with the first linear eigenvalue. In this case, Abaqus is unable to reach a converged solution at 

a given time increment, and the code execution stops. As shown in Figure 61, using the Riks method in 

combination with a very small loading increment resulted in both the case with no imperfection and the 

case with the left edge bow shape matching the case seeded with the first ten linear eigenvalue buckling 

modes very closely. The circumferential buckling modes also matched that of the seeded cases. All three 

of these cases predict the onset of buckling, evident as a knee in the load versus deflection curves, at a 

load level very close to the linear eigenvalue buckling load of 95,900 lb.

It should be noted that previous work by Myers et al. (2013) indicated that qualitatively similar 

imperfections with magnitudes on the order of that shown in Figure 58 did have a sizable impact on the 

buckling response of cylindrical arc-segment honeycomb sandwich panels. However, the 1/16th panels 

considered in that study were a different size than the FRF panel considered herein (5- by 3-ft versus 12-

by 7.67-ft ), and they buckled in a different mode. These differences are likely to account for the fact that, 

in present investigation, the imperfections were observed to have little effect on the Riks buckling 

analysis.

Figure 62 shows axial strain field and the radial displacement field predicted by the Riks method with 

the very small loading increment just after the buckling knee at an applied axial displacement of –0.19 in. 

Note that, because this shell-based model used homogenized effective stiffness terms to define the 

material (rather than ply definitions) the predicted strain fields are constant through the panel thickness. 

Note the similarity of Figure 62(a) (FRF panel) to Figure 55 (honeycomb sandwich panel) and the 

similarity of Figure 62(b) (FRF panel) to Figure 51 (honeycomb sandwich panel). The FRF and 

honeycomb sandwich plots are mirrored with respect to each other because, as can be seen from 

comparing Figure 62(b) to Figure 51, opposite edges were predicted to buckle outwards in the 

simulations.

Figure 60.—Abaqus nonlinear buckling analyses of a 12- by 7.67-ft FRF sandwich panel.

0

20,000

40,000

60,000

80,000

100,000

120,000

140,000

160,000

0 0.1 0.2 0.3 0.4 0.5

Applied Compressive Load (lb)

Applied Axial Compressive Displacement (in.)

Nonlinear Analysis, No Imperfection

Nonlinear Analysis, Seeded with 10 Modes @0.02 in.

Nonlinear Analysis, Left Bow

Nonlinear - No Imperfection Nonlinear - Left Bow

Negative Eigenvalues
NASA/TM—2013-216574 43

Figure 61.—Abaqus nonlinear buckling with Riks analyses of a 12- by 7.67-ft FRF sandwich panel.

Figure 62.—Fields predicted by the Abaqus nonlinear buckling analyses of a 12- by 7.67-ft FRF sandwich panel at an 

applied global displacement of –0.19 in., (a) axial strain field, and (b) radial displacement field.

0

20,000

40,000

60,000

80,000

100,000

120,000

140,000

160,000

0 0.1 0.2 0.3 0.4 0.5

Applied Compressive Load (lb)

Applied Axial Compressive Displacement (in.)

Nonlinear Analysis, No Imperfection

Nonlinear Analysis, Seeded with 10 Modes @0.02 in.

Nonlinear Analysis, Left Bow

Riks Nonlinear Analysis, No Imperfection, Small Increments

Riks Nonlinear Analysis, Left Bow, Small Increments

Linear Eigenvalue Buckling Load

Nonlinear - No Imperfection Nonlinear - Left Bow

Negative Eigenvalues

(a) (b)
NASA/TM—2013-216574 44

6.0 Conclusions

This report presents finite element based buckling analyses for 1/6th arc-segment panel of a full 33-ft 

diameter cylindrical barrel section of a payload fairing structure. Both linear and nonlinear analyses, shell 

and solid models, and honeycomb sandwich and fiber reinforced foam (FRF) sandwich configurations 

were considered using the ANSYS and Abaqus commercial finite element codes. The panel cross sections 

geometries considered were chosen to be representative of realistic Space Launch Systems (SLS) fairing 

barrel panels. Linear eigenvalue shell element based buckling analyses indicated that a smaller 12- by 

7.67-ft panel, four of which could be cut from the full 27- by 17.25-ft panel 1/6th arc-segment panel, could 

serve as a manageable compression test article, while also providing two key features in its expected 

buckling response. First, the initial predicted buckling mode shape was characterized by waves with

alternating radial directions on the panel sides. This is representative of a classical periodic full cylinder 

buckling mode, where waves with alternating radial direction progress circumferentially around the 

cylinder. Second, there was a large (approximately 50 percent) difference between the first and next 

lowest predicted buckling eigenvalue. A large spread between eigenvalues is desirable in terms of test 

versus prediction correlation because potential imperfections in the panel construction and the test set up 

are then less likely to cause a different actual buckling mode. Based on these results, the 12- by 7.67-ft

panel dimensions were recommended for large buckling test panels for both the honeycomb sandwich and 

FRF sandwich panel concepts. The shell-based linear buckling analyses performed using ANSYS and 

Abaqus were in very good agreement, while, qualitatively, the honeycomb and FRF sandwich 

configurations performed quite similarly. Of course, some quantitative difference was present due to the 

slightly lower effective stiffness and different dimensions of the two configurations.

Shell-based nonlinear buckling analyses of various honeycomb sandwich panel sizes showed that, 

once again, the buckling predictions of ANSYS and Abaqus were in very good agreement. In these cases, 

the first ten linear eigenvalues were used to seed the nonlinear analyses with geometric imperfection in 

both codes. Then, using ANSYS, a detailed nonlinear buckling analysis study was performed with a solid 

element model of the previously determined smaller 12- by 7.67-ft panel. First, this model was used to 

show good agreement with the ANSYS shell element based nonlinear model in terms of mode and 

buckling load. Then, detailed strain histories were examined at various locations over the panel area that 

represent potential strain gage locations for a test panel. By monitoring the strain histories on the front 

and back sides of the panel at a given location, it was shown that the point of a strain reversal could be 

used as an estimate of the onset of buckling.

Finally, shell element based nonlinear buckling analyses of the 12- by 7.67-ft FRF panel were 

performed using Abaqus. It was shown that, as in the case of the honeycomb sandwich configuration, 

when seeded with the first ten linear buckling modes, the FRF panel nonlinear predictions matched the 

mode and buckling load predicted by linear analysis closely. Further, the predicted load versus

displacement curve was qualitatively similar to that of the honeycomb sandwich. However, with no 

imperfection, Abaqus predicted a different axial (column-like) buckling mode with much higher buckling 

load. These results were characterized by negative eigenvalues while the panel remained in a linear 

regime, an indication of unreliable results. Similar unreliable results were obtained when including 

geometric imperfections, measured from an actual manufactured 12- by 7.67-ft FRF test panel, in the 

analysis. To overcome these issues in the nonlinear analysis, a Riks arc-length method analysis, wherein a 

very small loading increment size was enforced, was performed for the FRF panel with the realistic 

imperfections in Abaqus. These results matched the prediction for the FRF panel seeded with the first ten 

linear buckling modes and are expected to be a good prediction of the performance of the actual FRF test 

panel.

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