Evan J. Pineda, David E. Myers, Brett A. Bednarcyk, and Thomas M. Krivanek

Glenn Research Center, Cleveland, Ohio

A Numerical Study on the Effect of Facesheet-Core 

Disbonds on the Buckling Load of Curved

Honeycomb Sandwich Panels

NASA/TM—2015-218861

August 2015

Paper no. 1713
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Evan J. Pineda, David E. Myers, Brett A. Bednarcyk, and Thomas M. Krivanek

Glenn Research Center, Cleveland, Ohio

A Numerical Study on the Effect of Facesheet-Core 

Disbonds on the Buckling Load of Curved

Honeycomb Sandwich Panels

NASA/TM—2015-218861

August 2015

Paper no. 1713

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Glenn Research Center

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NASA/TM—2015-218861 1

A Numerical Study on the Effect of Facesheet-Core Disbonds on the 

Buckling Load of Curved Honeycomb Sandwich Panels

Evan J. Pineda, David E. Myers, Brett A. Bednarcyk, and Thomas M. Krivanek

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

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.

Introduction

The Space Launch Systems (SLS) payload fairing (PLF) is a key component of the next generation of 

heavy lift launch vehicles. The baseline design for the PLF involves separable honeycomb sandwich 

petals composed of aluminum honeycomb core and carbon fiber reinforced polymer (CFRP) facesheets. It 

serves to protect the vehicle payload from aerodynamic and acoustic loads during launch and must 

separate from the vehicle cleanly without re-contact. Pre-existing flaws, such as manufacturing defects or 

unexpected damage occurring during handling, assembly, payload encapsulation, and vehicle integration,

can jeopardize the expected life and performance of the PLF. Typically, any damage initiated in the 

structure prior to launch will be repaired, if detected through visual inspection or non-destructive 

inspection (NDI). However, it is possible that some damage is undetected and does not get repaired, or 

impact damage may occur during flight. As such, the PLF must be damage tolerant.

Low-speed impact such as tool drops or unintended contact with ground support equipment, is the 

most prevalent cause of post-manufacturing damage. For honeycomb sandwich structures, low-speed 

impact results in core crushing, delamination/disbonds, and matrix cracking (Refs. 1 and 2). Subsequent 

loading can lead to kink band formation (fiber microbuckling), indentation (core crushing growth), and 

delamination/disbond growth (Refs. 1 to 4). It is essential to understand how the damage, ensuing from a 

low-speed impact, affects the performance of the PLF during launch, and to develop a suitable damage 

tolerance approach (DTA) for the 10.0 m diameter PLF that can be used to assess the critical flaw size 

and location and includes damage mitigation, inspection, and repair.

The PLF structure is designed to buckle before strength failure occurs and at a relatively low strain 

state. Therefore, due to the large size of the PLF acreage and the nature of the on-pad and aerodynamic 

loads, it is expected that buckling will remain the dominant mode of failure, even when there is damage 

present. A thorough experimental and numerical investigation of the buckling of 1/16th arc segments of 

the full-scale PLF barrel has previously been conducted by NASA for four different panel configurations 

(Refs. 5 to 9). Also contained in these papers is a study on the effect of geometric imperfections on the 

buckling response of the curved sandwich panels. Farrokh et al. (Ref. 10) investigated the effect of a joint, 
NASA/TM—2015-218861 2

and subsequent damage to the joint resulting from impact, on the buckling load of the curved panel 

segment. This work is complimentary to other NASA work related to cylindrical shell buckling 

knockdown factors (Refs. 11 to 13). In the current PLF designs, the resulting strength margins are 

typically high relative to the buckling margins. However, if it was determined that designs utilizing 

pristine strength allowables are not conservative enough to ensure that buckling occurs well before 

strength failure, the open hole or compression after impact (CAI) allowables could be utilized. 

Unfortunately, numerous additional coupon tests would be needed to develop the CAI allowables. 

However, CAI allowables have been developed for comparable systems used in the design of other 

structures, as this is the dominant mode of damage induced failure for these heavier loaded structures 

(Ref. 14). Results from these studies can be leveraged during the design phase of the PLF.

As mentioned previously, various damage mechanisms arise during impact and may evolve during 

service. Here, attention is focused only on the reduction in the overall buckling load of the PLF structure 

due to disbonding between the composite facesheet and honeycomb core resulting from a low speed 

impact. One reason for focusing on facesheet-core disbonds is that the tools necessary to perform the 

analysis have been extensively validated in the open literature and are integrated into commercially 

available software packages. Additionally, others have focused on CAI (Refs. 1 to 4, 14 and 15) and 

fatigue damage growth (Refs. 16 to 18) (as these modes are pertinent to those structures, which included 

the Ares I launch vehicle interstage, the Orion Multipurpose Crew Vehicle Stage Adaptor, and other hatstiffened panels, and solid laminate skins) (Refs. 3, 14 to 18). Information and data from these studies are 

complimentary (as opposed to duplicative) to what will be obtained from this study. These results can be 

utilized to guide further development of a comprehensive DTA.

Herein, numerical analysis of 1/16th arc segments of the full-scale PLF barrel was conducted using 

the finite element method (FEM). The facesheet-core disbond was simulated by separating the facesheet 

nodes from the core nodes in a predetermined, circular, disbonded region. A surface-to-surface contact 

algorithm was utilized to eliminate interpenetration of the facesheet and core elements. 

Previous authors utilized FEM analysis to calculate the strain energy release rate at a delamination 

front in post-buckled laminates (Refs. 19 and 20). Reeder et al. (Ref. 21) used a similar technique to 

investigate the effect of delamination growth on the post-buckling response of solid laminates using the 

virtual crack closure technique (VCCT) (Ref. 22). In addition, numerous authors have modeled the effect 

of core crushing on the CAI strength of honeycomb sandwich panels (Refs. 3, 4, and 15). However in the 

present work, the facesheet-core disbonds are not allowed to propagate, and core crushing is not 

considered in order to isolate the effect of the presence of a facesheet-core disbond on the post-buckling 

response of the panel. Geometrically nonlinear, post-buckling analysis was performed by incrementally 

applying a uniform edge displacement to the curved panel. The buckling load of the curved panel was 

determined by monitoring changes in the slope of the resulting load-deflection curve. A parametric study 

was conducted by varying the diameter of the disbonded region and the post-buckling response and 

buckling loads were compared to experimental and numerical data for pristine panels.

Buckling Test of Pristine Curved Sandwich Panel

To provide baseline data, a buckling experiment on a 3 by 5 ft section of a 1/16th arc segment of a

10 m diameter PLF honeycomb sandwich barrel section without a disbond (pristine) was conducted. The 

complete details and results for this test are available in previous work (Refs. 5 and 23).

The 1/16th fabrication demo panel was constructed on a concave composite tool (5 m radius of 

curvature) using an automated tape laying process. The pre-impregnated (pre-preg) tape was composed of 

unidirectional fibers and was 6 in. wide. The pre-preg contained IM7 fibers and 977-3 epoxy. The 

stacking sequence of the facesheets was [45°/90°/–45°/0°]s. The facesheets were bonded to the 1 in. thick 

aluminum core using FM 300 film adhesive, 0.08 lbs/ft2

. The aluminum honeycomb core was curved, via 

machining, to match the radius of the tool. The facesheets and core adhesive were co-cured in a single 

autoclave cycle. The aluminum core was Alcore PAA-CORE 5052, 0.0007 in. thick with 0.125 in. cell 
NASA/TM—2015-218861 3

size, with a density of 3.1 pcf. An adhesive splice was needed to join discontinuous sections of the core 

because the 1/16th barrel section panel dimensions exceeded the size of the pre-manufactured core. A

core splice consisting of Hysol 9396.6 foaming adhesive was used to fill the gap between the two pieces 

of aluminum core.

One 3 ft wide by 5 ft long test specimen was machined from the manufacturing demo panel following 

non-destructive examination. End plates were employed in the testing. These were 1.0 in. thick aluminum 

plates and had a slot in the shape of the specimens cross section machined in the center. The slot width 

and length were such that, when centered, the specimen had a clearance of 0.5 in around the perimeter. 

After the specimen-end was centered in the slot and squared, it was potted with “UNISORB” V-100 

epoxy grout. When both ends were potted and following cure, specimen ends were machined flat and 

parallel. The potted dimensions of the panel are shown in Figure 1. Note that the final dimensions of the 

gage section were not exactly 3 by 5 ft. The panel width dimension plotted in Figure 1 represents the 

average arc-length of the outer mold line (OML) and inner mold line (IML) facesheets. In addition to the 

overall dimensions, Figure 1 shows the relative position of the core splice with respect to the panel ends.

Once potted, strain gages were affixed to the OML and IML surfaces of the panel. These gages were 

monitored and the strains were recorded during loading of the panel. The panel was tested at the

NASA Langley Research Center (LaRC) in a servo-hydraulic test frame. The panel was secured between 

two loading platens, with the bottom loading platen being fixed and the top platen allowed to move in the

y-direction. The panel was loaded in compression until catastrophic failure occurred. Additional 

instrumentation included three direct-current displacement transducers (DCDTs) used to measure the 

global axial deformation of the panel, and a load cell attached to the load platen to measure the applied 

load. Finally, the panels were painted using a speckle pattern for photogrammetric measurements. Visual 

image correlation (VIC) was employed to obtain full-field strain measurements during the test as well as 

high-speed video.

Figure 1.—Dimensions of pristine curved honeycomb 

sandwich buckling test specimen with end plates.

Dashed line indicates core splice locations.
NASA/TM—2015-218861 4

Determining the Buckling Load

Stable bifurcation of the panel can occur over a finite time (and applied edge displacement). 

Therefore, determining the buckling load of a panel from experimental data can be subjective. Various 

methods for determining the buckling load from experimental data is given in Singer et al. (Ref. 24). Ko

(Ref. 25) compared the force/stiffness (F/S) method and the Southwell method for predicting the buckling 

strength of tubular panels. The original F/S method, developed by Jones and Greene (Ref. 26), used a 

stiffness (force divided by global strain) versus displacement plot to obtain the buckling load.

In the present work, the F/S method was utilized to calculate the buckling load from both the tests and 

analysis. The stiffness measure used was taken to be the applied load P divided by displacement U (the 

average of all three DCDT measurements for the experimental value, or the end shortening for the 

numerical values). Rather than plot this stiffness measure against load, the normalized stiffness P/U

versus displacement curves were generated, as shown Figure 2. Referring to Figure 2, it can be observed 

that there is some initial settling in the test; after which, the critical buckling load Pcr is taken as the load 

at which the relationship between the decreasing stiffness and applied load becomes linear. This is 

marked by the intersection of a nearly vertical dashed line with the F/S curve on Figure 2 and corresponds

to 75.6 kips.

Note that the method used in this work varies slightly from the original F/S method (Ref. 26). In the 

original method, Pcr is taken as the load when the F/S curve crosses the load axis. However, this is well 

into the post-buckled regime. If there is local failure or instability, Jones and Greene (Ref. 26) 

recommend using the point of intersection of a straight line with the F/S curve to determine the buckling 

load. The slope of the straight line is obtained from the critical strain required to initiate local failure. For 

these analyses, this limiting strain was unknown. As mentioned before, these methods (as well as other 

methods in the literature) contain some bias. The method used in this work is no different, but as long as a 

consistent method is used to calculate the Pcr for all cases, comparisons can be made. For the datasets 

obtained from the experiment and generated from the FEM simulations, the described method provided 

the most consistent results with the least ambiguity.

Figure 2.—F/S curve of experiment showing edge load/displacement P/U versus compressive load P.
NASA/TM—2015-218861 5

Numerical Post-Buckling Analysis of Curved Sandwich Structures

Containing Static Disbonds

The post-buckling response of 3 by 5 ft. sections of 1/16th arc segments of a 10.0 m diameter PLF 

honeycomb sandwich barrel section, with and without facesheet-core disbonds, was conducted using the 

Abaqus/Standard FEM package (Ref. 27). The facesheets consisted of 8-ply, IM7/977-3 CFRP composite 

laminates with a [45°/90°/–45°/0°]s stacking sequence. The elastic properties for IM7/977-3 were 

obtained from the Orion materials database. The elastic properties for the aluminum (5052-T6 alloy),

3.1 pcf, honeycomb core were obtained from commercially available datasheet provided by Alcore and 

are presented in Table I.

The honeycomb core was modeled using reduced-order, quadratic solid elements (C3D20R) and the 

facesheets were modeled using reduced-order quadratic shell elements (S8R). The facesheets were 

attached to the core via tie constraints to allow for dissimilar meshes for the cases where a disbond was 

present. From previous analyses, including the experimental end fixtures in the model was deemed noncritical. Thus, they were omitted from these models. A uniform, compressive edge displacement in the

y-direction (see Figure 1) was applied to all nodes on the top edge of the panel, all other degrees of 

freedom (DOFs) on the top edge nodes were constrained, and the DOFs of all nodes on the bottom edge 

of the panel were constrained to simulate typical buckling experiment boundary conditions. The implicit 

dynamic FEM solver, along with geometric non-linearity, in Abaqus were used for the progressive 

collapse (post-buckling) simulations. A realistic displacement rate of 3.0E-4 in./s was used.

The FEM geometry and mesh used to model the pristine panel are presented in Figure 3. The geometry 

was created based on the gage dimension given in Figure 1, such that the average of the arc lengths of the 

OML and IML facesheets was 35.5 in. The core splice shown in Figure 1 was not included in the FEM 

model. A global seed, which controlled the size of the elements, of 1 in. was used to create the meshes for 

the core and facesheets. One element was used through the thickness in the core for all analyses in this 

work. In addition to the post-buckling analysis, a linear eigenvalue analysis was performed.

To simulate damage resulting from low-speed impact, a circular disbond with diameter D, was placed 

between the core and the inner surface of the OML facesheet. The FEM geometry, displaying a typical 

facesheet-core disbond is shown in Figure 4. The disbond area on the outer surface of the core and inner 

surface of the OML (highlighted in red) were not tied. Instead, surface to surface contact interaction was 

used with frictionless, finite sliding and a “hard” pressure-overclosure relationship. The penalty method 

was used as the constraint enforcement method (Ref. 27). The inner OML and outer core surfaces outside 

of the disbonded area (shown in gray in Figure 4) are still attached through a tie constraint, as are the 

entire outer IML and in inner core surfaces because there is no disbond between these surfaces. A 

demonstrative FEM mesh for a panel containing a 5 in. diameter disbond is presented in Figure 5. A 

global seed of 1 in. is still utilized. However, the meshes for the core and OML facesheet differ from the 

pristine model (Figure 3) because of the circular shape of the disbonded area. The IML facesheet remains 

the same as in Figure 3.

TABLE I.—ELASTIC PROPERTIES FOR

ALUMINUM HONEYCOMB CORE

Property, units Value

Ex, ksi 75

Ey, psi 21.28

Ez, psi 21.28

νxy 1.0E-05

νxz 1.0E-05

νyz 0.33

Gxy, ksi 22

Gxz, ksi 45

Gyz, psi 5.32

ρ, pcf 3.1
NASA/TM—2015-218861 6

The diameter of the disbonded area was parametrically reduced in this study (i.e., disband diameter 

decreased from 20 to 5 in., typically in 5 in. increments), post-buckling analyses were conducted, and the 

buckling loads calculated using the method described in the previous section. Disbonds below 5 in. 

diameter were not considered, since diameters smaller than 5 in. had very little effect on the critical 

buckling load and global buckling mode (presented in the next section). Secondly, smaller disbonds 

would have required a finer mesh in the disbonded region, which would have yielded inconsistent mesh 

densities for simulations with disbonds smaller than 5 in. in diameter as compared to those greater than or 

equal to 5 in.

 

Figure 3.—FEM meshes used in numerical analysis of pristine curved honeycomb sandwich panel. 

Left: FEM mesh of IML facesheet composed of S8R, reduced-order shell elements. Middle: FEM 

mesh of honeycomb core composed of C3D20R, reduced-order solid elements. Right: FEM mesh 

of facesheets composed of S8R, reduced-order shell elements.

 

Figure 4.—FEM geometry of curved honeycomb sandwich panel showing disbonded 

area (D = 5 in.) between OML facesheet and core highlighted in red. Left: Honeycomb 

core geometry. Right: OML facesheet geometry.
NASA/TM—2015-218861 7

 

Figure 5.—FEM mesh used in simulations of curved honeycomb sandwich 

panels containing facesheet core disbond (D = 5 in.). Disbonded area 

highlighted in red. Left: Honeycomb core mesh. Right: OML facesheet mesh. 

Note, IML facesheet mesh remains the same as that shown in Figure 3.

Results

Pristine Panel

Critical Buckling Load

Figure 6 shows the F/S curve from the pristine post-buckling analysis, as compared to experiment. It 

can be inferred from Figure 6 that critical buckling load, predicted by FEM, was 71.4 kips. This 

prediction was conservative and exhibiting a 5.6 percent error from the experimental value. The linear 

eigenvalue for this panel was calculated to be 0.19 and the corresponding buckling load was 77.4 kips 

(2.3 percent error). It should be noted that including geometric imperfections has been shown to improve 

the post-buckling results for this panel (Refs. 5 and 9). Such imperfections were not considered herein.

Buckling Mode

The buckling mode (post-buckled shape) of the pristine panel predicted using a linear eigenvalue 

analysis, post-buckling analysis, and obtained from the VIC of the experiment is displayed in Figure 7. 

The radial displacement of the IML surfaces of the panels are shown in Figure 7. Blue/purple indicates 

displacement toward the positive x (or OML) direction (out of the page), whereas red indicates 

displacement towards the IML (into the page). With the linear eigenvalue analysis, the direction in which 

the eigenvectors are applied is arbitrary. The VIC data shows the panel buckling towards the IML, but the 

direction of buckling predicted by the post-buckling analysis was incorrect (opposite). It has been shown 

that geometric imperfections are required to get the correct buckling direction in a post-buckling analysis

for this panel, and the buckling load is marginally dependent on the buckling direction and degree of 

imperfection (Refs. 5 and 9).
NASA/TM—2015-218861 8

Figure 6.—F/S curves showing edge load/displacement P/U versus compressive load P from

post-buckling simulation of pristine panel compared to experiment. 

 

Figure 7.—Post-buckled view of radial displacement IML surface curved honeycomb sandwich 

panels. Left: Eigenvalue analysis. Middle: Post-buckling analysis. Right: Experimental VIC.
NASA/TM—2015-218861 9

Figure 8.—Critical buckling load Pcr versus diameter of circular disbonded D area 

predicted by post-buckling FEM simulations. P* is the design-to buckling load based 

on the linear eigenvalue, a safety factor of 1.4, and a shell buckling knockdown factor 

of 0.65.

Damaged Panel

Critical Buckling Load

A plot of critical buckling load versus disbonded area obtained from the progressive collapse 

simulations is displayed in Figure 8. It can be observed that the presence of even a 5 in. diameter disbond 

affected the buckling load, albeit slightly (1.6 percent reduction in critical load). The influence of a 10 in.

diameter disbond was more pronounced and yielded nearly a 5 percent reduction in Pcr. As expected, the 

critical buckling load continued to decrease as the disbonded area was increased. However, it should be 

mentioned that the energy required to create a 10 in. diameter disbond is extremely high, and would most 

likely result in visible surface damage. Thus, damage resulting from this level of impact would be 

detected and repaired. For comparison, the buckling load from the experiment and from the linear 

eigenvalue analysis are plotted on Figure 8.

Typical aerospace designs utilize a factor of safety (FS) (Ref. 28). In addition, design-to loads for 

shell buckling also include a shell-buckling knockdown factor (SBKF) since buckling is extremely 

sensitive to a plethora of imperfections (Ref. 29). The margin of safety (MS) is calculated using the FS 

and SBKF

1 * − λ = FS

SBKF MS cr (1)

where λcr is the linear eigenvalue obtained from analysis. From this equation the maximum applied load 

P* allowed to meet SBKF and FS requirements based on the current design, can be calculated





 



 = λ

FS

SBKF P Pcr * (2)
NASA/TM—2015-218861 10

where is the buckling load obtained from the linear eigenvalue analysis. Using typical values

(FS = 1.4 and SBKF = 0.65 (Refs. 28 to 29)) P* for the pristine panel was calculated and plotted

(red line) on Figure 8. It can be readily seen the buckling loads of the damaged panels (even for the 

maximum damaged area of 20 in. in diameter) do not approach P*. Furthermore, a 5 in. diameter damage 

zone is well within the detectable range of modern non-destructive evaluation (NDE) techniques, and 

would be repaired upon inspection. This indicates that the SBKF is severely conservative with respect to 

damage, and previous work by Pineda et al. (Ref. 9) has shown similar findings regarding geometric 

imperfections. Although, these studies have not been performed on full cylinders, they support the efforts 

led by the NASA Langley Research Center to formulate a new, more physics-based SBKF to provide 

more design flexibility, and ultimately weight savings (Refs. 11 to 13).

Stiffness

The stiffness of the panels U U =0.1

P used in each simulation was calculated by dividing the applied 

compressive load evaluated at a displacement U of 0.1 in. The degradation in stiffness as a function of 

disbond size, along with stiffness of undamaged test panel, is given in Figure 9. The predicted stiffness of 

the pristine panel was 399 kips/in., which was 1.6 percent greater than the actual stiffness of the panel

(394 kips/in.). Figure 9 illustrates that the presence of a disbond has a noticeable effect on the stiffness of the 

panel. It is possible that this information could be used to detect disbonds or other damage in these panels.

Figure 9.—Stiffness evaluated at 0.1 in applied displacement versus 

diameter of disbonded area D from FEM simulations. Stiffness from experiment on 

undamaged panel is also displayed for comparison.

λ Pcr

U U =0.1

P
NASA/TM—2015-218861 11

 

Figure 10.—Global buckling modes predicted with FEM. (Left). No damage. 

(Middle). Diameter of disbond D = 5 in. (Right). Diameter of disbond D = 10 in.

Buckling Mode

The post-buckled shapes for the undamaged, D = 5 in., and D = 10 in. are presented in Figure 10 and 

the radial displacement contours are plotted on the deformed shapes. The buckling direction of the

D = 5 in. panel was towards the OML, which is the same as the undamaged prediction, but opposite the 

experiment. The out-of plane (radial) displacement of the D = 5 in. panel differs slightly from the pristine 

panel, and the disbonded area is apparent in the displacement field contour. With a disbond diameter of 

10 in. (left contour in Figure 10), the buckling direction switches towards the IML (as in the experiment), 

and facesheet wrinkling can be observed in the damage zone.

Figure 11 shows the progression of radial displacement (as a color contour) and the deformed shape, 

locally in the disbonded area obtained from the post-buckling simulation with a damaged area diameter of 

5 in. Not that the color scales for the contours in Figure 11 are not the same from one load level to the 

next. On the left, the ratio of the current load P to critical buckling load Pcr is given. Early in the 

simulation, at 26 percent of the buckling load, the facesheet started to indent the core, and the local 

wrinkling mode exhibited a half wave shape in each direction. It was not until the load reached 88 percent

of Pcr that the local mode transitioned to a full wave in the y-direction (height), and facesheet and core 

start to separate slightly (the displacement contours of the facesheet and core show some small variation). 

At 99 percent of Pcr the local winkling mode began transitioning towards a 1.5 wave shape, and the core 

has lost contact with the facesheet in a significantly large area. Finally in the post-buckled regime

(101 percent of Pcr) there is substantial loss of contact between the core and facesheet. The core has 

displaced noticeably towards the IML in a 1.5 wave shape, while the OML facesheet remains 

comparatively flat.

Similar plots are given in Figure 12 for the case of a 10 in. disbond, D = 10 in. Resembling the local 

deformation evolution of D = 5 in., the facesheet and core displaced into a half wave shape at 7 percent of 

the buckling. At 59 percent of the buckling load, the core and facesheet transitioned directly into a 1.5 

wave shape, yet contact was lost in the regions surrounding the peaks of the wave shape. The facesheet 

and core separated substantially at 97 percent of the buckling load. Finally at the buckling load, the 

facesheet remained in a 1.5 wave shape while the core deformed into a shape close to a half wave. 

Facesheet-core contact only persisted at the boundaries of the disbonded areas and the center of the panel. 

It can also be seen from Figure 12 that the separation between the facesheet and core progresses much 

faster, relative to Pcr, than when D = 5 in.
NASA/TM—2015-218861 12

P/Pcr Radial 

Displacement 

(in.)

OML 

Facesheet

Core OML

+

Core

0.26

0.88

0.99

1.01

Figure 11.—Local facesheet wrinkling in disbonded zone with diameter D = 5 in.
NASA/TM—2015-218861 13

P/Pcr Radial 

Displacement 

(in.)

OML 

Facesheet

Core OML

+

Core

0.07

0.59

0.97

1.00

Figure 12.—Local facesheet wrinkling in disbonded zone with diameter D = 10 in.
NASA/TM—2015-218861 14

Figure 13.—Representation of compressive response of 

aluminum honeycomb core. Xult is the compressive core 

ultimate strength, and Xcrush is the core crush strength.

Core Crushing

As mentioned in the introduction, another mechanism that may arise in a honeycomb sandwich panel, 

besides debonding of the core and facesheet, is core crushing. This may occur as a result of the impact 

causing the damage or subsequently when the facesheet starts wrinkling (local buckling and pressing on 

the core). The idealized stress-strain response of the honeycomb core is given in Figure 13 (Refs. 1 to 4).

The core behaves linearly until Xult is exceeded, as indicated with the blue line. After which, the core is 

crushed and the stress drops rapidly to Xcrush. Subsequent loading results in perfectly plastic deformation. 

After the core is crushed, the elastic stiffness of the core is reduced, marked by the dashed black line. Any 

elastic unloading and reloading will follow this path until the Xcrush is achieved. Then, the core will 

resume perfectly plastic behavior.

The through-thickness, compressive stress in the core for the undamaged panel did not exceed Xult and 

only surpassed Xcrush well into the post-buckled regime and only locally in a region near the potted ends of 

the panel. The end fixtures were not included in these models; however, it is expected that the potting 

would alleviate these stresses and core crushing would not be an issue for the pristine panel. Moreover, no 

core crushing was observed after the buckling test on the undamaged panel.

Nonlinear behavior in the core (as shown in Figure 13) was not included in the model. However, 

Figure 14 contains plots of the applied load when the through thickness, compressive stress at an 

integration point in a core element first exceed the core allowables as a function of disbond size. The blue 

line containing “*” markers represents the applied load at which the compressive stress in the core first 

exceed Xult, and the purple line with “x” markers shows what the applied load was when the compressive 

stress in the core was first above Xcrush. For reference, the non-linear buckling load Pcr as a function of 

disbond diameter and the knocked down linear eigenvalue load P* for the pristine panel are also 

presented in Figure 14.

It can be seen that Xult is not exceeded until the buckling load is exceeded, or very near the buckling 

load (within 10 percent) for all cases. Most likely, a controlled validation test for this analysis would 

involve placing a Teflon insert between the OML facesheet and core prior to adhering the facesheet to the 

core. Thus, based on Figure 14, core crushing should not be a concern.

However in a “real-world” scenario, impact would be primary the causes of the disbond and could result 

in core crushing. It can be inferred from Figure 14 that, with the current panel design, it is possible that the 

applied loads would surpass the load necessary to induce a nonlinear response in pre-crushed core.

It is not apparent what influence on the overall structure local core crushing will have or to what 

extent the core crushing would evolve. Furthermore, if the core is crushed during the impact event itself, 

the facesheet may also have a change in morphology due to damage, and the local facesheet wrinkling 

mode that occurs during subsequent axial loading might change and/or may not exert large enough 

stresses on the pre-crushed core to result in further core crushing.
NASA/TM—2015-218861 15

Figure 14.—Applied load at which the local through-thickness stress at an integration 

point in a core element first exceeds the core crush ultimate strength (305 psi) Pult

and core crush allowable (145 psi) Pcrush as a function of disbond size. Predicted 

buckling load from post-buckling analysis Pcr and knocked down (by FS and SBKF) 

buckling load from eigenvalue analysis P* also plotted for comparison.

Moreover, the large stresses in the core may only be a facet of the mesh-dependent stress 

concentration induced in the model at the disbond fronts. The core crush allowable was exceed at 

integration points in elements that were near the boundary of the disbonded zone, or the crack front, 

where the gradients are very steep. This mesh dependence may be why the panel with the 7 in. diameter

disbond exhibited a lower load necessary to induce local core crushing than the panels containing 10 in. 

or 15 in. diameter disbonds, as observed in Figure 14.

Lastly, it may be possible that not all of the core in the disbonded region has crushed. Thus, using the 

ultimate core strength (305 psi) to indicate core crush initiation may have been more appropriate.

Higher fidelity analysis, including better mesh refinement near the disbond boundary, or physical 

testing, is needed to assess if core crushing prior to buckling is a realistic concern. To model the 

mathematical singularity appropriately, elements containing double nodes at the crack front are needed. It 

may be possible to capture the stresses surrounding the crack front accurately by including cohesive zone 

elements in the models (Refs. 27, 30, and 31). If core crushing does affect the buckling response of the

panel, a non-linear constitutive model incorporating core crushing behavior can be included (Ref. 3), but a 

non-local consideration of stress in the vicinity of the crack front may be necessary due to the 

mathematical singularity. Furthermore, the evolution of the disbonded area can be included through the 

use of VCCT or cohesive zone elements. Future work will focus on developing models that encompass 

these mechanisms.

Facesheet Failure

The Tsai-Hill failure criterion (Ref. 32) was used to predict potential facesheet strength failure during 

the progressive collapse simulation. B-Basis stress allowables from the Orion database were used in the 

criterion. The failure criterion was satisfied after the panel had buckled for the cases considered. This

would indicate that facesheet failure, which could be due to mechanisms such as matrix cracking, or fiber 

microbuckling (kink band formation), may not influence the buckling load of the impact damaged panel, 

and may not be a primary concern. However, it is still possible that damage induced during the impact

serves as a nucleation site for premature damage growth during subsequent loading.
NASA/TM—2015-218861 16

Conclusions

The effect of impact damage on a large curved honeycomb sandwich structures was idealized as a 

disbond between the CFRP facesheet and aluminum honeycomb core. The panel segments considered are 

part of 10 m diameter SLS payload fairing barrel section. Since the design of the fairing is driven by 

failure due to buckling, not strength, the influence of the presence of a disbond on the buckling load of the 

panel was of primary concern.

The FEM was used to investigate the effect of disbonds on the buckling load of 3 by 5 ft segments of 

the fairing barrel section. Circular disbonded regions were simulated by separating the nodes of the 

facesheet from the core and utilizing a surface-to-surface contact relationship in this region. In a

parametric study, the diameter of the disbonded zone was varied from 5 to 20 in. Progressive collapse 

simulations were performed to predict the post-buckling response of the panels containing disbonds. The 

results from the post-buckling analysis were compared to experimental test data for a pristine 

(undamaged) panel and linear eigenvalue buckling load predictions.

The analysis results indicated that a disbond with a diameter of 10 to 15 in. (which is well within the 

detectable range of NDE techniques) would be required before the buckling load of the panel was 

decreased substantially (> 20 percent). Moreover, a factor of safety (FS) and a shell buckling knockdown 

factor (SBKF) is used when stiffness-driven curved shell structures are designed. The analysis predicted 

that a disbond greater than 20 in. in diameter would be required to violate the FS and SBKF.

The global buckling mode of the panels appeared insensitive to the presence of the disbond. However, 

local buckling, or wrinkling of the facesheet was observed when there was a disbond greater than 5 in.

Moreover, the predicted direction of buckling switched from towards the OML to towards the IML when 

the disbonds were greater than or equal to 10 in. in diameter.

The local facesheet wrinkling induced larger compressive stresses at integration points within core 

elements that were at the boundary of the disbond, or crack front. These stresses exceed the core crushing 

ultimate compressive allowable, provided by the vendor, prior to buckling, indicating that core crushing 

may possibly effect the buckling response of the panel. However, local failure of the facesheets, due to 

strength failure, did not appear to be a prevalent issue.

Future work will involve, the utilization of finer meshes, along with crack tip elements or cohesive 

elements to capture the stress concentrations at the disbond boundary that could affect the core crushing.

In addition, core crushing can be incorporated into the model through the use of a non-linear constitutive 

model for the core, and disbond growth can be captured through VCCT or with cohesive zone elements.

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