Bart F. Zalewski and William B. Dial

ZIN Technologies, Middleburg Heights, Ohio 

Brett A. Bednarcyk

Glenn Research Center, Cleveland, Ohio

Methods for Assessing Honeycomb 

Sandwich Panel Wrinkling Failures

NASA/TM—2012-217697

October 2012
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Bart F. Zalewski and William B. Dial

ZIN Technologies, Middleburg Heights, Ohio 

Brett A. Bednarcyk

Glenn Research Center, Cleveland, Ohio

Methods for Assessing Honeycomb 

Sandwich Panel Wrinkling Failures

NASA/TM—2012-217697

October 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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NASA/TM—2012-217697 1

Methods for Assessing Honeycomb 

Sandwich Panel Wrinkling Failures

Bart F. Zalewski and William B. Dial

ZIN Technologies

Middleburg Heights, Ohio 44130

Brett A. Bednarcyk

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

1.0 Introduction

This document presents the assessment of an accurate yet efficient method for predicting sandwich 

panel facesheet wrinkling using the wrinkling stress calculation developed by Heath (1960) and Vinson 

(1999) and the combined multiaxial failure criterion developed by Ley et al. (1999). These closed form 

predictions were verified using several finite element analyses with varying levels of detail described in 

this report. The closed form method was then implemented into the HyperSizer Structural Sizing 

Software (HyperSizer, 2012) as a user-defined failure criterion, and the implemented code was verified 

using hand calculations. The motivation for this work was the need for an accurate wrinkling solution that 

is also sufficiently computationally efficient so as to be implemented into sizing software, such as 

HyperSizer, for accurate and quick wrinkling assessment early in the structural design process.

Facesheet wrinkling, often referred to simply as wrinkling, is a failure mode that is commonly 

observed in sandwich panels with thin facesheets and lightweight cores, which do not provide a great deal 

of support to the facesheets. The failure is characterized by short wavelength buckling in one or both 

facesheets, as shown in Figure 1, and the failure is usually catastrophic soon after the onset of wrinkling

(Collier Research Corporation, 2011). It is thus critical to account for this failure mode when designing 

sandwich panels for structural applications.

2.0 Closed Form Wrinkling Equations

The existing method in HyperSizer treats sandwich panel facesheet wrinkling as buckling of a simple 

beam on elastic foundation and predicts wrinkling failure based on a modification of the equation 

developed by Hemp (1948), which is the same equation developed by Yusuff (1955). The average 

wrinkling stress is given by,

f c

c f

f E t

E t

sw = 0.82 E (1)

where Ef is the effective modulus of the facesheet, tf is the facesheet thickness, Ec is the core throughthickness modulus, and tc is the core thickness. The equation has shortcomings due to its beam 

assumption when it is used in an analysis of a plate, but it represents a good starting point approximation. 

The Heath equation is more appropriate for honeycomb panels and showed very good correlation with 

both simple and detailed finite element simulations (Vinson, 1999). The Heath (1960) equation was 

developed for an isotropic facesheet and is given by Vinson (1999) as,

( ) 2 3 1

2

c f

f c f

t ν

t E E

sw=

−

(2)
NASA/TM—2012-217697 2

Figure 1.—Schematic illustration of the facesheet wrinkling failure mode in a sandwich panel (Collier Research 

Corporation, 2011).

where νf is the effective Poisson ratio of the facesheet. This equation can be modified for an anisotropic 

facesheet as (Vinson, 1999),

( ) c xy yx

f c fx fy

t ν ν

t E E E

sw=

3 1−

2

(3)

where Efx is the effective modulus of the facesheet in the 0° (x) direction, Efy is the effective modulus of 

the facesheet in the 90° (y) direction, νxy is the xy Poisson ratio of the facesheet, and νyx is the yx Poisson 

ratio of the facesheet. Equation (3) is used in this report to predict the wrinkling stress.

Although Equation (3) is appropriate for the honeycomb construction and accounts for anisotropy of 

the facesheets, it does not consider a case where a plate is subjected to combined multiaxial loading. Real 

structures are almost always subjected to combined multiaxial loads, so they need to be accounted for in 

the wrinkling failure criterion. Ley et al. (1999) suggests two plane stress failure criteria developed by 

Sullins et al. (1969) and Bruhn (1973). The Sullins et al. (1969) wrinkling failure criterion can only be 

used if both of the principal stresses, in the plane of the facesheet, are compressive. The criterion is stated 

as (Ley et al., 1999),

1 σ σ 0 σ σ

1 2

2

2

3

1

1 + = < ≤ 













 , sw sw

(4)

where σ1 is the average major principal stress in the facesheet, σ2 is the average minor principal stress in 

the facesheet, sw1 is the corresponding average wrinkling stress allowable in the major principal direction, 

and sw2 is the corresponding average wrinkling stress allowable in the minor principal direction. The 

Bruhn (1973) wrinkling failure criterion places no restrictions on the stress components, aside from 

requiring one normal stress component to be compressive. This criterion is stated as (Ley et al., 1999):

1

σ σ τ

2 3 3 3

= 















+

+

K sw sw

xx yy xy (5)

where σxx is the average 0° stress in the facesheet, σyy is the average 90° stress in the facesheet, τxy is the 

average shear stress in the facesheet, and K is a factor dependent on the direction of the largest applied 

compressive stress. If the largest compressive stress is parallel to the core ribbon direction K = 1, 

otherwise, K = 0.95. The Sullins et al. (1969) failure criterion was modified to account for Equation (3), 

which does not differentiate the wrinkling stress between principal directions as,

1 σ σ 0 σ σ

1 2

2

3

1  + = < ≤ 

 



 , sw sw

(6)
NASA/TM—2012-217697 3

The Bruhn (1973) equation was also modified with a conservative approach of always setting 

K = 0.95, resulting in,

1

τ

0.95

σ σ

2 3 3 3

= 















+

+

sw sw

xx yy xy (7)

3.0 Uniaxial Verification of Closed Form Wrinkling Equations

Equation (3) was verified first using simple plate finite element models that were subjected to 

uniaxial compression. For this verification study, an 8 in. wide by 11 in. high flat sandwich panel was 

considered. The results were checked for sensitivity to the through thickness mesh density of the core. 

The facesheets were constructed using a quasi-isotropic [45/90/–45/0]s layup of IM7/977-3 plies and a 

1 in. thick core was considered as a 3.1 pcf 1/8-5052-.0007 aluminum honeycomb. The material 

properties of the facesheet plies are restricted and thus not provided herein. The core effective material 

properties are given in Table I.

TABLE I.—3.1 pcf 1/8-5052-.0007 ALUMINUM HONEYCOMB EFFECTIVE MATERIAL PROPERTIES

Property Symbol Value

Through-thickness tensile Young's modulus Et 0.075 Msi

Through-thickness compressive Young's modulus Ec 0.075 Msi

Transverse through-thickness shear modulus Gw 0.022 Msi

Longitudinal through-thickness shear modulus Gl 0.045 Msi

Through-thickness tensile strength Ftu 0.215 ksi

Transverse through-thickness shear strength Fsuw 0.09 ksi

Longitudinal through-thickness shear strength Fsul 0.154 ksi

Stabilized through-thickness compressive strength Fcus 0.215 ksi

Bare through-thickness compressive strength Fcub 0.2 ksi

Through-thickness compressive crushed stress Fcuc 0.13 ksi

The simple finite element models explicitly considered the through-thickness behavior, with the width 

of the panel (8 in.) being captured as the out-of-plane thickness of the elements. Since the facesheets in this 

analysis are quasi-isotropic, they were modeled using isotropic plates with equivalent, homogenized, elastic

properties. All models (see Figure 3) consisted of facesheets with eight linear plane stress plate elements

though the thickness, while the core was modeled using orthotropic plate elements to reflect the significant 

differences between axial and through-thickness properties. Both full and symmetric models were examined 

to observe the impact of the assumption of symmetry on the predicted wrinkling behavior. The boundary 

conditions are shown in Figure 2 and consisted of fixed constraints at the bottom edge, symmetry or free

constraints on the sides, and applied vertical displacement at the top, with other displacement components 

constrained. Linear buckling and geometrically nonlinear static analyses were performed to ensure the 

validity of both of these types of solution. As will be shown, in general, they corresponded very well. The 

six finite element models used in this comparative study are described in Table II.

The applied global displacement at which wrinkling was predicted by the finite element models is 

given in Table II. For the linear buckling analyses, these values were calculated from the first eigenvalue 

returned by the finite element solver. For the nonlinear static analyses wrinkling displacement was taken 

as the displacement right before the load began to decrease. It should be noted that the onset of wrinkling 

was insensitive to the number of elements used through the thickness of the core, the assumption of 

symmetry, and linear buckling versus nonlinear analysis method. Figure 3 shows wrinkling shapes 

obtained from FEA. It should be noted that these are exaggerated to clarify the wrinkling mode shape. 

Next, the closed form expression given by Equation (3) was used to perform a quick assessment of 

wrinkling for the honeycomb panel, and the result was compared with the wrinkling load predicted by
NASA/TM—2012-217697 4

Figure 2.—Boundary conditions employed 

in the two-dimensional planar honeycomb 

sandwich finite element model.

TABLE II.—TWO-DIMENSIONAL PLANAR HONEYCOMB SANDWICH FINITE ELEMENT MODELS

Model 

number

Type Elements through the 

facesheet thickness

Elements through 

the core

Analyses performed

1 Symmetric 8 20 Linear buckling

Nonlinear static

2 Symmetric 8 8 Linear buckling

3 Symmetric 8 4 Linear buckling

4 Symmetric 8 2 Linear buckling

5 Full 8 1 Linear buckling

Nonlinear static

6 Full 8 20 Linear buckling

Nonlinear static

TABLE III.—COMPARISON OF THE TWO-DIMENSIONAL PLANAR HONEYCOMB 

SANDWICH FINITE ELEMENT MODELS AND THE MARGIN OF SAFETY (MOS) 

PREDICTED USING THE CLOSED FORM EXPRESSION

Model Analysis type Wrinkling displacement,

in.

Closed form 

MOS

1 Linear buckling 0.1823088 0.0246

1 Nonlinear static 0.1925624 –0.0296

2 Linear buckling 0.1823032 0.0247

3 Linear buckling 0.1822670 0.0249

4 Linear buckling 0.1821046 0.0258

5 Linear buckling 0.1816686 0.0282

5 Nonlinear static 0.1900624 –0.0169

6 Linear buckling 0.1823088 0.0246

6 Nonlinear static 0.1930624 –0.0321
NASA/TM—2012-217697 5

each finite element model. The margin of safety (MOS) associated with the closed form expression was 

calculated as the closed form equation wrinkling stress divided by the finite element model wrinkling 

stress minus 1. Thus, perfect agreement between the closed form expression and a given finite element 

model would yield a MOS of zero. The results, shown in Table III, illustrate a very good correlation 

between the observed wrinkling in finite element solutions and the closed form formulation of 

Equation (3) for the considered uniaxial loading. Compared to all six linear finite element analyses, the 

closed form expression provided a positive MOS, meaning the closed form expression was slightly nonconservative. Compared to the three nonlinear finite element analyses, the closed form expression was 

slightly conservative, with a negative MOS. The good agreement reported in Table III indicates potential 

for using this equation for wrinkling stress assessment in combined loading situations in conjunction with 

an appropriate multiaxial failure criterion. 

Model Wrinkling shapes

1 

Linear

1 

Nonlinear

2 Linear

3 

Linear

4 

Linear

5 

Linear

5 

Nonlinear

6 

Linear

6 

Nonlinear

Figure 3.—Wrinkling shapes for each two-dimensional planar finite element analysis.
NASA/TM—2012-217697 6

4.0 Multiaxial Verification of Closed Form Wrinkling Equations

To assess wrinkling in a panel where combined loading is present, two detailed finite element models 

were constructed that explicitly model the cell-level details of the core aluminum honeycomb. Both 

models were symmetric in the 8 in. width direction of the panel (i.e., 4 in. width analyzed), see Figure 4. 

The boundary conditions consisted of a fixed bottom edge, symmetry constraint on one of the sides, and 

an applied vertical displacement at the top, with the other displacement components constrained. 

Although only a uniaxial displacement was applied, the corners of the panel experienced multiaxial state 

of stress as the panel’s Poisson effect expansion (y-direction) is constrained by the fixed boundary 

conditions at the top and the bottom. The facesheets were modeled using plate elements with explicit 

[45/90/–45/0]s layups, and the core was modeled using isotropic plate elements with Al 5052 properties

(E = 10.2 Msi, ν = 0.33). Four elements were used through the core thickness. The first model included 

nonlinear spring elements between core and facesheets to model the contribution of adhesive and account 

for the different core stiffness in tension and compression. The compressive spring stiffness was restricted 

from displacing more than half the facesheet thickness. The model was subjected to a nonlinear static 

analysis that consisted of 10 steps with the structural stiffness matrix being updated at each iteration. 

Additional details of the finite element model are shown in Figure 5.

In the first model, the stiffness of the springs used to model the adhesive was calibrated using a 2- by 

2-in. model, constructed similarly to the above model, such that the tensile and compressive stiffness 

matched through thickness tension and compression test results (Lerch, 2011). In these tests, the average 

core through thickness tensile and compressive Young's moduli were 44 and 80 ksi, respectively. The 

calibration procedure resulted in the following tensile (kt), compressive (kc), and shear (ks) spring 

stiffnesses,







× < −

≥ − =

= =

 

 

1.022 10 lb/in., 0.0212

1379 lb/in., 0.0212

774.8 lb/in.

6

z

z

c

t s

u

u

k

k k

(8)

The nonlinear progressive collapse analysis resulted in the through thickness displacements shown in 

Figure 6. The figure shows out-of-plane displacement contours and total displacement shapes at various 

global displacements, which are noted on the figure.

The model results show progressive collapse, which mimics a wrinkling type failure, as would be 

expected. The strains at the top and bottom corners of the panel were monitored to determine the 

displacement at which the panel wrinkled. Since wrinkling is a facesheet buckling phenomena, strains on 

the outside and on the core side of a single facesheet were considered. For global buckling one would 

need to consider strains on both facesheets. The strain data are plotted in Figure 7 for the corner point 

indicated with a circle in Figure 6.The point at which the panel wrinkled was determined as the point at 

which the principal strain reverses slope and remains reversed (Singer et al., 1998) (Figure 7).

The first detailed model took approximately 64 hr to run on a 32 bit Windows PC using the 

NASTRAN finite element code. 

The second detailed model took into account differences between the through thickness tensile and 

compressive properties of the core by modeling the elastic modulus of the core aluminum material as a 

nonlinear property calibrated to match the aforementioned experimental results. This was meant to 

alleviate a slight penetration of the facesheet into the core observed in the first detailed model. A tensile 

modulus of 3.438 Msi and a compressive modulus of 6.250 Msi resulted from the calibration procedure. 

The second detailed model was subjected to a nonlinear static analysis that consisted of 10 steps with the 

stiffness matrix being updated at each iteration. The model was the same as before except for the 

facesheet having coincidental nodes with the core, instead of facesheet and core being connected through 

nonlinear spring elements. 
NASA/TM—2012-217697 7

Figure 4.—Detailed finite element model of the honeycomb sandwich panel.
NASA/TM—2012-217697 8

Figure 5.—Details of the finite element model mesh. (a) Through thickness view with facesheet. (b)

Through thickness view without facesheet. (c) Isoparametric view with facesheet. (d) Isoparametric 

view with facesheet.

(a) (b)

(c) (d)
NASA/TM—2012-217697 9

Figure 6.—Progressive wrinkling deformation of the detailed model utilizing nonlinear springs to model the adhesive 

between the facesheet and core.

Figure 7.—Determination of wrinkling point in the first detailed finite element model.

0.111 in.
NASA/TM—2012-217697 10

The nonlinear progressive analysis resulted in the through thickness displacements shown in Figure 8. 

The figure shows out-of-plane displacement contours and total displacement shapes at various global 

displacements, which are indicated on the figure. Note that these global displacement levels are not the 

same as those shown in Figure 6, although both Figure 6 and Figure 8 include a figure at the displacement 

associated with the simulated buckling point. 

As in the first model, the second model results show progressive collapse, which mimics a wrinkling 

type failure, as would be expected. As before, the strains at the top and bottom of the panel were 

monitored to determine the displacement at which the panel wrinkled, see Figure 9. The point plotted in 

Figure 9 is indicated with the circle in Figure 8.

The second detailed model took approximately 68.5 hr to run on a 32 Windows PC using the 

NASTRAN finite element software. 

Since wrinkling is a local buckling phenomena, the method for wrinkling failure determination was 

used similar to that used for global buckling. The point at which the strain slope reversed was considered 

as failure (Singer et al., 1998) (Figure 9).

Figure 8.—Progressive wrinkling deformation of the detailed model utilizing coincident nodes between the facesheet 

and core and distinct tensile and compressive moduli for the core aluminum material.
NASA/TM—2012-217697 11

Figure 9.—Determination of wrinkling point in the second detailed finite element model.

5.0 HyperSizer Implementation

The two detailed models were compared to a laminated plate model with a layup consisting of the 

entire sandwich stack up, see Figure 10. A model like this can be used in HyperSizer to perform sizing 

analysis iterations and failure analyses. The model a fixed bottom edge with the applied vertical 

displacement at the top edge, while other displacement components at the top edge were constrained. Two 

displacements were applied, each corresponding to a wrinkling displacement from the detailed models, 

namely 0.111 and 0.112 in. The element forces were imported into HyperSizer to perform the built in

wrinkling analysis. 

The wrinkling margins predicted by HyperSizer using the existing Hemp (1948) Equation (1) for the 

applied displacements of 0.111 and 0.112 in. are displayed in Figure 11 and Figure 12.

The existing HyperSizer method based on the Hemp (1948) equation provides a highly negative 

margin of safety, indicating the high degree of conservatism compared to the detailed finite element 

model results. Equations (3), (6), and (7) were implemented into HyperSizer as a user defined failure 

criterion. This requires the user to provide a dynamically linked library (.dll) of the code that returns a 

margin of safety to HyperSizer. Thus, two margin of safety equations were developed based on 

formulation provided by Chambers (1995). The margins of safety (MOS) associated with Equations (6) 

and (7) can be written as,

1, σ σ 0

σ σ

1 MOS 1 2

2

3

1

− < ≤

 + 

 





=

sw sw

(9)

1

τ

0.95

σ σ

1 MOS

2 3 3 3

−

















+

+

=

sw sw

xx yy xy

(10)

0.112 in.
NASA/TM—2012-217697 12

Figure 10.—Laminate plate finite element model used for HyperSizer evaluation.

Ultimate MS γ LS Location – Analysis Description

–0.3376 (0) 1 Top Honeycomb Face Wrinkling, Eq. (2), Honeycomb or RCS Core, X, Y, and Interaction

–0.3376 (0) 1 Bottom Honeycomb Face Wrinkling, Eq. (2), Honeycomb or RCS Core, X, Y, and Interaction

Figure 11.—Wrinkling margins of safety for the first detailed model using the existing HyperSizer method.

Ultimate MS γ LS Location – Analysis Description

–0.3426 (0) 1 Top Honeycomb Face Wrinkling, Eq. (2), Honeycomb or RCS Core, X, Y, and Interaction

–0.3426 (0) 1 Bottom Honeycomb Face Wrinkling, Eq. (2), Honeycomb or RCS Core, X, Y, and Interaction

Figure 12.—Wrinkling margins of safety for the second detailed model using the existing HyperSizer method.
NASA/TM—2012-217697 13

Ultimate MS γ LS Location – Analysis Description

0.0652 (0) 1 Top Honeycomb Face Sandwich Face, User Defined #1

0.0652 (0) 1 Bottom Honeycomb Face Sandwich Face, User Defined #1

Figure 13.—Wrinkling margins of safety for the first detailed model using the newly 

implemented HyperSizer method.

Ultimate MS γ LS Location – Analysis Description

0.05423 (0) 1 Top Honeycomb Face Sandwich Face, User Defined #1

0.05423 (0) 1 Bottom Honeycomb Face Sandwich Face, User Defined #1

Figure 14.—Wrinkling margins of safety for the second detailed model using the 

newly implemented HyperSizer method.

The margins of safety calculated using this new method for the wrinkling displacements determined 

from the two detailed finite element models are given in Figure 13 and Figure 14. It is evident that the 

new failure margins of safety correspond very well with the detailed finite element analyses, although 

they are slightly non-conservative. This non-conservatism could be handled with a knock-down or 

correction factor when using this method in design. It is also noteworthy that the discrepancy between the 

new closed form method and the detailed finite element results for this case that includes multiaxial loads 

are approximately twice those observed in the pure uniaxial cases reported in Table III. The run time for 

the entire analysis was on the order of few minutes due to the computational efficiency of the HyperSizer 

process.

6.0 Conclusion

This report has presented the verification of a set of closed form equations that can be used to predict 

the facesheet wrinkling failure of sandwich panels. Based on the work of Sullins et al. (1969), Bruhn 

(1973), and Ley et al. (1999), these equations admit anisotropic facesheet properties and multiaxial 

loading. The equations were implemented within the HyperSizer Structural Sizing Software (HyperSizer, 

2012) such that they can be used in sizing structures to evaluate wrinkling failures in sandwich panel 

structural applications. The equations were compared to linear and nonlinear simplified plane finite 

element models for wrinkling of a realistic IM7/977-3 facesheet, aluminum honeycomb core sandwich 

panel for uniaxial loading, as well detailed finite element models of the panel that explicitly model the 

honeycomb structure for multiaxial loading. The new wrinkling assessment equations matched very well 

with the finite element results for uniaxial loading. For multiaxial loading, the equations were shown to 

be still in very good agreement with the detailed finite element results (although slightly nonconservative), and in much better agreement than the existing wrinkling assessment methodology within 

HyperSizer.

References

Bruhn E. F., “Analysis and Design of Flight Vehicle Structures,” Tri-State Offset Company, pp. C12.1-

C12.52, 1973.

Chambers, J. A., “Preloaded Joint Analysis Methodology for Space Flight Structures,” NASA TM–

106943, 1995.

Collier Research Corporation, “Methods and Equations for Sandwich Panel Facesheet Wrinkling,”

HyperSizer Structural Sizing Software Documentation, 2011.

Heath, W. G., “Sandwich Construction, Part 2: The Optimum Design of Flat Sandwich Panels,” Aircraft 

Engineering, 1960.
NASA/TM—2012-217697 14

Hemp, W. S., “On a Theory of Sandwich Construction,” ARC Technical Report R&M 2672, 1948.

HyperSizer Structural Sizing Software, Collier Research Corp., Newport News, VA, 

www.HyperSizer.com, 2012.

Lerch B. A., Personal Communication, NASA Glenn Research Center, 2011.

Ley, R. P., Lin, W., and Mbanefo, U.., “Facesheet Wrinkling in Sandwich Structures,” NASA/CR—1999-

208994, 1999.

Singer J., Arbocz, J., T. Weller, T. “Buckling Experiments: Experimental Methods in Buckling of ThinWalled Structures,” Wiley, 1998.

Sullins, R. T., Smith G. W., and Spier, E. E., “Manual for Structural Stability Analysis of Sandwich 

Plates and Shells,” NASA CR–1457, 1969.

Vinson, J. R., The Behavior of Sandwich Structures of Isotropic and Composite Materials, Technomic 

Publishing Co., Lancaster, PA, 1999.

Yusuff, S., “Theory of Wrinkling in Sandwich Construction,” Journal of the Royal Aeronautical Society, 

Vol. 59, pp. 30-36, 1955.

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This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

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14. ABSTRACT

Efficient closed-form methods for predicting the facesheet wrinkling failure mode in sandwich panels are assessed. Comparisons were made 

with finite element model predictions for facesheet wrinkling, and a validated closed-form method was implemented in the HyperSizer 

structure sizing software. 

15. SUBJECT TERMS

Wrinkling; Finite element method; Buckling; Nonlinearity; Failure; Honeycomb structures; Sandwich structures; Mathematical 

models; Structural design; Failure modes; Anisotropy; Compression loads; NASTRAN; Laminates; Simulation 

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