National Aeronautics and

Space Administration

IS20

George C. Marshall Space Flight Center

Marshall Space Flight Center, Alabama

35812
The NASA STI Program...in Profile

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientific and Technical Information (STI) 

Program Office plays a key part in helping NASA 

maintain this important role.

The NASA STI Program Office is operated by 

Langley Research Center, the lead center for 

NASA’s scientific and technical information. The 

NASA STI Program Office provides access to 

the NASA STI Database, the largest collection of 

aeronautical and space science STI in the world. 

The Program Office is also NASA’s institutional 

mechanism for disseminating the results of its 

research and development activities. These results 

are published by NASA in the NASA STI Report 

Series, which includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant 

phase of research that present the results of 

NASA programs and include extensive data 

or theoretical analysis. Includes compilations 

of significant scientific and technical data 

and information deemed to be of continuing 

reference value. NASA’s counterpart of peerreviewed formal professional papers but has less 

stringent limitations on manuscript length and 

extent of graphic presentations.

• TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or of 

specialized interest, e.g., quick release reports, 

working papers, and bibliographies that contain 

minimal annotation. Does not contain extensive 

analysis.

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientific and technical conferences, 

symposia, seminars, or other meetings sponsored 

or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientific, technical, 

or historical information from NASA programs, 

projects, and mission, often concerned with 

subjects having substantial public interest.

• TECHNICAL TRANSLATION. 

English-language translations of foreign 

scientific and technical material pertinent to 

NASA’s mission.

Specialized services that complement the STI 

Program Office’s diverse offerings include creating 

custom thesauri, building customized databases, 

organizing and publishing research results...even 

providing videos.

For more information about the NASA STI Program 

Office, see the following:

• Access the NASA STI program home page at 

<http://www.sti.nasa.gov>

• E-mail your question via the Internet to 

<help@sti.nasa.gov>

• Fax your question to the NASA STI Help Desk 

at 443 –757–5803

• Phone the NASA STI Help Desk at 

443 –757–5802

• Write to:

NASA STI Help Desk

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076–1320
i

NASA/TP—2011–216460

Comparison of Open-Hole Compression

Strength and Compression After Impact

Strength on Carbon Fiber/Epoxy Laminates 

for the Ares I Composite Interstage

A.J. Hodge, A.T. Nettles, and J.R. Jackson

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

March 2011

National Aeronautics and

Space Administration

Marshall Space Flight Center • MSFC, Alabama 35812
ii

Available from:

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076 –1320

443–757–5802

This report is also available in electronic form at

<https://www2.sti.nasa.gov/login/wt/>
iii

TABLE OF CONTENTS

1. INTRODUCTION ............................................................................................................. 1

2. MATERIALS AND TESTING .......................................................................................... 3

2.1 Compression After Impact ........................................................................................... 3

2.2 Open-Hole Compression .............................................................................................. 6

2.3 Testing of Simulated Delaminations and Unbonds ...................................................... 9

2.4 Comparison of Impact Damage, Open-Hole Compression, 

and Delamination ........................................................................................................ 10

2.5 Nondestructive Evaluation Size Correlation ................................................................ 10

3. CONCLUSIONS ................................................................................................................ 13

APPENDIX A—MECHANICAL TEST DATA .................................................................... 14

APPENDIX B—CROSS SECTION OF IMPACT DAMAGE ............................................... 18

REFERENCES ....................................................................................................................... 22
iv

LIST OF FIGURES

 1. Schematic of stress distribution near a hole with characteristic distance 

(do and ao) as applicable to the point and average stress criterion .............................. 2

 2. NDE indication width versus impact energy ............................................................. 4

 3. Sandwich edgewise compression ................................................................................ 4

 4. Residual compression strength versus impact energy ................................................ 5

 5. Residual compression strength versus NDE indication size ....................................... 6

 6. Northrop OHC fixture .............................................................................................. 7

 7. Boeing CAI fixture .................................................................................................... 7

 8. Laminate OHC results .............................................................................................. 8

 9. OHC normalized to 1/32-in-diameter hole ................................................................ 9

10. Comparison of all defect types .................................................................................. 11

11. Finite element analysis model of residual thermally induced stresses in 

(a) 12 lb/ft3 core and (b) 6.1 lb/ft3 core ...................................................................... 17

12. 1 ft-lb, 0.5-in-diameter impactor (no delamination) .................................................. 18

13. 2 ft-lb, 0.5-in-diameter impactor ............................................................................... 18

14. 2 ft-lb, 1.5-in-diameter impactor ............................................................................... 18

15. 3 ft-lb, 0.5-in-diameter impactor ............................................................................... 19

16. 3 ft-lb, 1.5-in-diameter impactor ............................................................................... 19

17. 4 ft-lb, 0.5-in-diameter impactor ............................................................................... 19

18. 4 ft-lb, 1.5-in-diameter impactor ............................................................................... 19

19. 6 ft-lb, 0.5-in-diameter impactor ............................................................................... 20
v

LIST OF FIGURES (Continued)

20. 6 ft-lb, 1.5-in-diameter impactor ............................................................................... 20

21. 8 ft-lb, 0.5 in-diameter impactor ................................................................................ 20

22. 8 ft-lb, 1.5 in-diameter impactor ................................................................................ 20

23. 10 ft-lb, 0.5 in-diameter impactor .............................................................................. 21

24. 10 ft-lb, 1.5-in-diameter impactor .............................................................................. 21
vi

LIST OF TABLES

1. Compression test results of sandwich defect coupons ................................................. 10

2. Comparison of IRT, UT, and cross-sectional analysis for different impact 

energies with a 0.5-in-diameter impactor .................................................................... 11

3. Comparison of IRT, UT, and cross-sectional analysis for different impact 

energies with a 1.5-in-diameter impactor .................................................................... 12

4. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results ........................................... 14

5. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results (batch 1) ............................ 15

6. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results (batch 2) ............................ 15

7. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results (batch 3) ............................ 16

8. 56-ply face sheet, 12 lb/ft3 core damage tolerance results ............................................ 16
vii

LIST OF ACRONYMS AND SYMBOLS

Al aluminum

CAI compression after impact

FEP fluorinated ethylene propylene

IRT infrared thermography

NDE nondestructive evaluation 

OHC open-hole compression

OML outer mold line

PAUT passed array ultrasonic testing 

UT ultrasonic test
viii

NOMENCLATURE

ao average stress criterion characteristic distance

do point stress criterion characteristic distance

EL longitudinal modulus of elasticity

ET transverse modulus of elasticity

GLT shear modulus

kT stress concentration factor

x impact energy

b0 regression constant

b1 regression coefficient

nLT Poisson’s ratio

σr residual stress
1

TECHNICAL PUBLICATION

COMPARISON OF OPEN-HOLE COMPRESSION STRENGTH AND COMPRESSION

AFTER IMPACT STRENGTH ON CARBON FIBER/EPOXY LAMINATES 

FOR THE ARES I COMPOSITE INTERSTAGE

1. INTRODUCTION

 

Mechanical testing of representative subelements of the Ares I composite interstage was performed as part of the development of damage tolerance strength allowables and a damage threat 

assessment. Properties obtained included compression after impact (CAI), notched compression, 

and compression with simulated delaminations. The results were used to assess the damage tolerance 

of the interstage design and establish thresholds for inspection. This Technical Publication compares 

the compression strength results from various impact energy levels and impactor geometries to the 

compression results of various notch and delamination sizes.

A failure prediction model of ‘notched’ composite laminates (i.e., laminates with a stress 

riser) has been an endeavor pursued for many years. Notches from holes or impact damage in a�  laminate are more detrimental to compression strength than the tensile strength.1–6 In the 1970s, Whitney 

and Nuismer provided the average stress and point stress failure criteria for notched laminates based 

upon fracture toughness criteria.7 Originally developed for notched tension, this method has successfully been applied to notched compression. The stress concentration factor kT is used in conjunction 

with a characteristic distance for the two criteria for the laminate to fit a curve to empirically derived 

data (fig.1). In the point stress criterion, failure occurs when the stress at a distance do away from 

the hole reaches the unnotched material strength σo. In the average stress criterion, failure occurs 

when the average stress over the distance ao reaches σo. It was originally hoped that the characteristic distance parameters would be universal constants for all fibrous material systems and all layups. 

However, it has been shown that these values are dependent upon many factors including hole size 

and stacking sequence.8 Notwithstanding, the simplicity of application of the Whitney-Nuismer 

model adds appeal to any user and this predictive model is described in most textbooks. In the 1990s, 

Soutis et al. developed a cohesive zone model for predicting open-hole compression (OHC) strength 

based upon the kink band formation in zero-degree plys within the laminate.9 The model was further 

refined to predict CAI strength. Other researchers have continued to refine these approaches. Chen 

et al. modified the Whitney average stress criterion to calculate a characteristic length for tension and 

compression.10 The model agreed well for all laminates tested except those with a high percentage of 

zero-degree plys. Callus compared the point and average stress criterion models with a cohesive zone 

model.11 The average stress criterion and cohesive zone model were almost identical for hole sizes 

zero to 0.25 in. 
2

o 

y

d x o ao

 o

σ

σ

σ

 σx

Figure 1. Schematic of stress distribution near a hole with characteristic 

distance (do and ao) as applicable to the point and average 

stress criterion.

Other researchers have made comparisons between the strength reductions of OHC and CAI. 

Byers (1980) found that impact damage size measured by ultrasonic through transmission was not 

a good predictor of residual strength and did not correlate well with OHC, while Williams (1984) 

found similar reductions in strength between OHC and CAI strengths for a given damage size.12,13

More recently, Chen modeled CAI residual strength as an eliptical hole on the assumption of similar 

failure mechanisms; however, Edgren et al. concluded that modeling impact damage as an open hole 

was not valid.14,15 Clearly, the study of compressive failure mechanisms of notched laminates is an 

area of intense study.

A major disadvantage of any of these failure criteria is the reliance on an accurate empirical 

measurement of unnotched strength. Unnotched strength is dependent upon test method, specimen geometry, and quality of specimen preparation. The models described above can alternately 

be utilized to predict unnotched strength when notched strength is known. This raises the question: “Why normalize to unnotched strength?” Strength allowables consist of reduction factors, or 

knockdowns, for discontinuities, notches, process variability, and damage such that a well-designed 

structure would never be loaded to pristine, unnotched strength values. Providing an unnotched 

strength allowable may in fact be detrimental to the design process by suggesting that such capability 

is achievable. Decisions and analyses become anchored to this irrelevant value rather than a value 

relevant to the design function of a composite structure.
3

2. MATERIALS AND TESTING

Face sheets in this study were composed of IM7/8552-1 140 g/m2 areal weight fiber placed 

ribbon. The acreage design of the Ares I upper stage interstage consisted of thin laminates with 

a (45,0,–45,0,90,0,0,90,0)s layup on each face sheet. Thick laminates represented buildups for high 

stress regions and were 40- and 56-ply quasi-isotropic. Acreage sections were co-cured onto 1.5-inthick 5052 aluminum (Al) honeycomb core with 1/8-in cell size on two densities of core—3.1 and 

6.1 lb/ft3. Processing problems were encountered with the lighter core material which included face 

sheet wrinkling and core crushing during autoclave cure. Thus, only the 6.1 lb/ft3 results are examined. Forty-ply face sheet laminates were only fabricated with 6.1 lb/ft3 core. Fifty-six-ply face sheets 

were co-cured onto 1.5-in-thick, 12 lb/ft3 Al core. It was later found that the large coefficient of thermal expansion mismatch between the 12 lb/ft3 core and the carbon fiber face sheets led to core node 

disbonds during cooldown on two of the three batches of core. Face sheets were bonded to the core 

with 0.08 lb/ft2 FM300K film adhesive at 355 °F with a 55 lb/in2 autoclave pressure. 

Three series of tests were conducted: CAI, OHC, and compression with simulated delamination. Delaminations were simulated with two layers of 0.001-in-thick fluorinated ethylene propylene 

(FEP) tape inserted within the face sheet during layup. The OHC specimens were not sandwich panels, but were instead monolithic laminates cured at the same pressure as the sandwich panels.

The raw test data for the mechanical tests are available in appendix A.

2.1 Compression After Impact

Impact testing was performed with three different impactor diameters (0.25, 0.5, and 1.5 in). 

Sandwich specimens (4 in wide by 6 in long) were impacted over a range of impact energies such as to 

cause damage ranges not visible to face sheet penetration. Post-impact damage sizes were obtained 

with nondestructive evaluation (NDE). Thin face sheet specimens were evaluated with infrared thermography (IRT) and thick face sheet specimens were evaluated with passed array ultrasonic testing 

(PAUT). The length and width of the NDE indication was noted for each specimen. Figure 2 illustrates the damage width indicated by NDE compared to the impact energy and impactor size. The 

0.25-in-diameter impacting tup began to penetrate at 10 ft-lb with full penetration of the face sheet 

at 12 ft-lb. The 0.5-in-diameter tup began to penetrate the face sheet at 15 ft-lb. The 1.5-in-diameter 

tup did not penetrate the face sheet at the impact energies applied and produced little visible damage 

other than a dent. It is apparent from figure 1 that damage size is a function of impactor size, especially between the 0.5- and 1.5-in-diameter impactors.
4

0 2 4 6 8 10 12 14 16

1.4

1.2

1

0.8

0.6

0.4

0.2

0

18-Ply Face Sheet, 0.25-in-Diameter Tup

18-Ply Face Sheet, 0.5-in-Diameter Tup

18-Ply Face Sheet, 1.5-in-Diameter Tup

Impact Energy (ft-lb)

IRT Width (in)

Figure 2. NDE indication width versus impact energy.

CAI specimens were then potted with epoxy into a 1-in-thick Al frame on each loading end. 

The loading ends were machined flat and parallel to within 0.001 in. Strain gages were bonded to 

the front and back surfaces away from the damage zone to monitor alignment during mechanical 

testing. Figure 3 is a schematic of the edgewise compression test.

Aluminum Frame

Test Specimen

Strain Gage

Potting Material

Figure 3. Sandwich edgewise compression.
5

In order to apply a uniform strain to both face sheets, the compression platens were fixed. 

The impacted sandwich specimens were loaded in edgewise compression at a constant crosshead 

rate until failure. Figures 4 and 5 illustrate the residual compression strength as compared to impact 

energy and damage width for the three impactor sizes used. It is apparent that the residual strength 

is dependent upon the impact energy and is independent of the impactor sizes used in this study. 

However, the residual strength appears to reach a plateau for a given critical impact energy, even if 

the damage size increases (≈45,000 psi at 12 ft-lb for the data presented here).

0 2 4 6 8 10 12 14 16

80,000

70,000

60,000

50,000

40,000

30,000

20,000

10,000

0

Impact Energy (ft-lb)

Stress (psi)

18-Ply Face Sheet, 0.25-in-Diameter Tup

18-Ply Face Sheet, 0.5-in-Diameter Tup

18-Ply Face Sheet, 1.5-in-Diameter Tup

Figure 4. Residual compression strength versus impact energy.
6

0 0.2 0.4 0.6 0.8 1 1.2 1.4

80,000

70,000

60,000

50,000

40,000

30,000

20,000

10,000

0

IRT Width (in)

Stress (psi)

18-Ply Face Sheet, 0.25-in-Diameter Tup

18-Ply Face Sheet, 0.5-in-Diameter Tup

18-Ply Face Sheet, 1.5-in-Diameter Tup

Figure 5. Residual compression strength versus NDE indication size.

2.2 Open-Hole Compression

OHC testing was performed on 18-ply flat laminates with the Northrop OHC test fixture and 

the Boeing CAI fixture. Figures 6 and 7 are schematics of the two fixtures. The Northrop fixture 

uses specimens 1 in wide by 3 in long. The Boeing fixture, also known as ASTM D7137, Standard 

Test Method for Compressive Residual Strength Properties of Damaged Polymer Matrix Composite 

Plates, utilized 4 in wide by 6 in long specimens.17

Three layups were tested: 18-ply orthotropic, 16-ply quasi-isotropic, and 32-ply unidirectional. 

The 18-ply and unidirectional laminates were tested in both the axial and transverse direction, making a total of five configurations. The 18-ply axially oriented coupons had the majority of fibers oriented in the loading direction with (56/22/22) representing the percentage of plies in the 0-, 45-, and 

90-deg orientations. The 18-ply transverse specimens had a layup of (22/22/56). Quasi-isotropic specimens (25/50/25) were tested for comparison to the highly oriented laminates. Specimens with hole 

sizes ranging from 0.03125 to 1.32 in were tested. The zero-degree unidirectional coupons initially 

failed by in-plane shear parallel to the fibers at the hole edges prior to compressive failure. Thus, the 

ultimate failure strength showed no notch sensitivity since the test is reduced to a net section compression test. Because of the dual failure mode, the zero-degree data are not included in the figures. 

The 90-deg unidirectional coupons exhibited a much smaller notch sensitivity than the laminates. 

Figure 8 illustrates the OHC strength results for 1-in-wide coupons. As noted, prior researchers 

found both the cohesive zone model and average stress criterion to adequately fit OHC strength. The 

average stress criterion is shown as the solid lines in figure 8. 
7

Figure 6. Northrop OHC fixture.16

Top Assembly

Test Specimen

Base Assembly 

Figure 7. Boeing CAI fixture.17
8

140,000

120,000

100,000

80,000

60,000

40,000

20,000

0

Hole Diameter (in)

Failure Stress (psi)

(45,0,–45,0,90,0,0,90,0)s

(–45,90,45,90,0,90,90,0,90)s

(–45,90,45,0)2s

(90)32

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

Figure 8. Laminate OHC results.

When considering fitting the OHC results to a predictive model, one must account for the 

parameters that may be adjusted to fit the data. Past researchers have adjusted the characteristic distance parameter to change the tensile model to a compressive model. The stress concentration factor 

was determined analytically from the elastic properties obtained through classical lamination theory 

and the following:

kT = 1+ 2

EL

ET

− νLT

⎛

⎝

⎜

⎞

⎠

⎟ +

EL

GLT

. (1)

A finite width correction factor as described in Carlsson and Pipes was employed.18 It was 

found that using the same characteristic distance for three different layups did not adequately model 

the data. Furthermore, the average stress criterion best fit all three layups when a constant stress 

concentration factor was used, rather than adjusting based upon elastic properties as calculated from 

equation (1). The only way all three layups could be fit to the average stress criterion was to adjust 

the stress concentration factor and the characteristic distance parameters. 

When one considers all of the OHC strength values are fit to a model based on an unnotched 

strength, the importance of this parameter becomes apparent. It was found that if the OHC strength 

results are normalized to an arbitrary notch size (1/32 in, for example) the model then becomes predictive for unnotched strength. By removing the unnotched data from the equation, the predictive 

model can better fit the notched results of the three layups using only one stress concentration factor 
9

and one characteristic distance (fig. 9). The line representing the average stress criterion utilizes a 

characteristic distance of 0.03 in and a stress concentration factor of 3.

1.6

1.4

1.2

1

0.8

0.6

0.4

0.2

0

Hole Diameter (in)

Failure Stress (Normalized to 1/32-in Hole)

(45,0,–45,0,90,0,0,90,0)s

(–45,90,45,90,0,90,90,0,90)s

(–45,90,45,0)2s

Average Stress Criterion

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

Figure 9. OHC normalized to 1/32-in-diameter hole.

As previously noted, normalizing the OHC strengths to the unnotched strength is dependent 

upon obtaining an accurate unnotched strength. A low test value for unnotched strength will provide 

low estimates for notched properties. 

CAI residual strength has been successfully fitted with a power law of the expression: log(σr) 

=b0+b1log(x), where σr is the residual stress, b0 and b1 are the regression constant and regression 

coefficient, respectively, from the population statistics, and x is the impact energy (or damage size in 

this study).19–21 Interestingly, when the OHC strengths from this study were fitted with the power 

law, the curve fit better than with the average stress criterion for the 18-ply laminate. However, if the 

stress concentration factor for the 18-ply laminate is allowed to be adjusted, the average stress criterion could provide a better fit.

2.3 Testing of Simulated Delaminations and Unbonds

Delaminations and adhesive unbonds within the sandwich were simulated with a double layer 

of FEP tape. These panels were fabricated with hand layup using 12-in-wide prepreg rather than 

fiber placement. Three geometries of FEP tape were placed within the laminate: (1) A 1-in-diameter 

disk, (2) a 1-in square, and (3) a 2-in-long by 1-in-wide rectangle, where the length is perpendicular 

to the loading direction. Two different depths were examined: (1) Between the face sheet and core on 
10

the bag side representing the outer mold line (OML) and (2) within the midplane of the face sheet on 

the OML side. The test results demonstrated that a delamination within the face sheet is much more 

detrimental to the compression strength than an unbond between the core and the face sheet. Table 1 

shows the results of the sandwich edgewise compression testing with simulated delaminations and 

adhesive unbonds.

Table 1. Compression test results of sandwich defect coupons.

Defect

Geometry

Defect

Location

Failure Stress

(lb/in2)

1-in disk

1-in disk

1-in square

1-in square

1- × 2-in rectangular

1- × 2-in rectangular

No defect

0.25-in open hole

0.25-in open hole

Between face sheet and core

Center of OML face sheet

Between face sheet and core

Center of OML face sheet

Between face sheet and core

Center of OML face sheet

NA

Center through hole

Center through hole

104,000

77,000

109,000

70,000

84,000

54,000

101,000

61,000

65,000

2.4 Comparison of Impact Damage, Open-Hole Compression, and Delamination

When comparing the residual compression strength of the three types of stress risers as a�  function of defect size, it is obvious that the OHC bounds the CAI and the simulated delaminations. Figure 9 illustrates: (1) CAI strengths for the 18-, 40-, and the 56-ply sandwich coupons, (2) compression 

strength for the 18-ply simulated delamination sandwich coupons, and (3) the OHC strengths for 

the 18-ply laminate coupons. One issue with a comparison between the OHC and CAI results is the 

difference in structure. The OHC coupons were flat laminates and the CAI coupons were sandwich 

structure. Typically, laminates have a higher cure pressure than sandwich structures. However, for 

this study, all samples had the same cure pressure. A limited set of OHC was performed on the sandwich. Results compared well with the OHC laminate (see table 1 and fig. 10). The initial undamaged, 

or pristine, compression strength for the sandwich was lower than the undamaged strength for the 

laminate. This can be attributed to factors such as the dimpling of the face sheet during co-cure as 

well as the different test geometry. However, once a hole or impact damage was present, the notch 

seemed to overwhelm other factors.

2.5 Nondestructive Evaluation Size Correlation

The majority of dimensional measurements for delamination size were obtained by measurement of the IRT indication size. Thicker laminates; namely, the 56-ply quasi-isotropic laminate, used PAUT for the measurement of delamination size. To compare measurement techniques, 

16 specimens were evaluated with both IRT and PAUT followed by cross-sectional analysis. The 

18-ply sandwich specimens were impacted at ranges from 1 to 10 ft-lb with 0.5- and 1.5-in-diameter 

impactors. The 0.5-in-diameter impactor produced more visible damage than the 1.5-in-diameter 

impactor. However, the NDE indication was typically larger for the larger diameter indenter. The
11

160,000

140,000

120,000

100,000

80,000

60,000

40,000

20,000

0

NDE Width or Hole Diameter (in)

Stress (psi)

18-Ply Laminate—OHC

18-Ply Sandwich—CAI

18-Ply Sandwich—FEP Tape

40-Ply Quasi Sandwich

56-Ply Quasi Sandwich

18-Ply Sandwich—OHC

0 0.5 1 1.5 2 2.5

Figure 10. Comparison of all defect types.

IRT and ultrasonic test (UT) methods correlated well as shown in tables 2 and 3. The cross-sectional 

measurement of damage size was typically the same or smaller than the damage size as measured by 

NDE methods. This can be explained because the cross section was often difficult to obtain at the 

exact center of the impact location, particularly for the lower energy impacts. The cross sections did 

show a difference in damage microstructure as illustrated in appendix B.

Table 2. Comparison of IRT, UT, and cross-sectional analysis for different 

impact energies with a 0.5-in-diameter impactor.

Impact Energy 

(ft-lb)

IRT Width 

(in)

UT Width 

(in)

Cross-Sectional Width 

(in)

0.8

1.8

2.6

3.6

5.4

7.2

9.1

0

0.53

0.62

0.66

0.75

0.75

0.89

0

0.56

0.59

0.65

0.80

0.89

1.00

0

0.425

0.55

0.55

0.74

0.84

0.9
12

Table 3. Comparison of IRT, UT, and cross-sectional analysis for different

impact energies with a 1.5-in-diameter impactor. 

Impact Energy 

(ft-lb)

IRT Width 

(in)

UT Width 

(in)

Cross-Sectional Width 

(in)

0.8

1.8

2.7

3.7

5.5

7.4

9.0

10.5

0

0.58

0.68

0.93

1.01

1.18

1.28

1.27

0

0.71

0.77

0.83

1.03

1.30

1.24

1.45

0

0.26

0.7

0.7

0.95

1.05

1.25

1.45
13

3. CONCLUSIONS

Given that the microstructure of damage induced by impact is dependent upon the size of 

the impactor, the validity of predicting residual strength primarily on damage area is questionable. 

Apparently, there are other factors influencing the residual strength other than the planar size of 

damage.

Impact energy was a better predictor of residual strength than NDE size. This conclusion 

holds regardless of the impactor sizes tested in this study. While the larger diameter impactor produced a larger NDE indication for a given impact energy, the residual strength was the same regardless of impactor size (see fig. 2). However, this affect also occurs where the residual strength begins 

to plateau. So, either method of strength prediction appear appropriate for the material examined in 

this study.

It was found that using a constant stress concentration factor of 3 provided better average 

stress criterion curve fits for the OHC strength data than when the calculated value from laminate 

elastic properties (eq. (1)) was used.

Some densities of core may be inappropriate for co-cure applications at the autoclave pressure 

used in this study. The lowest density core experienced core crushing in some instances. The highest 

density core exhibited node disbonds due to residual stresses induced during autoclave cooldown.
14

APPENDIX A—MECHANICAL TEST DATA

Table 4 includes the damage tolerance test results for the (45,0,–45,0,90,0,0,90,0)s 18-ply face 

sheet with 3.1 lb/ft3 Al core. Only one panel was tested due to face sheet wrinkling and core crushing. Due to processing anomalies, it is not recommended to use such a thin foil thickness core with 

co-cure applications.

Table 4. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results.

Sample Name

Tup Diameter 

(in)

Impact Energy 

(ft-lb)

IRT Width 

(in)

Failure Stress 

(psi)

AC-B1P3C3A-P1-A7-DT-1

AC-B1P3C3A-P1-A7-DT-2

AC-B1P3C3A-P1-A7-DT-3

AC-B1P3C3A-P1-A7-DT-4

AC-B1P3C3A-P1-A7-DT-5

AC-B1P3C3A-P1-A7-DT-6

AC-B1P3C3A-P1-A7-DT-7

AC-B1P3C3A-P1-A7-DT-8

AC-B1P3C3A-P1-A7-DT-9

AC-B1P3C3A-P1-A7-DT-10

AC-B1P3C3A-P1-A7-DT-11

AC-B1P3C3A-P1-A7-DT-12

AC-B1P3C3A-P1-A7-DT-13

AC-B1P3C3A-P1-A7-DT-14

AC-B1P3C3A-P1-A7-DT-15

AC-B1P3C3A-P1-A7-DT-16

AC-B1P3C3A-P1-A7-DT-17

AC-B1P3C3A-P1-A7-DT-18

AC-B1P3C3A-P1-A7-DT-19

AC-B1P3C3A-P1-A7-DT-20

AC-B1P3C3A-P1-A7-EC-10

AC-B1P3C3A-P1-A7-EC-11

AC-B1P3C3A-P1-A7-EC-12

AC-B1P3C3A-P1-A7-EC-13

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

1.53

3.49

4.74

6.32

8.81

9.51

3.75

5.49

7.34

9.24

11.08

13.85

3.33

5.64

7.38

9.10

11.10

13.84

1.71

9.00

3.60

3.68

8.78

9.15

0.2

0.27

0.33

0.37

0.44

0.44

0.18

0.55

0.59

0.69

0.69

0.83

0.4

0.55

0.7

0.9

0.99

1.08

0.3

0.77

0.4

0.44

0.44

0.46

69,548

49,308

46,686

43,576

38,421

41,494

54,882

52,870

46,869

45,029

45,163

45,486

64,533

50,023

52,364

48,152

46,659

50,426

64,358

 –

 –

 –

 –

 –

Tables 5–8 include the damage tolerance test results for the (45,0,–45,0,90,0,0,90,0)s 18-ply 

with 6.1 lb/ft3 Al core. Three sets of panels composed of three lots of core and three lots of film 

adhesive were fabricated in the same autoclave run. The face sheets for all three panels were from the 

same lot of material. 
15

Table 5. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results (batch 1).

Sample Name

Tup Diameter 

(in)

Impact Energy 

(ft-lb)

IRT Width 

(in)

Failure Stress 

(psi)

MDC-B1P1C1A-P1-A7-DT-1

MDC-B1P1C1A-P1-A7-DT-2

MDC-B1P1C1A-P1-A7-DT-3

MDC-B1P1C1A-P1-A7-DT-4

MDC-B1P1C1A-P1-A7-DT-5

MDC-B1P1C1A-P1-A7-DT-6

MDC-B1P1C1A-P1-A7-DT-7

MDC-B1P1C1A-P1-A7-DT-8

MDC-B1P1C1A-P1-A7-DT-9

MDC-B1P1C1A-P1-A7-DT-10

MDC-B1P1C1A-P1-A7-DT-11

MDC-B1P1C1A-P1-A7-DT-12

MDC-B1P1C1A-P1-A7-DT-13

MDC-B1P1C1A-P1-A7-DT-14

MDC-B1P1C1A-P1-A7-DT-15

MDC-B1P1C1A-P1-A7-DT-16

MDC-B1P1C1A-P1-A7-DT-17

MDC-B1P1C1A-P1-A7-DT-18

MDC-B1P1C1A-P1-A7-DT-19

MDC-B1P1C1A-P1-A7-EC-7

MDC-B1P1C1A-P1-A7-EC-9

MDC-B1P1C1A-P1-A7-EC-11

MDC-B1P1C1A-P1-A7-EC-13

0.25

0.25

0.25

0.25

0.25

0.25

0.5

0.5

0.5

0.5

0.5

0.5

1.5

1.5

1.5

1.5

1.5

1.5

NA

0.5

0.5

0.5

0.5

1.55

3.70

5.97

7.76

9.72

12.22

3.82

5.68

7.21

9.83

11.63

14.62

4.30

5.84

7.92

12.56

10.98

14.29

0.00

3.83

3.84

9.83

9.83

0.34

0.41

0.43

0.62

0.64

0.66

0.47

0.51

0.52

0.64

0.66

0.72

0.53

0.6

0.83

0.8

1.07

1.28

–

0.45

0.53

0.54

0.61

73,927

59,209

53,612

47,911

42,125

46,445

58,968

53,515

48,512

46,458

42,537

45,913

54,477

52,164

50,017

48,832

46,997

48,185

108,431*

 –

 –

 –

 –

* Failure by end brooming.

Table 6. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results (batch 2).

Sample Name

Tup Diameter 

(in)

Impact Energy 

(ft-lb)

IRT Width 

(in)

Failure Stress 

(psi)

MDC-B1P2C2A-P1-A7-DT-1

MDC-B1P2C2A-P1-A7-DT-2

MDC-B1P2C2A-P1-A7-DT-3

MDC-B1P2C2A-P1-A7-DT-4

MDC-B1P2C2A-P1-A7-DT-5

MDC-B1P2C2A-P1-A7-DT-6

MDC-B1P2C2A-P1-A7-DT-7

MDC-B1P2C2A-P1-A7-DT-8

MDC-B1P2C2A-P1-A7-DT-9

MDC-B1P2C2A-P1-A7-DT-10

MDC-B1P2C2A-P1-A7-DT-11

MDC-B1P2C2A-P1-A7-DT-12

MDC-B1P2C2A-P1-A7-DT-13

MDC-B1P2C2A-P1-A7-DT-14

MDC-B1P2C2A-P1-A7-DT-15

MDC-B1P2C2A-P1-A7-DT-16

MDC-B1P2C2A-P1-A7-DT-17

MDC-B1P2C2A-P1-A7-DT-18

MDC-B1P2C2A-P1-A7-DT-19

MDC-B1P2C2A-P1-A7-EC-10

MDC-B1P2C2A-P1-A7-EC-11

MDC-B1P2C2A-P1-A7-EC-12

MDC-B1P2C2A-P1-A7-EC-13

0.25

0.25

0.25

0.25

0.25

0.25

0.5

0.5

0.5

0.5

0.5

0.5

1.5

1.5

1.5

1.5

1.5

1.5

NA

0.5

0.5

0.5

0.5

1.63

3.72

5.97

7.76

9.77

12.22

3.81

5.67

7.64

9.82

12.62

14.49

4.02

6.03

7.92

13.61

13.66

14.41

0.00

3.82

3.76

9.83

9.78

0.3

0.42

0.44

0.63

0.69

0.74

0.52

0.64

0.58

0.67

0.72

0.79

0.55

0.72

0.82

1.04

1.26

1.23

–

0.51

0.55

0.5

0.61

61,829

55,937

52,557

49,044

45,132

42,855

56,301

50,026

50,190

44,798

42,685

45,913

55,949

53,276

50,290

48,704

45,102

43,615

72,086*

 –

 –

 –

 –

* Failure by end brooming.
16

Table 7. 18-ply face sheet, 3.1 lb/ft3 core damage tolerance results (batch 3).

Sample Name

Tup Diameter 

(in)

Impact Energy 

(ft-lb)

IRT Width 

(in)

Failure Stress 

(psi)

MDC-B1P3C3A-P1-A7-DT-1

MDC-B1P3C3A-P1-A7-DT-2

MDC-B1P3C3A-P1-A7-DT-3

MDC-B1P3C3A-P1-A7-DT-4

MDC-B1P3C3A-P1-A7-DT-5

MDC-B1P3C3A-P1-A7-DT-6

MDC-B1P3C3A-P1-A7-DT-7

MDC-B1P3C3A-P1-A7-DT-8

MDC-B1P3C3A-P1-A7-DT-9

MDC-B1P3C3A-P1-A7-DT-10

MDC-B1P3C3A-P1-A7-DT-11

MDC-B1P3C3A-P1-A7-DT-12

MDC-B1P3C3A-P1-A7-DT-13

MDC-B1P3C3A-P1-A7-DT-14

MDC-B1P3C3A-P1-A7-DT-15

MDC-B1P3C3A-P1-A7-DT-16

MDC-B1P3C3A-P1-A7-DT-17

MDC-B1P3C3A-P1-A7-DT-18

MDC-B1P3C3A-P1-A7-DT-19

MDC-B1P3C3A-P1-A7-EC-10

MDC-B1P3C3A-P1-A7-EC-11

MDC-B1P3C3A-P1-A7-EC-12

MDC-B1P3C3A-P1-A7-EC-13

0.25

0.25

0.25

0.25

0.25

0.25

0.5

0.5

0.5

0.5

0.5

0.5

1.5

1.5

1.5

1.5

1.5

1.5

NA

0.5

0.5

0.5

0.5

1.62

3.69

5.93

7.76

9.73

12.22

3.82

5.59

7.64

9.82

11.61

14.66

3.74

6.86

7.92

11.06

14.23

14.36

0.00

3.78

3.76

9.83

9.75

0.36

0.41

0.45

0.62

0.65

0.71

0.46

0.64

0.54

0.64

0.68

0.76

0.53

0.63

0.81

0.99

1.28

1.19

–

0.47

0.58

0.57

0.67

65,814

56,379

51,614

47,501

47,239

43,913

56,238

49,312

46,685

43,032

44,627

40,961

59,296

53,505

49,126

46,782

47,355

42,294

105,888*

 –

 –

 –

 –

* Failure by end brooming.

Table 8. 56-ply face sheet, 12 lb/ft3 core damage tolerance results.

Sample Name

Tup Diameter 

(in)

Impact Energy 

(ft-lb)

PAUT Width 

(in)

Failure Stress 

(psi)

HDC-B4P1C1A-P1-A7-EC-1

HDC-B4P1C1A-P1-A7-EC-2

HDC-B4P1C1A-P1-A7-EC-3

HDC-B4P1C1A-P1-A7-EC-4

HDC-B4P1C1A-P1-A7-EC-5

HDC-B4P1C1A-P1-A7-EC-6

HDC-B4P1C1A-P1-A7-EC-7

HDC-B4P1C1A-P1-A7-EC-8

HDC-B4P1C1A-P1-A7-EC-9

HDC-B4P1C1A-P1-A7-EC-10

HDC-B4P1C1A-P1-A7-EC-11

HDC-B4P1C1A-P1-A7-EC-12

HDC-B4P1C1A-P1-A7-EC-13

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

0.5

20.89

30.92

38.36

47.38

54.06

54.14

46.77

37.87

30.74

20.41

17.47*

11.40

67.60*

0.56

1.24

1.06

1.21

1.65

1.65

1.77

1.59

1.00

0.59

0.50

0.62

2.12

76,845

49,303

45,841

41,707

40,218

39,977

41,496

43,429

52,410

73,616

>78,000

>78,001

35,981

* Values estimated from drop height due to instrumentation error.

Table 8 includes the damage tolerance test results for the (45,90,–45,0)7s 56-ply face sheet with 

12 lb/ft3 Al core. Three sets of panels composed of three lots of core and three lots of film adhesive 

were fabricated in the same autoclave run. The face sheets for all three panels were from the same 

lot of material. Only one of the three panels was machined into coupons because two of the panels 
17

exhibited excessive core node disbonds as evidenced by NDE. Cross-sectional analysis confirmed the 

node disbonds. Subsequent stress analysis confirmed high residual stresses incurred within the Al 

core during autoclave cooldown as a result of coefficient of thermal expansion mismatch between 

the face sheet and core. Tensile stress on the core node bondline is a function of foil thickness. Finite 

element analysis predicted twice the tensile stress in the node bond for the 12 lb/ft3 core compared 

to the 6.1 lb/ft3 core. Average tensile stress estimates across the node bondline were 2,300 psi average 

for 12 lb/ft3 core and 1,100 psi for 6.1 lb/ft3 core. When modeled with a fillet at each bend in the core, 

the average tensile stresses were reduced. However, the discrepancy in magnitude between the different foil thicknesses remained. The failure illustrates the importance of process development as well 

as multiple batch testing—batch 1 core did not exhibit the disbonds. It is recommended that Al core 

with a foil thickness as great as the 12 lb/ft3 material not be used in co-cure applications with carbon 

fiber face sheets. Figure 11 illustrates the stress model on the node bondline.

(a) (b)

Figure 11. Finite element analysis model of residual thermally induced stresses in 

(a) 12 lb/ft3 core and (b) 6.1 lb/ft3 core.
18

APPENDIX B—CROSS SECTIONS OF IMPACT DAMAGE

Figures 12–24 are cross sections of the NDE correlation panel. The 18-ply sandwich panels 

were impacted at energy levels from 1 to 10 ft-lb. A replicate of each impact was conducted with both 

0.5- and 1.5-in-diameter impactors. Two important observations regarding the induced damage were 

observed. First, the larger diameter impactor created a longer delamination than the smaller diameter impactor. Second, the damage induced by the smaller diameter impactor appeared to be more 

widespread through the thickness of the laminate than the damage induced by the larger diameter 

impactor.

Figure 12. 1 ft-lb, 0.5-in-diameter impactor (no delamination).

Figure 13. 2 ft-lb, 0.5-in-diameter impactor.

Figure 14. 2 ft-lb, 1.5-in-diameter impactor.
19

Figure 15. 3 ft-lb, 0.5-in-diameter impactor.

Figure 16. 3 ft-lb, 1.5-in-diameter impactor.

Figure 17. 4 ft-lb, 0.5-in-diameter impactor.

Figure 18. 4 ft-lb, 1.5-in-diameter impactor.
20

Figure 19. 6 ft-lb, 0.5-in-diameter impactor.

Figure 20. 6 ft-lb, 1.5-in-diameter impactor.

Figure 21. 8 ft-lb, 0.5 in-diameter impactor.

Figure 22. 8 ft-lb, 1.5 in-diameter impactor.
21

Figure 23. 10 ft-lb, 0.5 in-diameter impactor.

Figure 24. 10 ft-lb, 1.5-in-diameter impactor.
22

REFERENCES

 1. Dorey, G.: “Impact and Crashworthiness of Composite Structures,” Structural Impact and Crashworthiness, Vol. 1, J. Morton (ed.), Elsevier Applied Science Publishers, London and New York, 

1984.

 2. Cantwell, W.; Curtis, P.; and Morton, J.: “Post-Impact Fatigue Performance of Carbon Fibre 

Laminates with Non-Woven and Mixed-Woven Layers,” Composites, Vol. 14, No. 3, pp. 301–305, 

July 1983.

 3. Dorey, G.: “Impact Damage in Composites—Development, Consequences and Prevention,” 

Proceedings, Sixth International Conference on Composite Materials, Second European Conference on Composite Materials, Vol. 3, pp. 3.1–3.26, 1986.

 4. Morton, J.; and Godwin, E.W.: “Impact Response of Tough Carbon Fibre Composites,” Compos. Struct., Vol. 13, pp. 1–19, 1989.

 5. Bishop, S.M.: “The Mechanical Performance and Impact Behavior of Carbon-Fibre Reinforced 

PEEK,” Compos. Struct., Vol .3, pp. 295–318, 1985.

 6. Cantwell, W.J.; Curtis, P.T.; and Morton J.: “An Assessment of the Impact Performance of CFRP 

Reienforced with High-Strain Carbon Fibers,” Compos. Sci. Technol., Vol. 25, pp. 133–148, 1986.

 7. Whitney, J.M.; and Nuismer, R.J.: “Stress Fracture Criteria for Laminated Composites Containing Stress Concentrations,” J. Compos. Mater., Vol. 8, pp. 253–265, 1974.

 8. Pipes, B.R.; Wetherhold, R.C.; and Gillespie, J.W., Jr.: “Notched Strength of Composite Materials,” J. Compos. Mater., Vol. 13, pp. 148–160, April 1979.

 9. Soutis, C.; Fleck, N.A.; and Smith, P.A.: “Failure Prediction Technique for Compression Loaded 

Carbon Fibre-epoxy Laminate With Open Holes,” J. Compos. Mater., Vol. 25, pp. 1476–1498, 

November 1991.

10. Chen, P.; Shen, Z.; and Wang, J.Y.: “Prediction of the Strength of Notched Fiber-Dominated 

Composite Laminates,” Compos. Sci. Technol., Vol. 61, No. 9, pp. 1311–1321, July 2001.

11. Callus, P.J.: “The Effects of Hole-size and Environment on the Mechanical Behaviour of 

a Quasi-isotropic AS4/3501-6 Laminate in Tension, Compression and Bending,” Austrailian 

Government Department of Defense, DSTO-TR-2077, November 2007.
23

12. Byers, B.A.: “Behavior of Damaged Graphite/Epoxy Laminates Under Compressive Loading,” 

NASA Contractor Report 159293, Marshall Space Flight Center, AL, August 1980.

13. Williams, J.G.: “Effect of Impact Damage and Open Holes on the Compressive Strength of 

Tough Resin/High Strain Fiber Laminates,” NASA Technical Memorandum 85756, Marshall 

Space Flight Center, AL, February 1984.

14. Chen, P.; Shen, Z.; and Wang, J.-Y.: “A New Method for Compression After Impact Strength 

Prediction of Composite Laminates,” J. Compos. Mater., Vol. 36, No. 5, pp. 589–610, 2002.

15. Edgren, F.; Asp, L.E.; and Bull, P.H.: “Compressive Failure of Impacted NCF Composite 

Sandwich Panels—Characterization of the Failure Process,” J. Compos. Mater., Vol. 38, No. 6, 

pp. 495–514, 2004.

16. Adams, D.F.: “Open Hole Compression Testing,” High Perform. Compos., pp. 12–13, March 2005.

17. ASTM D7137, Standard Test Method for Compressive Residual Strength Properties of Damaged Polymer Matrix Composite Plates, 2007.

18. Carlsson, L.A.; and Pipes, B.R.: Experimental Characterization of Advanced Composite Materials, 2nd Ed., Technomic Publishing Company, Inc., pp. 129–141, 1997.

19. Caprino, G.: “Residual Strength Prediction of Impacted CFRP Laminates,” J. Compos. Mater.,

Vol. 18, pp. 508–518, 1984.

20. Jenq, S.T.; Wang, S.B.; and Sheu, L.T.: “A Model for Predicting the Residual Strength of GFRP 

Laminates Subject to Ballistic Impact,” J. Reinf. Plast. Compos., Vol. 11, pp. 1127–1141, 1992.

21. Nettles, A.T.; and Jackson, J.R.: “Compression After Impact Testing of Sandwich Composites 

for Usage on Expendable Launch Vehicles,” J. Compos. Mater., Vol. 44, pp. 707–738, 2010.
24

REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining 

the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing 

this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operation and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid 

OMB control number.

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY) 2. REPORT TYPE 3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S) 5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION

 REPORT NUMBER

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITOR’S ACRONYM(S)

11. SPONSORING/MONITORING REPORT NUMBER

12. DISTRIBUTION/AVAILABILITY STATEMENT

13. SUPPLEMENTARY NOTES

14. ABSTRACT

15. SUBJECT TERMS

16. SECURITY CLASSIFICATION OF:

a. REPORT b. ABSTRACT c. THIS PAGE

17. LIMITATION OF ABSTRACT 18. NUMBER OF

 PAGES

19a. NAME OF RESPONSIBLE PERSON

19b. TELEPHONE NUMBER (Include area code)

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18

Comparison of Open-Hole Compression Strength and Compression

After Impact Strength on Carbon Fiber/Epoxy Laminates 

for the Ares I Composite Interstage

A.J. Hodge, A.T. Nettles, and J.R. Jackson

George C. Marshall Space Flight Center

Marshall Space Flight Center, AL 35812

National Aeronautics and Space Administration

Washington, DC 20546–0001

Unclassified-Unlimited

Subject Category 24 

Availability: NASA CASI (443–757–5802)

Prepared by the Materials and Processes Laboratory, Engineering Directorate

M–1309

Technical Publication

NASA/TP—2011–216460

composites, sandwich structures, impact damage, open-hole compression, residual strength

01–03–2011

UU 32

NASA

U U U

Notched (open hole) composite laminates were tested in compression. The effect on strength of various sizes of through holes was examined. Results were compared to the average stress criterion model. 

Additionally, laminated sandwich structures were damaged from low-velocity impact with various impact 

energy levels and different impactor geometries. The compression strength relative to damage size was 

compared to the notched compression result strength. Open-hole compression strength was found to 

provide a reasonable bound on compression after impact.

STI Help Desk at email: help@sti.nasa.gov

STI Help Desk at: 443–757–5802
The NASA STI Program...in Profile

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientific and Technical Information (STI) 

Program Office plays a key part in helping NASA 

maintain this important role.

The NASA STI Program Office is operated by 

Langley Research Center, the lead center for 

NASA’s scientific and technical information. The 

NASA STI Program Office provides access to 

the NASA STI Database, the largest collection of 

aeronautical and space science STI in the world. 

The Program Office is also NASA’s institutional 

mechanism for disseminating the results of its 

research and development activities. These results 

are published by NASA in the NASA STI Report 

Series, which includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant 

phase of research that present the results of 

NASA programs and include extensive data 

or theoretical analysis. Includes compilations 

of significant scientific and technical data 

and information deemed to be of continuing 

reference value. NASA’s counterpart of peerreviewed formal professional papers but has less 

stringent limitations on manuscript length and 

extent of graphic presentations.

• TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or of 

specialized interest, e.g., quick release reports, 

working papers, and bibliographies that contain 

minimal annotation. Does not contain extensive 

analysis.

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientific and technical conferences, 

symposia, seminars, or other meetings sponsored 

or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientific, technical, 

or historical information from NASA programs, 

projects, and mission, often concerned with 

subjects having substantial public interest.

• TECHNICAL TRANSLATION. 

English-language translations of foreign 

scientific and technical material pertinent to 

NASA’s mission.

Specialized services that complement the STI 

Program Office’s diverse offerings include creating 

custom thesauri, building customized databases, 

organizing and publishing research results...even 

providing videos.

For more information about the NASA STI Program 

Office, see the following:

• Access the NASA STI program home page at 

<http://www.sti.nasa.gov>

• E-mail your question via the Internet to 

<help@sti.nasa.gov>

• Fax your question to the NASA STI Help Desk 

at 443 –757–5803

• Phone the NASA STI Help Desk at 

443 –757–5802

• Write to:

NASA STI Help Desk

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076–1320
National Aeronautics and

Space Administration

IS20

George C. Marshall Space Flight Center

Marshall Space Flight Center, Alabama

35812