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i

NASA/TP—2010–216434

Simplification of Fatigue Test Requirements 

for Damage Tolerance of Composite 

Interstage Launch Vehicle Hardware

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

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

June 2010

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>

Acknowledgments

This work was supported by NASA under the auspices of the Upper Stage Program Office 

at Marshall Space Flight Center.
iii

TABLE OF CONTENTS

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

2. BACKGROUND ............................................................................................................... 3

2.1 Effects of Fatigue ......................................................................................................... 4

2.2 Load Spectrum ............................................................................................................ 7

2.3 Limit Load (Stress) Value in Design and the Need for Compression Fatigue Data ...... 8

3. EXPERIMENTAL ............................................................................................................ 10

3.1 Static Testing ............................................................................................................... 10

3.2 Cyclic Testing ............................................................................................................... 11

3.3 Materials and Specimens ............................................................................................. 11

3.4 Nondestructive Evaluation Testing .............................................................................. 11

4. TEST RESULTS ................................................................................................................ 13

4.1 Impact Testing ............................................................................................................. 13

4.2 Fatigue Testing ............................................................................................................. 13

4.3 Residual Strength Testing ............................................................................................ 17

5. DISCUSSION ................................................................................................................... 20

6. CONCLUSIONS ............................................................................................................... 23

REFERENCES ...................................................................................................................... 24
iv

LIST OF FIGURES

1. Damage tolerant coupon test schematics ................................................................. 1

2. Schematic of impact damage growth showing three phases during fatigue testing ... 6

3. Load spectrum for Ares I interstage ........................................................................ 7

4. Simplified load spectrum for the Ares I interstage ................................................... 8

5. Residual strength (CAI) curve for the Ares I interstage acreage material ................. 10

6. Typical IRT indication of impact-damaged laminate .............................................. 12

7. Examples of the visual damage caused by impacts .................................................. 15

8. Plot of results from table 5 ...................................................................................... 17

9. IRT results of a typical specimen (no growth) and the two specimens 

 that exhibit damage growth after 10,000 cycles ........................................................ 18

10. Fatigue life data from this and other studies ............................................................ 20
v

LIST OF TABLES

1. Summary of compression fatigue after impact data from the open literature ............. 5

2. Impact results for specimens to be fatigue tested ........................................................ 14

3. B-Basis summary of the first series of fatigue after impacts tests: 1,000 cycles 

per stress amplitude ................................................................................................... 15

4. A-Basis summary of the first series of fatigue after impacts tests: 1,000 cycles 

per stress amplitude ................................................................................................... 16

5. Summary of the second series of fatigue after impacts test ........................................ 16

6. Static compressive breaking stress of specimens that survived all fatigue series ......... 19
vi

LIST OF ACRONYMS AND SYMBOLS

BVID barely visible impact damage

CAI compression after impact

IRT infrared thermography

LL limit load

MSFC Marshall Space Flight Center

NDE nondestructive evaluation

S-N stress versus number of cycles to failure curve

TP Technical Publication

UL ultimate load
vii

NOMENCLATURE

n design load limit to flight

sAVG breaking stress

sLL limit load stress

sUL ultimate load stress
viii
1

TECHNICAL PUBLICATION

Simplification of Fatigue Test Requirements for Damage Tolerance

of composite interstage Launch Vehicle Hardware

1. INTRODUCTION

One of the fracture control requirements for composite hardware on manned launch vehicles is that damage will not propagate and result in failure. MSFC–RQMT–3479, Fracture Control 

Requirements for Composite and Bonded Vehicle and Payload Structures, requires coupon level 

curves to be generated for (a) residual compression strength as a function of damage size for a multiple number of load cycles, (b) life curves for multiple damage sizes, and (c) threshold compression 

strain for multiple damage sizes.1 Examples of these three types of curves are shown in figure 1.

(a) Residual Strength Curves (b) Life Curves

(c) Threshold Strain Curves

Residual Compression Strength 

Damage Growth 

‘No-Growth Threshold Strain’ 

for Largest Damage Tested 

Increasing Damage Size 

‘No-Growth Threshold Strain’ 

for Smallest Damage Tested 

Applied Cyclic Strain for 1 Million Cycles 

Flaw/Damage Size Cycles to Failure (life) 

Increasing Damage Size 

Compression Load Level 

Increasing Number 

of Load Cycles 

Figure 1. Damage tolerant coupon test schematics.1
2

Figure 1(a) is based on the assumption that fatigue loading will decrease the residual compression strength of a composite laminate for a given damage size. For example, the upper curve may 

represent the static case most often published in literature. The curves below that are plots of compression strength versus damage size after the specimens have been fatigue loaded and are assumed 

to be lower than the static case (i.e., it is assumed fatigue will be detrimental to the compression 

strength of a laminate with a flaw).

Figure 1(b) illustrates the typical stress versus number of cycles to failure (S-N) curve found 

in the literature for metallic materials in tension. These curves are based on the assumption that, as 

the cyclic compressive load increases, the number of cycles to failure will decrease and the degradation rate is the same regardless of damage size.

Figure 1(c) determines the compressive strain level at which damage will not grow after 

1 million cycles. The value of 1 million is arbitrarily chosen as runout, after which the component 

is expected to be retired from service. This figure assumes a very long lifetime of the part (1 million cycles at a significant strain). Furthermore, this figure assumes that damage will grow and that 

growth can be measured with some sort of nondestructive technique before catastrophic failure.

To generate these curves, a large number of specimens needs to be tested under fatigue, which 

is time consuming. Since composite laminates behave differently in fatigue than metallic structures, 

it is worth noting the differences to examine if the construction of these curves can be simplified, 

thus reducing testing time and costs. In addition, specifics related to launch vehicles (as opposed to 

aircraft) will be noted so that they can be used to further simplify obtaining the necessary information to design a damage-tolerant launch vehicle structure.
3

2. BACKGROUND

The effects of cyclic loading on the structure’s material and the life cycle of the structure must 

be taken into account when the structure is being assessed for fatigue. Due to shorter operational 

times, expendable rocket components will experience fewer cyclic loads than fixed wing aircraft components, and fixed wing aircraft will experience fewer cycles than rotating hardware on rotorcraft. 

This will determine when runout is considered to have occurred and will be discussed more in section 2.2. The next section presents a review of the compressive cyclic loading aspects from past studies in the open literature and what may be expected in this Technical Publication (TP).

Since the terms ‘breaking,’ ‘ultimate,’ and ‘limit’ loads are used differently according to industry practice, the terms used in the remainder of this TP will be defined below. The terms ‘stress’ and 

‘strength,’ as used in this study, will also be defined.

Breaking load is simply the load at which catastrophic failure of the laminate occurs. For 

a test series with 15 specimens, there will be 15 unique and distinct breaking loads.

Average breaking load is the statistic used as an estimator for population mean and is equal 

to the sum of individual breaking loads divided by the number of individual loads. The confidence 

band on the estimate of mean improves with increasing replications and reducing variance. The average breaking load can also be referred to as ‘expected’ breaking load.

Ultimate load (UL) is a statistically based value above which a certain percentage of specimens are expected to survive with a given confidence level. Typically, a ‘B-Basis value’ means a load 

at which there is a 95% confidence level that at least 90% of the specimens will survive. This value 

is at the lower tail of the breaking load probability distribution. An ‘A-Basis value’ means a load at 

which there is a 95% confidence level that at least 99% of the specimens will survive.

Limit load (LL) is the maximum load expected on a structure during its design life. In order 

for a structure to have a positive margin of safety, the LL must be less than the UL divided by a factor of safety. For the aerospace components mentioned in this study, the factor of safety is 1.4. Thus, 

LL must not exceed 71.4% of the UL.

Stress is load divided by the initial cross-sectional area of a specimen. The cross-sectional 

area of all specimens in this study was the same, so transforming between load and stress is straightforward. The symbols for average breaking stress, ultimate stress, and limit stress are σAVG, σUL, and 

σLL, respectively.

In this TP, ‘strength’ is the σAVG of a specimen and, more specifically, compression after 

impact (CAI) strength is the compressive σAVG of a specimen that has experienced a foreign object 

impact event.
4

2.1 Effects of Fatigue

A review of the literature was performed on previous work regarding the detection, measurement, and consequences of damage growth due to compressive fatigue loading. Tension fatigue and/

or the use of artificially induced delaminations, rather than impact damage, where not considered in 

order to keep the literature review as close to the planned experimental part of this study as possible.

Table 1 summarizes the pertinent results from a variety of compression-fatigue-after-impact 

test programs conducted in the past. The majority of past studies examined specimens with barely 

visible impact damage (BVID). Most components are required and designed to carry the UL with 

BVID. Therefore, past studies have set BVID as the threshold for allowable damage size. Table 1 

presents compression load as a percentage of the average (or expected) breaking load of a sample 

with the same level of damage so a direct comparison between studies could be made. In some cases, 

this percentage had to be calculated from the data since compression fatigue amplitude as a percentage of average strength was not always explicitly given and is detailed in the footnotes below the 

table. The cycles to failure column presents the highest and lowest number of cycles to failure, which 

typically corresponds to the lowest and highest loads in the previous column. It was of interest to 

see if other experimental programs found damage growth (using any nondestructive technique), so 

a ‘growth detected’ column was added.

Since two of the studies used variable amplitude fatigue spectrums, the maximum amplitude 

in the spectrum is presented. The corresponding cycles are the numbers of cycles to this maximum 

load, while no other load excursions are counted. The results are presented as the number of ‘programs’ that survived. Each program had 200 excursions to the maximum amplitude load and the 

necessary conversions were made to fit the data into table 1.

Only three studies in table 1 published data for compression stress amplitudes <60% of the 

predicted static CAI stress. In these three studies, the specimens all experienced no deleterious effects 

when fatigued <60% of the static strength value. This is indicative of the good fatigue properties of 

composites in which the fibers have a crack retardation effect for the multitude of matrix cracks that 

form during fatigue as demonstrated by Reifsnider19,20 and Kunz.21 One of the earliest realizations 

noted over two decades ago was that ≈60% of static CAI failure stress is the ‘lower boundary of 

a fatigue failure regime.’16 The lower bound of 60% of static CAI failure stress was also suggested in 

reference 8, which found that, in addition to this lower bound, compressive stress ranges (difference 

between consecutive stress levels) on the order of 20% could also be eliminated from the fatigue load 

spectrum since they have no deleterious effects. Also of note are the results of fatigue data published 

in which damage initiated from the free edge of impact damaged test specimens22 or cutouts23 before 

any growth was seen in the impact damage area, indicating that free edges may be more of a fatigue 

concern than impact damage.

When examined on a basis of damage severity, the fatigue life of laminates tends to improve 

(i.e., the S-N curve gets flatter as more damage is induced).2,13,15 The static compressive strength 

drops with increasing damage but the effects of fatigue are diminished. This leads to the conclusion 

that the undamaged case will be the most conservative when examining the fatigue life of expendable 

structures in compression.
5

Table 1. Summary of compression fatigue after impact data from the open literature.

Type Specimen

Compression 

Amplitude Range* Cycles to Failure 

Compression Strength 

if Survived Fatigue* Growth Detected

48-ply AS4/3501-62 88–97 496,509–533 N/A No

8-ply sandwich AS4/3501-62 73–112** ∞***–5 (∞= 60,000) ≈109† 6 of 17

48- and 42-ply AS4/3501-6 (blunt impact)3 71–81 ∞ (∞= 1 million) N/A 1 of 4

48- and 42-ply AS4/3501-6 (penetration impact)3 64–73 ∞ (∞= 1 million) N/A 4 of 4

48-ply IM7/HTAC (thermoplastic)4 60–78 2 million–200 N/A No

48-ply IM7/BMI4 71–90 1.5 million–1,000 N/A Yes

16-ply IM7/PEEK5 60 ∞ (∞= 1 million) ≈100 No

16-ply IM7/epoxy5 60 ∞ (∞= 1 million) ≈110 3 of 17

T50/934 6-ply tubes6 75 ∞–8,931 (∞= 1 million) 100 No

48-ply IM7/8551-77 68–75 301,000–7,050 N/A N/A

32-ply AS4/3501-68 30–80 ∞–136 (∞= 1 million) 105 Yes above 70%

21-ply T300/9149 67–97 781,321–2 N/A Yes near failure

16-ply AS4/855210 75–85 ∞–≈1,000 (∞= 1 million) N/A Yes at 85%

4-ply sandwich carbon/epoxy fabric11 20–90 ∞–≈2 (∞= 150,000) 70–115 N/A

48-ply HTA/637612 60–90 ∞–120 (∞= 1 million) N/A Yes near failure

24-ply UT500/QU135-197A13 79–98 ≈5,000–150 N/A Yes

32-ply AS4/PEEK13 66–98 ≈900,000–60 N/A Very Little

2-mm-thick T300/vinylester fabric14 60–80 ≈1 million–300 N/A No

16-ply carbon/epoxy5‡,15 62–80 ∞–9,000 (∞= 1 million) N/A No

56-ply XAS/914C16 88§ 107,600–6,800 N/A Yes

56-ply XAS/914C17 88§ 30,000–600 80–89 Yes

32-ply IM7/3501-618 74–84 ∞ –33,500 (∞= 1 million) 121–134 No

32-ply IM7/8551-718 33–78 ∞ –4,330 (∞= 1 million) 90–109 No

 *Values given as a percentage of the average static CAI strength.

 **Values can be >100% due to scatter in the static data not being accounted for.

***∞ implies runout.

 †Growth detected on all samples tested post fatigue.

 ‡Details on material not given.

 §Maximum value of spectrum loading.

In the only study that saw any damage growth at a load value <60% of the static strength 

value, the specimens were first preloaded in a block loading at high fatigue amplitude >60% in order 

to initiate damage growth.22 However, according to reference 1, loading was on a ‘failed’ specimen.

Another interesting result seen in table 1 is that the strength after fatigue was the same or 

higher than the expected static CAI strength in most of the studies in which the residual compression 

strength of fatigued specimens were tested. This was most pronounced in reference 18 for a brittle 

matrix composite. The authors in this study did not conclude a physics based reason for this; however, they did mention that such results were found for tension testing of open hole laminates.24,25

These strength improvements due to fatigue were seen even when damage growth had occurred.
6

For the studies in which damage growth did occur, the growth was seen to occur near the 

end of the load spectrum with damage rapidly increasing as failure was reached. This indicates that 

determination of a ‘no-growth’ threshold strain will be difficult to determine. A thorough treatment 

of the evolution of impact damage growth is given by Chen et al. in which the damage growth is seen 

to occur in three distinct phases as shown in figure 2.10 By utilizing real-time acoustography, more 

detailed information about the damage size could be made without unloading and reloading specimens. For constant amplitude fatigue loading at amplitude 75% or less of average CAI strength, no 

growth occurred before runout, which was considered 1 million cycles. For specimens that did demonstrate damage growth (80% and 85% of average CAI strength) a three-phase growth pattern was 

observed. Phase 1 shows rapid growth before slowing dramatically. During this period, the increase 

in impact damage size is due to a change in shape of the impact damage in which the planar area of 

damage simply achieves smoother edges.

Phase 1 Phase 3

20

10 

Percentage Growth of Damage Width

Percentage of Life 

10 90 100

Phase 2 

Figure 2. Schematic of impact damage growth showing three phases during fatigue testing.

Phase 2 consists of the vast majority of the life of the specimen and is a period of very slow 

growth until phase 3 is reached. At this point, the growth of the impact damage area becomes rapid 

near the end of the life of the specimen. Phase 3 growth, which is easier to detect using more conventional nondestructive evaluation (NDE) techniques, has been seen in other studies that do not have 

the fidelity to observe the phase 2 slow growth.8,9,16 A study that used digital speckle photometry 

also showed three regions of damage growth like that illustrated in figure 2.12 These results indicate 

that determination of a ‘no-growth’ threshold strain is more difficult to define than it is for metals. 

‘No detrimental growth’ is a more accurate term for composites since rounding (the formation of 

smoother areas of the planar area of damage) of the damage area can be beneficial to residual compression strength of impact damaged laminates.2,5,8,11,18

Reference 1 does not mention tension or in-plane shear after impact for launch vehicle 

hardware that is driven by either of these two types of loading. It has been shown that static 
7

compression after impact strength degrades more than static tension after impact strength for 

a given level of damage severity;26–31 therefore, compression of a damaged laminate will be of more 

criticality than tension of a laminate with the same level of impact damage. The tensile aspects of 

a structure designed primarily for tensile loads, such as a rocket motor case, should be considered 

since tensile and compressive failure modes are different and independent in composite laminates.32

2.2 Load Spectrum

The number of cycles, amplitude, and R-ratio at which to test specimens after impact will 

obviously depend upon the application and hardware in question. Virtually all of the current knowledge and literature on this subject is from the aircraft and rotorcraft industry. Conservative measures 

are used because the manufacturers of airplanes and helicopters do not know the specific load profile 

of each flight or how long one flight may last. Since expendable launch vehicles have a value of one 

flight/life, requiring four lifetimes to account for material variability and a given number of excursions to design limit load per flight (n) results in 4n cycles at LL to demonstrate fatigue life. Since the 

part is not to be used again, phase 2 nondetrimental damage growth is of no practical consequence.

The order of applied loads is not critical for spectrum fatigue loading of composites.8,16,17

Thus, to facilitate fatigue testing, the highest loads are tested first to accelerate the testing process. If 

the highest blocks of loads do not cause degradation or failure, the lower loads are removed from the 

spectrum.33

As an example, the load spectrum for the composite interstage for the Ares I launch vehicle 

is shown in figure 3. As can be seen, the structure will experience one cycle at LL and two cycles at 

a value >50% of LL, thus simplifying the fatigue spectrum as shown in figure 4. Compressive Stress 1/2 LL

1/4 LL

1/6 LL

All Loads Compressive

560,000 Cycles

1 Day Fueled on Launch Pad

One Lifetime

Liftoff Ascent

70

10 1047

Jettison

LL

Cycles

Figure 3. Load spectrum for Ares I interstage.

A launch vehicle structure will typically only experience a few cycles near LL values. Furthermore, it has been demonstrated that detrimental damage growth is not observed at load levels in 
8

1/2 LL

All Loads Compressive

LL

Cycles

One Lifetime

Figure 4. Simplified load spectrum for the Ares I interstage.

a structure designed to accommodate damage. Therefore, it is not necessary to fully develop the 

curves shown in figure 1 for single-use launch vehicles. By treating the structure as statically loaded, 

the curves in figure 1 can be greatly simplified.

2.3 Limit Load (Stress) Value in Design and the Need for Compression Fatigue Data

The structural LL is the maximum load that is predicted for the structure to withstand during 

operating conditions. Design teams use this value, coupled with the material capability, to develop 

hardware sizing information. Typically, the structural LL is used, along with the required safety factor, to determine the UL for use in design assessments or hardware test programs. The design UL is 

then compared to the material UL capability as the primary measure for structural capability.

Alternatively, a value for material limit stress (or load) may be developed and used as 

a direct comparison to the structural limit stress (or load). The material LL is the measure of ultimate strength for the required safety factor and will be designated as sLL in this TP. For manned 

spaceflight hardware, the required ultimate factors are 1.4 and 2 for net section and discontinuities, 

respectively. Ideally, the design is optimized to minimize weight such that the structural limit stress is 

only slightly lower than the material limit stress.

The material sUL calculated is usually a ‘basis’ value at which only 10% of the specimens 

(B-Basics) are expected to fail. For the composite laminate planned for the acreage of Ares I Upper 

Stage Interstage, this basis value for CAI loading is ≈86% of the average CAI strength (for any damage severity).34 Thus, the material limit stress (sLL) for net section acreage capability in this particular case is calculated by

sLL=0.86/1.4 (CAIAVG)=61.4% of average static CAI strength .	 (1)
9

Once the safety factors and basis values are accounted for, the sLL is 61.4% of the average 

static CAI strength, and stresses (loads) above this value result in a negative margin of safety. If one 

accepts the premise that compression fatigue load values <60% of the average static CAI strength are 

not significant for an impact-damaged laminate,8,16 then it appears that compression-compression 

fatigue testing becomes unnecessary. Moreover, reference 10 indicates no damage growth prior to 

1 million cycles at constant amplitude compression fatigue loads below 75% of the average CAI 

static strength. These data are especially relevant for expendable launch vehicles that experience 

a limited number of fatigue cycles. Exceptions may exist for rotating parts or hardware subjected to 

high vibratory frequencies resulting in a substantial number of fatigue cycles (>1 million cycles).

While most of the CAI strength data are from tests on coupon-size specimens, the static 

residual strength has been shown to be insensitive to specimen size and curvature.35 No observed 

growth on numerous subscale and two full-scale test articles of the Boeing 777 empennage structure 

were reported.36
10

3. EXPERIMENTAL

The premise that fatigue testing will reduce to the static case for compression-dominated 

loads on expendable launch vehicles was examined using a series of tests at NASA Marshall Space 

Flight Center (MSFC). CAI strength was the property of interest in this study since (1) compression 

strength values are most affected by impact damage and (2) the hardware in question, an interstage, 

is a heavily compression-dominated structure.

3.1 Static Testing

A CAI map was generated in a previous study in which static residual strength as a function 

of various impact parameters was presented.37 The results showed that damage width, as measured 

by NDE, and incident energy of impact were two parameters that gave the most reliable residual 

compression strength data. The correlation between these measures of damage severity and visual 

damage is ongoing as BVID is a nonquantitative, subjective measurement. BVID must be defined for 

a given structure. However, this study is only concerned with post impact fatigue behavior that can 

be presented as a function of static residual strength for any given level of damage. The results from 

the static strength study are presented in figure 5 for reference purposes.

Predicted Breaking Stress

CAI=100.5(IE)–0.424 

Material Ultimate 

Stress (B-Basis) 

Material Limit 

Stress (B-Basis) 

IM7/8552 18-Ply 

70

60

50

40

30

20

10

0

0 2 4 6 8 10

Compression Strength (ksi)

Impact Energy (ft-lb)

Figure 5. Residual strength (CAI) curve for the Ares I interstage acreage material.
11

As figure 5 shows, the static compressive load-carrying capability of the sandwich structure 

drops with increasing impact energy. The testing in this study examined the effects of fatigue on 

the life of the sandwich panel where the fatigue stress amplitude was measured as a percentage of 

expected sUL.

3.2 Cyclic Testing

Constant amplitude cyclic loads were established for a given damage size based on a percentage of B-Basis UL or LL. In this study, the B-Basis load was found to be 86% (A-Basiswas 75%) of 

the expected failure load. Thus, B-Basis UL is 86% of static CAI strength and B-Basis LL is 61% of 

static CAI strength for any given impact energy. It was decided to test the impact damaged laminates 

above the B-Basis LL in order to produce a level of conservatism. Since previous work found that a 

difference of 20% between peak compression amplitudes had no effect on damage size or residual 

strength,8 a much larger compression fatigue load difference of 90% was chosen in an attempt to 

force a change in damage size and residual strength.

3.3 Materials and Specimens

The materials and specimens used in this study were identical to those used for the static testing. They will be briefly reviewed here.

The specimens consisted of composite sandwich panels that were manufactured by a cocure 

process. The honeycomb core used was perforated 5052 aluminum with a density of 3.1 lb/ft3 and 

a thickness of 1.125 in. The face sheets consisted of IM7/8552 carbon/epoxy and the film adhesive 

used to bond the face sheet to the core was FM 300K with an areal density of 0.08 lb/ft2. The layup 

of the face sheet was (+45,0,–45,0,90,0,0,90,0)S.

Each face sheet had a thickness of 0.115 in. The panels were manufactured as 24-× 24-in

square sandwich plates. All panels were subjected to NDE using infrared thermography (IRT) testing to ensure no flaws were present before specimens were machined from the panels. From the 

24-× 24-in panels, 4-× 6-in specimens were machined for impact and subsequent compression testing. 

This gave each specimen a cross-sectional area of 0.92 in2, which was used to calculate the breaking 

stresses.

After impact, the specimens were inspected with IRT and then were prepared for end loading 

compression testing. In order to avoid premature failures due to end brooming, the loaded ends of 

the specimens were potted in an aluminum frame and machined to a tolerance of ±0.001-in parallelism. Details of the specimen geometry are published elsewhere.34

3.4 Nondestructive Evaluation Testing

The damage in the specimens was assessed with IRT. IRT is an NDE technique that uses 

a sensitive infrared camera to monitor heat dissipation from a surface induced with a flash of heat 

from a quartz lamp. Any areas of damage will dissipate heat at a different rate than undamaged 

material and the results seen are termed ‘indications.’ A typical indication for the type of specimen 
12

used in this study is shown in figure 6. This two-dimensional indication assesses the planar area of 

damage with little information about the through thickness characteristics of the damage. Thus, 

damage growth in this study was measured by any increase in the planar size of the indication.

Damage Size

0°

45°

90°

Figure 6. Typical IRT indication of impact-damaged laminate.
13

4. TEST RESULTS

Testing consisted of the following three phases: (1) Inducing and measuring impact damage 

size by IRT, (2) fatigue testing and reassessing the damage size, and (3) residual compression strength 

after fatigue.

4.1 Impact Testing

A total of 30 specimens were impacted at various levels of impact energies for this fatigue 

study. An impactor with a diameter of 0.5 in was used for all impacts utilizing a drop weight tower 

that is described in more detail elsewhere.34 Table 2 presents the impact results of the 30 specimens 

used for this fatigue study. The solid line in figure 5 represents the expected sUL.

An example of the visual damage that the specimens incurred is shown in figure 7, where 

specimens at the low and high ends of impact energies used are presented.

4.2 Fatigue Testing

The maximum compressive fatigue load amplitude was chosen to be more than 100% of the 

B-Basis LL for a given damage size. For this study, constant amplitude fatigue cycles were applied 

with the minimum compressive load always set at 10% of the maximum compressive load for an 

R-ratio equal to 0.1 (a load range of 90%). The loading frequency was 5 Hz. The testing was performed in two series. The first series cycled 15 of the specimens 1,000 times. One thousand cycles is 

extreme for the Ares I interstage, given that only a few cycles above 50% of LL are expected (fig. 3). 

If a specimen survived the prescribed 1,000 cycles, it was checked with IRT for damage growth. One 

thousand additional cycles were applied to the specimen with an additional 10% increase in stress 

amplitude. After two cyclic stress amplitude increases for a total of 3,000 load cycles, the specimen 

was tested for static residual compression strength to ascertain any strength degradation due to the 

severe cyclic loading conditions. The results from the first series of tests are summarized in table 3 for 

B-Basis sUL and sLL and table 4 for A-Basis sUL and sLL. 

All 15 specimens survived the 3,000 cycles and no growth of the damage area could be detected 

during any of the three loading phases. Ten percent of the specimens were expected to fail at one 

cycle when the fatigue stress amplitude was 100% of sUL, since this is the B-Basis failure stress. 

Given that the composite interstage structure sees about three load excursions of significance and 

since the structural LL will be less than or equal to the material LL, no further fatigue testing is necessary to validate the fatigue life of the impact-damaged material of which the composite interstage 

is comprised. Despite this early observation, this study progressed to even more aggressive fatigue 

testing to push the limits since the validity of the design envelope can only be demonstrated by testing outside of the envelope.
14

Table 2. Impact results for specimens to be fatigue tested.

Specimen I.D.

Impact Energy

(ft-lb)

Damage Width

(in)

Expected Ultimate Stress 

(ksi)*

F-18-1 2.6 0.44 57.6

F-18-2 3.4 0.45 51.4

F-18-3 3.4 0.49 51.4

F-18-4 3.1 0.56 53.5

F-18-5 3.1 0.49 53.5

F-18-6 4.1 0.56 47.5

F-18-7 4 0.51 48

F-18-8 3.9 0.6 48.5

F-18-9 4 0.56 48

F-18-10 4 0.6 48

F-18-11 9.7 1 33

F-18-12 9.9 0.93 32.7

F-18-13 9.8 0.98 32.8

F-18-14 9.8 0.89 32.8

F-18-15 9.8 0.89 32.8

F2-18-1 3.7 0.71 49.6

F2-18-2 4.4 0.74 46.1

F2-18-3 5.7 0.96 41.3

F2-18-4 7.9 0.93 36

F2-18-5 8.7 0.98 34.5

F2-18-6 8.6 0.92 34.7

F2-18-7 9.8 0.98 32.8

F2-18-8 11.1 0.97 31.1

F2-18-9 11.8 0.94 30.3

F2-18-10 8.1 0.9 35.6

F2-18-11 5.4 0.87 43.3

F2-18-12 6.8 0.89 38.3

F2-18-13 12.2 0.85 29.9

F2-18-14 12.2 0.9 29.9

F2-18-15 12.1 0.74 30

*As determined by reference 37

For 15 additional specimens, the number of fatigue cycles was increased to 10,000. Only one 

stress amplitude was used since some specimens did fail during the fatigue testing. A summary of 

these specimens is presented in table 5.

In this series of fatigue tests, 5 of the 15 specimens failed before 10,000 cycles. Two of these 

failed specimens were at the two highest load amplitudes tested. Damage growth was detected on 

only two of the surviving specimens; however, the two specimens exhibiting damage growth maintained a residual strength >96% of the expected static strength. The results are plotted in figure 8 

to observe trends. The number next to each datum that demonstrated failure is the impact energy in 

foot-pounds that the specimen was damaged with.
15

4.4 ft-lbs 12.1 ft-lbs

Figure 7 Examples of the visual damage caused by impacts.

Table 3. B-Basis summary of the first series of fatigue after impacts tests: 1,000 cycles 

per stress amplitude.

Specimen I.D.

First Cyclic 

Stress Amplitude 

(ksi)

Percentage 

of σAVG/

σBUL/ σBLL

Second Cyclic 

Stress Amplitude 

(ksi)

Percentage 

of σAVG/

σBUL/ σBLL

Third Cyclic 

Stress Amplitude 

(ksi)

Percentage 

of σAVG/

σBUL/ σBLL

Percent Growth 

in Area of 

Delamination

F-18-1 49.9 75/87/122 54.9 82/95/133 59.9 89/104/146 0

F-18-2 43.6 73/85/119 47.9 80/93/130 52.3 88/102/143 0

F-18-3 45 76/88/123 49.3 84/98/137 54 90/105/147 0

F-18-4 46.5 75/87/122 51.2 83/96/134 55.9 90/105/147 0

F-18-5 42.2 68/79/111 46.4 75/87/122 50.7 82/95/133 0

F-18-6 34.7 63/73/102 38.2 69/80/112 41.6 76/88/123 0

F-18-7 40.9 73/85/119 45 81/94/132 49 88/102/143 0

F-18-8 38.7 69/80/112 42.6 76/88/123 46.4 83/96/134 0

F-18-9 39.8 71/83/116 43.8 78/91/127 47.8 86/100/140 0

F-18-10 39.8 71/83/116 43.8 78/91/127 47.8 86/100/140 0

F-18-11 20.5 53/62/87 22.5 58/68/95 24.6 65/75/105 0

F-18-12 25 65/76/106 27.5 72/84/118 30 79/92/129 0

F-18-13 23.5 66/77/108 28.2 74/86/120 30.7 81/94/132 0

F-18-14 23.5 62/72/101 28.2 74/86/120 30.7 81/94/132 0

F-18-15 25 65/76/106 27.5 72/84/118 30 78/91/127 0
16

Table 4. A-Basis summary of the first series of fatigue after impacts tests: 1,000 cycles 

per stress amplitude.

Specimen I.D.

First Cyclic 

Stress Amplitude 

(ksi)

Percentage 

of σAVG/

σAUL/σALL

Second Cyclic 

Stress Amplitude 

(ksi)

Percentage 

of σAVG/

σAUL/σALL

Third Cyclic 

Stress Amplitude 

(ksi)

Percentage 

of σAVG/

σAUL/σALL

Percent Growth 

in Area of 

Delamination

F-18-1 49.9 75/100/140 54.9 82/109/153 59.9 89/119/166 0

F-18-2 43.6 73/97/136 47.9 80/107/149 52.3 88/117/164 0

F-18-3 45 76/101/142 49.3 84/112/157 54 90/120/168 0

F-18-4 46.5 75/100/140 51.2 83/111/155 55.9 90/120/168 0

F-18-5 42.2 68/91/127 46.4 75,100/140 50.7 82/109/153 0

F-18-6 34.7 63/84/118 38.2 69/92/129 41.6 76/101/142 0

F-18-7 40.9 73/97/136 45 81/108/151 49 88/117/164 0

F-18-8 38.7 69/92/129 42.6 76/101/142 46.4 83/111/155 0

F-18-9 39.8 71/95/133 43.8 78/104/146 47.8 86/115/161 0

F-18-10 39.8 71/95/133 43.8 78/104/146 47.8 86/115/161 0

F-18-11 20.5 53/71/99 22.5 58/77/108 24.6 65/87/121 0

F-18-12 25 65/87/121 27.5 72/96/134 30 79/105/148 0

F-18-13 23.5 66/88/123 28.2 74/99/138 30.7 81/108/151 0

F-18-14 23.5 62/83/116 28.2 74/99/138 30.7 81/108/151 0

F-18-15 25 65/87/121 27.5 72/117/134 30 78/104/146 0

Table 5. Summary of the second series of fatigue after impacts test.

Specimen I.D.

Max Compressive 

Stress 

(ksi)

Percentage of 

σAVG/σBUL/σBLL

Percentage of 

σAVG/σAUL/σALL

Number of Cycles 

to Failure

Percent Growth in Area 

of Delamination

F2-18-1 43.1 75/87/122 75/100/140 3,339 N/A

F2-18-2 37.5 70/81/113 70/93/131 9,147 N/A

F2-18-3 27 56/65/91 56/75/105 Survived 104 0

F2-18-4 28 67/78/109 67/89/125 Survived 104 0

F2-18-5 25.7 64/74/104 64/85/119 Survived 104 0

F2-18-6 28.8 71/83/116 71/95/133 Survived 104 0

F2-18-7 26.8 71/82/115 71/95/133 Survived 104 0

F2-18-8 26.4 73/85/119 73/87/136 Survived 104 0

F2-18-9 28.6 81/94/132 81/108/151 Survived 104 47

F2-18-10 33.9 82/95/133 82/109/153 8,223 N/A

F2-18-11 35.9 73/85/119 73/97/136 Survived 104 0

F2-18-12 34.2 77/89/125 77/103/144 Survived 104 0

F2-18-13 29.3 83/98/137 83/111/155 Survived 104 93

F2-18-14 29.8 86/100/140 86/115/161 8,312 N/A

F2-18-15 30.9 89/103/144 89/119/166 3,341 N/A
17

Failed in Fatigue

Runout

Runout With Growth

3.7

12.2

12.2

8.1

4.4

120

100

80

60

40

20

0

0 2,000 4,000 6,000 8,000 10,000 12,000

Cycles to Failure

Expected Compression Strength (%)

Figure 8. Plot of results from table 5.

From this plot, the only basic conclusion that can be drawn is that, in general, higher fatigue 

loads are more detrimental. It also appears that the level of impact damage has no effect when the 

fatigue stresses are based as a percentage of predicted CAI strength. Although the evaluation and 

quantification of any possible trends would require additional specimens, the number of cycles experienced by expendable launch vehicles is very low with stress amplitudes generally <60% of σUL.

The IRT results of a typical specimen and the two specimens that exhibit damage growth are 

shown in figure 9. The dotted oval in each left image is superimposed over the corresponding image 

on the right to give a qualitative measurement of damage growth.

The IRT technique is not as precise at quantifying damage size as some types of NDE. However, IRT is the most likely candidate for use in the field, so it was employed here. 

4.3 Residual Strength Testing

Table 6 shows the results from the static compression strength testing of the specimens in this 

study that survived the fatigue loading.

The conclusion drawn from this testing is that fatigue tends to increase the compression 

strength of damaged material. This phenomena has been observed in other studies that examined
18

Specimen F2-18-7

Specimen F2-18-9

No Cycles

No Cycles

10,000 Cycles

10,000 Cycles

Specimen F2-18-13 No Cycles 10,000 Cycles

Figure 9. IRT results of a typical specimen (no growth) and the two specimens 

that exhibit damage growth after 10,000 cycles.

the residual strength of specimens that had survived a fatigue regimen.2,5,6,8,11,18 Only one reference 

could be found where fatigue decreased the CAI strength of the laminates.17 In that reference, the 

high amplitude loads were held for 120 s, perhaps indicating a creep rupture phenomena that may 

need to be explored for parts experiencing long dwell times at UL.
19

Table 6. Static compressive breaking stress of specimens that survived all fatigue series.

Specimen I.D.

Residual 

Compression Strength 

(ksi)

Percent of Expected Residual 

Compression Strength (σAVG)

F-18-1 73.7 111

F-18-2 74.6 115

F-18-3 71.2 116

F-18-4 64.9 117

F-18-5 69.1 113

F-18-6 59.9 108

F-18-7 62.1 105

F-18-8 64.7 122

F-18-9 61.4 110

F-18-10 63 119

F-18-11 49.7 136

F-18-12 49.1 127

F-18-13 48.2 130

F-18-14 38.3 96

F-18-15 46.6 117

F2-18-3 46.8 98

F2-18-4 41.6 100

F2-18-5 38.1 95

F2-18-6 39.5 98

F2-18-7 41.1 108

F2-18-8 38.1 105

F2-18-9* 33.9 96

F2-18-11 45.4 92

F2-18-12 39.5 89

F2-18-13* 36.6 105

*Specimen exhibited damage growth.
20

5. DISCUSSION

In assessing the damage tolerance approach to human-rated expendable launch vehicles 

according to reference 1, it is evident that the required coupon-level testing reduces to a static case. 

The aspects of repeated loading will have no effect on the composite laminate when compressive 

in-plane loading is involved. Using conservative assumptions, no more than three cycles to LL are 

expected for the Ares I composite interstage. Thus, the number of fatigue cycles used in this test program where high by at least two orders of magnitude.

All evidence from this and past studies indicate that compression fatigue loading will increase 

the residual strength of the material. Also, once cyclic loading has ceased to increase the residual 

compression strength, the residual compression strength reverts back to the static case. Thus, static 

residual strength is a lower bound for laminates used in expendable load vehicles and development 

of figure 1(a) is unnecessary.

Figure 1(b) can be simplified since the diagram will consist of data in unrealistic regions 

of stress amplitude and/or number of cycles to LL as shown schematically in figure 10. Figure 10 

combines some of the values given in the references in table 1 with the data in this study. No growth 

or failures occurred after 10,000 cycles at loads as high as 60% of the expected average failure load, 

which represents the approximate greatest allowable LL for a given damage size.

Data From This Study

Data From References

UL

LL

Too Many Cycles

Loads Too High

Design Space for

Launch Vehicles

Cycles to Failure (10x)

Static CAI Strength (%)

120

100

80

60

40

20

0

0 1 2 3 4 5 6 7

Figure 10. Fatigue life data from this and other studies.
21

The necessity of constructing figure 1(c) is dubious, as specimens that exhibited damage 

growth were stronger than equivalent ones that were not fatigued. In addition, detection of damage 

growth is difficult and has been shown to occur either at the early stages of a specimen’s fatigue cycle 

(which is associated with an increase in strength) or near the end of a specimen’s life (when failure is 

catastrophic), and the number of cycles to failure cannot be predicted with any semblance of accuracy due to the large amount of scatter in the data. Because of the rapid growth at the end of life, the 

number of cycles to failure is difficult to predict making end-of-life growth measurements difficult. 

The first two of three phases of fatigue life illustrated in figure 2 were demonstrated in this study. 

Phase 1 was demonstrated by the 1,000-cycle specimens getting stronger and phase 2 was demonstrated by the 10,000-cycle specimen’s strength remaining the same.
22
23

6. CONCLUSIONS

From this study, it appears that all fatigue load cases can be considered static for expendable launch vehicles subjected to in-plane compressive loads. The number of cycles is low and the 

maximum applied stress is just at the level needed for fatigue failure. Tight corners, ramps, and 

other discontinuities may be affected by fatigue loading but the in-plane CAI properties are not. If 

a subcomponent or full-scale static test shows out-of-plane failure, then fatigue loading may be of 

a concern; however, if an out-of-plane failure occurs, the part may need to be redesigned. Thus, 

fatigue testing is unnecessary above the coupon level.

The data from the experiments in this study corroborate with those from other studies in 

which a fatigue limit above 60% of CAI strength exists for laminates. In this particular case, since 

runout was reduced to 10,000 cycles, the fatigue limit was >80% of CAI strength, which should never 

occur in a launch vehicle structure since structural LL is, at most, equal to the material UL reduced 

by a safety factor that makes the maximum allowed load about 61% of average CAI strength. Other 

conclusions from this study include the following:

• Setting a prescribed number of loads and load amplitudes based upon the load spectrum of the 

component to test fatigue life should take place on a case-by-case basis to provide more flexibility 

than setting fixed numbers such as presented in reference 1.

• Undamaged specimens have a steeper S-N curve and indicate that specimens that are impact 

damaged are less affected by fatigue. Thus, any fatigue tests that are warranted should occur on 

undamaged specimens to obtain a ‘worst case’ life curve.

• Impact damage growth cannot be used as a clear degradation parameter since most laminates that 

showed growth had improved or exhibited no change in CAI strengths.

• A full-scale fatigue test is not needed for a composite structure provided that the development and 

full-scale static tests are successful This same conclusion was reached by those who originally were 

responsible for developing the load enhancement factor approach for composite laminates.38
24

REFERENCES

 1. NASA Marshall Space Flight Center Fracture Control Requirements for Composite and Bonded 

Vehicle and Payload Structures, MSFC–RQMT–3479, Marshall Space Flight Center, AL, June 

2006.

 2. Rosenfeld, M.S.; and Gause, L.W.: “Compression Fatigue Behavior of Graphite/Epoxy in the 

Presence of Stress Raisers,” Fatigue of Fibrous Composite Materials, ASTM STP 723, American 

Society for Testing and Materials, pp. 174–196, 1981.

 3. Ramkumar, R.L.: “Effect of Low-Velocity Impact Damage on the Fatigue Behavior of Graphite/Epoxy Laminates,” Long-Term Behavior of Composites, ASTM STP 813, T.K. O’Brien (ed.), 

American Society for Testing and Materials, pp. 116–135, 1983.

 4. Griffin, C.F.; and Becht, G.J.: “Fatigue Behavior of Impact Damaged BMI and Thermoplastic 

Graphite Composites,” Proceedings, 36th International SAMPE Symposium and Exhibition, San 

Diego, CA, Part 2, pp. 2197–2209, April 15–18, 1991.

 5. Ong, C.-L.; Sheu, M.-F.; Liou, Y.-Y.; and Hsiao, T.-J.: “The Study of the Fatigue Characteristics 

of Composite After Impact,” Proceedings, 36th International SAMPE Symposium and Exhibition, San Diego, CA, pp. 912–923, April 15–18, 1991.

 6. Krafchak, T.M.; Petra, J.-M.; Davidson, B.D.; and Chen, G.-S.: “Effect of Impact Damage on 

the Compression Fatigue Behavior of Composite Tubes,” AIAA Paper 93-13999-CP, Proceedings, AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Material Conference, La Jolla, CA, pp. 859–866, April 19–22, 1993.

 7. Portanova, M.A.; Poe, C.C.; and Whitcomb, J.D.: “Open Hole and Post-Impact Compression 

Fatigue of Stitched and Unstitched Carbon/Epoxy Composites,” NASA/TM—1990–102676, 

Langley Research Center, Hampton, VA, June 1990.

 8. Mitrovic, M.; Hahn, H.T.; Carman, G.P.; and Shyprykevich, P.: “Effect of Loading Parameters 

on the Fatigue Behavoir of Impact Damaged Composite Laminates,” Comp. Sci. and Tech., 

Vol. 59, pp. 2059–2078, 1999.

 9. Symons, D.D.; and Davis, G.: “Fatigue Testing of Impact-Damaged T300/914 Carbon-FiberReinforced Plastic,” Comp. Sci. and Tech., Vol. 60, pp. 379–389, 2000.

10. Chen, A.S.; Almond, D.P.; and Harris, B.: “In Situ Monitoring in Real Time of Fatigue-Induced 

Damage Growth in Composite Materials by Acoustography,” Comp. Sci. and Tech., Vol. 61, 

pp. 2437–2443, 2001.
25

11. Shyprykevich, P.; Tomblin, J.; Ilcewicz,L.; et al.: “Guidelines for Analysis, Testing and Nondestructive Inspection of Impact-Damaged Composite Sandwich Structures,” DOT/FAA/

AR-02/121, Federal Aviation Administration Technical Center, Atlantic City, NJ, March 2003.

12. Melin, L.G.; and Schon, J.: “Buckling Behaviour and Delamination Growth in Impacted Composite Specimens Under Fatigue Load: An Experimental Study,” Comp. Sci. and Tech., Vol. 61, 

pp. 1841–1852, 2001.

13. Uda, N.; Ono, K.; and Kunoo, K.: “Compression Fatigue Failure of CFRP Laminates With 

Impact Damage,” Comp. Sci. and Tech., Vol. 69, pp. 2308–2314, 2009.

14. Saito, H.; and Kimpara, I.: “Damage Evolution Behavoir of CFRP Laminates Under PostImpact Fatigue With Water Absorption Environment,” Comp. Sci. and Tech., Vol. 69, pp. 847–

855, 2009.

15. Gower, M.R.L.; and Shaw, R.M.: “Assessment of the Applicability of Compression-AfterImpact (CAI) and Open Hole Tension (OHT) Methods for Use Under Fatigue Loading,” 13th 

European Conference on Composite Materials, Stockholm, Sweden, June 2–5, 2008, <http://

extra.ivf.se/eccm13_programme/program.htm>, Accessed March 3, 2010.

16. Clark, G.; and van Blaricum, T.J.: “Load Spectrum Modification Effects on Fatigue of ImpactDamaged Carbon Fibre Composite Coupons,” Composites, Vol. 18, No. 3, pp. 243–251, 

July 1987.

17. Saunders, D.S.; and van Blaricum, T.J.: “Effect of Load Duration on the Fatigue Behavior of 

Graphite/Epoxy Laminates Containing Delaminations,” Composite., Vol. 19, No. 3, pp. 217–

228, July 1988.

18. Swanson, S.R.; Cairns, D.S.; Guyll, M.E.; and Johnson, D.: “Compression Fatigue Response 

for Carbon Fiber With Conventional and Toughened Epoxy Matrices With Damage,” J. Eng. 

Mater. Technol., Vol. 115(1), pp. 116–121, January 1993.

19. Masters, J.E.; and Reifsnider, K.L.: “An Investigation of Cumulative Damage Development in 

Quasi-Isotropic Graphite/Epoxy Laminates,” Damage in Composite Materials, ASTM STP 775, 

K.L. Reifsnider (ed.), American Society for Testing and Materials, pp. 40–62, 1982.

20. Highsmith, A.L.; and Reifsnider, K.L.: “Stiffness-Reduction Mechanisms in Composite Laminates,” Damage in Composite Materials, ASTM STP 775, K.L. Reifsnider (ed.), American Society for Testing and Materials, pp. 103–117, 1982.

21. Kunz, S.C.; and Beaumont, P.W.R.: “Microcrack Growth in Graphite Fiber-Epoxy Resin Systems During Compressive Fatigue,” Fatigue of Composite Materials, ASTM STP 569, American Society for Testing and Materials, pp. 71–91, 1975.
26

22. Han, H.T.; Mitrovic, M.; and Turkgenc, O.: “The Effect of Loading Parameters on Fatigue of 

Composite Laminates: Part III,” DOT/FAA/AR-99/22, Federal Aviation Administration, Washington, DC, June 1999.

23. Potter, R.T.: “The Significance of Defects and Damage in Composite Structures,” AGARD 

Conference on Characterization, Analysis, and Significance of Defects in Composite Materials, 

AGARD-CP-355, January 1983.

24. Simonds, R.A.; and Stinchcomb, W.W.: “Response of Notched AS4/PEEK Laminates to 

Tension/Compression Loading,” Advances in Thermoplastic Matrix Composite Materials, 

ASTM STP 1044, G.M. Newaz (ed.), American Society for Testing and Materials, pp. 133–145, 

1989.

25. Bakis, C.E.; and Stinchcomb, W.W.: “Response of Thick, Notched Laminates Subject to 

Tension-Compression Cyclic Loads,” Composite Materials: Fatigue and Fracture, ASTM STP 

907, H.T. Hahn (ed.), American Society for Testing and Materials, pp. 314–334, 1986.

26. Dorey, G.: “Impact and Crashworthiness of Composite Structures,” in Structural Impact and 

Crashworthiness, Vol. 1, J. Morton (ed.), G.A.O. Davies (ed.), Elsevier Applied Science Publishers, London, England and New York, New York, pp. 155–192, 1984.

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

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

305, 1983.

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

Proceedings, Sixth International Conference on Composite Materials, Second European Conference on Composite Materials, E.J. Matthews, N.C.R. Buskell, J.M. Hodgkinson, and J. Morton 

(eds.), Vol. 3, pp. 3.1–3.26, 1986.

29. Morton, J.; and Godwin, E.W.: “Impact Response of Tough Carbon Fiber Composites,” Comp. 

Struct., Vol. 13. pp. 1–19, 1989.

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

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

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

CFRP Reinforced With High-Strain Carbon Fibers,” Comp. Sci. and Tech., Vol. 25. pp. 133–

148, 1986.

32. Puck, A.; and Schurmann, H.: “Failure Analysis of FRP Laminates by Means of Physically 

Based Phenomenological Models,” Comp. Sci. and Tech., Vol. 58, pp. 1045–1067, 1998.

33. Nyman, T.; Ansell, H.; and Blom, A.: “Effects of Truncation and Elimination on Composite 

Fatigue Life,” Comp. Struct., Vol. 48, pp. 275–286, 2000.
27

34. Nettles, A.T.; and Jackson, J.R.: “Developing a Material Strength Design Value Based on Compression After Impact Damage for the Ares I Composite Interstage,” NASA/TP—2009–215634, 

Marshall Space Flight Center, AL, January 2009.

35. Moody, R.C.; Harris, J.S.; and Vizzini, A.J.: “Scaling and Curvature Effects on the Damage 

Tolerance of Impacted Composite Sandwich Panels,” J. Sand. Struct. Mater., Vol. 4, pp. 71–82, 

2002.

36. Fawcett, A.; Trostle, J.; and Ward, S.: “777 Empanage Certification Approach,” Proceedings, 

11th International Conference on Composite Materials (ICCM-11), Austrailia, pp. 178–199, 

July 14–18, 1997.

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

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

38. Whitehead, R.S.; Kan, H.P.; Cordero, R.; and Saether, E.S.: “Cerification Testing Methodology 

for Composite Structures,” DOT/FAA/CT-86/39, Federal Aviation Administration Technical 

Center, Atlantic City, NJ, October 1986.
28

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a. REPORT b. ABSTRACT c. THIS PAGE

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 PAGES

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Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18

Simplification of Fatigue Test Requirements for Damage Tolerance 

of Composite Interstage Launch Vehicle Hardware

A.T. Nettles, A.J. Hodge, 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–1283

Technical Publication

NASA/TP—2010–216434

composites, damage tolerance, fatigue, requirements, damage growth

01–06–2010

UU 40

NASA

U U U

The issue of fatigue loading of structures composed of composite materials is considered in a requirements document 

that is currently in place for manned launch vehicles. By taking into account the short life of these parts, coupled with 

design considerations, it is demonstrated that the necessary coupon level fatigue data collapse to a static case. Data 

from a literature review of past studies that examined compressive fatigue loading after impact and data generated 

from this experimental study are presented to support this finding. Damage growth, in the form of infrared thermography, was difficult to detect due to rapid degredation of compressive properties once damage growth initiated. 

Unrealistically high fatigue amplitudes were needed to fail 5 of 15 specimens before 10,000 cycles were reached. Since 

a typical vehicle structure, such as the Ares I interstage, only experiences a few cycles near limit load, it is concluded 

that static compression after impact (CAI) strength data will suffice for most launch vehicle structures.

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