October 2013

NASA/TM–2013-218050

Material Property Characterization of 

AS4/VRM-34 Textile Laminates

Ray W. Grenoble

Langley Research Center, Hampton, Virginia

William M. Johnston

Science and Technology Corporation, Hampton, Virginia
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 plays a key part in helping NASA maintain 

this important role.

The NASA STI program operates under the 

auspices of the Agency Chief Information Officer. 

It collects, organizes, provides for archiving, and

disseminates NASA’s STI. The NASA STI 

program provides access to the NASA Aeronautics 

and Space Database and its public interface, the 

NASA Technical Report Server, thus providing one 

of the largest collections of aeronautical and space 

science STI in the world. Results are published in 

both non-NASA channels and 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 counterpart of peerreviewed formal professional papers, but 

having 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 missions, 

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 also include organizing 

and publishing research results, distributing 

specialized research announcements and feeds, 

providing information desk and personal search 

support, and enabling data exchange services.

For more information about the NASA STI 

program, see the following:

 Access the NASA STI program home page 

at http://www.sti.nasa.gov

 E-mail your question to help@sti.nasa.gov

 Fax your question to the NASA STI 

Information Desk at 443-757-5803

 Phone the NASA STI Information Desk at 

443-757-5802

 Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076-1320
National Aeronautics and 

Space Administration

Langley Research Center 

Hampton, Virginia 23681-2199

October 2013

NASA/TM–2013-218050

Material Property Characterization of 

AS4/VRM-34 Textile Laminates

Ray W. Grenoble

Langley Research Center, Hampton, Virginia

William M. Johnston

Science and Technology Corporation, Hampton, Virginia
Available from:

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076-1320

443-757-5802

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration.
. 

1

Abstract 

Several material properties (modulus, strengths, and fracture 

toughness) of a textile composite have been evaluated to provide input 

data to analytical models of Pultruded Rod Stiffened Efficient Unitized 

Structure (PRSEUS). The material system is based on warp-knitted 

preforms of AS4 carbon fibers and VRM-34 epoxy resin, which have 

been processed via resin infusion and oven curing. Tensile, compressive, 

shear, and fracture toughness properties have been measured at ambient 

and elevated temperatures. All specimens were tested in as-fabricated 

(dry) condition. Specimens were tested with and without throughthickness stitching. 

Introduction

Development work on the Pultruded Rod Stiffened Efficient Unitized Structure 1,2 (PRSEUS) concept has recently 

progressed via collaborative work between NASA and Boeing on the Hybrid Wing Body aircraft (HWB). The 

HWB is a design for a future airliner, based on a flying wing configuration. The width of the aircraft’s body center 

section results in a pressurized cabin with a rectangular cross-sectional shape. Whereas a conventional (elliptical 

cross-section) fuselage carries most internal pressure-driven loads through skin tension, the flat sides of the HWB 

pressure cabin will carry those loads primarily via bending. 

PRSEUS leverages lessons learned in previous NASA-sponsored textile composites development activities such as 

the Advanced Composites Technology (ACT) program3,4. ACT focused on development of carbon fiber textilebased composites, which incorporated through-thickness stitching to increase damage tolerance. The development 

of innovative processing methods to reduce fabrication costs was also pursued. The program culminated with the 

fabrication (using a resin film infusion process) and testing of a representative wing box structural component with 

embedded impact damage. 

The primary features distinguishing PRSEUS are: 

 Skins and stiffening members composed of warp-knitted, multidirectional, dry carbon fiber 

preforms 

 Foam-filled frames covered in one of the aforementioned preforms 

 Stringers incorporating a pre-cured, unidirectional carbon/epoxy rod in the upper flange 

 Through-thickness stitching between the stringer/frame base flanges and the skin 

 Out-of-autoclave processing (via a vacuum-assisted resin infusion process) using low-cost foam 

assembly tooling and oven cure. 

The unidirectional rod in the stringer flanges provides the structure with very high bending stiffness in the 

longitudinal direction. This attribute makes the concept attractive for application to HWB’s pressurized cabin. The 

stitching at the stiffener/skin bondline is intended to suppress delamination of that critical interface well into a postbuckling loading regime5

. 

During the fabrication process, dry stiffener and frame preforms are assembled upside-down in foam/plastic tooling. 

The tear straps and skins are placed over the stiffening elements and the constituents are then stitched together 

through the thickness from what will be the tool side (skin side) of the part. The position of the precured rods in the 

stringers relative to the skin is established by the holes through the foam cores of the frames. Once the stitched 

assembly is removed from the tooling, it is effectively self-supporting. This attribute greatly reduces the amount of 

tooling required during the remainder of the manufacturing process. An illustration of a stringer-frame intersection, 

which shows how the structure is assembled, is provided in Figure 1. 
. 

2

To date, little material property data for PRSEUS structure have been available in the open literature. This deficit 

has prevented development of accurate finite element models of this structural concept, and has limited the ability of 

others outside NASA and Boeing to evaluate the suitability of PRSEUS for their own applications. The present 

study is intended to address this lack of data by developing a set of elastic material properties for use as input to 

analytical models of PRSEUS acreage (skin and stiffener webs). Coupons are tested to provide data on the tensile, 

compressive, and shear properties. All coupons were tested in as-fabricated condition. Testing was performed at 

72°F and 180°F. In addition, Mode I and Mode II interlaminar fracture toughness testing was performed to provide 

data for analysis of debonding and delamination. 

Materials   This study evaluates the mechanical properties of balanced, symmetric laminates, fabricated from warp-knitted 

[0/90]T preforms of AS4 fiber, and infused with VRM-34 epoxy resin (both from Hexcel Corporation). The 

thickness of the panels fabricated varied with the requirements of each specimen type, as shown in Table 1. Boeing 

Research and Technology (Huntington Beach, CA) fabricated the panels and shipped them to NASA Langley 

Research Center. 

Some panels incorporated through-thickness stitching to represent the stitched bondline between the skins and 

stiffeners. Stitches were formed from 1600 denier Vectran thread at 5.08 stitches per inch. In PRSEUS structure, 

stitches are placed in pairs. The first stitch enters the skin near the center of the stiffener element lower flange half, 

and is oriented normal to the skin surface. A second, inclined, stitch is placed on the bag side of the part, adjacent to 

the stiffener web and exits on the tool side under the stiffener web. Figure 2 illustrates the stitch placement on a 

typical stringer. Placement of stitches within stitched specimens tested in this work followed a similar scheme. The 

stitches were placed in pairs; one stitch was oriented normal to the skin surface and one was inclined at a 56° angle 

relative to the tool surface. All stitch lines were oriented along the length of the specimens. Table 1 shows the 

complete test matrix. 

Characterization

Each specimen set contained six replicates. All specimens were tested in servo-hydraulic test stands. Tensile 

specimens were loaded through hydraulic wedge grips. Compression, shear, and fracture toughness specimens were 

tested in test fixtures meeting their respective test standard recommendations. All data analysis was performed in 

the manner recommended by the relevant ASTM test standards. 

Table 1. Test matrix 

Test Type ASTM Test

Standard

Test

Temperatures,

°F

Unstitched Stitched

Specimen

Dimensions‡,

in.

Extensometer / Strain Gages

Tension D3039 72, 180 Y Y 12 x 1 x 0.063 back‐to‐back  /

back‐to‐back at 90°

Compression D6418 72, 180 N Y 6 x 0.5 x 0.157 none / back‐to‐back axial

Inplane Shear D7078 72, 180 Y Y 2.2 x 3 x 0.157 none / back‐to‐back at 45°

Mode I Fracture

Toughness D5229 72, 180 Y N† 7 x 1 x 0.314 none

Mode II Fracture

Toughness ENF Draft 72, 180 Y Y 7 x 1 x 0.314 none

†

 Test found inapplicable due to failure mode exhibited 

‡

 All laminates were symmetric and balanced, and were based on 0/90 and 90/0 preforms of 0.0157 in. thickness.om 
. 

3

Specimen Preparation

In general, specimens were cut from their precursor panels using a waterjet. Tensile specimens were tabbed with a 

balanced, symmetric layup of the same AS4/VRM-34 material used for the specimens. Tabs were 3.5 in. long x 

0.063 in. thick, and their ends were tapered at a 6° angle toward the gage section. The tab material was secondarily 

bonded to the precursor panel using Hysol EA9394TM adhesive and was cured at 150°F for one hour. Individual 

tensile specimens were then cut from the panel with a diamond-grit grinding wheel mounted on a surface grinder. 

Axial strains were monitored with a pair of extensometers (MTS Type 632), calibrated over a 0.12-in. length. 

Transverse strains were monitored via back-to-back transverse strain gages in the center of the gage section. 

The edges and ends of compression test specimens were ground flat and parallel by trimming their waterjet-cut 

edges on a surface grinder, using a diamond-grit cutting wheel. Tabs were made from glass/epoxy laminate (G10), 

1/4 in. in thickness. They were secondarily bonded to the panel in a manner identical to that of the tensile 

specimens. Axial strains were monitored during testing with a pair of back-to-back strain gages, 0.25 in. long, 

centered in the specimens’ 0.5-in. long gage length. 

The inplane shear specimens were of the V-notch rail shear configuration. These specimens require geometric 

tolerances in the gage region that could not be maintained by the waterjet cutter. The final geometry of the notch on 

these specimens was established with a diamond grit endmill in a milling machine. Shear strains within the gage 

section were monitored with back-to-back strain gages mounted at +/- 45° angles to the applied load. Stitched shear 

specimens were cut in two configurations to assess the effects of shear loading along the stitch line. The first 

configuration placed the line of normal-direction stitches through the center of the notch, as seen in Figure 3(a). The 

second configuration centered the notch between the two stitch lines, which is shown in Figure 3(b). 

Mode I fracture toughness specimens were of the double cantilever beam (DCB) geometry. Load introduction was 

through adhesive-bonded aluminum hinges. Both types of fracture toughness specimen included an initial 

delamination, 2.5 in. long, created by inclusion of a 0.0005 in.-thick Teflon film in the specimen midplane. Mode II 

fracture specimens were tested in an end notched flexure (ENF) configuration. They required no preparation beyond 

waterjet cutting and applying the appropriate markings to the specimen edges. Crack front location accuracy was 

enhanced on both fracture toughness specimens by applying a thin coat of flat white spray paint to the specimen 

edges. 

Results

Tension Testing

Failure in all tensile specimens (stitched and unstitched) was preceded by tab delamination. Delamination typically 

initiated at the tip of the tapered tab end, adjacent to the gage section. When the behavior was first observed, 

clamping pressure in the test stand hydraulic grips was identified as a probable cause. The initial specimens were 

gripped with a pressure of 3000 psi, a value that past experience had shown to be appropriate for a specimen of this 

type. To assess if the grip pressure was playing a role in the observed failure mode, four dummy specimens were 

cut from the panel trim area and their edges were highlighted with a thin coat of white paint. These coupons were 

tested in tension at several grip pressures; 2500 psi, 1500 psi, 1000 psi, and 500 psi, 500 psi being the lowest 

controllable pressure that the test stand’s hydraulic grip controller could provide. All of the dummy specimens 

delaminated at the tab edge without slipping between the grips and tab surfaces. Figure 4 shows one edge of each 

specimen from the grip pressure experiment. Yellow arrows in the photos indicate the location of the delaminations 

in each dummy specimen. After completion of the grip pressure trials, the tabs on the first tensile test specimen 

were removed by inserting a thin blade under the edge of the tab and prying the tab off. The fracture surfaces were 

examined under a low-power microscope, and a photomicrograph of one tab surface is shown in Figure 5. The tab 

surface clearly has no traces of adhesive present, except for a small amount along the thin edge of the tab. This 

indicates that the adhesive did not bond to the tab surface. Based on the two sets of findings, grip pressure was 

eliminated as a critical factor in the observed failure mode, and the precise cause of the delamination is unknown. 

However, tab surface contamination and faults with the adhesive are plausible explanations for the bond failure. All 

other test variables had been consistent with ASTM D3518 recommendations, so the remaining specimens were 

tested with 3000 psi of grip pressure. 
. 

4

The measured tensile properties are compiled in Table 2. The undesired failure mode calls into question the ultimate 

tensile strengths reported in this work, and they may differ from those of similar specimens that fail strictly in the 

gage section. However, there appears to be sufficient data to calculate reliable modulus data for all of the 

specimens. Specimen #5 in the 72°F / unstitched set had a tensile strength of only 64.4 ksi, only 57% of the strength 

of the three strongest specimens. If the strength of #5 is taken as an outlier, the remaining five specimens average 

103.7 ksi tensile strength with a coefficient of variation (COV) of 9.46%. Similarly to the 72°F data set, the 

unstitched specimen #5 tested at 180°F exhibited an unexpectedly low strength. Considering it an outlier, the 

average ultimate strength increases to 107 ksi, and the COV decreases to 6.27%. Alternative values of average, 

standard deviation, and COV are provided for both sets of unstitched specimens in Table 2. In both specimen sets, 

modulus and Poisson’s ratio averages are comparatively unaffected by the inclusion of specimen #5. Examining the 

strength averages in Table 2, if we neglect specimen #5 in the unstitched specimen sets, we can conclude that the 

stitches reduced the tensile strength of the material by 13% and increased the scatter in the strength data. 

Post-test photographs of the tensile specimens are provided in Figure 6 - 10. Yellow arrows indicate the location of 

the ultimate tensile failure for each specimen. The figures clearly show that 22 of 24 tensile specimens failed 

directly under the tabs. Other damage visible in the specimens is primarily attributable to spring-back of the 

specimen after final fracture. 

Compression

All of the compression specimens exhibited acceptable failure modes, generally near the center of the gage section. 

The specimen set tested at 72°F displayed brooming failures in four of six specimens. The specimens tested at 

180°F all exhibited a buckling failure at an angle to the specimen loading direction. The compression test data are 

compiled in Table 3. The test specimens are shown in Figure 10 and 12. The increase in test temperature from 72°F 

to 180°F decreased compressive strength and compressive modulus by 7.5% and 4.2%, respectively, and increased 

the scatter in those properties. 

Inplane Shear

The inplane shear test data are tabulated in Table 4 – 6. Similar to the other specimen types, data from 72°F tests 

are provided in part (a) of each table. Data from testing at 180°F are provided in part (b). Examination of the data in 

Table 4 indicates that the specimens tested at 72°F behave consistently, with little scatter. However, the first two 

specimens tested at 180°F displayed unexpectedly low modulus and strength. Post-test examination of the 

specimens revealed no indications of anomalies in the specimens, and the data file indicates temperatures close to 

180°F. Because no reason could be found for the difference in properties, the shear property data in Table 4(b) are 

presented two ways, one includes all data from the specimen set, the other assumes that specimens #1 and #2 are 

outliers. 

If we base our assessment of the unstitched material only on the last four specimens, Table 4 indicates that the 

increase in test temperature from 72°F to 180°F caused a 3% decrease in ultimate shear strength and no decrease in 

modulus or 5%-offset shear strength. This indicates that the shear properties of the unstitched material are relatively 

insensitive to temperature over that temperature range. In contrast to that conclusion, Table 5 clearly indicates that 

the shear properties of the specimens with stitching through the root of the notch were significantly degraded by the 

elevated test temperature. Shear modulus decreased by 18% with the increase in test temperature. Ultimate and 

5%-offset shear strength decreased by 15% and 20%, respectively, at 180°F. Scatter in the data increased at 180°F 

as well. The specimens with stitching around the root of the notch displayed similar sensitivity to elevated 

temperature. It is clear from Table 6 that shear modulus decreased by 18% at elevated temperature. Ultimate shear 

strength was reduced by nearly 19% and the 5%-offset shear strength was reduced by more than 27%. The reason 

for the high sensitivity of the stitched-around specimens is unclear. However, Table 6(a) shows that shear strength 

in the room temperature specimens scattered to a larger degree than is seen in either the unstitched or stitchedthrough specimen sets tested at that temperature. 

The previous discussion compared the effects of test temperature on otherwise identical specimen sets. Table 7 

compares the effects of both stitching and temperature on shear properties across specimen sets. The knockdowns 

were calculated for all shear specimen sets relative to unstitched material tested at 72°F. Presented in this manner, it 
. 

5

is clear that stitching has a small effect on room temperature properties, but at elevated temperatures, significant 

decreases in strength and modulus can be expected.

The unstitched inplane shear specimens displayed subtle indications of surface cracking throughout the region 

between the specimen notches. This damage was only visible under 10X magnification. Specimens tested at room 

temperature and at 180°F appeared identical. Damage within the specimens with stitching through the root of the 

notch appeared to be concentrated along the stitch line. Close examination of these specimens reveals a small 

wrinkle through the gage section directly over the central stitch line. This wrinkle may be attributable to 

microbuckling of 90° fiber bundles between resin rich regions around the stitches and shear deformation within 

those regions. The specimens tested at 180°F have a more prominent wrinkle. Similar to the other specimens, this 

wrinkle is only visible under the smooth covering of strain gage adhesive. Damage in the specimens with stitch 

lines spanning the notch region was mostly seen between the two lines of stitches. The most prominent sign of 

damage in specimens tested at 72°F and 180°F was small striations oriented between notches. These can only be 

seen in regions covered with a thin, smooth layer of strain gage mounting adhesive. Examination of the edges in the 

notch region reveals no signs of delamination in any of the shear specimens. The lack of scoring in the grip sections 

of the specimens show that all specimens were held immobile in the grips and that load was applied in the manner 

expected during testing, which lends confidence to the shear strength data generated in this work. Post-test 

photographs of shear specimens were not included in this report because the macroscopic signs of damage were too 

subtle to be captured via that method. 

Mode I Fracture Toughness

The load, displacement, and crack length data from the DCB tests were analyzed by the Modified Compliance 

Calibration method of the test standard. The resultant R-curves for DCB specimens are compiled in Table 8. 

The critical fracture toughness for each specimen is listed in Table 9. To avoid variability due to geometric 

variation in the Teflon insert, GIC is taken as the apparent fracture toughness at a point at which the crack front 

propagated 5 mm. from the insert edge. Note that these data are for specimens without through-thickness stitching. 

Testing of stitched DCB specimens proved to be impracticable. 

Testing of the first stitched DCB specimen revealed uncertainties related to the test specimen and analysis method. 

Because of the difference in stitch angles relative to the laminate midplane, the vertical stitches carried more of the 

crack opening load than the angled stitches. As a result, as the crack began to extend from the Teflon insert, the arms 

of the specimen exhibited noticeable twisting and the crack front became highly skewed. The crack propagated 

several stitches before stopping. Increasing loads eventually caused the stitches to fracture, and the crack to 

propagate the full length of the specimen. The non-uniformity of the applied load and the highly skewed crack front 

during crack extension were inconsistent with the assumptions of the test specification. However, the observed 

behavior is consistent with analytical predictions of stitched composite behavior6,7. Further DCB testing was 

deferred until in-depth examination of the test results for the first specimen can evaluate the validity of the test 

results. In the end, time constraints prevented completion of stitched DCB specimen testing, so those results will 

not be included in this report. 

Mode II Fracture Toughness

The results of Mode II fracture toughness testing are tabulated in Table 10. Fracture toughness values are provided 

for the first loading (non-precracked) and the second loading (precracked). During compliance calibration loadings 

of the stitched specimens, some variation in the applied load was caused by the unevenness of the stitched IML 

surface. Fitting least-square linear curves to the stitched specimen data was therefore less accurate than with the 

specimens without stitching. These uncertainties result in more scatter in apparent fracture toughness than the 

unstitched specimens. Even with the additional scatter, the data indicate that the stitches provided an increase in 

Mode II fracture toughness. 

Concluding Remarks

Several material properties (modulus, strengths, and fracture toughness) of a textile composite have been 

characterized to provide input data to analytical models of PRSEUS structure. Tensile, compressive, shear, and 

fracture toughness properties have been measured at ambient and elevated temperatures, and with and without 
. 

6

through-thickness stitching. Tensile specimens exhibited tab-related failure modes, which may have resulted in 

overly conservative estimates for tensile strength. A test temperature of 180°F had little effect on tensile properties 

in comparison to similar specimens tested at ambient temperature. Stitching reduced tensile strength by 

approximately 13%. Compressive properties were reduced by the elevated test temperature by 4% – 8%. Elevated 

test temperature had no significant impact on shear properties of unstitched material. Similarly, the presence of 

stitching alone had only small effects (less than 6% reductions) on shear properties at ambient temperature. 

However, the presence of stitching at elevated test temperatures resulted in reductions in shear modulus of up to 

17% and reductions in shear strength of up to 23% relative to unstitched, room temperature strength. The effect of 

stitching could not be assessed on Mode I fracture toughness because of unsteady crack growth. Mode II fracture 

toughness was measured on both stitched and unstitched specimen, and the stitching clearly increased GIIC but 

caused an increase in the variability of that property. 

References

1. Velicki, Alex and D. Jegley. PRSEUS Development For The Hybrid Wing Body Aircraft, 11th AIAA 

Aviation Technology, Integration, and Operations (ATIO) Conference; AIAA-2011-7025. 

2. Jegley, Dawn C, and A. Velikci. Status of Advanced Stitched Unitized Composite Aircraft Structure, 51st 

AIAA Aerospace Sciences Meeting; AIAA-2013-0410. 

3. Dow, Marvin. and H B. Dexter. Development of Stitched, Braided, and Woven Composite Structures In 

the Act Program and at Langley Research Center (1985 to 1997) – Summary and Bibliography. NASA TP97-206234, November 1997, 

http://ntrs.nasa.gov/?N=123&Ntk=All&Ntt=Dow%2C%20marvin&Ntx=mode matchallpartial

4. Dexter, H. Benson. and B. A. Stein. Advanced Composite Materials for Airframe Structures, Fiber-TEX 

1987- First Conference on Advanced Engineering Fibers and Textile Structures for Composites, NASA CP 

3001. 

5. Jegley, Dawn C. Behavior of Frame-Stiffened Composites Panels With Damage. 54th AIAA/ASME/ASCE/ 

AHS/ASC Structures, Structural Dynamics, and Materials Conference; AIAA 2013-1738. 

6. Glaessgen, E.H. and I. S. Raju, Debonding of Stitched Composite Sandwich Structures, 41st Structures, 

Structural Dynamics, and Materials Conference and Exhibit; AIAA-2000-1614. 

7. Glaessgen, E.H., I. S. Raju, and C. C. Poe. Fracture Mechanics Analysis of Stitched Stiffener-Skin 

Debonding, 39th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials 

Conference and Exhibit; AIAA-98-2022. 

Figure 1. PRSEUS structure stringer-frame intersection
. 

7

Figure 2. Cross-section of typical PRSEUS skin, tear strap and stringer 

Figure 3. Stitching scheme used in V-notched shear specimen preparation 
. 

8

Figure 5. Tensile specimen tab fracture surface after delamination 

Figure 4. Specimens used in the grip pressure experiment, showing typical disbond failure as seen in all tensile test 

specimens. Arrows indicate tab delaminations. Similar tab failures were observed in all tensile specimens. 
. 

9

Table 2. Tensile test results 

(a) 72°F

Specimen

Unstitched Stitched

Strength,

ksi

Modulus,

Msi

Poisson's  

Ratio

Strength,

ksi

Modulus,

Msi

Poisson's

Ratio

1 90.2 8.36 0.057 91.9 8.34 0.057

2 96.4 8.23 0.052 89.3 8.05 0.052

3 109.9 8.58 0.060 92.6 7.92 0.060

4 111.8 8.41 0.057 91.1 7.95 0.057

5 62.4 8.16 0.051 87.4 8.17 0.051

6 110.4 8.67 0.053 80.9 7.90 0.053

Average 96.8 8.40 0.055 88.9 8.05 0.055 [103.7]b

Standard Deviation 19.0 0.20 0.003 4.33 0.17 0.003 [9.81]b

COVa 19.6% 2.36% 6.19% 4.87% 2.12% 6.19% [9.46%]b

(b) 180°F

Specimen

Unstitched Stitched

Strength,

ksi

Modulus,

Msi

Poisson's  

Ratio

Strength,

ksi

Modulus,

Msi

Poisson's

Ratio

1 100.0 8.41 0.056 97.4 8.18 0.056

2 112.8 8.34 0.046 90.7 8.06 0.040

3 111.0 8.44 0.043 84.4 8.56 0.057

4 100.1 8.31 0.050 97.7 8.01 0.062

5 57.7 8.20 0.050 96.3 7.69 0.041

6 113.0 7.95 0.047 94.0 7.93 0.050

Average

99.1 8.27 0.049 93.4 8.07 0.051 [107.4]b

Standard Deviation 21.2 0.18 0.005 5.14 0.29 0.009 [6.74]b

COVa 21.4% 2.19% 9.35% 5.50% 3.58% 18.1% [6.27%]b

Note: Tab delamination on all tensile specimens may have induced premature tensile failure. The reported tensile 

data may underestimate actual material strength and modulus. 

a

 Coefficient of variation 

b

 Considering specimen #5 of both unstitched specimen sets an outlier 
. 

10

Table 3. Compression test results, stitched material 

Specimen

72°F 180°F

Strength,

ksi

Modulus,

Msi

Strength,

ksi

Modulus,

Msi

1 90.2  ‐‐  91.5 7.52

2 90.7  ‐‐  78.5 7.51

3 88.5  ‐‐  81.6 6.65

4 85.7 7.65 82.7 7.30

5 78.6 7.46 71.3 7.30

6 93.1 7.56 81.4 7.17

Average 87.8 7.56 81.2 7.24

Standard Deviation 5.14 0.09 6.54 0.32

COVa 5.85% 1.19% 8.05% 4.42%

-- No strain data recorded

a

 Coefficient of variation
. 

11

Table 4. Inplane shear test results, unstitched material 

(a)  72°F

Specimen Chord Modulus,

Msi

Ultimate

Strength,

ksi

5% Offset

Strength,

ksi

1 0.73 13.0  ‐‐a

2 0.75 14.2 13.1

3 0.73 14.4 13.1

4 0.67 14.6 13.1

5 0.73 14.5 13.0

6 0.72 14.6 12.9

Average 0.72 14.4 13.1

Standard Deviation 0.03 0.17 0.04

COVb 3.67% 1.15% 0.33%

(b)  180°F

Specimen Chord Modulus,

Msi

Ultimate

Strength,

ksi

5% Offset

Strength,

ksi

1 0.54 11.2 9.05

2 0.60 11.2 10.1

3 0.74 14.3 13.2

4 0.71 14.1 13.4

5 0.75 13.4 12.9

6 0.72 14.1 13.0

Average 0.68 13.1 11.9

[0.73]c [13.99]c [13.11]c

Standard Deviation 0.09 1.34 1.70

[0.02]c [0.368]c [0.211]c

COVb 13.1% 10.2% 14.3%

[2.59%]c [2.63%]c [1.61%]c

a Strain gage failure prior to this point 

b Coefficient of variation 

c With specimens #1 and #2 considered to be outliers 
. 

12

Table 5. Inplane shear test results, stitched through the notch 

(a)  72°F

Specimen Chord Modulus,

Msi

Ultimate

Strength,

ksi

5% Offset

Strength,

ksi

1 0.76 14.4 13.0

2 ‐‐a ‐‐a ‐‐a

3 0.73 14.0 12.9

4 0.73 13.7 12.9

5 0.74 13.2 12.7

6 0.70 13.8 12.7

Average 0.73 13.7 12.8

Standard Deviation 0.01 0.27 0.13

COVb 1.87% 2.00% 0.99%

(b)  180°F

Specimen Chord Modulus,

Msi

Ultimate

Strength,

ksi

5% Offset

Strength,

ksi

1 0.61 11.8 10.3

2 0.59 11.5 10.3

3 0.61 11.5 10.1

4 0.60 11.9 10.4

5 0.61 12.1 10.5

6 0.57 10.9 9.5

Average 0.59 11.6 10.3

Standard Deviation 0.02 0.40 0.19

COVb 2.71% 3.44% 10.3%

a

 Data acquisition failure 

b

 Coefficient of variation 
. 

13

Table 6. Inplane shear test results, stitched around notch 

(a)  72°F

Specimen Chord Modulus,

Msi

Ultimate

Strength,

ksi

5% Offset

Strength,

ksi

1 0.76 14.6 14.6

2 0.76 15.9 12.8

3 0.75 14.4 12.8

4 0.76 13.6 12.9

5 0.76 15.5 14.8

6 0.78 14.9 13.2

Average 0.76 14.9 13.8

Standard Deviation 0.01 0.80 1.00

COVa 1.37% 5.38% 7.22%

(b)  180F

Specimen Chord Modulus,

Msi

Ultimate

Strength,

ksi

5% Offset

Strength,

ksi

1 0.63 11.8 10.1

2 0.61 11.5 10.0

3 0.62 11.6 9.9

4 0.64 12.1 10.2

5 0.63 13.0 10.1

6 0.63 12.0 10.1

Average 0.63 12.0 10.0

Standard Deviation 0.01 0.53 0.10

COVa 1.17% 4.43% 1.00%

a

 Coefficient of variation

Table 7. Knockdown on shear properties due to stitching and elevated test temperature. All knockdowns are relative to 

unstitched material tested at 72°F, and considering specimens #1 and #2 in Table 4(b) to be outliers. 

Property Test

Temperature

Stitching

US SA ST

Shear Modulus 72°F ‐‐  5.9% 0.8%

180°F 1.3%  ‐13.2%  ‐17.4%

Ultimate Shear

Strength

72°F ‐‐  2.8%  ‐5.5%

180°F ‐3.2%  ‐16.6%  ‐23.2%

5% Offset Shear

Strength

72°F ‐‐  ‐5.5%  ‐2.2%

180°F 0.3%  ‐23.2%  ‐21.1%

US = unstitched; SA = stitched around notch; ST = stitched through notch
. 

14

Table 8. Mode I fracture toughness testing as a function of crack length at ambient temperature 

Specimen #1 Specimen #2 Specimen #3 Specimen #4 Specimen #5 

Crack Length, GI, Crack Length, GI, Crack Length, GI, Crack Length, GI, Crack Length, GI, 

in in∙lbf/in2 in in∙lbf/in2 in in∙lbf/in2 in in∙lbf/in2 in in∙lbf/in2 

1.59 0.17 1.84 2.61 1.81 2.14 1.86 2.60 1.79 2.06 1.68 0.42 1.86 2.69 1.83 2.36 1.88 2.92 1.86 2.28 1.71 2.39 1.91 2.89 1.86 2.44 1.94 3.06 2.05 2.90 1.75 2.56 2.08 3.42 2.11 3.05 1.99 3.09 2.25 3.35 1.79 2.67 2.31 3.93 2.24 3.52 2.16 3.41 2.40 3.36 1.97 3.22 2.49 3.86 2.46 3.93 2.37 3.51 2.56 3.17 2.13 3.40 2.67 3.73 2.72 3.65 2.52 3.17 2.73 2.98 2.31 3.28 2.85 3.78 2.88 3.72 2.72 2.99 2.94 3.12 2.58 3.28 3.06 3.78 3.07 3.38 2.93 3.05 3.01 3.21 2.78 3.09 3.28 3.53 3.26 3.36 3.25 3.35 3.22 3.32 2.98 3.27 3.46 3.26 3.47 3.41 3.30 3.01 3.34 3.13 3.19 3.34 3.50 3.21 3.64 3.52 3.45 3.14 3.65 3.16 3.37 3.50 3.68 3.01 3.70 3.34 3.68 3.03 3.65 2.76 3.55 3.34 3.69 2.99 3.78 3.25 3.71 2.77 3.70 2.81 3.60 2.98 3.72 3.04 3.81 3.30 3.73 2.82 3.74 2.79 3.63 2.86 3.80 3.23 3.83 3.29 3.78 2.64 3.81 2.80 3.65 2.75 3.84 3.23 3.81 2.67 3.86 2.83 3.67 2.75 3.86 2.77 3.76 2.79 3.84 2.72
. 

15

Table 9. Precracked GIC, unstitched material tested at 72°F. 

Specimen GIC,

lbf‐in/in2

1 2.24

2 2.36

3 1.98

4 2.28

5 2.26

6 2.29

Average 2.23

Standard Deviation 0.13

COV† 5.83% †

   Coefficient of variation

Table 10. Mode II fracture toughness of PRSEUS laminates, tested at 72°F and 180°F, with and without throughthickness stitching. 

Test

Temperature,

°F

Specimen

Specimen GII,

in‐lbf/in2

Average GII (SD / COV)†,  

in‐lbf/in2

Non‐

Precracked Precracked Non‐Precracked Precracked

Unstitched

72

1 1.96 1.53

2.11

(0.37 /

17.5%)

1.63

(0.110 /

6.75%)

2 2.30 1.69

3 2.22 1.46

4 1.99 1.74

5 1.55 1.69

6 2.66 1.69

180

1 2.01 1.57

2.03

(0.100 /

4.93%)

1.69

(0.130 /

7.69%)

2 2.07  ‐‐ 

3 2.06 1.87

4 1.97 1.59

5 2.18 1.77

6 1.89 1.62

Stitched

72

1 3.06 3.85

2.49

(0.550 /

22.1%)

3.02

(0.710 /

23.5%)

2 2.61 2.51

3 2.37 2.05

4 1.63 3.46

5 3.07 2.66

6 2.24 3.58

180

1 2.85 2.91

2.97

(0.419 /

14.1%)

3.66

(0.739 /

20.2%)

2 2.91 2.80

3 ‐‐  ‐‐ 

4 3.19 4.30

5 3.51 4.29

6 2.38 3.98

†

 (Standard deviation / Coefficient of variation) 

-- Missing or faulty data
. 

16

Appendix: Post‐test photographs of test specimens

Figure 6. Tensile specimens with final failure location marked, no stitching, tested at 72°F. 

Figure 7. Tensile specimens with final failure location marked, no stitching, tested at 180°F. 
. 

17

Figure 8. Tensile specimens with final failure location marked, stitched material, tested at 72°F. 

Figure 9. Tensile specimens with final failure location marked, stitched material, tested at 180°F. 
. 

18

Figure 10. Edge-on view of unnotched compression specimens tested at 72°F 

Figure 11. Edge-on view of unnotched compression specimens tested at 180°F 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Material Property Characterization of AS4/VRM-34 Textile Laminates 

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Grenoble, Ray W.; Johnston, William M.

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

NASA Langley Research Center

Hampton, VA 23681-2199

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-20276

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 01

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

Several material properties (modulus, strengths, and fracture toughness) of a textile composite have been evaluated to provide 

input data to analytical models of PRSEUS structure. The material system is based on warp-knitted preforms of AS4 carbon 

fibers and VRM-34 epoxy resin, which have been processed via resin infusion and oven curing. Tensile, compressive, shear, 

and fracture toughness properties have been measured at ambient and elevated temperatures. All specimens were tested in 

as-fabricated (dry) condition. Specimens were tested with and without through-thickness stitching. 

15. SUBJECT TERMS

Composites; Hybrid wing body; PRSEUS; Stitching; Textiles

18. NUMBER

 OF 

 PAGES

23

19b. TELEPHONE NUMBER (Include area code)

(443) 757-5802

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

 ABSTRACT

UU

Prescribed by ANSI Std. Z39.18

Standard Form 298 (Rev. 8-98)

3. DATES COVERED (From - To)

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

699959.02.22.07.01.01 

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2013-218050

16. SECURITY CLASSIFICATION OF:

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 Operations 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)

01- 10 - 2013