NASA/TM–2009-215955

Experimental Behavior of Fatigued Single

Stiffener PRSEUS Specimens

Dawn C. Jegley

Langley Research Center, Hampton, Virginia

December 2009
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NASA/TM–2009-215955

Experimental Behavior of Fatigued Single

Stiffener PRSEUS Specimens

Dawn C. Jegley

Langley Research Center, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-2199

December 2009
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.

Available from:

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076-1320

443-757-5802
Abstract

NASA, the Air Force Research Laboratory and The Boeing Company

have worked to develop new low-cost, light-weight composite structures

for aircraft. A Pultruded Rod Stitched Efficient Unitized Structure

(PRSEUS) concept has been developed which offers advantages over

traditional metallic structure. In this concept a stitched carbon-epoxy

material system has been developed with the potential for reducing the

weight and cost of transport aircraft structure by eliminating fasteners,

thereby reducing part count and labor. By adding unidirectional carbon

rods to the top of stiffeners, the panel becomes more structurally

efficient. This combination produces a more damage tolerant design.

This document describes the results of experimentation on PRSEUS

specimens loaded in unidirectional compression in fatigue and to failure.

Introduction

NASA, the Air Force Research Laboratory and The Boeing Company have worked to develop new lowcost, light-weight composite structures for aircraft. A Pultruded Rod Stitched Efficient Unitized Structure

(PRSEUS) concept has been developed which offers advantages over traditional metallic structure (ref 1-

4). In this concept a stitched carbon-epoxy material system has been developed with the potential for

reducing the weight and cost of commercial transport aircraft structure. By stitching through the

thickness of a dry carbon-epoxy material system, the labor associated with panel fabrication and assembly

can be significantly reduced. When stitching through the thickness of pre-stacked skin, stringers,

intercostals and spar caps, the need for mechanical fasteners is almost eliminated. This manufacturing

approach reduces part count, and therefore, cost of the structure. In addition, stitching reduces

delamination and improves damage tolerance, allowing for a lighter structure with more gradual failures

than traditional composites without through-the-thickness reinforcement. However, the PRSEUS concept

is relatively new and the behavior of PRSEUS specimens must be evaluated to understand both fatigue

and failure response.

The PRSEUS concept consists of carbon-epoxy panels fabricated from dry components and then infused

using high temperature and vacuum pressure only. No autoclave is required. Skins, flanges and webs are

composed of layers of graphite material forms that are prekitted in multi-ply stacks using Hercules, Inc.

AS4 fibers. Several stacks of the prekitted material are used to build up the desired thickness and

configuration. Specimens are stitched together using Vectran fibers. Stiffener flanges were stitched to

the skin and no mechanical fasteners are used for joining. To maintain the panel geometry during

fabrication, first stiffeners and then the skin are placed in a stitching tool for assembly prior to moving to

a curing tool for consolidation in the oven. The stiffeners running in the axial direction consist of webs

with a bulb of unidirectional carbon fiber rods at the top of the web. AS4 carbon fiber overwraps

surround the bulb. The stiffeners in the lateral direction are foam filled hats.
Rod

Stiffener

In the current study each prekitted stack had a [45/-45/0 2/90/02/-45/45]T laminate stacking sequence.

Stack thickness was approximately 0.052 inches. For this study specimens were cut from a large PRSEUS

panel that had seven rod-stiffeners with 6-inch spacing and three frames with 20-inch spacing. This panel

was fabricated under a NASA contract to The Boeing Company entitled "Damage Arresting Composites

for Shaped Vehicles." A photograph of this panel is shown in figure 1. All specimens were made from

AS4 fibers and HexFlow VRM 34 resin, as described in ref. 1. The rods were Toray unidirectional T800

fiber with a 3900-2B resin and frame stiffeners were filled with Rohacell foam. A sketch of the

intersection of rod and frame elements is shown in figure 2.

Figure 1. Cured panel prior to extraction of test specimens.

Figure 2. PRSEUS stiffener intersection.

Test Specimen Description

Two specimens with a single axial stiffener, herein identified as rod-stiffened specimens R3 and R4, and

two specimens with a central hat-stiffener representative of a fuselage frame, herein identified as frame

specimens F3 and F4, were examined in this study. Sketches of the cross section of each type of

specimen are shown in figure 3 and photographs of each specimen are shown in figure 4. Nominal skin

2
thickness was 0.052 inches. Rod-stiffeners had a 3.4-inch wide flange, a 0.104 inch-thick stiffener web

and a 1.5-inch tall stiffener. Flange thickness was half the blade or web thickness. One stack of

additional material was added under each flange as tear straps or stiffener caps. These reinforcing layers

covered the same area of skin as the stiffener flange. Frame stiffeners had a 4.0-inch wide flange and a

6.0-inch tall stiffener. Rod-stiffened specimens were 18 inches tall and 6 inches wide while frame

specimens were 16.6 inches tall and 10 inches wide. Prior to testing, each end of the specimen was potted

in 1.0-inch-deep epoxy compound and the ends were ground flat and parallel to each other to ensure

uniform load introduction. Rod-stiffened specimens contained no frames while frame specimens

contained one frame in the loading direction and two rod-stiffeners which were six inches apart,

perpendicular to the frames.

a) Rod stiffener.

b) Frame stiffener.

Figure 3. Stiffener geometry. Dimensions are in inches.
b) Rod specimen R4

a) Rod specimen R3

c) Frame specimen F3

Figure 4. Test specimens.
d) Frame specimen F4

Figure 4. Test specimens-continued.

Detailed measurements of skin and stiffener thicknesses are shown in figure 5. Skin thickness of

specimen R3 range from .052 to .061 inches. Skin thickness of specimen R4 range from .049 to .060

inches. Skin thickness of specimen F3 range from .044 to .056 inches. Skin thickness of specimen F4

range from .047 to .060 inches. Rod region diameter averaged .485 inches and frame width averaged .665

inches. Prior to attaching strain gages, the unstiffened surface, herein referred to as the Outer Mold Line

(OML), each specimen was scanned using a Coordinate Measurement Machine to document the geometry

of the unstiffened surface. The result of these scans is shown in figure 6. Surface measurements were

found to range from 0.015 to -.025 inches in the rod specimens and -.008 to .005 inches in the frame

specimens. These imperfections correspond to approximately 25% of the skin thickness for the rod

specimens and 10% of the skin thickness for the frame specimens.

Left edge

skin

Right edge

skin

Stiffener

R3 0.0569 0.0607 0.4891

0.0519 0.04845 0.4825

0.0565 0.05735 0.4882

R3 average 0.0551 0.0555 0.4866

R4 0.05670 0.05600 0.48445

0.04905 0.05370 0.48930

0.05790 0.05925 0.47915

R4 average 0.05455 0.05631 0.48430

F3 0.05615 0.04755 0.66365

0.05430 0.04450 0.6660

0.05645 0.04635 0.6618

F3 average 0.05563 0.04613 0.66381

F4 0.05410 0.04705 0.66285

0.05475 0.0552 0.66315

0.06015 0.0514 0.67210

F4 average 0.05633 0.05121 0.66603

Figure 5. Measured thicknesses. Dimensions are in inches.
b) Frame specimens

Figure 6. Measured out-of-plane imperfections prior to testing. Dimensions are in inches.

6
Prior to delivery to NASA the OML of specimen F4 was damaged during the machining process. A

photograph of this damage is shown in figure 7. The dent is approximately circular with a 1-inch

diameter and the maximum depth is approximately 0.1 inches. No damage at this location is visible on

the stiffened or Inner Mold Line (IML) side of the specimen.

Figure 7. Machining damage to specimen F4.

Test Procedure and Instrumentation

All specimens were loaded in axial compression. The applied spectrum of 55,000 cycles used for the

applied fatigue load. The 55,000 cycles was broken into 11 blocks of 5000 cycles with the maximum

loads and load rates shown in table 1 for each cycle. Specimens were loaded in compression only and

load was not removed between cycles. After the completion of the fatigue loading, each specimen was

loaded to failure in axial compression. Fatigue loading was typically applied at the rate of 1 cycle per

second while failure loading was applied at the rate of 5,000 to 10,000 lb/min. Selected intermediate

fatigue cycles were performed at a rate of 4,000 to 10,000 lb/min.

Displacement and strain gage data were recorded at the rate of once every second as load was applied

during the initial cycle, selected fatigue cycles and during the failure (post-fatigue) test as shown in table

2. Displacements were measured using three displacement transducers measuring end-shortening, and

one measuring stiffener motion. An additional transducer measuring out-of-plane motion of the skin at

the midlength location was included for the frame specimens. Nine strain gages were added to each rod

specimen and ten strain gages were added to each frame specimen in the pattern shown in figures 8. The

additional gage on the frame specimens was located on the frame web. Back-to-back strain gages were

used to monitor strains in the skin, flanges and stiffeners of each specimen. Metal supports were added to

the unloaded edges to restrain out-of-plane motion of the skin while allowing the specimen to deform inplane. These edge supports for the rod specimens supported IML and OML the full length of the

unloaded edge while the lateral rod-stiffeners prevented the edge supports from being continuous along

the full length of the frame specimen on the IML. Specimens in the test machine are shown in figure 9.

Buckling and failure behavior were noted for each specimen.

An optical measurement system was used to obtain full field displacement and strain results for the OML

of each specimen during the post-fatigue test. A photograph showing the Vision Image Correlation

(VIC)4 system is shown in figure 10. A black and white speckle pattern painted on the specimen surface

allows the VIC system to track surface motion. Two cameras positioned at different angles to the

specimen surface simultaneously photograph the specimen at set intervals during to test. In this case, the

specimen was photographed every 5 seconds, resulting in approximately 150 time steps of data. The VIC

7
Rod specimen Frame specimen

Figure 8. Strain gage patterns. All dimensions are in inches.

Figure 8. Specimens in test machine.

Figure 9:. Vision Image Correlation system.
*0,000

20,000

10,000

system also records the load from the test machine so images can be related to the corresponding load.

The VIC system compares the photographic images and produces three-dimensional displacements and

strains at each time step. Since the full-field results are dependent upon the cameras having an

unobstructed view of the surface, some local areas where strain gages or wires are removed from the plot

or identified as not representing surface behavior in some figures.

Results and Discussion

Displacements, strains and failure modes are presented herein, followed by a comparison with data

obtained previously for specimens subjected only to static loading. Full field results show displacement

and strain distributions as the specimen approached failure.

Displacements and Loads

A comparison of measured displacement as a function of load for end-shortening and out-of-plane

deformation from the pre-fatigue loading to 20,500 lb and the post-fatigue loading to failure of rod

specimens R3 and R4 is shown in figures 1 1a and 1 1b, respectively. Measured displacements for the prefatigue loading to 41,000 lb and the post-fatigue loading to failure of frame specimens F3 and F4 is

shown in figures 12a and 12b, respectively. Data for the pre-fatigue loading are shown as solid lines and

data for the post-fatigue loading are shown as dashed lines. These results indicate that there is no

noticeable change in behavior due to the fatigue cycling.

50,000

End-shortening

40,000 1___N

Lateral web \

01

Load, lb motion

Pre-fatigue

Post-fatigue

0 1 1 oI I I I 1 1

-0.04 -0.02 0.00 0.02 0.04 0.06 0.08 0.10

Displacement, inches

a) Specimen R3

Figure 11. Pre-fatigue and post-fatigue displacements for rod specimens.

9
Lateral web

50,000 motion

`,

40,000

Load, lb y

30,000 i

20,000

Transducer

10,000 slipped

End-shortening

Pre-fatigue

Post-fatigue

0

-0.04 -0.02 0 0.02 0.04 0.06 0.08 0.10

Displacement, inches

b) Specimen R4

Figure 11. Pre-fatigue and post-fatigue displacements for rod specimens-continued.

100,000

Lateral web

Out-of-plane motion

80,000 — !

Load, lb

60,000

(0,000

^^ I

20,000

Pre-fatigue 1

Post-fatig ue tt

0

-0.15 -0.10 -0.05 0.00

Displacement, inches

MAI

Mea:E rement

-loc o sn

0.05 0.1 0

End-shortening

J°

}}

a) Specimen F3

Figure 12. Pre-fatigue and post-fatigue displacements for frame specimens.

10
/ End-shortening

^l

ill

Pre-fatigue

Post-fatigue

70,000 Lateral web

motion

60,000

Load, lb

50,000

40,000 Out-ofplane

30,000 ^

20,000

10,000

0

-0.050 0.0 0.050 0.10

Displacement, inches

b) Specimen F4

Figure 12. Pre-fatigue and post-fatigue displacements for frame specimens-continued.

Rod specimens failed at loads of approximately 45,000 and 4B,654 lb. Frame specimens failed at loads of

85,152 and 64,999 kips. Measured displacements are shown in figures 13 and 14 for the post-fatigue

tests of rod and frame specimens, respectively. Data for specimens R3 and F3 are shown as solid lines

and data for specimens R4 and F4 are shown as dashed lines. End-shortening results indicate a uniform

displacement was applied to the end of specimens R3 and R4. For example, end-shortening results

among the three transducers differ by no more than 2% for a load of 4,000 lb and a load of 5% for a load

of 40,000 lb for specimen R4. Similar uniform shortening results are seen for specimen F3 however some

discontinuities in he slope of he end-shortening results for specimens F3 and F4 indicate some slippage in

the transducers or in the platen for loads greater than 40,000 lb. This discontinuity could not be

reproduced for a dummy specimen using the same test arrangement.

11
*0'000

Load ' lb

)0'000

40'000

Endshortening

50,000

Buckling

40,000 End-shortening

Lateral

Load, lb web

motion

30,000

b

20,000 4

m ,

10,000 ,

++ , Specimen R3

Specimen R4

0 1

v

-0.04 -0.02 0.00 0.02 0.04 0.06 0.08 0.10

Displacement, inches

Figure 13. Post-fatigue displacements for rod specimens.

100'000

Out-of-plane Lateral web

motion

20'000

— Specimen F3

— Specimen F4

0

-0.15 -0.10 -0.05 0.00 0.05 0.10

Displacement' inches

Figure 19. Post-fatigue displacements for frame specimens.

12
The transducer on the stiffener indicates rolling of the web of the rod specimen at approximately 42,000

lb. The frame of specimens F3 and F4 did not roll. The out-of-plane displacement measurement at the

axial centerline of the frame specimens indicates that the edge supports did not fully restrain out-of-plane

motion at the unloaded edges. These restraints consisted of individual aluminum pieces held at the top

and bottom of the specimen but they were not continuous on the stiffened side because the rod stiffeners

went to the edge of the specimen (see figure 4). In addition, these restraints were held in place by screws

and the vibration of the system during cycling caused the screws to back out. The screws were tightened

after every 2-3 blocks and no damage was evident based on this flexibility.

Full-field results for out-of-plane deformation at approximately the maximum applied load are shown in

figure 15 and 16 for rod and frame specimens, respectively. These results indicate rod specimens buckled

prior to failure but frame specimens did not. In the rod specimens the centerline where the rod stiffener

was located did not move out-of-plane during loading while the unloaded edges did move out of plane

despite the edge restraints. A buckle pattern of approximately three half waves developed in each rod

specimen. In each rod specimen one edge moved forward in the positive z direction while the other edge

moved in the negative z direction about the same magnitude. In the full-field results, regions where strain

gages and wires obscure the specimen surface are removed from the plot. In the undamaged frame

specimen the out-of-plane displacements were localized to the corners and midlength location. Bending

between the lateral support of the potting and rods is evident. The midlength location is half way between

rod stiffeners. The primary deformation seen in the damaged frame specimen is in the region of the

damage. The displacements shown are the difference between the surface prior to loading and just before

failure.

Figure 15. Full-field out-of-plane displacement for rod-stiffened specimens immediately prior to failure.

Dimensions are in inches.

13
a) Frame specimen F3 immediately prior to failure

b) Frame specimen F4 immediately prior to failure

Figure 16. Full-field out-of-plane displacement change from pre-fatigue condition for frame specimens. Dimensions

are in inches

14
Strain

The load-strain relationship for skin and stiffener locations for the first cycle in the fatigue spectrum and

the post-fatigue failure test are shown in figures 17 and 18 for rod and frame specimens, respectively.

Strain gage locations are shown in the figures. Data for the pre-fatigue loading are shown as solid lines

and data for the post-fatigue loading are shown as dashed lines with. A comparison of the skin, flange

and rod strains before and after fatigue cycling indicates that the fatigue cycling had no noticeable effect

on the strain behavior of specimens R3 and R4. A comparison of the skin, flange and frame strains before

and after fatigue cycling indicates that the fatigue cycling had no noticeable effect on the strain behavior

of specimens F3 and F4, despite the visible damage in specimen F4 prior to loading.

The load-strain relationship for post-fatigue failure tests of rod and frame specimens are shown in figures

19 and 20, respectively. Strain gage locations are shown in the figure and the green dot in the center of

the sketch indicates the location of the pre-test damage in specimen F4. The Strains indicate nonlinear

behavior initiates in the OML gages (1 and 2) at approximately 23,000 lb for specimen R4 and 38,000 lb

for specimen R3. Buckling is not indicted prior to approximately 38,000 lb in either specimen. Buckling

is indicated at approximately 38,000 lb for R3 and 42,000 lb for R4. Even though flange/skin gages 6 and

7 are back-to-back, they diverge at load initiation. Gages 8 and 9 show no divergence and little nonlinear

behavior in either rod specimen until failure.

Results for gages away from the pre-test damage site in specimen F4 and the equivalent location on

specimen F3 are shown in figure 20a whicle gages near the pretest damage site in specimen F4 and the

comparable loctions in specimen F3 are shown in figure 20b. These results indicate that the damage has

little impact on strains for loads less than 41,000 lb, which is the maximum load the specimens were

subjected to duing the fatigue cycling. However for loading greater than this level, the damage had

significant influence on strain behavior in the region of damage, causing loss of load-carrying ability at

the damaged skin at a load of approximately 50,000 lb rather than the 83,000 lb found in the undamaged

specimen.

Full-field results for axial strains at approximately the maximum applied load are shown in figures 21 and

22 for rod and frame specimens, respectively. The buckle pattern of three half waves in each direction is

consistent with the strain distribution shown for the rod specimens. Maximum strain is approximately

0.006 in./in. in magnitude in rod specimen R3. Peak strains are located above the axial centerline where

the maximum strain magnitude for specimen R3 is approximately 0.01 in./in. in tension in and

approximately -0.01 in/in. in compression at the same location in specimen R4. This difference in

direction is representative of the center buckle in one specimen moving toward the camera and in the

other specimen moving away from the camera.

In frame specimen F3 the peak axial strains in the OML are located near the specimen corner. In frame

specimen F4 peaks in axial strain also occur at the damage site. Since the VIC system only monitors the

OML, initiation of failure on the IML could not be captured in this manner.

15
50,000

40,000

Load, lb

30,000

20,000

10,000

OML corners IML flange

gages 1 and 2 gage 4

`

r OML gage 3

Pre-fatigue

Post-fatigue

0

-.010 -.008 -.00* -.004 -.002 0 .002

Strain, in./in.

a) Far field gages specimen R3

9ML gage 8

Rod gage 9

9ML gage 5 \

50,000 IML flange

gage 7 9ML gage 6

40,000

Load, lb _ \

30,000 ^ 1

20,000 ^5

10,000

Pre-fatigue

0

Post-fatigue

-.010 -.008 -.006 -.004 -.002 0 .002

Strain, in.ln.

b) Centerline gages specimen R3

Figure 17. Pre-fatigue and post-fatigue strains for rod specimens.

16
20,000

OML corners

gages 1 and 2

50,000

OML gage 3

IML flange

40,000 gage 4 ^

Load, lb

30,000 -

1

=^ ,

l

4-

10,000

Pre-fatigue

Post-fatigue

0

-.010 -.008 -.006 -.004 -.002 0 .002

Strain, in./in.

c) Far field gages specimen R4

IML flange OML gage 8 and

gage D Rod gage B

50,000 / OML gage 5

_411 OML gage 6

40,000

Load, lb

30,000

20,000

10,000

Pre-fatigue

0

Post-fatigue

-.010 -.008 -.006 -.004 -.002 0 .002

Strain, in.An.

d) Centerline gages specimen R4

Figure 17. Pre-fatigue and post-fatigue strains for rod specimens-continued.

17
Pre-fatigue

Post-fatigue OML corners

gages 1 and 2

100,000 r OML gage 3 \

80,000

Load, lb

60,000

40,000

IML flange

gage 4 \

20,000

0 1 VA 1

-0.,0050 -0.0025 0 0.0025 0.005

Strain, in.An.

a) Far field gages specimen F3

I9L flange

gage A \

100,000

80,000

Load, lb

60,000

40,000

20,000

Pre-fatigue

89L gage 8 Post-fatigue

frame web gage 10

89L gage D

89L gage 6

89L gage 5

A

0 1 A\1

-0.0050 -0.0025 0 0.0025

Strain, in./in.

0.005

b) Centerline gages specimen F3

Figure 18. Pre-fatigue and Post-fatigue strains for frame specimens.

18
Pre-fatigue OML gage 3 Post-fatigue

75,000 OML corner gage 1

OML corner gage 2

Load, lb

50,000

IML flange 't^^ ^^// gage 4 A 1^`/ rJ:lt I I IL 1-L,

25,000

0 1 v

--0.0050 -0.0025 0 0.0025 0.005

Strain, in./in.

c) Far field gages specimen F4

Pre-fatigue

Post-fatigue Frame web gage 10

OML gage 6

75,000

/ OML gage 9

/ OML gage 5

Load, lb ' -^ ^ ,,f

OML gage 8

50,000 ^^ r

s

IML flange

gage 7

25,000 SL9L

0 1 It

-0.0050 -0.0025 0 0.0025 0.005

Strain, in./in.

d) Centerline gages specimen F4

Figure 18. Pre-fatigue and Post-fatigue strains for frame specimens-continued.

19
Rod top

gage 9

50,000

i \+ 4

40,000 -

- ^ 5

` OML center

Load, lb `r ^ t

gage 8 t

30,000

20,000 OML corners

gages 1 and 2

10,000

Specimen R3

- - - - Specimen R4

0 1 1 1 •

-.008 -.006 -.004 -.002 0

Strain, in./in.

Figure 19. Post-fatigue strains for rod specimens.

OML corners

100,000 , gages 1 and 2

80,000 /

Load, lb

60,000 `I'i 1'

OML center

40,000 ^ 1 1

Frame

gage 8

k.^

' ti 1

gage 10

20,000

Specimen F3

Specimen F4

0 1 X%\ N

-.0050 -0.0025 0 0.0025 0.005 0.0075

Strain, in./in.

a) Away from damage site

Figure 20. Final test strains for frame specimens.

20
0.0075

OML 1 inch

from center

gage 6

100,000

80,000

Load, lb

60,000

40,000

20,000

V/gag

OML edge

gage 5

\^ e

OML center

8

Specimen F3

Specimen F4

OML center

0 gage 8

-0.0050 -0.0025 0 0.0025 0.0050

Strain, in./in.

b) Damaged region

Figure 20. Final test strains for frame specimens-continued.

Figure 21. Full field axial strains for rod specimens. Distance shown is in inches. Scale is in in./in.

21
Figure 22. Full field axial strains at failure for frame specimens. Distance shown is in inches. Scale is in

in./in.

Failure

The failure loads the four fatigued specimens as well as nominally identical specimens previously tested

under the Boeing contract are shown in table 3 and figure 23. Photographs of failed fatigued rod

specimens are shown in figure 24 and failed frame specimens are shown in figure 25.

Rod-stiffened specimens failed midlength across the full width of the specimen. Minimal delamination

can be seen between the flange and skin. Overwrap damage can be seen around the rod. Specimens R3

and R4 failed in the same manner. Frame specimen F4 was damaged prior to fatigue cycling but still

withstood the cycling with no indication of damage progression and failed at a load equal to 78 percent of

the failure load of specimen F3. Failure at the specimen corners and at the top of the location where the

rod overwrap intersects the frame is evident in both specimens F3 and F4.

22
V1s1ble dama\e pr1or to load1n

100,000 n 1 cycle to fa1lure n 55,000 cycles pr1or to fa1lure

80,000

Load, lb n 

\

1

60,000

40,000

20,000

0

Rod-st1ffened Frame spec1mens

Figure 23. Failure loads.

23
Figure 24. Rod specimen failure.

a) Specimen F3

b) Specimen F4

Figure 25. Frame specimen failure.

24
Comparison to past results

Four specimens that were nominally identical to the specimens described herein were loaded in a single

test to failure (ref 1). These specimens were cut from the same large panel as the fatigue specimens

described herein and tested under the Boeing contract. The failure loads for the NASA-tested and

Boeing-tested specimens are shown in table 3 and figure 23 for all eight specimens. Both fatigued rod

specimens failed at loads greater than one unfatigued specimen but less than the other unfatigued

specimen. Similarly, frame specimen F3 failed at a load greater than one unfatigued specimen but less

than the other unfatigued specimen. The fatigued specimens failed in the same manner as the unfatigued

specimens. There is no evidence the fatigue cycling had any effect on the failure loads of rod stiffened or

frame specimens. Photographs presented in reference 1 show the same failures as seen in specimens R3,

R4 and F3. No single-stiffener specimens discussed in reference 1 were damaged prior to loading so no

direct comparison to specimen F4 can be made.

Concluding Remarks

Based on experimental evaluation of nominally identical single-stiffener compression-loaded specimens

subJected one or 55,000 cycles of loading, fatigue cycling has no influence on failure load or mode.

Visible damage to the skin of a specimen where the frame was loaded in compression reduced the failure

load to 78 P of the failure load of the undamaged specimen.

References

1. Velicki, A., "Damage Arresting Composites for Shaped Vehicles," NASA CR-2009-215932 Sep., 2009.

2. Velicki, A., "Advanced Structural Concept Development Using Stitched Composites." Presented at the 49th

AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, Chicago Ill, April,

2008.

3. Jegley, D. C, Velicki, A., and Hansen, D. A., "Structural Efficiency Of Stitched Rod-stiffened Composite Panels

With Stiffener Crippling." Presented at the 49th AIAA/ASME/ASCE/AHS/ASC Structures, Structural

Dynamics and Materials Conference, Chicago Ill, April, 2008.

4. Air Vehicle Technology Integration program (AVTIP), Delivery Order 0059: Multi-role Bomber Structural

Analysis, AFRL-VVA-WP-TR-2006-3067, Krishna Hoffman, May 2006, Final Report for 14 December 2004-

08 May 2006, AFRL-VA-WP-TR-2006-3067

25
Table 1. Fatigue Spectrum for a single block *

Load (lb)

Rod-stiffened

Load (lb)

Frame-stiffened

Cycle 1 20,500 41,000

Cycle 2-14 19,500 39,000

Cycle 15-5000 18,500 37,000

*

5000-cycle blocks repeated 11 times for each specimen

Table 2 Load rate and data recorded through fatigue loading

Load Rate Data Recorded

Specimen R4

Block 1-11 Cycle

1-14

4000 lb/min Cycle 1 each block

Block 1-11 Cycle

15-5000

Approx. 20,000

lb/min

none

Specimen R3

Block 1-11

Cycle 1

5000 lb/min Cycle 1 for blocks 1,5,9

Block 1-11

Cycle 2-5000

Approx 20,000

lb/min

none

Specimen F3

Block 1-11 Cycle

1-14

10,000 lb/min Cycle 1 blocks 1, 2, 6

Block 1-11 Cycle

15-5000

Approx. 40,000

lb/min

none

Specimen F4

Block 1-11

Cycle 1

10,000 lb/min Cycle 1 for blocks 1, 6

Block 1-11

Cycle 2-5000

Approx 40,000

lb/min

none

Table 3 Specimen failure loads *

Specimen Test facility Test type Failure load (lb)

Rod 1 Boeing static 48,500

Rod 2 Boeing static 41,400

R3 NASA Fatigue/static 45,000

R4 NASA Fatigue/static 47,654

Frame 1 Boeing static 86,000

Frame 2 Boeing static 79,500

F3 NASA Fatigue/static 85,152

F4 NASA Fatigue/static

(damaged)

64,999

Data for specimens at the Boeing test facilty are presented in reference 1.

26
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01-12 - 2009 Technical Memorandum

4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER

Experimental Behavior of Fatigued Single Stiffener PRSEUS Specimens

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S) 5d. PROJECT NUMBER

Jegley, Dawn C.

5e. TASK NUMBER

5f. WORK UNIT NUMBER

699959.02.08.07.02.02

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

NASA Langley Research Center REPORT NUMBER

Hampton, VA 23681-2199

L-19782

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National Aeronautics and Space Administration NASA

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12. DISTRIBUTION/AVAILABILITY STATEMENT

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Availability: NASA CASI (443) 757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

NASA, the Air Force Research Laboratory and The Boeing Company have worked to develop new low-cost, light-weight

composite structures for aircraft. A Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept has been

developed which offers advantages over traditional metallic structure. In this concept a stitched carbon-epoxy material system

has been developed with the potential for reducing the weight and cost of transport aircraft structure by eliminating fasteners,

thereby reducing part count and labor. By adding unidirectional carbon rods to the top of stiffeners, the panel becomes more

structurally efficient. This combination produces a more damage tolerant design. This document describes the results of

experimentation on PRSEUS specimens loaded in unidirectional compression in fatigue and to failure.

15. SUBJECT TERMS

Composites; Structurally efficient; Lightweight; Graphite-epoxy; Hybrid wing body; Out-of-autoclave; VARTM; Stitched;

Buckling

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