- 

NASA Contractor Report 178409 

Composite Transport Wing 

Technology Development 

Dr. Ram C. Madan 

Douglas Aircraft Company 

McDonnell Douglas Corporation 

Long Beach, CA 90846 

Contract NAS1-17970 

February 1988 

(N& S A-CB- 7 78U 09 1 

TBCHNOLOGY DEVELOPMENT (Douglas Aircraft 

C OH POSITE TEAlQSPORT YI #G R 89- 20 8 42 

CO,), 92 p CSCL OIC 

Unclas 

63/05 0222714 

National Aeronautics and 4 I 

Space Administration 

Langley Research Center 

Hampton, Virginia 23665 
NASA Contractor Report 178409 

Composite Transport Wing 

Technology Development 

Dr. Ram C. Madan 

Douglas Aircraft Company 

McDonnell Douglas Corporation 

Long Beach, CA 90846 

Contract NAS1-17970 

February 1988 

National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23665 
ABSTRACT 

This report deals with design, fabrication, testing, and analysis of stiffened wing cover panels to assess 

damage tolerance criteria. The damage tolerance improvements were demonstrated in a test program 

using full-sized cover panel subcomponents. The panels utilized a “hard” skin concept with identical 

laminates of 44percent @degree, 44-percent k 45-degree, and 12-percent Wdegree plies in the skins 

and stiffeners. The panel skins were impacted at midbay between the stiffeners, directly over the stiffener, and over the stiffener flange edge. The stiffener blades were impacted laterally. Impact energy 

levels of 100 ft-lb and 200 ft-lb were used. NASTRAN finite-element analyses were performed to 

simulate the nonvisible damage that was detected in the panels by nondestructive inspection. A closedform solution for generalized loading was developed to evaluate the peel stresses in the bonded structure. 

Two-dimensional delamination growth analysis was developed using the principle of minimum potential energy in terms of closed-form solution for critical strain. An analysis was conducted to determine 

the residual compressive stress in the panels after impact damage, and the analytical predictions were 

verified by compression testing of the damaged panels. 

PRECEDING PAGE BLANK NOT FILMED 

... 

Ill 
FOREWORD 

This final report was prepared by Douglas Aircraft Company of the McDonnell Douglas Corporation, 

at Long Beach, California, under NASA Contract NASI-17970. It covers work performed between July 

1985 and May 1987. The program was sponsored by the National Aeronautics and Space Administration, Langley Research Center (NASA-LRC), Hampton, Virginia. Dr. Mark Shuart and Dr. J. G. 

Davis, Jr., were the NASA-LRC technical representatives for this project. 

Uses of commercial products or names of manufacturers in this report do not constitute official 

endorsement of such products or manufacturers, either expressed or implied, by the National Aeronautics and Space Administration. 

This report was prepared by the project manager, Dr. Ram C. Madan of Douglas Aircraft Company. 

The following Douglas personnel were principal contributors to the program: 

Program and Technical Management Structural Design and Analysis 

Dr. H. C. Schjelderup 

M. Klotzsche 

Project Manager 

Dr. R. C. Madan 

B. L. Bunin* 

Business Management 

T. E. Alvarado-Lee 

M. E. Eike 

Stress Analysis 

J. 0. Sutton 

K. A. Walker 

B. A. Hanson 

*Initially headed project. 

* 

Dr. R. C. Madan 

J. 0. Sutton 

D. A. Barkey 

R. E. Sagui 

Materials and Process Engineering 

M. F. Murphy 

Structural Testing and Coordination 

M. F. Murphy 

Dr. R. C. Madan 

J. 0. Sutton 

E. E. Hayman 

Technical Advisors 

J. 0. Sutton 

M. Ashizawa 

A. V. Hawley 

J, Genicki 
CONTENTS 

Section Page 

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

2 BASELINE WING DESIGN .............................................. 3 

2.1 Material Selection .................................................... 3 

2.2 Conceptual Design ................................................... 7 

3 TEST SPECIMENS AND PROCEDURES .................................. 13 

3.1 Specimen Description ................................................. 13 

3.2 Specimen Fabrication ................................................. 19 

3.2.1 Stringer Fabrication ...... .................................... 19 

3.2.2 Skin Fabrication .............................................. 20 

3.2.3 Skin-Stringer Assembly ........................................ 20 

3.2.4 Demonstration Panels ......................................... 21 

3.3 Test Procedures ..................................................... 22 

3.3.1 Ancillary Tests ................................................ 22 

3.3.2 Development Tests ............................................ 24 

3.3.3 Test Plan for Demonstration Panels .............................. 30 

4.1 NASTRAN Model ................................................... 31 

4.1.1 Analysis ..................................................... 32 

4.1.2 Refined Analysis .............................................. 33 

4.1.3 Specimen and Support Conditions ................................ 36 

4.2 Delamination Growth - An Approximate Solution ........................ 36 

4.3 Peel Stress Analysis .................................................. 43 

4.3.1 Analysis ..................................................... 43 

4.3.2 Results ...................................................... 48 

4.4 Residual Strength Analysis ............................................ 52 

4.4.1 Effect of a Finite-Width Panel ................................... 55 

4.4.2 Stress Concentration Factor ..................................... 55 

4.4.3 Residual Strength ............................................. 56 

4.4.4 Analysis Method .............................................. 57 

5 TEST RESULTS AND DISCUSSION ...................................... 59 

5.1 Ancillary Tests ...................................................... 59 

5.1.1 Tension and Compression Tests .................................. 59 

5.1.2 Open-Hole Compression Tests ................................... 59 

5.1.3 Compression-Bearing Tests ..................................... 60 

5.1.4 One-Dimensional Delamination Tests ............................. 60 

5.1.5 Two-Dimensional Delamination Tests ............................ 61 

5.1.6 Edge Delamination Test ........................................ 62 

5.1.7 Hinged Double Cantilever Beam Test ............................. 62 

5.1.8 Compression-After-Impact Test ................................. 62 

5.2 Development Test Specimens .......................................... 65 

5.2.1 Experimental Results .......................................... 65 

5.2.2 Predicted Strengths ............................................ 72 

5.3 Demonstration Panels ................................................ 75 

6 CONCLUSIONS ........................................................ 79 

ACKNOWLEDGMENT .................................................. 81 

REFERENCES .......................................................... 81 

4 ANALYTICAL DEVELOPMENT ......................................... 31 
ILLUSTRATIONS 

Figure Page 

1 

2 

3 

4 

5 

6 

7 

8 

9 

10 

11 

12 

13 

14 

15 

16 

17 

18 

19 

20 

21 

22 

23 

24 

25 

26 

27 

28 

29 

30 

31 

32 

33 

34 

35 

36 

37 

38 

39 

40 

41 

42 

43 

44 

45 

46 

D-3304-2 Baseline Aircraft ................................................. 1 

Baseline Wing Structural Arrangement ....................................... 2 

NASA ST-1 Tests - Residual Strength (RoomTemperature. Dry) ................ 4 

NASA ST-1 Tests - Impact Damage Size (Room Temperature. Dry) .............. 5 

Damage After Impact (2Gft-lb Drop) for Composite Wing Material Selection ....... 5 

Wing Box Shear Stiffness Requirement ....................................... 7 

Wing Box Extensional Stiffness Requirement .................................. 8 

Shear Stiffness Distribution ................................................ 8 

Extensional Stiffness Distribution ........................................... 9 

Extensional Stiffness Distribution in Skin and Stringers (7-inch Pitch) .............. 9 

Composite Wing Cover Panel Concept ....................................... 11 

Modified Concept for Blade Stringers ........................................ 12 

Typical Curing Cycle for Wing Skin and Stringer ............................... 13 

Compression-Delamination Specimens ....................................... 16 

Peel (a) and Pull-Off (b) Specimen Configurations 

Stringer Tooling Concept .................................................. 20 

Demonstration Panel (Five-Stringer Panel) ................................... 22 

One-Dimensional Delamination Coupon Test Setup ............................ 23 

Unnotched Compression Test Setup ......................................... 24 

Support Conditions for Impact Tests 

Impact Panel Configurations and Impact Locations ............................ 25 

Strain Gage and Impact Locations on Panels .................................. 29 

Compression Test Setup ................................................... 30 

Typical NASTRAN Model of Impact Panels .................................. 31 

Tension and Compression Bearing Test Specimen .............................. 16 

17 

TrialImpactSpecimens .................................................... 18 

Damage Screening and Damage Tolerance Compression Panels ................... 18 

NDI of Panel D2After Impact (Visible Damage - 100-ft.lb. 1/4-in . Impactor) ...... 21 

.............................. 

......................................... 25 

NASTRAN Model Analysis for Flange Edge Impact (100 ft-lb) 

Comparison of Interlaminar Stresses in Thin (18-Ply) and Thick (36-Ply) Flange 

Peel Moment Versus Pull-Off Load Failure Envelope for Delaminations in Skin at 

................... 32 

Blade Panels ........................................................ 34 

Stringer Interface .................................................... 35 

NASTRAN Model with Adhesive Layer ...................................... 35 

Elliptical Delamination Geometry ........................................... 37 

Sensitivity of G, to Material Property G ..................................... 40 

Model Used for Analysis ................................................... 43 

Peel Stress Distribution .................................................... 49 

Sheer Stress in Adhesive Layer .............................................. 49 

Axial Stress in Skin (Upper Plate) ............................................ 50 

Out-of-Plane Deflection of Skin (Upper Plate) ................................. 50 

Axial Displacement of Skin (Upper Plate) ..................................... 51 

Axial Displacement of Flange (Lower Plate) ................................... 51 

Axial Stress in Flange (Lower Plate) .......................................... 52 

Interlaminar Tensile Stress Contours Derived from Refined NASTRAN Model ...... 

Variation of G, and G, with a and b .......................................... 

Sensitivity of Critical Delamination Growth Strain to the Material Property G,, ...... Sensitivity of Parameter G, to the Depth of the Delamination Plane ................ 

36 

39 

41 

42 
ILLUSTRATIONS 

(Continued) 

Figure Page 

47 

48 

49 

50 

51 

52 

53 

54 

55 

Axial Stress Distribution for a Plate with a Hole and Characteristic Length. do 

(Point Stress Criterion) ................................................ Critical Stress Intensity Factor and Inherent Flaw Size. c,. for Different Laminates ... Residual Stress Distribution Versus Damage Size for Various Values of do (Point 

Stress Criteria) at KT = 3.5 ............................................. Stress Distribution for Stiffened Panel ....................................... 1808I/IM6 Fracture Toughness ............................................. Three-Stringer Panel B 1 After Compression Test (Midbay Impact) ................ Three-Stringer Panel C1 After Compression Test (Midbay 200 ft.lb. 

l/CInchImpactor) ................................................... Compression-After-Impact Test - Three-Stringer Panels C1 and C1' ............. Five-Stringer Panel 5D- 1 Compression-After-Impact Test ....................... 

53 

54 

56 

57 

63 

68 

69 

74 

76 

TABLES 

Table Page 

1 

2 

3 

4 

5 

6 

7 

8 

9 

10 

11 

12 

13 

14 

15 

16 

17 

18 

19 

20 

21 

22 

23 

24 

25 

26 

27 

Materials Selection ....................................................... 4 

6 

Quality ControlTest Data ................................................. 6 

Average Laminate Per Ply Thickness 

Selection Criteria for Compression Panel Configuration ......................... 10 

Composite Wing Technology Ancillary Tests .................................. 14 

Composite Wing Technology Development Tests .............................. 15 

Composite Wing Technology Demonstration Tests ............................. 15 

Stiffened Panels .......................................................... 19 

1Wft-lb Impact Tests and Results for Two-Stringer Panels ....................... 1Wft-lb Impact Tests and Results for B-Type Three-Stringer Panels 

l Wft-lb Impact Tests and Results for C-Type Three-Stringer Panels 

Tension and Compression Test Results ....................................... 59 

Open-Hole Compression Test Results ........................................ 60 

Compression Bearing Test Results ........................................... 60 

One-Dimensional Compression Delamination Test Results ....................... Two-Dimensional Compression Delamination Test Results ...................... 62 

ST-2 Edge Delamination Tension Test Data ................................... 63 

ST-5 Hinged Double Cantilever Beam Test Data ............................... 64 

Compression After Impact Test Data ........................................ 64 

Strain Gage Readings for Panel B1 ........................................... 65 

Strain Gage Readings for Panel B2 ........................................... Strain Gage. Readings for Panel C1' .......................................... 66 

Strain Gage Readings for Panel C2 .......................................... 67 

Strain Gage Readings for Panel C3 .......................................... 67 

Strain Gage Readings for Panel C4 .......................................... 68 

Residual Strength Analysis ................................................. 78 

......................................... 

27 

27 

28 

............... 

............... 

61 

66 

Summary of Compression-After-Impact Test for Three-Stringer Panels ............ 70 

X 
NOMENCLATURE 

Semiaxis lengths of a delaminated idealized ellipse 

Impact damage area detected by C-scan 

A, in-plane stress matrix; refer to delaminated plies 

S-matrix of delamination 

B, coupling matrix; refer to delaminated plies 

Ratio of residual strengths of unstiffened and stiffened panel 

Constant represented by Equation 1 

Half length of a notch 

Characteristic zone length for inherent flaw and point stress criteria 

D, bending matrix; refer to delaminated plies 

where i =j = 1, 2, 6 for matrices A, B, and D 

Bending rigidity of skin and stringer flange 

Young's modulus of the original laminate 

Membrane stiffness 

Young's modulus of adhesive layer 

Young modulus of skin and stringers, respectively 

Functions of matrices A, B, D, a, b, and v 

Yield strength of material in compression 

Strain energy release rate with respect to a when b is constant 

Strain energy release rate with respect to b when a is constant 

Strain energy release rate 

Strain energy release rate, Mode I 

Critical stress intensity factor 

Bending curvatures 

Stress intensity factor 

Stress concentration factor for finite and infinite width, respectively 

Applied force per unit width in x direction 

Applied shear force, bending moment, and axial force 

Radius of a circular hole in a specimen 

Total surface area including delaminated region 

xi 
t 

U1 

u2 

U 

x, Y, z 

Y 

CY 

YC 

EC 

Ecr 

€0, 

Excr 

Ex, Ey, Yxy 

V 

a 

(Jcr 

UCS 

ON 

7, 

Subscripts 

1,2, c 

Thickness of laminate before any damage 

Thickness of adhesive layer 

Thickness of delaminated sublaminate 

Skin thickness 

Stringer flange thickness 

Ply thickness 

(t, + tc)/D,, (t2 + tCP, 

Axial displacement of skin 

Axial displacement of stringer 

Total strain energy of the region 

Displacement components in x, y, and z directions, respectively 

Width of the laminate or panel 

Rectilinear coordinate axes 

Finite-width factor 

a/b aspect ratio 

Shear strain in adhesive layers 

Failure strain 

Compressive strain before or after impact at failure 

Applied strain in x-direction 

Critical compressive strain at onset of delamination 

In-plane strain components 

Poisson’s ratio of the parent laminate 

Total potential energy 

Compressive stress before or after impact at failure 

Compressive residual strength of stiffened panel 

Ultimate strength of unnotched and notched specimen 

Shear stress in adhesive 

Skin, stringer flange, and adhesive layer, respectively 

x ii 
SECTION 1 

INTRODUCTION 

NASA programs have provided the aircraft manufacturer, the FAA, and the airlines with the experience 

and confidence needed for extensive use of composites in secondary and medium-primary structure in 

future aircraft. Secondary and control-surface structures made of composites are already in airline service on production aircraft, and composite medium-primary structures have been introduced for flight 

service evaluation. 

While these applications have produced worthwhile weight savings, the use of composite materials in 

wing and fuselage primary structures offers a far greater opportunity for saving weight since these structures comprise approximately 75 percent of the total structural weight of a large transport aircraft. 

The specific objective of this Composite Transport Aircraft Wing Technology Development program 

was to design, manufacture, and test composite wing cover panel structure representative of a commercial transport aircraft that met all strength, aeroelastic, and damage tolerance requirements at the lowest 

possible cost. 

Douglas selected a baseline transport aircraft (Figure 1) with the advanced engineering designation 

D-33042 for the composite wing technology development program. The vehicle considered is an 

intermediate-range, 150-passenger commercial transport aircraft with propfan engines, planned for 

introduction into airline service in the mid-1990s. Its high-technology wing (Figure 2) features a supercritical airfoil and a high-aspect ratio for low-drag performance. Design integration studies previously 

conducted by Douglas indicated that synergistic benefits can be achieved to reduce both drag and structural weight when high-aspect-ratio wings are constructed of lightweight advanced composite materials. 

bp-138 FT 2 IN.---

CHARACTERISTICS 

WING AREA 

ASPECT RATIO 

WING SWEEP 

1,100 FT* 

10.8 

24 DEG 

109 FT 7.5 IN. 

MULTILOBE 

153 IN. WIDE 

BB I1 an 

17 FT 8 IN. 

Figure 1. D-3304-2 Baseline Aircraft 

1 
X 'I 

6 IN. 

Figure 2. Baseline Wing Structural Arrangement 

The results of the Composite Transport Aircraft Wing Technology Development program are described 

in the following sections. Section 2 deals with material selection and conceptual design of the baseline 

composite wing. Section 3 provides details on specimen fabrication and test procedures; the impacttesting of the development specimens is also described. Section 4 gives an account of analytical 

developments, including the NASTRAN modeling of impact damage, a two-dimensional closed-form 

solution for delamination growth, a peel stress analysis, and a residual strength analysis. Section 5 

discusses experimental and analytical results. Finally, Section 6 summarizes achievements resulting from 

this program. 

2 
SECTION 2 

BASELINE WING DESIGN 

The baseline wing for the composite wing technology program is the D-33042 being developed by the 

Douglas Advanced Engineering department. The D-33042 is an intermediate-range, 15@passenger, 

high-technology transport aircraft with propfan engines, designed for introduction into commercial 

service in the mid-1990s. The wing features a supercritical airfoil and a high aspect ratio for improved 

low-drag performance. A general arrangement of the D-3304-2 is shown in Figure 1. The structural 

arrangement of the wing box shown in Figure 2 indicates the location of spars and ribs, together with the 

principal external load attachment features. The internal substructure is not substantially changed for 

the composite design. 

The structural arrangement, external loads, and criteria for damage tolerance, and fail-safe requirements were selected, based on the following overall design criteria for the baseline composite wing: 

Damage Tolerance - Ultimate load with Mil-A-Prime “just visible” damage. 

Fail-Safety - Limit load with visible damage (broken fibers). 

Residual Strength (Completion of Flight) - With penetration damage 2 inches in diameter, the aircraft should be able to complete the flight mission. 

Wing Fuel Tankage - All tankage: 9-g crash load. With this load, the front spar should be able to 

resist the pressure. 

Access Doors - @percent effective. The door will be designed to withstand all pressure and 

mechanical loads, while the adjoining structure will be designed in such a way that the door is not 

subjected to any load. 

The damage tolerance criteria were selected for defining damage levels at the threshold of visibility and 

for obvious discrete source damage in accordance with FAR 25.571(e) and AC 20-107A. The new MILA-Prime specification for damage tolerance of composite structures, developed under U.S. Air Force 

contract, served as a guideline. The load requirements for the stiffened cover panel tests were based 

upon these criteria. 

2.1 MATERIAL SELECTION 

Extensive testing was conducted to evaluate candidate materials under Douglas in-house IRAD programs. Several materials (see Table 1) from different manufacturers were evaluated for damage 

tolerance. Many toughened residfiber systems introduced by many suppliers of preimpregnated 

materials have been evaulated. These systems match high-elongation fibers with toughened ductileepoxy resins. The NASA ST-1 compression-after-impact test (see Reference 1) was conducted on 11 

potential candidate material systems. The residual compressive strength and damage area after impact 

were determined for all of the materials. The results shown in Figures 3 and 4 indicate that, except for 

the APC-2/AS4 thermoplastic material, 1808I/IM6 is the most damage-resistant composite material 

system of currently available materials. The superior damage tolerance of the 1808I/IM6 system results 

from a thermoplastic adhesive layer (or interleaf) located between plies in the laminate. An example of 

this damage tolerance is illustrated in Figure 5. The cross sections of 1808I/IM6 and 3501-6/AS4 

laminates following 20 ft-lb impacts are shown in the figure, and the 1808I/IM6 laminate is clearly less 

damaged than the 3501-6/AS4 laminate. 

3 
c 

0 

I 

Io 

z 

w 

a: 

w 

rn 

a: 

u_ 

b 

B 

8 

a 

0 

B 

L 

n 

5: 

2 

3 

a: 

80 

60 

40 

20 

0 

Table 1 

Materials Selection 

AMERICAN CYANAMID 1808111M6 IC1 APC-2IAS4 

FERRO 9011IIM6 

NARMCO 5245CIC12K 

NARMCO 5245CIAS6 

NARMCO CELSTARIC12K 

NARMCO 100-11C12K 

ClBA R636ICHS 

NARMCO 100D-lIC12K 

ClBA RX74-21-lIlM6 

HEXCEL F58411M6 

UNION CARBIDE ERLX1962/T40 

HERCULES 855111M7 

IMPACT ENERGY = 20 FT-LB 

- 

rn 

I 

Q 

P 

0 

W 

a: 

Figure 3. NASA ST-1 Tests - Residual Strength (Room Temperature, Dry) 

4 
IMPACT ENERGY = 20 FT-LB 

Figure 4. NASA ST-1 Tests - Impact Damage Size (Room Temperature, Dry) 

L - ... - 

HERCULES 

3501 -61AS4 

m .. ___ - 

AMERICAN CYANAMID 

1808111M6 

Figure 5. Damage After Impact (20-ft-lb Drop) for Composite Wing Material Selection 

5 
Several observations have been made regarding the processing, handling, and overall quality of the 

1808IAM6 material. Mechanical and chemical quality control tests indicate that the material will meet 

the requirements of Douglas process specifications. Preliminary results indicate that the prepreg 

material seems to exhibit considerable variation from batch to batch, as indicated by the data in Tables 2 

and 3. This is particularly the case for ply thickness, which is shown in Table 3 to range from 0.00613 to 

0.00637 inch. Later, material supplied by the vendor to fabricate the skin for five-stringer panels resulted 

in plies that were each 0.0083 inch thick. This poses a design constraint on configurations for which 

structural components must rely on some laminate thickness tolerance. This material is not procured to 

a specific specification. The ply thickness would be more consistent if a specification were invoked. 

~ 

GEL TIME AT 350°F 

(MIN:SEC) 

Table 2 

Quality Control Test Data 

~~ ~ ~~ ~ ~~ 

1. 1255 14:OO 11:12 657 1356 

2. 16:32 13:lO 750 8:56 12:14 

3. - 13:39 9:41 11:00 12:45 

AVG 14:43 13% 9:34 8:58 1258 

8110 B111 Bl12 8113 8114 I I ROLLNUMBER 

RESIN CONTENT (Yo) 

VOLATILES (Yo) 

THERMOPLASTIC 

CONTENT(%) 

1. 

2. 

3. 

AVG 

32.42 

31.25 

32.48 

32.05 

1. 

2. 

3. 

AVG 

1.04 

1.06 

1.09 

1.06 

32.0 

31.8 

31.2 

31.7 

0.87 

0.81 

0.75 

0.81 

28.55 

28.98 

29.63 

29.05 

28.70 

30.77 

32.38 

30.62 

1.09 

1.06 

1.03 

1.06 

1.22 

1.21 

1.16 

1.20 

32.1 

29.6 

30.4 

30.7 

0.91 

0.97 

0.87 

0.92 

1. - 8.0 6.67 7.39 6.00 

2. - 7.7 6.57 6.26 7.20 

3. - 7.4 5.61 6.72 7.3 

AVG - 7.7 6.28 6.79 6.8 

FIBER AREAL WEIGHT 

(gmlm2) 

1. - 145.4 153.7 150.1 148.4 

2. - 146.0 151.5 150.7 148.2 

3. - 149.2 151.6 146.7 147.6 

AVG - 146.9 153.3 149.2 148.1 

Table 3 

Average Laminate Per Ply Thickness 

16-PLY PSEUDO-ISOTROPIC 

18-PLY WING-SKIN 

24-PLY PSEUDO-ISOTROPIC 

27-PLY WING-SKIN 

45-PLY WING-SKIN 

54-PLY WING-SKIN 

THICKNESSPLY (IN.) 

0.00634 

0.00636 

0.00637 

0.00614 

0.00626 

0.00613 

ROLL NUMBER 

B110 

81 10 

61 10 

8112 

8111 

B112 

6 
Monolayer properties and strength of a lamina established for the test prediction (only) for the 

1808I/IM6 carbon-epoxy composite material system at room temperature, dry, under an in-house 

IRAD program are given below. 

Elastic Properties of a Lamina 

E, = 18.5 x 106psi YLT = 0.33 

ET = 1.09 x 106psi t, = 0.0062 in. 

GLT = 0.70 x 106psi F,, = 18.4 ksi 

Strength of a Lamina 

FLT = 265 ksi 

FLC = 185 ksi 

F, = 6.5 ksi 

FTC = 34 ksi 

2.2 CONCEPTUAL DESIGN 

The stiffness distribution requirements for the high-aspect-ratio wing of the D-3304-2 are shown in Figures 6 and 7. For this stiffness-critical design, if aeroelastic effects are not considered, the structural 

weight is purely a function of the extensional modulus of the cover panels. Therefore, the cover skin 

stacking sequence was selected as the highest possible percentage of 0-degree plies that would not 

adversely affect the repairability of the structure. High percentages of 45-degree plies were also desirable 

to meet torsional stiffness requirements. Ultimately, a (0/45/0/ - 45/0/45/0/ - 45/90) pattern was 

selected as a baseline (see Figures 8, 9, and lo), and a design strain of 4,500 4n.h. was established. 

These curves result from tailoring the laminate while keeping the approximate percentage of plies the 

same. 

In general, stringer spacing is a function of extensional (Et) stiffness as shown in Figures 6 and 7. The 

stringer spacing of 7 inches was decided by realistic design constraints and resulted primarily from the 

requirement for large access doors along the span. Since the doors are roughly 12 inches wide, only one 

cut stringer at each door is required. A panel with wider spacing would not carry adequate load, while a 

panel with narrower spacing would require two cut stringers at each cutout. 

0 100 200 300 400 700 

SPANWISE STATION, X (IN.) 

Figure 6. Wing Box Shear Stiffness Requirement 

7 
X 

SPANWISE STATION, X (IN.) 

Figure 7. Wing Box Extensional Stiffness Requirement 

2.0 

s 

3 

z 

8 1.0 

t 

6 

4 

l1.5 

cri 

5 

K 

I 

v) 0.5 J 

W 

z 

d 

n " 

0 100 200 300 400 500 600 700 

SPANWISE STATION, X (IN.) 

Figure 8. Shear Stiffness Distribution 

8 
0 100 200 300 400 600 700 

SPANWISE STATION, X (IN.) 

Figure 9. Extensional Stiffness Distribution 

8 

ACTUAL (SKIN + STRINGER) 

6- 

0 I I I I I I 

0 100 200 300 400 500 600 700 

SPANWISE STATION, X (IN.) 

Figure 10. Extensional Stiffness Distribution in Skin and Stringers (7-inch Pitch) 

9 
The conceptual design effort for this program focused on the detailed design of the stiffened panel test 

specimens. Various concepts for the details of the stiffened panel design were developed and the most 

promising candidates were selected for further development. Variables such as configuration, stiffener 

type and spacing, rib spacing, etc., were considered. All concepts were evaluated in terms of structural 

efficiency, cost, durability, and repairability. Weight comparison was considered by comparing the 

selected composite design concept with a conventional aluminum baseline sized to the same load and 

stiffness requirements. 

Given the internal loads and strain level limitations resulting from the stiffness-critical configuration of 

the baseline wing, various cover panel concepts were evaluated for their ability to meet all design requirements. Early evaluations of compression panel configurations shown in Table 4 studied the relative 

Table 4 

Selection Criteria for Compression Panel Configuration 

ISOGRID 

ACED STITCHING 

COMBINED DAMAGE 

TOLERANCE AND 

FAIL SAFE APPROACH ZEBRA CLOTH PREFERRED EASY TO FABRICATE. PROPAGATION OF USING BOTH "SOFT 

TO HAND LAYUP OF STRIPS 

SKIN" AND "PLANKED 

SKIN IBI"C0NCEPTS 

FAIRLY EASY TO REPAIR. 

EXTENSIVE FILLING OF EXTERIOR SURFACE. 

POSITIVE DAMAGE ARRESTMENT. 

10 
merits of several different concepts. Initial studies were centered on a “soft skin” concept (Le., low axial 

stiffness) for enhanced damage tolerance, similar to the approach developed and adopted by McCarty 

and Whitehead (Reference 2). However, the relatively low stiffness (EA) of these concepts resulted in a 

substantial weight penalty for the baseline wing structure. Thus, with the strain level limitations imposed 

for aileron effectivity and with some of the promising properties demonstrated by new material systems, 

a “hard skin” approach (Le., high axial stiffness) was selected. The concept shown in Figure 11 

represents the baseline wing cover panel design. Several stringer sections were considered, but the 

various concepts resulted in roughly equal cover panel weights for the stiffness-critical design while 

maintaining adequate resistance to local and general instabilities. The blade section stringer was selected 

on the basis of minimum complexity and lower manufacturing costs. The stringers were bonded to the 

wing skins for cost effectiveness. It was initially assumed that stringers would be secondarily bonded to 

the skin and that stitching or bolts would only be added when absolutely necessary. The bond line interface is subjected to transverse tension forces due to fuel pressure and other effects. It was found that 

severe impact loads caused the stringers to separate from the skin in some cases. This problem was 

resolved by modifying the stringer to incorporate thinner flanges, as shown in Figure 12. 

4 

* - 0.43 (72 PLIES) 

2.0 

0.34 (54 PLIES) 0.23 (36 PLIES) 

-- 7 

The costs associated with developing, fabricating, and assembling composite aircraft structures are of 

major, if not primary, importance to a production commitment for primary structure applications. This 

program develops the technology for a wing cover panel structure that would be readily adaptable to 

automated, low-cost manufacturing methods, without sacrificing structural integrity. 

ALL DIMENSIONS IN INCHES 

PLY PATTERN IN SKIN, STRINGERS [44/44/12] (%0/45/90) 

Figure 11. Composite Wing Cover Panel Concept 

11 
7 (0145lOl- 45lOl45101- 451901- 7 I 

2.0 

I- ALL DIMENSIONS IN INCHES 2.4-4 ' 

Figure 12. Modified Concept for Blade Stringers 

12 
SECTION 3 

TEST SPECIMENS AND PROCEDURES 

Specimens were fabricated from 1808I/IM6 using the manufacturer's recommended procedures. A 

typical autoclave cure cycle is shown in Figure 13. Tooling for large composite wing parts was considered in terms of tooling materials and tool design and fabrication. Tool material selection was based 

on dimensional stability, durability, and a predictable tool life to establish acceptable recurring and 

overall costs. 

400 

TEMP, 

(OF) 

T 

350 

300 

250 

200 

150 

100 

\ 

\ \ \ 

\ 

\ 

\ 

\ 

\ 

AVERAGE PRESSURE = 100 PSI 

MATERIAL = 1808111M6 

I I I 

0 60 120 180 240 

0 

TIME, t (MINUTES) 

Figure 13. Typical Curing Cycle for Wing Skin and Stringer 

3 

3.1 SPECIMEN DESCRIPTION 

Tables 5,6, and 7 summarize all the tests performed under this program. Specimens were divided intothree categories as given in each table. Ancillary test coupons shown in Table 5 were designed to determine the elastic properties (tension and compression), bearing, toughness characteristics (compressionafter-impact, double cantilever beam, edge delamination), and delamination onset stresses (one- and 

two-dimensional delamination) of the laminates. The dimensions and lay-up of NASA test specimens 

are given in Reference 1. Bearing and delamination test coupon dimensions and lay-up are shown in 

Figures 14 and 15, respectively. 

Where holes were required in specimens, they were drilled using carbide-tipped twist drills and were 

reamed to tolerance at low speeds. During the drilling process, the thermoplastic material melted and 

filled the flutes of the drill bit. However, this was not a serious problem, even for the very thick 

laminates. It is interesting to note that, when the laminates were drilled with the interleaf material on the 

back surface, the thermoplastic worked quite well as a backup ply, aiding in the prevention of delamination on the back surface. 

13 
TYPE OF TEST 

Table 5 

Composite Wing Technology 

Ancillary Tests 

NO. SPECIMENS 

UNNOTCHED TENSION 

UNNOTCHED COMPRESSION 

OPEN HOLE COMPRESSION 

COMPRESSION-BEARING 

DOUBLE CANTILEVER BEAM 

EDGE DELAMINATION 

COMPRESSION-AFTER-IMPACT 

2-D DELAMINATION 

1-D DELAMINATION 

c 

0 

* 

1 , I I I 

3 REFERENCE 1 

5 REFERENCE 1 

3 REFERENCE 1 

3 FIGURE 14 

3 REFERENCE 1 

6 REFERENCE 1 

5 REFERENCE 1 

3 FIGURE 15 

18 FIGURE 15 

14 
PANEL TYPE 

Table 6 

Composite Wing Technology 

Development Tests 

NUMBER OF 

SPECIMENS TYPE OF TEST 

TRIAL IMPACT 

CRIPPLING SPECIMEN 

STRINGER PEEL TEST 

SCREENING TESTS WITH 

SELECTED LEVELS OF 

IMPACT DAMAGE 

PANEL TYPE 

DISCRETE SOURCE DAMAGE 

VISIBLE IMPACT DAMAGE 

I 1 I 

t 

2 

n 

Table 7 

Composite Wing Technology 

Demonstration Tests 

TRIAL IMPACT DAMAGE 

NDI AND EVALUATION 

COMPRESSION STRENGTH 

PEEL TEST 

RESIDUAL COMPRESSION STRENGTH 

NO. SPECIMENS TYPE OF TEST 

c 

c RESIDUAL COMPRESSION STRENGTt 

RESIDUAL COMPRESSION STRENGTt 

15 
-3161N. HOLE WITH BOLT 

1 

- UMINATE 

I! 

TEFLON PIECE (0.005 IN.) 

2 IN.- 

I 1 

I 

LAMINATE B 

1 I I1 

LAMINATEA: 1fO/4!W-45),~s ,/" 

LAMINATE 8: IlWW - 46)fll1, ONYol- 4W), 

Figure 14. Tension and Compression Bearing Test Specimen 

1 1"'t- I c 

- 0 34 IN. 

I 

9 IN. 

(a) 0NE.DIMENSIONAL DELAMINATION SPECIMEN 

AGE LENGTH 

71.0-IN. DIA 

I I I 

16 
Table 6 summarizes the development tests, and Figures 16,17, and 18 show the dimensions (see Table 8) 

of test specimens. These trial impact, compressive strength, and residual compressive strength tests were 

designed to determine the residual strength for various impact damages. The peel stress tests gave interlaminar stresses in the adhesively bonded structure. The demonstration panel specimens are given in 

Table 7. These panels were fabricated and impact-damaged at Douglas, and then shipped to NASA 

where the compression-after-impact tests were conducted. 

Attachments to or through the stringer blades were avoided because of the bonded assembly of skins and 

stringers on the baseline composite wing. This approach was adopted in order to minimize the possibility 

of developing critical interlaminar tension stresses at the bond line interface. The most critical load 

conditions that will produce interlaminar tension stresses between skins and stringers result from 

internal fuel pressure in combination with some other load case. This phenomenon is described in Figure 16, where the tension pull-off test is compared to the actual conditions found in the baseline design. 

After evaluating several different concepts for this specimen, two configurations shown in Figure 16 

were selected. The test consisted of a single-stringer panel, roughly square in plan view, which was 

loaded in a four-point bending with the loads applied on the skin side of the panel. The deflections 

induced under these loads were similar to those resulting from internal fuel tank pressures in the actual 

wing. The tensile stresses developed at the interface were not load-induced, but were deflection-induced, 

and resulted in critical peel stresses at the edge of the stringer flange. 

P 

I 

I- 7.438 IN. 1 1.719 IN. 

3.18 IN. 

0.39 IN. 

SKIN AND STRINGER 

LAY-UP 

[01451OI - 451901 - 451014510], I ji ! 

42.4 IN.+ 

2.0,". 

PANEL PANEL TYPE t, L(IN.) W(IN.) 

IV PEEL 36 PLIES 4.0 9.0 

V PEEL 18 PLIES 4.0 9.0 

VI PULL-OFF 36 PLIES 2.0 4.0 

Figure 16. Peel (a) and Pull-Off (b) Specimen Configurations 

17 
SINGLE-STRINGER PANEL 

(a) THREE-STRINGER PANEL I 

TWO-STRINGER PANEL -wp4 

I tW 

Ow = DEPTH OF STRINGER 

P1 = STRINGER SPACING 

(b) 

SEE TABLE 8 FOR DIMENSIONS 

Figure 17. Trial Impact Specimens 

7 

1 

OF ST 

(b) FIVE-STRINGER PANEL 

- 

i- ll ts t, ll 

'RINGER 

1 - 9A 

':\ IMPACT LOCATION '-' 

SEE TABLE 8 FOR DIMENSIONS W 

tW 

Figure 18. Damage Screening and Damage Tolerance Compression Panels 

18 
Table 8 

Stiffened Panels 

(Nominal Dimensions in Inches) 

1 

5.0 

20.0 

2.0 

2.4 

- 

1.3 

0.11 

0.46 

0.35 

DESCRIPTION 2 

14.0 

36.0 

2.0 

2.4 

7.0 

2.3 

0.23 

0.46 

0.35 

PANEL WIDTH 

PANEL LENGTH 

STRINGER DEPTH 

FLANGE LENGTH 

STRINGER SPACINGS 

STRINGER EDGE DISTANCE 

FLANGE THICKNESS 

STRINGER WEB THICKNESS 

SKIN PANEL THICKNESS 

3-STRINGER 

C-TYPE 

20.0 

18.0 

2.0 

2.4 

7.0 

2.5 

0.1 1 

0.46 

0.35 

B-TYPE 

21.0 

15.0 

2.0 

2.4 

7.0 

3.0 

0.23 

0.46 

0.35 

NUMBER OF 

STRINGERS 

- 

5 

33.0 

56.0 

2.0 

2.4 

7.0 

2.5 

0.1 1 

0.43 

0.43 

3.2 SPECIMEN FABRICATION 

Techniques for fabricating stringer and skins, and for assembling all the test specimens are given in detail 

in this section. 

3.2.1 Stringer Fabrication 

A tooling concept for the fabrication of blade stringers (see Figures 17 and 18) was selected, based on inhouse tooling concept development. The selected approach, illustrated in Figure 19, uses matched metal 

tooling made from 6061-T6 aluminum. The parts consist of a V-block, two matched mandrels, blade 

stops, and an upper caul plate. During in-house development work, stringer sections 'made with this 

tooling demonstrated accurate section geometry, excellent dimensional stability, and a superior surface 

finish. In addition, this concept may be used in either a press or autoclave cure. 

The composite lay-up takes place directly on the mandrels with the stringer section in halves. The part is 

debulked, chilled, and released from the tool surface. Release film is applied at the interface between the 

tool and the part, which is then placed between the mandrels with the required blade thickness set by the 

mechanical stops. The mandrels are linked vertically by a set of pins, which pass through the stop. The 

mandrels are placed in the V-block, the upper caul plate is fitted into place, and the entire upper surface 

of the tool is sealed with a vacuum bag. During cure, autoclave pressure forces down the upper caul 

plate, which in turn forces the mandrels further into the V-block. This transfers pressure into the blade 

section of the stringer until the stop gap has been reached. 

All the stringers for single-stringer and multistringer panels were fabricated using the above tooling concepts. The fabrication of Qfoot stringer tools for use in fabricating the large five-stringer cover panel 

specimens was completed, and the stringers were subsequently bonded to the skin panels. The nominal 

dimensions of all the stiffened panels are given in Table 8, and those of the compression test and trial 

impact panels are shown in Figure 17. Flanges were constructed from 36 plies of 1808I/IM6 material for 

19 
'STOP 

V-BLOCK 

Figure 19. Stringer Tooling Concept 

single-, double-, and three-stringer panels (see Figure 11). Later, however, an 18-ply flange (Figure 12) 

was used for the five-stringer panels. This concept was superior to the 36-ply-flange concept in resisting 

impacts without disbonding of the skin. 

3.2.2 Skin Fabrication 

Skin panels for single- and multistringer panels were fabricated with 54 plies of 1808I/IM6 material as 

given in Table 8. Except for the five-stringer panels, the thickness of the skins was 0.35 inch with 

[(45/0/ - 45/0)/90],slaminate configurations. The five-stringer panels were 0.43 inch thick because the 

interleaf film material received from CYCOM was thicker than that received earlier. The skin for all 

panels was manually laid up and cured with the standard curing cycle and was within the maximum 

allowable porosity level of 2 percent. 

3.2.3 Skin-Stringer Assembly 

When the stringers were placed over the skin panel in order to bond them together, gaps of different sizes 

were found between the stringers and skin along the centerline normal to the stringers (see Figure 18b). 

The gap was 0.03 inch for stringer 8A, 0.018 inch for stringer 9A, 0.019 inch for stringer 10B, and 0.012 

inch for stringer 10A. Such gaps at full pressure would produce areas of undesirable porosity in the bond 

line. A special nonadhereing epoxy film, FM641, known as Verifilm Green has been developed to check 

for the presence of gaps at full pressure and temperature. A single ply of the 0.015-inch-thick FM641 was 

sandwiched between the two plies of 0.001-inch-thick Mylar. The usual bagging procedure was 

employed using the FM641 instead of the FM300. 

All bond lines were sealed using 1-inch aluminum foil tape, which was faired up to 90 degrees from the 

skin, over the flange of the stringer. A Coreprene dam was placed over this foil to mate with and soften 

the flange edges. Once the panel was on the caul plate, four layers of Coreprene dam with the foil were 

laid on the ends of the stringers. All dams were sealed to the part with the pressure-sensitive Mylar tape. 

20 
Before the Verifilm was applied to the panel, quite a few tests at the coupon level were conducted. These 

tests demonstrated that the presence of the FM641 layer improved the bond line strength to 4,911 psi, 

while the through-ply arrangement of FM300 gave 2,919 psi. After bagging, the panel was cured at 

290°F and 50 psi for 15 minutes, and then cooled down to 150°F for inspection. The bond lines 

measured 0.013-0.015 inch for 8A, 0.014-0.016 inch for 9A, 0.013-0.014 inch for 10B, 0.012-0.015 inch 

for 8B, and 0.012-0.014 inch for 10A. In all cases, the Verifilm was uniform across the bond lines except 

where pressure was lost directly under the blade in three separate 1.5-inch areas. The Verifilm determines 

the size of gaps to be closed by adhesives and is not an adhesive itself. 

3.2.4 Demonstration Panels 

Two five-stringer panels have been fabricated using the same cure cycle. The skin panels were made from 

a recently procured material with a thickness of 0.0083 inch per ply. These large panels, which were 

impact-damaged at Douglas, were compression-tested at NASA Langley. The test plan sent to NASA 

shows that one panel will have visible damage and the other will have discrete source impact damage at 

midbay. To achieve visible damage, several trial impact tests were conducted on several undamaged 

parts of already impacted two-stringer panels. Finally, both five-stringer panels were impacted at midbay with 100 ft-lb of impact energy using a 17-inch support span. No NDI-detectable damage was produced when the first panel was impacted with a 1-inch-diameter impactor. The second panel was 

impacted with a 1/4-inch-diameter impactor, and visible damage resulted. The C-scan for this panel is 

shown in Figure 20. Both panels were impacted using a 17-inch center support span to compare the 

Figure 20. NDI of Panel D2 After Impact (Visible Damage - lOO-ft-lb, 11441. Impactor) 

21 ,R!S!P<!AL PACE is 

OF POOR QUALITY 
Figure 21. Demonstration Panel (Five-Stringer Panel) 

results with the three-stringer panels. The 1/4inch impactor penetrated halfway into the skin. Two 

aluminum ribs were fabricated from angle sections for each panel and attached to the skin with bolts at 

37-inch center spans (see Figure 21). The panels only differed from each other by about 1/4 inch in 

length. All other dimensions are the same for both panels, as shown in Table 8. 

3.3 TEST PROCEDURES 

3.3.1 Ancillary Tests 

All the tests except compression-delamination and bearing tests were conducted in accordance with Reference 1. Each part of the specimens was C-scanned before and after assembly. The specimens were 

instrumented with strain gages before testing. These tests were performed in a 1. l-million-pound-capacity machine with a 0.05-in./min controlled stroke. The compression test was conducted in the setup 

shown in Figure 22. All the test specimen ends and sides were kept parallel to each other to within a 

0.005-inch tolerance. Thickness and width of the test coupon parts were recorded by taking the average 

of more than seven places on the part using a micrometer. Strain measurements, where applicable, were 

recorded at different load increments. Load-deflection history and failure load were recorded for each 

specimen. Depending on the test, the specimens were C-scanned after the test to determine the extent of 

the damage. 

Two-dimensional compression delamination specimens were tested in the ST- 1 (Reference 1) fixture 

with antibuckling plates. One-dimensional delamination tests were performed in a setup similar to that 

shown in Figure 23. 

22 

ORIGINAL PAGE 

BLACK AND WHITE PHGTCGRAPH 
MTS 1 

(1 00-Kl PS CAPACITY) 

Figure 22. One-Dimensional Delamination Coupon Test Setup 

23 
APPLIED 

LOAD 

1 

0.5 IN 

BONDED 

DOUBLERS 

FM73 (HIGH TEMP) 

FM300 (ROOM TEMP) 

--L 

c --t = 0.088IN. 

. 

-.3t-* -- \TEST SECTION 

z 

I 

BONDED 

DOUBLERS 

FM73 (HIGH TEMP) 

FM300 (ROOM TEMP) 

--L 

c --t = 0.088IN. 

. 

-.3t-* -- \TEST SECTION 

z 

3 

r LATERAL RESTRAINTS 

Fiaure 23. Unnotched ComDression Test SetuD - e- -- 

3.3.2 Development Tests 

The development tests concentrated on three areas. The first test in this group involved impacting twostringer panels at midbay, stringer flange, stringer blade, and laterally on the stringer blade. These 

panels were impacted at various energy levels ranging from 20 to 100 ft-lb. After impact damage was 

inflicted, the panels were nondestructively inspected to determine the severity of damage. The results 

were used to determine the amount of impact to be applied to three-stringer panels. All of these panels 

were supported at their opposite edges similar to the configuration shown in Figure 24a. 

Stringer pull-off tests were also conducted. These tests were designed to simulate wing skin flexing due to 

aerodynamic loads or pressurization of the fuel tank. The tests examined the strength of the bond 

between the stringer and the skin. Four-point beam bending loads were applied in such amanner that the 

stringer flanges were peeled from the skin without the application of direct tension of the stringer blade. 

The impact locations of two- and three-stringer panels are shown in Figure 25. The structural test plan 

was drafted with NASA consultation to test three-stringer damage tolerance panels after they were 

impact-damaged with 100 ft-lb of energy at different locations. One of the panels was impacted with 

200 ft-lb of energy, which resulted in visible damage. 

24 
34N. SQUARE TUBULAR SUPPORT BLOCKS, BASE PLATE CUT OUT AT STRINGERS 

SIMPLY SUPPORTS END 112 IN. 

OF STRINGER BLADE 

SIMPLY SUPPORTS END 1/2 IN. 

OF SKIN AND STRINGER FLANGE 

(b) SIMPLY SUPPORTED ENDS (c) SIMPLY SUPPORTED ENDS 

36.0 IN. 

2.0 IN. TYP 

AND BOTTOM TO PREVENT OVERTURNING 

SIMPLY SUPPORTS END 112 IN. OF PANEL 

(d) SIMPLY SUPPORTED ENDS 

Figure 24. Support Conditions for Impact Tests 

- 14.0 IN. I 

PANEL I 

135.6 N-rn (100 FT-LB) IMPACT LOCATIONS 

1. OVERBLADE 2-STRINGER PANELS 

2. OVER FLANGE EDGE 2-STRINGER PANELS 

3. MIDBAY 2-STRINGER PANELS 

4. LATERAL BLADE 2-STRINGER PANELS 

5. OVERBLADE 3-STRI NG ER PANELS 

6. OVER FLANGE EDGE 3-STRINGER PANELS 

7. MIDBAY 3-STRINGER PANELS 

8. 45-DEG LATERAL BLADE 3-STRINGER PANELS 

ALL SKIN AND STRINGEG LAY-UPS 

[0145101- 4510145101 - 451901, (SyM, 

1- 20.0 IN. 

.IES 

6-TYPE PAN EL C-TYPE PANEL 

t, = 36 PLIES 

L = 151N. L = 18.0IN. 

1, = 18 PLIES 

Figure 25. impact Panel Configurations and Impact Locations 

25 
3.3.2.1 Impact Tests - The impact test requirements currently proposed under AFWAL contract 

(Reference 2) call for ultimate strength to be maintained following an impact level that just causes visual 

damage, up to a maximum of 100 ft-lb. The laminates fabricated with interleaved material had no visual 

damage when tested below the maximum impact energy level. Therefore, all panels were impacted at 100 

ft-lb. This impact was attained by dropping a 1-inch-diameter impactor (attached to a 21.8-pound 

weight) from a height of 55 inches. 

Locations for external skin impacts were midbay between stringers, over the blade, and over the stringer 

flange edge. Internal impact tests were applied laterally on the stringer blade. Tables 9 through 11 

describe each test individually, while Figures 16, 24, and 25 detail the associated impact locations and 

test boundary conditions. 

The test panels were inspected by C-scan. Panels that were impacted midbay between stringers had only 

local skin delaminations directly beneath the impact. Panels impacted over the blade or over the flange 

edge were more likely to delaminate in the first few plies of the skin adjacent to the stringer interface. 

Such delaminations occurred at the stringer either beneath the impact or near the panel supports. The 

C-scans of the lateral blade impacts showed delaminations that were strictly confined to the central 

region of the stringer by the blade, beneath the lateral impact site. In all cases, C-scan aberrations were 

attributed to delaminations of the composite and not adhesive bond failure. 

The smaller peel and pull-off interaction specimens were tested statically until failure. Failure was ultimately caused by skin delamination beneath the edge of the flange. 

Two-Stringer Panels - The two-stringer panel impact results are summarized in Table 8. In test 1 A, the 

delamination occurred in the first few plies at the base of the stringer, above the adhesive layer. This 

result was not anticipated, because the damage to the panel was sustained at each end, while the impact 

site was relatively unaffected. It was obvious that the delaminations were caused by excessive interlaminar normal forces generated at each end of the panel. The test setup indicated that the manner in 

which the panel was supported was conducive to this type of failure under the bending loads resulting 

from the impact. 

Three-Stringer Panels - The impact results for three-stringer panels with a flange thickness of 0.23 inch 

are given in Table 10 (B-type). These panels were impacted at midbay and at the middle of the panel over 

the centerline of the stringer from the cover side. In this case, the panels were supported at the two ends 

just like clamped edges. In the case of the midbay impact, more damage resulted than for the twostringer panel. Midstringer impact unexpectedly disbonded the stringer from the skin with peeling of two 

plies from the skin. Approximately 60 percent of the bonded stringer was separated from the skin. This 

damage was caused because of the stiffer blade and the rigid boundary conditions. 

Four three-stringer C-type specimens (Table 11) were impact-tested, and the damage status of the panels 

was as follows: 

1. Panel C1 was impacted over the middle of the center stringer at the outer surface of the panel with 

100 ft-lb of energy. The panel was simply supported at the outer edges of the blades as shown in Figure 24b. In this panel, no damage was found either by visual or NDI techniques. To demonstrate 

the effect of visible (front surface) impact damage on compressive strength, panel C1 was subsequently impacted at 45 degrees with 200 ft-lb of energy using a l/Cinch-diameter impactor and was 

redesignated as panel C1' . No trace of any delamination or disbonding was found. The impact 

with this impactor at midbay caused the desired visible damage, penetrating into the skin. 

26 
- 

TEST 

ID 

62 

1A 

3-STRI NG ER, 

36-PLY FLANGE 

2A 

- 

3A 

4A 

- 

- 

TEST 

ID 

Table 9 

100-ft-lb Impact Tests and Results for Two-Stringer Panels 

C-SCAN OF CENTER STRINGER PANEL IMPACT SUPPORT 

:ONFIGURATION LOCATION CONDITIONS 

REF FIG. 25 REF FIG. 25 REF FIG. 24 

1 

IN.$ 

I 

3.0 6.0 9.0 12.0 i5.oia.a 

l..l..l..l..l..~ 

2 

OVER2-STRINGER, FLANGE A 

36-PLY FLANGE EDGE CLAMPED ENDS 7 

IMPACT 

NO STRINGER DELAMINATION A 3 

P-STRINGERI MIDBAY CLAMPED ENDS ONLY LOCAL SKIN 

36-PLY FLANGE 1 IMPACT 1 BENEATH IMPACT 

4 

36-PLY FLANGE BLADE CLAMPED ENDS 

2-STRINGER, LATERAL A 

IMPACT 

Table 10 

100-ft-lb Impact Tests and Results for B-Type Three-Stringer Panels 

PANEL 

CONFIGURATION 

B1 

I REFFIG.25 

3-STRI NG ER, 

36-PLY FLANGE 

IMPACT 

LOCATIO N 

REF FIG. 25 

5 

OVER-BLADE 

IMPACT 

7 

MIDBAY 

I M PACT 

SUPPORT 

CONDITIONS 

REF FIG. 24 

A 

CLAMPED 

ENDS 

A 

CLAMPED 

ENDS 

C-SCAN OF CENTER STRINGER 

I-* X AXIS 

IN. E 2.5 5.0 7.5 

2.4 IN. 

7- 

NO STRINGER DELAMINATION 

ONLY LOCAL MATRIX DAMAGE 

BENEATH IMPACT 

(SIMILAR TO 3A AND C2) 

27 
- 

rEST 

ID 

c1 

- 

c3 

- 

c2 

- 

c4 

- 

C1' 

2. 

3. 

4. 

Table 11 

100-ft-lb Impact Tests and Results for C-Type Three-Stringer Panels 

PANEL 

CONFIGURATION 

FIG. 25 

3-STRINGER, 

1 8-PLY FLANGE 

3-STRINGER, 

ia-PLY FLANGE 

3-STRINGER, 

18-PLY FLANGE 

3-STR I N G E R , 

1 8- PLY F LA N G E 

-DOIMPACT 

LOCATION 

FIG. 25 

5 

OVER-BLADE 

IMPACT 

6 

FLANGE EDGE 

IMPACT 

OVER7 

MIDBAY 

IMPACT 

8 

45-DEG LATERAL 

BLADE IMPACT 

~ 

MIDBAY 

I M PACT 

SUPPORT 

CONDITIONS 

FIG. 24 ~~ ~ 

B 

SIMPLE SUPPORT 

C 

SIMPLE SUPPORT 

C 

SIMPLE SUPPORT 

D 

SIMPLE SUPPORT 

C 

SIMPLE SUPPORT 

C-SCAN OF CENTER STRINGER 

NO DAMAGE 

1- X AXIS 

IN. i I 3;0, ,6;0, , gi0 

1 

I I I 

- I_ . ... 

~~~~ ~~ 

I NO STRINGER DELAMINATION 

ONLY LOCAL SKIN DAMAGE , I BENEATH IMPACT I 

II 

T 

VISIBLE DAMAGE FIG. 26 

Panel C2 was impacted at midbay with 100 ft-lb of energy and supported as shown in Figure 24c. 

There was a visible 2-square-inch delamination on the far side of the panel. 

Panel C3 was impacted at the outer side of the panel near the edge of the middle stringer flange. The 

panel was supported as shown in Figure 24c. The impact energy of 100 ft-lb produced an approximately l-l/Zinch delamination and peeling of the flange from the skin near its edge. 

Panel C4 was impacted at the middle of the blade of the center stringer in an inclined (45-degree) 

fixture, as shown in Figure 24d. The panel was simply supported on its two sides. The panel was 

also retained at two top corners in the direction normal to the panel to prevent any overturning at 

the time of impact. The three-stringer panel was impacted at 45 degrees with 200 ft-lb of energy. 

The panel was debonded about 3.5 inches along the skin and stringer bond line. 

28 
3.3.2.2 Mechanical Tests - The strain gage and impact locations of all six panels are shown in Figure 26. The locations of gages around the damage areas are offset from the centerline to avoid interference. Three major types of damage (disbonding of inner plies, delamination, and separation of the 

stringer from the skin) were observed on these panels. 

The tests for six three-stringer damage tolerance screening specimens were conducted in the Baldwin test 

machine (Figure 27) with its capacity of 1.1 million pounds and by uniform end shortening at a rate of 

0.02 inch per minute. The data from all strain gages were recorded simultaneously. The results are 

discussed in Section 5. 

1,2.. .12 = STRAIN GAGE LOCATIONS 

t = IMPACTSITES 

B,, C,, ETC. = PANEL 

Figure 26. Strain Gage and Impact Locations on Panels 

29 
Figure 27. Compression Test Setup 

3.3.3 Test Plan for Demonstration Panels 

Both panels were impact-tested, and aluminum ribs were attached. The damage was evaluated, and the 

panels were shipped to NASA Langley, where they were tested. 

The panels were potted and machined to correct sizes at the NASA facility. NASA applied 40 gages to 

the panel with nonvisible damage and 46 gages to the panel with visible damage. The panels were tested 

under compression to their ultimate strength. The strain gage data acquisition was performed by NASA 

personnel using a multichannel computer. 

ORIGINAL PAGE 

BLACK AND WHITE PHOTVGRAPH 

30 
SECTION 4 

ANALYTICAL DEVELOPMENT 

Considerable effort was devoted to developing analytical techniques in the areas of interlaminar stress, 

delamination growth, peel stress, and residual strength. To understand the disbonding due to impact on 

stiffened panels for a composite wing, a NASTRAN finite-element model was created, and the interlaminar stresses due to impact loading were determined. The NASTRAN analysis with experimental 

results has been used to improve the blade flange design. 

Delamination due to impact loading and disbonding of the adhesively bonded skidstringer in the stiffened panel is the vital phenomenon observed in the experimental studies. It is imperative to investigate 

the delamination growth mechanism under compressive loading and to explore the peel stresses in the 

skidstringer interface. A two-dimensional closed-form solution was developed to predict delamination 

growth using the principle of minimum potenhl energy. Most of this work was accomplished under an 

in-house IRAD program. In this case, elliptical delamination was assumed and general plate theory, 

including bending-extensional coupling, was applied to determine the strain energy release rate components. The effects of various parameters on these components were studied. 

A peel stress analysis technique was developed to determine closed-form solutions for peel stress, axial, 

and out-of-plane displacements. A method based on the point stress criteria was developed to determine 

the residual strength of an impact-damaged stiffened panel. 

4.1 NASTRAN MODEL 

For the case of the impact panels studied here, NASTRAN models were created to determine the 

resulting internal peel moments and pull-off loads. A typical model is illustrated in Figure 28 and is 

relatively simple. Figure 29 shows a flange edge impact NASTRAN model. The flange and skin are 

Figure 28. Typical NASTRAN Model of Impact Panels 

31 
Figure 29. NASTRAN Model Analysis for Flange Edge Impact (100 ft-lb) - See Table 10, 

Test ID C3 

modeled as one element, thereby making no allowance for the bond line. These models were intended to 

produce internal loads resulting from the beam action of the stringer resisting the applied impacts. 

The impact loads were simulated in NASTRAN by using an equivalent static load (Reference 3). This 

load was determined’ by the following equation, where w is the actual weight dropped from height, h, 

onto a panel with stiffness, k. 

Equivalent Static Load = [ 1 + (1 + 2hk/~)”~]w 

The stiffness, k, was determined by first analyzing a unit load case for the NASTRAN models, from 

which k = (unit load)/(deflection for unit load) at the impact location. The equivalent static loads were 

then calculated and used for stress analysis by NASTRAN. 

The NASTRAN models were validated by comparing the NASTRAN deflection predicted for twostringer panel impact test to that measured during a high-speed motion picture of the test. The 

NASTRAN model predicted a 0.19-inch out-of-plane deflection, which compares well to the 0.21-inch 

deflection actually measured from the film. 

4.1.1 Analysis 

Failure Mechanism - There are two types of failure mechanisms - concussive damage and stringer 

delamination. 

32 
Concussive Damage - The first type of impact damage seen is the local matrix cracking associated with 

the direct absorption of the impact energy. This occurred during the midbay impacts, where skin alone 

had to absorb the impact force. For the impacts over the flange and over the blade, the flange and skin 

acted together to provide a greater thickness for absorbing the direct impact energy. Therefore, delaminations due to energy absorption did not occur at those locations. 

Stringer Delamination - This section focuses on the stringer delaminations caused by the internal loads 

acting in the panel during flexure. These delaminations are seen at various locations along the stringer 

(Tables 9 through 11). 

The analysis of stringer delamination requires identification of the loads and their interactions. The skin 

delaminations are caused by interlaminar tension stress at the stringer and skin interface. The load that 

would cause this interfacial tension is a direct pull-off load applied from the stringer blade and a peel moment applied from the skin (Reference 4). A secondary mode of delamination may occur in the radius of 

the stringer elbow. The responsible load here is primarily the lateral moment at the root of the stringer 

blade. These loads and their effect on interlaminar stresses are illustrated in Figure 30. 

From the NASTRAN internal loads output, pull-off loads and peel moments were determined. Graphs 

were constructed plotting peel moment versus pull-off load at half-inch intervals along the length of the 

critical stringer. By considering the separation of delaminated and undamaged data points, an interaction between peel moment and pull-off load was determined. This data point became the intercept for 

the peel moment axis. 

The peel and pull-off behavior identified in Figure 30 provides insight into the shape of the failure 

envelope. The pull-off load causes high interlaminar stresses beneath the stringer blade, while the peel 

moment has negligible effect there. Thus, a failure line predicting delamination beneath the blade should 

be horizontal, involving only pull-off. At the flange edge, however, both the pull-off load and the peel 

moment have an effect. Thus, a second failure line predicting delamination beneath the flange edge 

should be sloped to indicate an interaction. The slope should illustrate decreasing pull-off strength with 

increasing peel stress. (This shape is in accordance with Reference 5.) 

The shape of the failure envelope varies with the geometry of the cross section, as shown in Figure 30. 

(This is supported by Reference 6.) The data were therefore divided into two groups to define separate 

envelopes for the 36-ply flange and the 18-ply flange cross sections. The delaminated data points 

identified on the peel versus pull-off charts, and their C-scans were used to help determine whether the 

delaminations initiated at the flange edge or beneath the blade. The final envelopes derived are shown in 

Figure 31. 

4.1.2 Refined Analysis 

As an additional method to support the trends presented here, a much more detailed set of NASTRAN 

models was created (Figure 32). These models separated the skin and stringer flange into separate plate 

elements, connecting them by a model of the adhesive layer. To model the adhesive, a series of shear 

panels containing shear properties were interspersed with rod elements containing tensile properties. 

The tensile stresses predicted by these adhesive rods also represent the interlaminar tension in the first 

few plies adjacent to the adhesive. The magnitudes of the tensile stresses are based on approximate 

adhesive properties and therefore may be inaccurate, but they qualitatively reinforce the trends 

previously discussed. 

Examples of the NASTRAN interlaminar tensile stress solutions are shown in Figure 33. They correlate 

well with the critical stress locations given by the peel versus pull-off analysis method, either peaking at 

33 
INTERNAL 

LOAD 

1 t T I 

t 

P 

S: SUPPORTS I 

EFFECT OF A 

UNIFORMLY STIFF FLANGE 

P 

HIGH INTERLAMINAR TENSION 

UNDER BLADE AT 

FLANGE EDGE 

HIGH INTERLAMINAR TENSION 

AT FLANGE EDGE 

HIGH INTERLAMINAR RADIAL 

TENSION IN THE ELBOW 

EFFECT OF A 

UNIFORMLY FLEXIBLE FLANGE 

P 

HIGHER INTERLAMINAR TENSION 

UNDER BLADE ONLY 

REDUCED INTERLAMINAR 

TENSION AT FLANGE EDGE 

SINCE FLEXIBLE FLANGE 

ATTRACTS LESS LOAD 

SIMILAR RESPONSE CAUSING 

RADIAL ELBOW TENSION 

Figure 30. Comparison of lnterlaminar Stresses in Thin (18-Ply) and 

Thick (36-Ply) Flange Blade Panels 

34 
6- 

5- 

4- 

3- 

2- 

1- PU L L-OFF 

LOAD, PP 0- 

(LBIIN. x lo3) -1 

-2- 

-3- 

-4- 

-5- 

-6-1 

PEEL MOMENT, Mp (IN.-LBIIN.) 

DELAMINATION 54 PLIES 

BENEATH BLADE 

- 0 STATIC PEEL TESTS 

0 PULL-OFF INTERACTION CHECK 

FOR CURVE A 

A = 36-PLY FLANGE(t1) 

B = 18-PLY FLANGE(tf) 

I, I I I I I I I II I I I I I I, 

Figure 31. Peel Moment Versus Pull-Off Load Failure Envelope for Delaminations in 

Skin at Stringer Interface 

CENTER STRINGER MODELED SEPARATELY FROM SKIN, 

ADHESIVE LAYER MODELED TO CONNECT THE TWO 

SKIN SECTION 

WITH 

07-DP-8210 

Figure 32. NASTRAN Model with Adhesive Layer 

35 
(a) Test - Bl 

Delaminations Occurred at Each End, 

Presumably Initiating at Flange Edge 

(Reference Table lo) 

@ INDICATES IMPACT 

LOCATION 

Test - C3 

Delaminations Occurred at End 

Beneath Blade and in Center at 

Flange Edge (Reference Table 11) 

-x - x 

II I I I I I I I I I IN. ' 

-9-8 -6 -4 -2 0 2 4 6 89 

DISTANCE ALONG CENTER STRINGER, X 

-G 

I 

-xJ-x I 

G 

Figure 33. lnterlaminar Tensile Stress Contours Derived from Refined NASTRAN Model 

(Reference Figure 32) 

the flange edge or beneath the blade. The highest interlaminar tension stresses also occur in the regions 

that failed during testing. A series of such models suggests that 18,850 psi would be the interlaminar 

tension strength for the panels modeled here. The details of the analysis can be found in Reference 7. 

4.1.3 Specimen and Support Conditions 

A simplified NASTRAN investigation of the specimen's bending response, and that of the wing structure, indicated that the test support conditions (Figure 24a) were unreasonably overrestrained. The joint 

bending rigidity offered by the rib supports is quite small in the actual aircraft structure, with less than 

half of the fixed-end moment being restrained. An investigation of the design revealed that the most 

rigid (shortest) rib bay could be simulated with a pin-ended specimen 17 to 18 inches long between supports, depending on the impact location. The three-stringer screening specimens were redesigned based 

on this information. 

From NASTRAN analytical investigations, it was clear that the bending stiffness of the flange bonded 

to the skin plays an important role in the skidstringer disbonding mechanism. From the results of trial 

impact tests and NASTRAN models, it was decided to reduce the blade flange thickness from 36 plies 

(Figure 11) to 18 plies (Figure 12). 

4.2 DELAMINATION GROWTH - AN APPROXIMATE SOLUTION 

This work is based on a thin-film model that has a delamination with an elliptical clamped boundary 

(Figure 34). As the delamination grows, it maintains its elliptical shape, but with a different aspect ratio. 

The delamination portion of the laminate on the sublaminate is thin compared with the parent laminate; 

36 
I' 

I 

I 

I 

z 

> 

SEC A-A 

Figure 34. Elliptical Delamination Geometry 

i.e. , the parent laminate does not bend. The in-plane deflections at the boundary of the delamination are 

determined by the strains in the parent laminate. The out-of-plane deflection of the elliptical plate is 

expressed as: 

(1) w = aC/,(1 -~2--2 2 Y) 

where 

- 

x = x/a, 7 = y/b, and C/, is an arbitrary constant. 

The in-plane displacements are given by: 

dx = cox ra where [r = (1 - 72)1/2] 

dy = eOysb where [s = (1 - X2)'/2] 

E,, = critical buckling strain of ellipse. 

Eoy = -v€ox 

The in-plane membrane deflections u and v are functions of w. 

2 

u = (eox + Clr6) xa 

v = (-v + C;a2s6) yb wherea = a/b 

37 
The total strain energy U of the region S is 

(ETA€ + 2eTBK + KTDK)dxdy 

2 

J-l cr + (S - a ab) (A,, - 2v A,, + ",A2,) ~02~ (4) 

A, B, and D refer to the constituent matrices of the delaminated sublaminate. S is the total area of the 

~ laminate including delamination. 

~ 

The strain energy release rate for changes in the a and b dimensions of the ellipse are defined as 

fori = a,b 

where 

g = [9 ff - 16 f4( f2cox + fl 11% 

c4 1 u = -, f4 + c; f3 + c: (f2EOX + f, ) + - f,€2, 2 2 

1 

2 

+ - (S - aab) (A,, - 2vA12 + v2A2,) 

f, = aab(Al1 - 2vA12 + v2AI2) 

f, = 20.617(D1, + 0.667a2DI2 + a4D2, + 1.333 a2DM)/a 

f, = 1.718ab[A1, - v a2A2, + (a2 - v 1 AI21 

f, = 14.085aa (BIZ - 2B66) 

f4 = ab[1.31624(AI1 + a4A2,) + 1.6684 dA,, + 3.883 (r2AM] 

(9) 

38 
Minimum potential energy is obtained by differentiating Equation (4) with respect to C; and setting the 

resulting relation to zero. This relation is the characteristic equation that is used to calculate the critical 

compressive strain, eXcr 

Cl = 0.32(g - 3f,)/f4 

Plots of G, and G, for a typical laminate (Figure 35) show that for near-circular delaminations Gb is 

greater than G,, Le., the tendency for the delamination to progress laterally is greater than its tendency 

to progress in the direction of the load. Also note that for small delaminations, as b/a increases beyond 

2, G, becomes greater than Gb, which tends to recircularize the delamination. But at the same time, as a 

(or the size of the delamination) becomes larger, G, approaches zero and Gb approaches a constant 

value. Thus, the delamination growth is stable. 

The effect of the material G,, on the parameter Gb and on the critical strain at which the delamination 

will propagate is shown in Figures 36 and 37. Clearly, it is advantageous to have a high value of G,,in the 

laminate. 

The sensitivity of the parameter Gb to the depth of the delamination plane is shown in Figure 38. The 

effects can vary depending on the size of the initial delamination. However, for the size of impact 

damage likely to occur in the new toughened system, the deeper delaminations are generally less critical 

than those toward the surface. 

1 

0.8 

n- 0.6 

z 

z < 

5 0.4 

0- 

0.2 

0 

SUBLAM [ f 45190/01 

STRAIN = 0.004 INAN. 

POISSON’S 

RATIO =0.5 

3501 -6/AS4 

4 

1 

0.8 

N0.6 i 

z + 

f - 

n‘ 0.4 

0 

0.2 

(I 

1 2 3 

a (IN.) 

SUBLAM [It 45190101 

STRAIN = 0.004 INAN. 

POISSON’S 

3501-6/AS4 

RATIO =0.5 

I I I 

2 4 6 8 10 

b (IN.) 

Figure 35. Variation of Ga and Gb with a and b 

39 
1.6 

1.5 

1.4 

1.3 

1.2 

1.1 

1 

0.9 

0.8 

G,,(IN..LB/IN.*) 

0.7 

0.6 

0.5 

0.4 

0.3 

0.2 

0.1 

0 

a=1 I i 

SUBLAM [ f 45/90/0] up = 0.5 

1',2*,3',4' SUBLAM [0/45/0/-451 up = 0.415 

0.004 INAN. 

up = POISSON'S RATIO OF PARENT LAMINATE 

3501-llAS4 G,, = 0.9 IN.-LBIIN. 

'1808111M6 G,, = 2.1 IN.-LB/IN.* 

0 2 4 6 8 10 

b (IN.) 

Figure 36. Sensitivity of Gb to Material Property G,, 

40 
15 

14 

13 

12 

11 

10 

9 

z -9 z 

88 = 

z 

d 

5' 

s 

0 

b6 K 

0 

5 

4 

3 

2 

1 

0 

CRITICAL DELAMINATIOY GROWTH STRAIN FOR DIFFERENT vP 

Y = 0.6' 

P 

Y = 0.4' 

Y = 0.2' - P 

SUBLAMINATE [451- 45190101 

'1808111MG G, = 2.1 IN-LB/lN.* 

3501-61AS4 G,, = 0.9 IN:LB/lN.* 

I I I I I I I 

2 4 6 

INITIAL DELAMINATION AREA (IN.2) 

Figure 37. Sensitivity of Critical Delamination Growth Strain to the Material Property G,, 

41 
1.8 

/ VARIATION OF G, WITH DELAMINATION DEPTH 

LAMINATE [(0145101- 45),90]3, 

Y = 0.415 ex = 0.004 

P 

0 

INITIAL CIRCULAR DELAMINATION AREA (IN.') 

Figure 38. Sensitivity of Parameter G, to the Depth of the Delamination Plane 

42 
4.3 PEEL STRESS ANALYSIS 

Bonded stiffened panels are being used more widely in the aerospace industry because of their lighter 

weight and fewer parts. Various authors have developed methods for analyzing the stresses in the 

adhesive between two bonded plates (References 8 and 9). However, a simpler method was needed to 

calculate these stresses in bonded stiffened panels, so that the designer could evaluate the relative importance of adhesive and adherend properties in determining the stresses. 

Under this program, a closed-form solution was obtained for the stresses and displacements of two 

bonded beams. A system of two fourth-order and two second-order differential equations with the 

associated boundary equations was determined using a variational work approach. A FORTRAN computer program was devised to solve for the eigenvalues and eigenvectors of this system and calculate the 

coefficients from the boundary conditions. The results were then compared with NASTRAN finiteelement solutions and were shown to agree closely. 

4.3.1 Analysis 

The model analyzed is shown in Figure 39. The upper and lower members are beams of unit width, and 

the effect of the stringer blade is disregarded. In this analysis, subscripts 1 and 2 refer to upper and lower 

beams, respectively, and subscript c refers to the adhesive. In addition to assuming beam behavior and 

small deflections, the adhesive is assumed to be linearly elastic and of constant thickness. The restraint 

on the adhesive provided by the two beams prevents Poisson's contraction, making the assumption of 

linear elasticity a reasonable one. 

PZ t', I ADHESIVELAYER 

tz I STRINGERFLANGE 

Figure 39. Model Used for Analysis 

43 
An expression for shear stress in the adhesive, 7,, which takes into account the differences in rotation 

and elongation between the two beams, is: 

where 

TI = (tl + t,)/D, T2 = (t2 + t,)/D2 

Using this expression, strain energy of the adhesive shear (U,) can be added to the contributions of 

strain energies of adhesive stretching (UE), beam stretching of skin and stringer (Us], Usd, and beam 

~ 

bending of skin and stringer (UBI, UBJ. 

For total potential energy: 

V = - (Plwl + P2w2 + Mldwl/dx + M2dw2/dx + Flu1 + F2~2) I x=L 

L 

+ 1/2s (Fl(dwl/dx)2 + F2(d~2/dx)~)dx 0 

Using the principle of minimum potential energy: 

Substituting ifl = wIDI, W2=w2D2, Ti, =uIEltl, andE2=u2E2t2yields the following expression 

for 6~: 

I where 

El = Eltl, E2 = E2t2 

44 
Integration by parts to eliminate differentials from the variational terms and grouping according to 

variational terms yield 12 boundary conditions and a system of two fourth-order and two second-order 

differential equations. 

Boundary Conditions 

At the centerline (x = 0), the stringer flange is clamped while the skin is free to translate but not rotate. At 

the free edge of the skin and stringer flange, the applied forces and moments are related to displacement 

as follows: 

(24) 

diil /dx = F1 dii2/dx = F2 d2F, /dx2 = M 1 d2E2/dx2 = M2 (24) 

System of Differential Equations 

Differentiating Equations (29) and (30) once with respect toxand substitutingu, = du,/dx, Ti4 = du2/dx, 

W3 = d2W1/dx2, and W4 = d2w2/dx2gives a system of six second-order equations, which can be written in 

matrix form: 

- - 

00 1 0 00 

00 0 1 00 

--A B CTI +FI/D1 CT2 CF -CG 

A -B DTI DT2+F2/D2 DF -DG 

00 ET1 ET2 EF -EG 

00 -ET1 -ET2 -EF EG - - 

45 
I where 

A = Ec/tcD 

D = GcT2D2/4tc 

G = 1/B2 

B = Ec/tcD2 

E = Gc/2tc 

I The characteristic equation for this system is defined by IM - AI1 = 0, where M is the matrix of coefficients of the system and I is the identity matrix. For this case, the characteristic equation is: I 

0 = A4 - [CTl + DT2 + F1/D1 + F2/D2 + E(F + G)]A3 

+ [A + B + CTlF2/D2 + DT2Fi/Dl + FI/DI*F~/D~ 

+ E(F + G)(Fl/Dl + F2/D2)]A2 - [(A + B + FI/DI*F2/D2)E(F + G) (32) 

+ (C + D)(AT2 + BT,) + BFI/Dl + AF2/D2]/2. + E(F + G)(AF2/D2 + BFl/Dl) 

I 

I This gives two cases, one with in-plane forces and one without. For the first case, the solutions to the I characteristic equation are (Al, A,, A,, A, 0,O). For the second case, the solutions are (Al, A,, A, 0,O ,O). 

Case 1 (with In-Plane Force) 

For this case, W1 can be expressed as follows: 

where pi = ,&- 

Putting this expression into Equations (27) through (30), it can be shown that the x3 and x* terms disappear and W1, W,, Ti1, and Ti, can be expressed as: 

46 
C12 = FC11/G and C14 = [(T1 + T2A/B)C9 + FC13]/G (40) 

Case 2 (without In-Plane Force) 

For this case, W1 can be expressed as follows: 

Putting this expression into Equations (27) through (30), it can be shown that the x5 and x4terms disappear and w 1, w2, u and u2 can be expressed as: 

47 
For both cases: 

Using these expressions for W ,, W,, U ,, and U2 and the boundary conditions resulting from integrating 

the expression for 6.lr by parts, the coefficients can be evaluated. Axial stresses in the skin (a,) and 

stringer (ad and the peel stress (ad are determined using Equations (37), (39), (49, (43, and (52). 

4.3.2 Results 

A FORTRAN computer program, PSTRESS, has been written to evaluate the coefficients described 

above for various loading conditions. The deflections calculated by this program are within 3 to 5 percent of those calculated using a NASTRAN finite-element analysis of an equivalent structure (Reference 7). The results were further corroborated by correlating the experimental results of an impact investigation with peel stresses calculated by PSTRESS (Reference 10). The PSTRESS results in this case 

agreed closely with NASTRAN finite-element analyses in predicting skidstringer deflections due to 

impact over the stringer blade. 

Figures 40 through 46 show plots of various stresses and displacements for a typical case with a composite skin and stringer. The skin lay-up is [0/45/0/ - 45/0/45/0 - 45/90],, and the lay-up of the flange 

is [0/45/0/ - 45/0/45/0/ - 45/90] s. The properties and loading used are as follows: 

E, =4,713 ksi E, =4,594 ksi 

D, = 3,685 lb-in. D, = 4,764 lb-in. 

t, =0.224 in. t, =0.224in. 

E, = 500 ksi G, =45 ksi 

t, =0.005 in. L =1.2in. 

Load Case 1: PI = 100 lb/in. 

Load Case 2: M , = 80 lb-in./in. 

Load Case 3: F, = 1,600 lb/in. 

Load Case 4: PI = 100 lb/in., M, = 80 lb-in./in., F, = 1,600 lb-in./in. 

P, =F2=M2=0 

Parametric studies using PSTRESS (Reference 1) have demonstrated the relationships between the 

adhesive and adherend properties and the peel and shear stresses in the adhesive. The peel stress at the 

flange free edge (x = L) is largely dependent on the applied moment and in-plane load, while the stress 

under the stringer blade (x = 0) is primarily the result of the applied shear. Peak peel stress at the free edge 

is generally considered the critical cause of failure. Peel stress can be reduced in a number of ways. 

Decreasing rigidity, modulus, or thickness of the flange relative to the skin will decrease peak peel stress 

at the free edge, but will also increase the stress under the blade in cases with applied shear. Choosing an 

4a 
11 

10 

9 

8 

7 

6 

5 

cT,(KSI) 4 

3 

2 

1 

0 

-1 

-2 

0 0.2 0.4 0.6 0.8 1 1.2 

X (IN.) 

Figure 40. Peel Stress Distribution 

0 SHEAR 

0 MOMENT 

0 AXIAL FORCE 

A COMBINED x 

I I 1 I I I I I I 

0 0.2 0.4 0.6 0.8 1 1.2 

X (IN.) 

Figure 41. Sheer Stress in Adhesive Layer 

49 
8 

2l 1 

0 SHEAR 

0 MOMENT 

0 AXIAL FORCE 

A COMBINED 

I I I I I I I I I I I 

0.002 

0.001 

0 

- 0.002 

- 0.003 

- 0.004 

0 0.2 0.4 0.6 0.8 1 1.2 

X (IN.) 

Figure 42. Axial Stress in Skin (Upper Plate) 

AXIAL FORCE 

A COMBINED 

0 0.2 0.4 0.6 0.8 1 .o 1.2 

X (IN.) 

Figure 43. Out-of-Plane Deflection of Skin (Upper Plate) 

50 
u1 (1N.x 

8 

7 

u2 (IN. x lo-' 

4 

3 

2 

1 

0 

0 0.2 0.4 0.6 0.8 1 .o 1.2 

X (IN.) 

Figure 44. Axial Displacement of Skin (Upper Plate) 

0 SHEAR 

0 MOMENT 

0 AXIAL FORCE 

A COMBINED 

0 0.2 0.4 0.6 0.8 1 .o 1.2 

X (IN.) 

Figure 45. Axial Displacement of Flange (Lower Plate) 

51 
4.0 

3.5 

3.0 

2.5 

2.0 

1.5 

1 .o 

0.5 

0 1 

0 0.2 0.4 0.6 0.8 1 .o 1.2 

X (IN.) 

Figure 46. Axial Stress in Flange (Lower Plate) 

adhesive with a lower Young’s modulus (EJ will also decrease peak peel stress, but shear stress at the free 

edge will be increased. Adhesive shear modulus (Gc), however, has relatively little impact on the peak 

peel stress. Increasing adhesive thickness (tJ is more effective than the above methods because it 

decreases peak peel stress at the free edge and at the blade and decreases shear stress. 

4.4 RESIDUAL STRENGTH ANALYSIS 

The analysis method in this section predicts the initial static strength of panels containing single damage. 

The residual strength values obtained using this method may require adjustment to incorporate the 

effects of multiple-flaw interaction if dealing with very closely spaced damage and dynamic effects if the 

structure is loaded at the time of impact. 

Many structural configurations provide multiple load paths and cannot be represented as monolithic 

sheets when assessing strength degradation from impact damage. Wing cover panels often consist of 

skin with mechanically fastened or integral stiffeners. The stiffeners can provide crack-arrestment capability because the stiffening member can accept load transferred from the skin and can resist crack opening displacement. In this way, it can significantly improve the residual strength of damaged structures. 

Nuismer et a1 (References 11 and 12) suggested that the average stress criteria can be applied to both compression- and tension-loaded panels. The test results obtained by the authors showed that tension failure 

criteria can be replaced by substituting allowable compression stress for tensile stress. With the same reasoning, the point stress criteria are applied here for compression failure cases. 

52 
Point stress criterion (Reference 11) is applied to determine the failure mechanism of a damaged panel. 

The criterion can be expressed as: 

It 11' ob X 

I do 1 

1 1'1 i t 

where the characteristic distance (do) is a material property, and the normal stress (ay) at distance do 

from the edge of the hole equals the unnotched compressive strength of the material (a,,). Figure 47 

shows the approximate distribution of u,, in the region do for x> c (or R), and Equation (54) is derived 

from Reference 13 by introducing the stress concentration factor (KT) of the laminate. 

where p = c/x and c = half-length of a notch or radius of a hole. Substituting Equation (54) in Equation 

(50) gives 

where p = c/(c + da and uN is the ultimate strength of the notched specimen. 

Reference 11 gives a relation between characteristic distances for point stress do and inherent flaw c, 

criteria as 

C, = 2d, (56) 

The characteristic distance c, is constant (References 12, 14, and 15) for a given composite material 

system and is computed by the test resutls as follows: 

C, = c/( (UJU&'- 1) and Kq = U,[T(C + CJ]'.~ (57) 

Critical stress intensity factor, Kq, and inherent flaw size, c,, for different laminates for carbon-epoxy 

systems (Reference 15) are shown in Figure 48. 

53 
1.60 1 I I I 1 1 . /* (%01% f 451%90) 

1.40 

1.20 

(0133167) 

1.00 ' -0- CRITICAL STRESS 

INTENSITY FACTOR 

ULTIMATE STRESS I o.80 ' (67/3310) 0 

- KQ (IN.)112 Io 0.60 - (62.5125/12.5) 

I (12.512W62.5) 

0.40 - I 

uO 

0.20 

0.00 A 1 1 I 

Figure 48. Critical Stress Intensity Factor and Inherent Flaw Size, co, for Different Laminates 

I (0/80/20)) ,r 

'(0150150) &@ / 

I 

I' 

' 

+---- 

I 

I 

/ 

(2018010)/<l 118910) 

(25150125) -/ e 

0 

4' 

0 

--e#@ JC-, 

1(5015010)1 

I 1 1 L 

54 

0.70 

0.60 

0.50 

INHERENT FLAW 

SIZE co (IN.) 0.40 

0.30 

0.20 

0.10 

0.00 

I 1 1 I 1 1 

0' ,0 (%O/% +451%90) 

v 

(0180120) 

0 (0/10010)~ 

0 f 

I 

' ,0 1, 

0" / 

(0133167) p:VW/0* I I 

---+-, 

I 

(6713310) / 

0 ' 1 118910) (25/50125) (2018010) 0 / 

(62.5125112.5) a 0 

I 

(12.5125162.5) 0 

' 1 I I 
4.4.1 Effect of a Finite-Width Panel 

Stress Concentration Factor 

The effect of a finite-width panel on the stress concentration factor for hole damage is described in 

Reference 1 1. 

For a hole, the finite-width factor Y is expressed as 

Y = KT/KF= (2 + (1 - 2~/~)3)/3( 1 - ~c/w) 

with radius c = 0.1875 in. and w = 2 in., Y = 1.04 

For a notch, this factor is expressed as 

Y = KT/KF= J2/(7rc) *tan (dw) 

withc = R = 0.1875in.andw = 2in. 

(59) 

Y = 1.02 (60) 

Stress Intensity Factor 

The finite-width factor, Y for Kq is given as 

Y = sec(nc/w) 

with c = 0.1875 in. and w = 2 in., Y = 1.022 

From Reference 14: 

Y = (1 - 0.5~ + 0.37~2- O.O44p3)/( 1 - p) 

with p = C/W = 0.1875/2 = 0.09375, Y = 1.005 

4.4.2 Stress Concentration Factor 

The KF stress concentration factor (Reference 14) for an orthotropic infinite-width plate is given as 

KF= 1 + &42z) [~--A,2+(A,,A,-A,22)/2A,1 (63) 

where the Aij components are in-plane laminate stiffness determined from laminated plate theory. The 

subscript 1 denotes the direction parallel to the applied stress. The following values of Aij are computed 

using an in-house program for 54-ply skin (0/45/0/ - 45/90/0/45/0/ - 45),, for the 1808I/IM6 material 

system. 

A,, = 3.397(10)6 

A,, = 0.6678(10)6 

A,, = 1.609(10)6 

A, = 0.7809(10)6 

E, = 10.07(10)6 

Using Equation (63), K 7 = 3.464 (64) 

The residual strength of laminate (0/45/0/ - 45/90/0/45/0/ - 45),, is plotted in Figure 49 for material 

1808I/IM6 with K, equal to 3.5 for different do values. 

55 
1 .o 

0.9 

0.8 

0.7 

0.6 

aN1aO 

0.5 

0.4 

0.3 

0.2 

MATERIAL - 1808111M6 

[(0/45/01 - 45)2/90],, 

I I I I I I I I I I I I I I I I I 

0 0.2 0.4 0.6 0.8 1 .o 1.2 1.4 1.6 1.8 2.0 

R 

Figure 49. Residual Stress Distribution Versus Damage Size for Various Values of do (Point 

Stress Criteria) at KT = 3.5 

4.4.3 Residual Strength 

uC = Residual strength of finite-width unstiffened panel in compression with damage 

oCs = aC for stiffened panel 

Cexp = (~c'(~cs 

Cexp is determined by a method similar to that developed for metal structure by T.R. Porter (Reference 

16). Cexp is related to stiffener area, skin gauge, stiffener spacing, and yield strengths of skin and stiffeners. The value of (T, (unstiffened) is determined from fracture mechanics techniques for notched 

(damaged) panels by using the critical stress intensity factor in compression. 

Failure Stress of Unstiffened Panel 

Failure Stress of Stiffened Panel Cexp = 

For metallic structures, an empirical relation has been plotted in Figure 50 for C,,depending on RFty for 

riveted stiffened panels. 

RFty = (EAStif/EASkin) (Fey stif/Fcyskin) (67) 

where Fcyis yield strength of the material, It is assumed that bonded stiffened structure behaves similarly 

to riveted metallic structure in compression. 

56 
In the case of IM6/18081 material with (44/44/12) percentage of (0/45/90)-ply laminate used for stiffeners and skin structures, 

where EASti,and EA,,, are the membrane stiffnesses of stiffener and skin. 

4.4.4 Analysis Method 

The values of c,, (T,, and Kq for a given laminate are determined by conducting unnotched and openhole compression tests using Equation (57). KF is determined using Equation (63), and Equation (56) is 

applied to determine d,. K,is found using Equation (58) or (59), depending on the type of damage, and 

Equation (67) is used to determine R,,,. CeXpis determined from Figure 50, and, finally, acSis determined 

from Equation (65). The application of this method is illustrated with actual cases in Section 5. 

1 .o 

0.8 

0.6 

Cexp 

0.4 

0.2 

0 

CURVE 

FAILURE STRESS OF UNSTIFFENED PANEL 

‘EXP = FAILURE STRESS OF STIFFENED PANEL 

Ft STIFFENER A IF ENER X RFty = (tsK,N)(STS:FEFNERSPACING) y~ty SKIN 

I I I I I I 

0 0.1 0.2 0.3 0.4 0.5 0.6 

RFty 

Figure 50. Stress Distribution for Stiffened Panel 

57 
SECTION 5 

TEST RESULTS AND DISCUSSION 

A summary of the tests for this program is presented in Tables 5 through 7. A series of ancillary tests 

(Table 5) to generate basic data required for the residual strength analysis of stiffened panel test 

specimens were conducted. Table 6 describes the development test program, also conducted at Douglas. 

These tests were included to assess the selected cover panel design concepts. Table 7 contains the large 

demonstration panel specimens to be tested at NASA Langley Research Center. 

The trial impact specimen (Table 6) was used to determine the extent and visibility of damage at various 

locations on the cover panel structure for specific impact energy levels. Single-stringer specimens were 

tested to evaluate stringer pull-off strengths and static crushing/crippling strengths. Screening tests 

determined the residual strength of the selected cover panels after they sustained impact damage at 

critical locations. These tests also determined the most critical impact location for the Douglas cover 

panel concepts. These tests were essential to the concept evaluation process. 

The large five-stringer panels shipped to the NASA facility for tests will demonstrate the damage tolerance capabilities of the selected concept for the most critical impact damage under compression loads. 

With five stringers and representative rib attach points, these tests will provide a realistic assessment of 

the final design. 

5.1 ANCILLARY TESTS 

5.1.1 Tension and Compression Tests 

Test results for the unnotched tension and compression tests are shown in Table 12. The tension test 

results were reasonably consistent for both strength and stiffness. The average ultimate tensile stress for 

the [(0/45/0/- 45)d90],,, laminate was roughly 133,000 psi, with a modulus value of approximately 

10,100,OOO psi. Unfortunately, four of the five unnotched compression test specimens suffered 

premature failures at the ends. The only specimen that appeared to exhibit a true compression failure, 

with a failure stress of 107,030 psi, is listed in Table 12. The test setup used is shown in Figure 23. The 

premature failures were characterized by a brooming or crushing of the specimens at the load introduction end. It is suspected that these failures resulted from insufficient quality in the surface grinding of the 

specimen ends. 

Table 12 

Tension and Compression Test Results 

SPECIMEN 

UTBl 

UTB2 

UTB3 

AVG 

AREA (IN.*) 

0.1150 

0.1151 

0.1145 

UCBl 

ULT LOAD (LB) 

15,020 

15,340 

15,460 

15,273 

- 8,830 

I 1 

FAILURE STRAIN (IN.) 

0.0013 

0.00133 

0.00131 

0.00133 

FAILURE STRESS (PSI) 

130,608 

133,275 

135,021 

132,968 

I -107,030 I 

5.1.2 Open-Hole Compression Tests 

Tests of open-hole compression specimens were conducted in accordance with NASA standards (Reference 1). The results of the tests are presented in Table 13. The results were generally as expected. In 

-- 

PRECEDING PAGE BLANX NOT FILMED 
WIDTH DIAMETER THICKNESS FAILURE GROSS FAILURE 

SPECIMEN W d t LOAD STRESS STRAIN 

NUMBER (IN.) (IN.) (IN.) wld (LE) (PSI) plNJIN. 

1 2.0 0.375 0.33 5.33 32,600 49,390 4,615 

2 2.0 0.375 0.33 5.33 32,150 48,710 4,552 

3 2.0 0.375 0.33 5.33 33,300 50,450 4,715 

MATERIAL: 1808111M6 

PLY THICKNESS: APPROXIMATELY 0.0062 IN.1PLY 

LAMINATE PAlTERN: (0145101 - 4510145101 - 45/90), 

Ktc. 

1.64 

1.67 

1.61 

‘Ktc VALUES BASED ON FCu = 100 KSI 

WIDTH 

SPECIMEN W 

NUMBER (IN.) 

1 2.0 

2 2.0 

3 2.0 

Table 13, the stress concentration factors for the open-hole tests (based on an unnotched compression 

allowable of 100 ksi) are compared to the computed elastic-isotropic stress concentration factors. A 

comparison of these values shows that the results are similar to previous test programs, and the test 

results indicate that the notch sensitivity exhibited by these specimens is typical of most carbon-epoxy 

material systems. The material was found to be resistant to impact damage because of the presence of the 

interleaf adhesive layer. 

ULTIMATE BEARING 

DIAMETER SPLICE ULTIMATE YIELD’ BEARING YIELD FAILURE 

d THICKNESS LOAD LOAD STRESS STRESS STRAIN 

(IN.) (IN.) (LE) (W (PSI) (PSI) (PERCENT) 

0.375 0.28 17,100 14,100 138,180 113,940 0.6135 

0.375 0.28 17,400 14,400 140,610 116,360 0.6086 

0.375 0.28 18,ooo 15,300 145,450 123,640 0.6239 

5.1.3 Compression-Bearing Tests 

The compression-bearing tests were designed to provide bearing stress allowables for the design and 

analysis of the bolted repair panel. The results shown in Table 14 were fairly consistent with an average 

ultimate bearing stress of roughly 140 ksi and an average “yield” stress of about 118 ksi. 

Table 14 

Compression Bearing Test Results 

MATERIAL: 1808111M6 

PLY THICKNESS: APPROXIMATELY 0.0061 IN.IP8 

FAILURE MODE: BEARING FAILURE IN CENTRAL MEMBER 

LAMINATE PATTERN: (0/45/0/- 45/0/45/0/-45/90) , SKIN THICKNESS = 0.33 IN. 

‘YIELD POINT DEFINED AS INITIAL POINT OF NONLINEARITY 

5.1.4 One-Dimensional Delamination Tests 

Eighteen one-dimensional delamination compression specimens were tested. The load was supplied by 

two specially prepared f 0.0005-inch oil-hardened stainless steel loading blocks approximately 0.380 

inch thick. Each was fastened to 0.1- by 2- by 5.2-inch aluminum sheet stock with 0.005-inch doublebacked tape. Finally, 0.020- by 2- by 3-inch aluminum shim stock was butted up to the loading blocks to 

provide a glide path for the specimen. This entire assembly was placed inside 0.85-inch-thick steel angle 

plates and bolted down (see Figure 22). 

60 
In each case, a strain gage was mounted on the laminate surface directly over the insert. It was assumed 

that, when the inserted delamination fractured, the outer plies at the delamination site would no longer 

carry load and the gage would indicate a sudden drop in strain level. This was the case in each test. It 

should be noted that these test results, shown in Table 15, were surprisingly consistent (free of scatter) 

compared to the results obtained by other researchers. 

I I 

1 SPECIMEN DELAMINATION ONSET ULTIMATE FAILURE I 

DEFECT ’ GAGE I ’ WIDTH THICK DEPTH, LOAD STRAIN STRESS I LENGTH LOAD STRAIN STRESS I IDNO. I (IN.) (IN.) SIZE (LE) (pIN./IN.) (PSI) I (IN.) (LE) (pIN./IN.) (PSI) ’ 507-1 I 1.00 0.34 PLlES5,6 21,550 6,350 63,382 27,050 6,868 79,559 I 507-2 I SIZE = 21,400 5,810 62,941 3.r 26,300 6,887 77,353 

I 507-3 I 0.5 IN. 22,150 7,400 65,147 1 27,300 8,868 80,294 

I 509-1 I PLIES 5,6 10,000 2,931 29,411 I 3.00 25,850 6,941 76,026 

I 509-2 I SIZE = 9,000 2,750 26,471 I 4 26,150 7,007 76,912 

I 509-3 I 1.0 IN. 9,000 2,701 26,471 I 25,900 7,205 76,176 

I 511-1’ I PLlES5,6 4,000 1,018 11,765 I 2.50 27,500 7,735 80,882 

I 511-2 I SIZE = 5,000 1,462 14,706 I 28,650 9,179 84,265 

I 511-3 I 1.5 IN. 4,000 1,053 11,765 ! 1 29,200 8,950 85,882 ‘ 513-1 I PLIES 3,4 12,000 3,662 35,294 I 3.00 29,350 8,358 86,324 

I 513-2 I SIZE = 11,OOO 3,443 32,353 28,850 8,521 84,853 

I 513-3 I 0.5IN. 12,500 3,841 36,765 1 28,450 8,683 83,676 ’ 515-1 I PLlES3,4 4,000 1,135 11,765 I 3.00 28,800 6,979 84,706 

I 515-2 I SIZE = 4,000 1,089 11,765 26,150 7,602 76,912 

I 515-3 I 1.0IN. 4,000 1,234 11,765 1 27,400 7,244 80,588 

517-1” I PLlES3,4 2,100 667 6,176 I 2.50 31,100 8,717 91,471 

26,650 8,207 78,382 

1.5 IN. 2,000 552 5,882 24,950 5,519 73,382 

517-2 I SIZE = 1,500 

I 

1 ’ 517-3 I 

In each case, initial growth of the inserted delamination was observed and noted. It was not clearly evident at what load level so-called “fast fracture” occurred, since the delamination growth to the specimen ends was somewhat sporadic. 

MATERIAL: 1808111M6 LAMINATE: [(0/+45/0/ -45),/90],, LENGTH: 9 IN. 

ROOM TEMPERATURE, DRY 

*LENGTH OF SPECIMEN IS 8.5 IN. 

5.1.5 Two-Dimensional Delamination Tests 

Two-dimensional delamination tests were conducted on three specimens with circular inserts between 

plies 3 and 4. The panels failed in compression across the free section (approximately 0.25-inch gap) 

above the antibuckling plates. Some damage propagated around the insert to the ninth ply just prior to 

total compressive failure. The opposite surface also buckled along the ninth ply, and this delamination 

also propagated into the compression failure area. Fracture along the central plies was also evident. 

The specimens were tested in the ST-1 (Reference 1) fixture with antibuckling plates. The insert buckled 

on the specimen surface at an average of about 40,000pounds. The buckle was initially the same shape as 

the insert. As the load increased, the delamination grew laterally toward the antibuckling plates. 

61 
SPECIMEN DIMENSIONS INITIAL DELAMINATION ULTIMATE FAILURE 

DEFECT GAGE 

ID NO. (IN.) (IN.) LOCATION (LE) (p INAN.) (PSI) (IN.) (LE) (plN./IN.) (PSI) 

WIDTH THICK SIZE, (IN.) LOAD STRAIN STRESS LENGTH LOAD STRAIN STRESS 

501-1 5.00 0.333 1-IN. DIA 45,000 2,850 24,024 9.00 125,840 8,431 75,504 

PLIES 3,4 

501-2 5.00 0.333 1-IN. DIA 40,000 2,375 24,024 9.00 120,120 7,890 72,072 

PLIES 3,4 

501-3 5.00 0.333 1-IN. DIA 40,000 2,198 24,024 9.00 124,300 7,650 74,580 

PLIES 3,4 

MATERIAL: 1808111M6 LAMINATE: [(O/+ 4510/-45),/90], 

ROOM TEMPERATURE, DRY 

i 

Seconds before the ultimate compression failure occurred in the free section, the delamination grew 

under the antibuckling plates. Test results for the three successful two-dimensional delamination tests 

are shown in Table 15. Three tested coupons of two-dimensional delamination tests (see Table 16) 

showed consistent results, which were considered very good for 1-inch-diameter inserts between plies 3 

and 4. 

5.1.6 Edge Delamination Test 

The results of the edge delamination tension tests (Reference 1) for the 1808IAM6 material are shown in 

Table 17. The values for fracture toughness, G,, are typical of most high-toughness material systems. 

The results, shown in Figure 51, were also consistent between the 8-ply and 11-ply tests. 

5.1.7 Hinged Double Cantilever Beam Test 

The hinged double cantilever beam test (Reference 1) results are presented in Table 18 and Figure 51. 

The average value for Glc of 1.71 in.-lb./in.2 is consistent with the results of the in-house material 

evaluation tests. The G 1c values were calculated using the energy-area integration method. Calculations 

using the modified direct beam equation method would give slightly higher G,, values. 

5.1.8 Compression-Af ter-Impac t Test 

Compression-after-impact tests (Reference 1) were run on five panels. The results of these tests are 

shown in Table 19. One panel was impacted at 20 ft-lb and two panels were impacted at 50 ft-lb with a 

1-inch hemispherical impactor. 

Two remaining panels were impacted at an energy level of 26.4 ft-lb using a 10-pound weight with a 

1/2-inch hemispherical impactor. The 26.4 ft-lb impacts left barely visible damage on the surface of the 

laminate, with no visible damage on the back surface. The 50 ft-lb impacts were slightly more visible on 

the impact surface, and a slight split in the 0-degree ply on the laminate back surface was visible. X-rays 

taken of each panel revealed minimal internal damage for the 26.4 ft-lb impacts, with a slight increase in 

damage area for the 50-ft-lb impacts. 

62 
Table 17 

ST-2 Edge Delamination Tension Test Data 

SPECIMEN 

ID 

ST-221 

ST-222 

ST-223 

AVERAGE: 

LAMINATE ORIENTATION: ( f 35/0/90), E,am = 9.97 x 106PSI E' = 8.13 X 106PSI 

= 10.45 X 1 O6 PSI E( * 35Ws TEST CONDITION: 75°F DRY I 

THICKNESS WIDTH 

(IN.) (IN.) 

0.0689 1.500 

0.0690 1.499 

0.0690 1.498 

0.0690 1.499 

DELAMINATION ONSET 

STRAIN' (IN./IN.) 

0.0047 

0.0040 

0.0048 

0.0045 

FAILURE 

STRAIN 

(IN./IN.) 

0.0057 

0.0058 

0.0056 

0.0057 

1 INTERLAMINAR 

FRACTURE 

TOUGHNESS 

DELAMINATION ONSET 

STRAIN. (IN./") (IN.) (IN AN.) 6, ('y') 

0.0064 

0.0070 

0.0072 

0.0069 

1.92 

2.30 

2.46 

2.23 

ST-21 1 0.0509 

ST-21 2 0.0510 

ST-21 3 0.0515 

AVERAGE: 0.051 1 

1.495 

1.495 

1.497 

1.496 

I LAMINATE ORIENTATION: ( f 30/ f 30/90/90), Elam = 8.38 x 106PSI E' = 5.32 x 10'PSI 

= 6.89 x 106PSI E( f 

TESTCONDITION: 75°F DRY I 

INTERLAMINAR 

FRACTURE 

TOUGHNESS 

G, (I?') 

2.32 

1.69 

2.43 

2.15 

TENSILE 

MODULUS 

(PSI) 

7.60 x 106 

7.59 x lo6 

7.62 x lo6 

7.60 x lo6 

'STRAIN AT FIRST DEVIATION FROM LINEAR STRESS-STRAIN CURVE 

MATERIAL: 1808111M6 

[=j [ f 35/0/90], ED 

3 

cu2s 

5 

5 

P 

8 

1 

0 

cui 

z 

g 

$ 

1- 

!- 

a 

e 

8 

1.71 

--a I 1- 

Figure 51. 1808111M6 Fracture Toughness 

63 
Table 18 

ST-5 Hinged Double Cantilever Beam Test Data 

COUPON a, 61 Pl a2 6, p* a3 63 

ID (IN.) (IN.) (IN.) (IN.) (IN.) (IN.) (IN.) (IN.) 

ST5-1 1.00 0.14 21 2.03 0.19 14 3.05 0.32 

ST5-2 0.98 0.13 22 2.00 0.20 14 2.97 0.33 

ST5-3 0.98 0.12 21 2.09 0.22 14 3.00 0.34 

LAMINATE ORIENTATION: (0)" 

TEST CONDITION: 75°F DRY 

GIC* 

p3 

(IN.) (IF) 

7 1.62 

6 1.91 

5 1.73 

~~ 

*Glc VALUES CALCULATED BY ENERGY-AREA INTEGRATION METHOD 

Table 19 

Compression After Impact Test Data 

SPECIMEN 

ID 

CAI-1 

CA 1-2 

CAI-3 

CAI-4 

CAI-5 

THICKNESS WIDTH 

(IN.) (IN.) 

0.3297 5.002 

0.3309 4.962 

0.3312 5.095 t 0.3303 5.096 

0.3310 5.93 

LAMINATE 0 

TEST CONDI' 

MATERIAL: 1 

MAXIMUM 

WIDTH OF 

IF 

T 

f 

3IENTATION: (+4! 

ION: 75°F DRY 

08111M6 h = 0.00 

VISUAL 

IMPACT DAMAGE 

YES NO 

YES NO 

YES NO 

YES YES* 

YES YES' 

'01 - 45/! 

I 

i25 IN./I 

IMPACT 

AREA 

(IN.*) 

2.84 

0.196 

2.47 

5.13 

7.7 

- 

3 

3 

FAILURE 

LOAD 

(KIPS) 

80.41 

105.16 

90.42 

76.89 

79.97 

FAILURE 

STRESS 

64.05 

53.58 

45.68 

47.44 

FAILURE 

STRAIN 

(pIN./IN. 

4,920 

6,731 

5,522 

4,625 

4,843 

COMPRESSlDh 

MODULUS 

(PSI) 

9.91 x io6 

9.52 x lo6 

9.70 x lo6 

9.88 x lo6 

9.80 x lo6 

'SLIGHT C L iPLlT IN 0-DEGREE PLY ON BACK S JRFACE 

64 
5.2 DEVELOPMENT TEST SPECIMENS 

There were two groups of panels - B-group and C-group. The panels were designated as B1 and B2, and 

as C1, C2, C3, and C4. The B-group stiffeners had 36-ply flanges, while the C-group stiffeners had 

18-ply flanges. During test, both groups of panels behaved quite differently. Panels B1 and B2 separated 

at the skidstringer bond line before reaching failure load. The skin panel alone reached the buckling 

stage and finally fractured in bending, leaving the stringer undamaged. This type of behavior was 

expected as the impact test on the B1 panel had caused the skin to separate extensively from the stringers. 

Panels C1, C2, and C3 were fractured along the centerline of the panels, without any sign of separation 

of skin from stringers. It appears that the crack started from the impact damage area and propagated 

across the width of the panel, simultaneously fracturing both skin and stringer. Complete details of the 

test, results, and the analysis are presented below. 

5.2.1 Experimental Results 

The strain gage readings for all locations of the six panels for compression tests at various load levels are 

summarized in Tables 20 through 25. Photographs of two panels following the compression tests are 

shown in Figures 52 and 53. Table 26 summarizes the results of the compression tests in terms of failure 

stress and strain. The average strain is computed by dividing the crosshead displacement (A) by the 

length of the panel (P). The stresses are computed on the basis of the gross cross-sectional area of the 

panel with nominal dimensions. In the test column of the above table, Ai-j defines the difference in 

failure strains at locations i and j in pin./in. 

Table 20 

Strain Gage Readings for Panel 61 

LOAD 

(KIPS) 

0 

50 

100 

150 

200 

250 

300 

320 

340 

360 

363 * 

1 

-2 

419 

873 

1,347 

1,815 

2,292 

2,773 

2,964 

3,157 

3,369 

3,401 

- 

2 

3 

465 

91 1 

1,381 

1,850 

2,326 

2,811 

3,002 

3,199 

3,418 

- 

3,451 

- 

3 

1 

421 

86 1 

1,334 

1,804 

2,285 

2,769 

2,960 

3,152 

3,351 

- 

3,381 

STRAIN 

4 

0 

442 

884 

1,360 

1,833 

2,317 

2,807 

3,001 

3,201 

3,430 

3,464 

5 

-2 

445 

871 

1,332 

1,793 

2,268 

2,751 

2,946 

3,145 

3,389 

3,426 

rGE READING ( 

6 

-1 

550 

985 

1,443 

1,898 

2,363 

2,830 

3,015 

3,198 

- 

3,373 

- 

3,399 

- 

7 

-3 

416 

81 2 

1,242 

1,660 

2,068 

2,462 

2,612 

2,730 

2,683 

2,676 

- 

- 

WNJlh 

8 

- 

- 

1 

475 

905 

1,373 

1,837 

2,297 

2,744 

2,913 

3,079 

3,195 

- 

3,212 

- 

9 

-1 

521 

927 

1,379 

1,835 

2,307 

2,787 

2,978 

3,175 

3,435 

3,474 

10 

2 

536 

945 

1,394 

1,851 

2,319 

2,787 

2,972 

3,158 

3,331 

3,357 

11 

1 

462 

898 

1,375 

1,874 

2,419 

3,007 

3,253 

3,558 

4,255 

4,360 

'LINEARLY EXTRAPOLATED TO FAILURE LOAD 

65 
Table 21 

Strain Gage Readings for Panel 82 

9 

-1 

973 

1,919 

2,849 

3,333 

3,817 

4,127 

STRAIN GAGE READING (- fi INAN 

10 

-2 

1,041 

2,065 

3,068 

3,593 

4,120 

4,457 

LOAD 

(KIPS) 4 

-2 

878 

1,874 

2,831 

3,340 

3,853 

4,181 

- 

11 

-3 

862 

1,836 

2,803 

3,327 

3,859 

4,199 

5 

- 

-1 

879 

1,906 

2,947 

3,478 

3,999 

4,332 

- 

- 

12 

0 

94 1 

1,911 

2,853 

3,346 

3,833 

4,145 

10 

0 

562 

1,071 

1,613 

2,174 

2,729 

3,007 

1 

11 

0 

554 

1,048 

1,577 

2,124 

2,729 

3,032 

2 3 6 7 8 

0 

935 

1,935 

2,893 

3,388 

3,878 

4,192 

3 

870 

1,812 

2,724 

3,205 

3,686 

3,994 

3 

806 

1,740 

2,633 

3,094 

3,540 

3,825 

-2 

960 

1,934 

2,891 

3,405 

3,926 

4,259 

-1 

898 

1,874 

2,825 

3,327 

3,831 

4,154 

0 

100 

200 

300 

350 

400 

432' 

-1 

839 

1,758 

2,635 

3,094 

3,550 

3,842 

'LINEARLY EXTRAPOLATED TO FAILURE LOAD 

Table 22 

Strain Gage Readings for Panel C1' 

STRAIN GAGE RE ,DING (- PINAN.) 

71819 12 

31 21 3 1 

I 50 I 505 574 1 521 1 537 1 523 1 612 58 1 

1,069 1,025 1,036 1,020 1,107 

1,581 1,547 1,550 1,537 1,623 1,398 1,453 1,561 

I 100 I 980 1,093 

1,650 

1,891 1,950 2,090 2,246 

I 250 I 2,459 2,353 2,442 2,623 2,883 

1 275* 1 2,708 - 

3,202 

'LINEARLY EXTRAPOLATED TO FAILURE LOAD 

66 
Table 23 

- 

Strain Gage Readings for Panel C2 

LOAD 

(KIPS) 

- 

0 

50 

100 

150 

200 

250 

300 

320 

360 

3 70 

380 

390 

- 

399 

'LIF 

- 

LOAD 

KIPS) 

- 

0 

50 

100 

150 

200 

250 

300 

340 

360 

400 

460 

480 

- 

484' 

STRAIN GAGE READING (-jt INJIN.) 

- 

12 

-3 

521 

1,050 

1,576 

2,126 

2,692 

3,304 

3,571 

4,233 

4,468 

4,734 

5,003 

5,245 

- 

- 

- 

- 

4 

3 

472 

988 

1,489 

2,010 

2,526 

3,050 

3,261 

3,683 

3,797 

3,917 

4,030 

4,132 

- 

FAILU 

- 

2 

5 

- 

5 

0 

48 1 

997 

1,498 

2,019 

2,539 

3,071 

3,285 

3,725 

3,849 

3,975 

4,095 

4,203 

i 

- 

LOAC 

6 

-2 

- 

7 

-3 

- 

8 

-1 

496 

1,014 

1,519 

2,043 

2,564 

3,101 

3,321 

3,765 

3,883 

- 

4,008 

4,123 

- 

4,227 

9 

-1 

517 

1,009 

1,491 

1,992 

2,489 

3,000 

3,210 

3,635 

3,753 

3,877 

3,991 

4,094 

- 

- 

- 

11 

0 

50 1 

1,031 

1,552 

2,098 

2,653 

3,247 

3,501 

4,094 

4,306 

4,508 

4,747 

- 

4,962 

1 

0 

444 

947 

1,434 

1,935 

2,433 

2,937 

3,139 

3,543 

3,647 

3,757 

3,859 

3,951 

ARLY E 

3 

3 

10 

-2 

545 

1,036 

1,521 

2,020 

2,519 

3,031 

3,234 

3,629 

3,730 

3,832 

3,924 

- 

4,007 

475 469 537 444 

979 994 1,027 930 

1,476 

1,985 

2.495 

1,507 

2,035 

2,562 

1,499 

1,985 

2,471 

1,404 

1,883 

2,357 

3,015 

3,224 

3,100 

3,317 

2,965 

3,161 

3,541 

2,823 

3,005 

3,643 3,744 3,322 

3,753 

3,871 

3,852 

3,968 

3,633 

3,731 

3,816 

3,376 

3,442 

3,980 4,070 3,480 

4,078 

- 

rRAPO 

4,162 

- 

3,892 

ATED T 

3,514 

Table 24 

Strain Gage Readings for Panel C3 

STRAIN GAGE READING (-jt 

- 

INAN.) 

- 

8 

2 

514 

1,057 

1,556 

2,060 

2,577 

3,101 

3,513 

3,718 

4,135 

4,736 

4,912 

- 

4,932 

- 

- 

2 

3 

457 

997 

1,494 

1,993 

2,507 

3,028 

3,443 

3,650 

4,078 

4,732 

4,954 

- 

4,979 

- 

- 

3 

4 

426 

969 

1,462 

1,959 

2,465 

2,979 

3,383 

3,586 

3,999 

4,621 

4,827 

- 

4,863 

- 

- 

4 

6 

455 

1,021 

1,541 

2,063 

2,597 

3,142 

3,575 

3,794 

4,243 

4,933 

5,173 

- 

5,200 

- 

- 

5 

3 

461 

1,001 

1,500 

2,004 

2,522 

3,052 

3,481 

3,698 

4,158 

4,912 

5,222 

5,255 

- 

- 

6 

2 

540 

1,061 

1,536 

2,012 

2,501 

2,995 

3,386 

3,581 

3,978 

4,561 

4,746 

- 

4,767 

- 

7 

-1 

495 

1,035 

1,535 

2,045 

2,569 

3,109 

3,540 

3,758 

4,215 

4,950 

5,227 

5,258 

- 

- 

- 

9 

1 

53 1 

1,044 

1,520 

2,006 

2,507 

3,018 

3,423 

3,625 

4,041 

4,677 

4,888 

- 

4,912 

1 10 11 12 

1 

41 7 

953 

1,444 

1,934 

2,433 

2,939 

3,339 

3,539 

3,940 

4,564 

4,773 

4,796 

-1 

506 

1,092 

1,648 

2,233 

2,860 

3,528 

4,099 

4,411 

5,118 

6,555 

7,283 

- 

7,358 

0 

517 

1,050 

1,546 

2,048 

2,567 

3,098 

3,519 

3,730 

4,169 

4,850 

5,084 

5,110 

-1 

547 

1,060 

1,544 

2,035 

2,543 

3,057 

3,466 

3,672 

4,095 

4,737 

4,950 

- 

4,974 

LOAD 

67 

*LINEARLY EXTRAPOLATED TO FAlLUl 
Table 25 

- 

Strain Gage Readings for Panel C4 

LOAD 

[KIPS) 

- 

0 

50 

100 

150 

200 

300 

400 

500 

550 

600 

610 

620 

637 

- 

*Lir 

1 

0 

450 

939 

1,427 

1,923 

2,094 

3,912 

4,950 

5,485 

6,034 

6,143 

6,265 

6,472 

ARLY E 

2 

1 

58 1 

1,099 

1,641 

2,194 

3,312 

4,467 

5,657 

6,279 

6,931 

7,070 

7,223 

7,483 

TRAPC 

3 

4 

477 

989 

1,510 

2,040 

3,105 

4,219 

5,407 

6,058 

6,811 

6,994 

7,217 

7,596 

ATED 7 

STRAIN GAGE READING (-p INAN.) 

4 

-3 

565 

1,097 

1,647 

2,204 

3,322 

4,472 

5,637 

6,234 

6,859 

7,010 

7,198 

7,518 

FAILU 

- 

- 

- 

5 

1 

51 0 

1,020 

1,543 

2,073 

3,134 

4,233 

5,365 

5,958 

6,565 

6,690 

6,817 

7,033 

I LOAD 

- 

6 

- 

0 

612 

1,116 

1,627 

2,148 

3,195 

4,275 

5,380 

5,952 

6,534 

6,652 

6,773 

6,979 

- 

7 

2 

544 

1,063 

1,600 

2,146 

3,244 

4,386 

5,573 

6,193 

6,799 

6,895 

6,953 

7,052 

8 

-3 

49 1 

1,005 

1,526 

2,055 

3,107 

4,174 

5,240 

5,752 

6,177 

6,224 

6,233 

6,248 

- 

- 

- 

9 

1 

569 

1,071 

1,596 

2,130 

3,196 

4,295 

5,417 

5,991 

6,568 

6,678 

6,789 

6,978 

- 

10 

-1 

527 

1,011 

1,493 

1,984 

2,963 

3,972 

5,008 

5,536 

6,057 

6,151 

6,242 

6,397 

- 

- 

11 

2 

530 

1,050 

1,579 

2,120 

3,203 

4,328 

5,492 

6,096 

6,692 

6,794 

6,877 

7,018 

- 

12 

-1 

565 

1,082 

1,622 

2,169 

3,268 

4,406 

5,579 

6,182 

6,777 

6,881 

6,967 

7,113 

- 

Figure 52. Three-Stringer Panel 61 After Compression Test (Midbay Impact) 

68 ORIGINAL PAGE 

BLACK AND WHlfc PHOTOGRAPH 
03’G1NAL P,4GE 

BLACK AND WHITE PtiOTOGRAPH 

Figure 53. Three-Stringer Panel C1 After Compression Test (Midbay 200 ft-lb, 

1N-inch Impactor) 

69 
Table 26 

Summary of Compression-After-Impact Test 

for Three-Stringer Panels 

32.56 

51.86 SKIN 

38.69 

61.71 SKIN 

46.48 

I 

4,333 

- 

4,000 

5,000 

MIDBAY 

EDGE OF 

BLADE 

DELAM AND 

LOCAL 

DISBOND 

DELAM AND 

DISBOND 

STRINGER 

SKIN61.68 

26.63 

6,888 

3,333 

Ai - j = - El (j.4 I NJI N.) 

STRAIN DlFF 

DAMAGE 

SIZE 

FAILURE 

LOAD (KIPS) 

GLOBAL STRAIN 

FAILURE 

STRESS (KSI) /L INJIN. 

DAMAGE 

LOCATION IMPACT I LOCATION 

PANEL 

B1 MIDBAY 4 IN.2 363 7-8 = 536 

5-6 = 9-10 = 0 

DISBOND 

AND LOCAL 

DELAM 

LENGTH OF 

STRINGER 

432 3-4 = 656 

5-6 = 73 

9-10 = 330 

82 

c1 

SKINSTRINGER 

480 

(NOT 

FAILED) 

1-2 = 280 

9-10 = 250 

5-6 = 170 

2 IN2 399 38.63 5-6 = 390 

7-8 = 700 

11-12 = 280 

c2 NIA 

c3 

I 

1.5 IN. 

LONG 

484 5-6 = 500 

3-4 = 340 

7-8 = 280 

SIDE OF 

BLADE 

637 1-2 = 1,000 

7-8 = 800 

5-6 = 0 

c4 

C1' 

DELAM AND 

DISBOND 

STRINGER 

SKIN3.5 IN. 

LONG 

1.5 IN. 

INSIDE 

DISBOND 

4.5 IN.' 

DELAM 

275 A 

1-2 = 200 

9-10 = 120 

3-4 = 0 

MIDBAY PEN ETRA- 

(VISIBLE 

DAMAGE) DELAM 

70 
Panel B1- The impact separated the skin of the panel from the stringer at the damage location. In this 

panel, the maximum strain occurred close to the impact-damaged location. From A7-8 (536 pinh.), it 

can be seen that the plate failed in the flexural buckling mode. While A5-6 is very small, it was observedthat plate and stiffener failed separately in bending. In this case, the skin side of the panel was bent outward before failure, while the blade was bent inward, as they were no longer bonded together. 

Moreover, just before failure, the plate was observed to be buckled near strain gage No. 7, and the 

damage initiated by impact had propagated across its width. Two computed values of stress are 

therefore shown in Table 26, one for the skidstringer assembly and the second for the skin only. This 

type of failure sequence (stiffener disbonding, skin buckling, and catastrophic failure) has been found in 

cases with thick (36-ply) flanges. The failure stress for the impact-damaged skin panel is given in 

Table 26. 

Panel B2 - As in the above panel, the impact damage separated the skin from the stringer at midstringer, as shown in Figure 52. A3-4 shows the difference in strain at the two opposite surfaces of the 

skin, which indicates plate bending. Again, the small values of A5-6 show that skin and stringers are 

bending separately and away from each other. The effect of the impact damage to the skin panel in this 

case was less severe than in panel B1. Figure 52 shows that the stringers did not fail with the skin. The 

failure strength of 61.71 ksi was computed for the skin with this type of damage. 

Panel C1- This panel was impacted with 100 ft-lb at midstringer without any apparent damage. The 

panel was loaded to 480 kips and was unloaded for later investigation. Table 26 shows that the compression loading induced bending in the panel, as is obvious from the values of A3-4 and A9-10. At a load of 

480 kips, the panel was stressed to 46.48 ksi with a maximum strain level of 5,289 pin./in. at the impact 

location. 

Panel C1‘ - To demonstrate the effect of visible (front surface) impact damage on compressive 

strength, Panel C1’ was subsequently impacted with 200 ft-lb of energy using a V4inch-diameter 

impactor. The impact with this impactor at midbay caused the desired visible damage, penetrating into 

the skin. 

The strain gage locations on Panel C1’ were similar to those shown for Panel C2. The strain gage readings for all locations of the panel for compression tests at various load levels are summarized in Table 22. 

As described earlier, C-type panels failed with skin and stringer as an integral structure in compression, 

as shown in Figure 53. Just before failure, the panel was observed to be in a flexural mode - buckledwith a mode shape characterized by a single half-wave along the length and across the width (see readings 

of gages No. 7, 8, 11, and 12). Then the crack propagated across the width from the initial damaged 

location, normal to the direction of the applied load. In general, the C-group panels’ failure sequence 

was panel buckling followed by catastrophic failure, except for panel C4, where crippling occurred at 

midstringer before the panel buckled. 

Panel C2 - This panel had nonvisible impact damage. The impact produced a 2-square-inch delamination in this panel at the far surface. It failed in a flexural buckling mode at a load of 399 kips, which produced a 38.63-ksi stress and 5,003-pin./in. maximum strain. The strain difference values of 390 pin./in. 

and 700pin./in. at midbay (Table 26) indicate obvious bending of skin and stringer together. The crack 

propagated from the impact site at midbay, across the width of the panel. The damage analysis of the 

panel is presented in Section 5.2.2. 

Panel C3 - The impact at the edge of the flange produced a 1.5-inch-long delamination and disbond 

along the bond line. There was no damage in skin or blade, only in the bonding area. Although the skin 

and stringer were not acting together in the disbonded area, the structure failed in a flexural buckling 

mode similar to Panel C1’ . This is verified by the values of the strain difference on the two sides at 

midstringer, A5-6 = 500pin./in. The compression failure strength of 46.86 ksi with a maximum strain of 

71 
7,283 pin./in. near the damaged area shows the improved damage tolerance of the panel under this 

impact condition. 

Panel C4 - This panel was impacted with 200 ft-lb at a 45-degree angle. The impact disbonded the 

stringer for a distance of 3.5 inches along the bond line and created a 1.5-square-inch delamination in the 

skin. The panel was compressed to a load of 637 kips, which resulted in a 61.68-ksi stress and a maximum 

strain of 7,223 pin./in. A Al-2 of l,000pin./in. and a A7-8 of 800pinAn. indicate a flexural behavior of 

the panel except at the midstringer location. At this location, skin and stringer are separated from each 

other, as shown by the small value of A5-6. A C-scan before impact showed a slight disbond inside the 

skin at the lower end, close to midstringer, which induced crippling prior to flexural buckling. Thus, the 

panel failure initiated with local buckling before its final failure. 

5.2.2 Predicted Strengths 

The residual stress analysis given in Section 4.4 has been employed to determine the damaged strength of 

stiffened panels that have been impact-tested. The criteria were based on the size of the damage throughthe-thickness, lateral to the compressive loading, without taking into account the effects of velocity, 

weight, size, and drop distance of the impactor. The correction factor is applied in calculations where the 

panel was impacted with high-energy (Panel C1 ’). The analysis is strictly applicable for damages due to 

fiber breakage. The results may be erroneous if the analysis is conducted on the panels with matrix 

cracks and delaminations. It is assumed that the damages resulting from impact at midbay are equivalent 

to a sharp-edged through-the-thickness crack. 

Case 1 - Panel B1 

Since in this case the midstringer was debonded under impact damage, the stiffener distance is 14 inches. 

Flange 36-Ply[(O/45/O/-45)2/90]2, t, = 0.234in. 1, = 2.4in. A, = 0.54i11.~ 

Skin Same laminate as in Case 1 t, = 0.34 in. 1, = 19 in. A, = 6.41 in.2 

Blade Same laminate as in Case 1 tb = 0.45 in. 1, = 1.77 in. Ab = 0.798 in.2 

where subscripts f, s, and b stand for flange, skin, and blade, respectively. 

Asti, = Ab -I- A, = 1.34 in.2 RFty = 1.34/(14 X 0.34) = 0.28 

A,, = Astif + A, = 7.75 in.2 (assuming only one stringer is intact with skin) 

From Section 4.4, Cexp = 0.695 

Lateral damage normal to compression loading, 2c = 2.2 in. 

c = 1.1 in. do = 0.025 in. KF= 3.464 oo = 107.03 ksi 

Using Equation (58), KT = 3.56 

From Equation (59, oN/o0 = 0.30, oN = 32.1 ksi 

Equation (65) gives ocs = uN/Cexp = 46.2 ksi 

Average strain E,,~ = aJEL = 46.2/(10.07 x 103) = 4,588 pin./in. 

where E, is the longitudinal modulus of the laminate. 

72 
Experimental results: 

P = 363 kips, ucs = PIAT,, = 46.84 ksi. 

Average strain cavg = A/P =- = 4333 pin./in. 0 065 

15 

where A = global displacement P = length of the panel 

Percentage difference between analytical and experimental results in terms of ucs 

= (46.2 - 46.84)/46.2 = -1.3% 

Case 2 - Panels C1‘ , C2 

Flange 18-ply [(0/45/0/ - 45),/90], tf = 0.117 in. 4 = 2.4 in. Af = 0.27 in? 

Blade 72-ply [(0/45/0/ -45),/90],, t, = 0.45 in. l,, = 1.88 in. & = 0.85 in? 

Skin 54-ply [(0/45/0/ -45),/90],, $ = 0.34 in. 5 = 19 in. A, = 6.41 in.’ 

Astif = Af + & = 1.12 in? hoT = 3 Astif + 19 x 0.34 = 9.77 in.’ 

From Equation (55) 

bty = 1.12/(7 X 0.34) = 0.474 

From Section 4.4, Cexp = 0.58 

Panel C1 I 

Lateral damage normal to compression loading, 2c = 3.7 in. 

do = 0.025 in. c = 1.85 in. KF = 3.464 

Using Equation (58), IC, = 5.13 

FromEquation(55), %/ao = 0.213 uo = 107.03 ksi = 22.8 ksi 

From Equation (65), ucs = %/CeXp = 39.3 ksi 

Average strain cvg = a,,/% = 39.3/(10.07 x 103) = 3,902 pin./in. 

where 5, laminate modulus in longitudinal direction (10.07 x 106psi) 

Since the panel was impacted with 200 ft-lb of energy with a 1/4-inch-diameter impactor, it is assumed 

that the damage corresponds to the limit load condition with a factor of 1.5. Thus, for this panel residual 

strength, uCs = 39.3/1.5 = 26.2 ksi. 

Experimental results: 

P = 275 kips 

ucs = 275/9.77 = 28.15 ksi 

AToT = 9.77 in? 

73 
Percentage difference between analytical and experimental results in terms of uCs 

= (26.2 - 28.15)/26.2 = 7.4% 

Figure 54 shows typical load versus displacement curves for Panels C1 and C1' , where A is the displacement of the panel. 

Global Strain G~~ = A/&' = 0.06/18 = 3,333 pin./in. 

50C 

40(: 

30( 

LOAD 

(KIPS) 

20c 

1 oc 

480,000 LB I 

10 

DISPLACEMENT 

(IN.) 

0.0 0.05 0.i o 

Figure 54. Compression-After-Impact Test - Three-Stringer Panels C1 and C1' 

Panel C2 

Lateral damage size normal to compression loading, 2c = 2.5 in. 

c = 1.25 in. do = 0.025 in. Kf'= 3.464 

Using Equation (58), K, = 4.05 

From Equation (55), uN/u0 = 0.27 uN = 28.9 ksi 

74 
From Equation (65), the residual strength of the stiffened panels, a,,, is 

u,, = uN/Cexp = 28.9/0.58 = 49.8 ksi 

Average strain, eavg = u,,/EL = 49.8/(10.07 x 109 = 4,945 pin./in. 

where EL is longitudinal modulus of laminate. 

Experimental results:. 

P = 399 kips AT,, = 9.77in.2 

u, = 399/9.77 = 40.84 ksi 

Percentage difference between analytical and experimental results in terms of ucs = (49.8 - 44.73)/49.8 

= 10.1 percent. 

5.3 DEMONSTRATION PANELS 

Two five-stringer panels with 17-inch center support spans (designated D1 and D2) were fabricated as 

described in Section 3. These large panels were impact-damaged at Douglas and subsequently 

compression-tested at NASA Langley. Panel D1 had discrete source impact damage at midbay. To 

achieve visible damage, several trial impact tests were conducted on several undamaged parts of already 

impacted two-stringer panels. No NDI-detectable damage was produced when the first panel was 

impacted with a 1-inch-diameter impactor. The second panel was impacted with a 1/4-inch impactor, 

and visible damage resulted. Two aluminum ribs were fabricated from angle sections for each panel and 

attached to the skin with bolts at 37-inch center spans (see Figure 21). Figure 55 is a photograph of the 

panel that was compression-tested at NASA. 

The residual strength analysis technique developed in Section 4.4 has been employed to predict the 

ultimate compression stress for both impact-damaged five-stringer panels. This analysis has not considered the effects of two ribs on each panel or the effect of the difference in per-ply thickness of the skin 

panels on mechanical properties of the material. 

Panels D1 and D2 

Flange 18-ply [(0/45/0/-45),/90], t, = 0.llOin. 1, = 2.4in. A, = 0.264in.2 

Blade 72-ply [(0/45/0/-45),/90], tb = 0.438 in. 1, = 1.89in. Ab = 0.828 in.2 

Skin 54ply [(0/45/0/-45)2/90],, t, = 0.439 in. 1, = 33 in. A, = 14.487in.2 

From Section 4.4 C,, = 0.62 

Panel D1 

The panel showed no apparent damage. A lateral damage normal to compression loading (assumed), 

2c = 0.8 in. 

do = 0.025, KY = 3.464, and u, = 107.03 ksi 

75 
OE’GINAL PAGE 

BLACK AND WHITE PWOTW9APa 

Figure 55. Five-Stringer Panel 50-1 Compression-After-Impact Test 

76 
Using Equation (58), KT = 3.51 

From Equation (59, uN/u, = 0.353 and uN = 37.78 ksi 

Equation (65), ucs= uN/Cexp = 60.93 ksi 

Average strain eavg = u,JEL = 5,695 pin./in. 

where E, is longitudinal modulus of the laminate (10.07 x 106psi) 

Experimental results: 

It is assumed that three stringers were disbonded at the time of panel failure. The failure compressive 

stress, ucs = P/(AToT - 3 Astif) where P = 741 kips. 

u,, = 74U16.89 = 43.87 ksi 

eavg = A/&? = 0.23/56 = 4,100tJn./in. 

Percentage difference between analytical and experimental results in terms of ucs 

= (60.93 - 43.87)/60.93 = 28percent. 

In this case, as mentioned earlier, there was no apparent damage on the panel due to 100 ft-lb impact at 

midbay. The comparison of the predicted and experimental failure stresses of the panel showed that the 

assumed lateral damage of 0.8 inch was very small. The impact might have caused matrix crack damage 

near the skidstringer bond line, which resulted in low residual strength. Moreover, the residual strength 

analysis is not accurate for small damages in the panel. 

Panel D2 

Lateral damage normal to compression loading, 2c = 2.5 in. 

do = 0.025 in. KY = 3.464 uo = 107.03 ksi 

Using Equation (58), KT = 4.05 

From Equation (53, uN/u, = 0.272 ffN = 28.9 ksi 

Equation (65), ucs= uN/Cexp = 47.07 ksi 

Average strain eavg = uCJEL = 4,674 pin./in. 

where E, is longitudinal modulus of the laminate (10.07 x 106psi). 

Experimental results: 

It is assumed that only two stiffeners were effective at the time of panel failure as the stiffeners were 

disbonded before failure. The compressive failure stress, ucs 

= P/(A,oT- 3AStif) 

77 
where P is the failure compression load (755 kips) 

ucs = 755/(20.16 - 3 x 1.09) = 44.7 ksi 

- 4,240 pin./in. 0.238 

56 E,,* = A/[ = - - 

Percentage difference of failure stress from analytical and experimental results 

= (47.07 - 44.7)/47.07 = 5% 

Of the damages predicted, the poor correlation is with the midbay impact, Panels C2 and Dl. The 

damage in these cases is much less representative of a through-the-skin crack than the other damage, 

which involves broken fibers. However, analysis approaches that model the damage as a circular cutout 

yield unconservative results with even greater error. The method proposed here is accurate for damage 

that results in broken fibers, and conservative for nonvisible damage that only involves matrix damage. 

A summary of all of the above results is given in Table 27. 

Table 27 

Residual Strength Analysis 

LATERAL 

DAMAGE SIZE 

PANEL ID 

I 3-81 I 2.2 

5-D1 0.80 

5-D2 2.50 

0.28 I 3.56 I 0.30 I 0.695 I 46.2 

0.474 5.13 0.213 0.58 39.3 

0.474 4.05 0.27 0.58 49.8 

0.355 1 3.51 I 0.353 1 ::6: 1 60.93 

0.355 4.05 0.27 47.07 

EAVG (pIN./IN.) 

ANALYSIS I TEST 

4,588 I 4,333 + 3,333 

4,945 

5,695 4,100 

78 
SECTION 6 

CONCLUSIONS 

The objective of the Composite Transport Wing Technology Development program was to design, fabricate, and test composite panels representative of commercial transport aircraft wing cover panels and 

capable of meeting all strength, aeroelastic, and damage tolerance requirements at the lowest possible 

cost. 

The design goals were accomplished by adopting the load intensity, stiffness, and damage tolerance 

requirements of an in-house technology-driver development aircraft as a baseline. A cover panel configuration was evolved that satisfied the aeroelastic requirements through the use of a high-modules “hard 

skin” approach. The damage tolerance and durability criteria were satisfied through the use of a unique 

combination of high-strength, high-modulus carbon fibers and a two-phase matrix material, which 

effectively controlled the amount of damage resulting from foreign object damage, while retaining 

excellent postdamage strength. The synergistic interaction of the aeroelastic requirement for high torsional and extensional stiffness in the covers, combined with the loading intensities and geometry of the 

substructure, led to a design that made use of simple compact prismatic reinforcing elements adhesively 

bonded to a homogeneous skin panel. This configuration was judged to be the most producible arrangement possible with existing composite fabrication techniques capable of exploiting the benefits of 

automated tape lay-up for large skin panels, pultrusion for constant section stiffeners, and the manual 

labor cost of adhesively bonded assemblies. 

Test panels were manufactured using conventional fabrication methods to minimize tooling costs, and 

excellent part quality was achieved. 

Preliminary tests to determine the susceptibility of the design to impact damage revealed the existence of 

complex peeling failure modes in the baseline design. Analysis tools were developed that allowed an 

understanding of the basic failure mechanism. This in turn led to a simple redesign of the structure, 

which eliminated the peeling defect sensitivity without compromising the producibility of the design. 

Panel compression tests performed on three- and five-stringer damaged specimens demonstrated that 

the design configuration met the damage tolerance requirements. Analysis methods were developed to 

allow prediction of the postdamage strength of the panels. These methods gave good correlation for the 

most critical types of damage (broken fibers) and conservative results for less critical damage cases. 

The program thus accomplished all of its major goals in demonstrating a practical, low-cost, structurally 

acceptable composite wing cover panel concept for commercial aircraft usage. 

79 
ACKNOWLEDGMENT 

The work described in this report was supported by NASA under contract NAS1-17970. The authors 

wish to thank Dr. Mark J. Shuart, the technical program manager of this project, Dr. John G. Davis, 

Jr., and the technical staff at NASA-LRC for their cooperation and support throughout the program. 

REFERENCES 

1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

10. 

11. 

12. 

13. 

14. 

15. 

16. 

Standard Tests for Toughened Resin Composites, NASA Reference Publication 1092, July 1983. 

McCarty, J., and Whitehead, R. C., Damage Tolerance of Composites, AFWAL Contract 

F33615-82-C3213, 5th Interim Report for the Period September 1984 to February 1985. 

Roark, R. J., Formulas for Stress and Strain, McGraw Hill Book Company, fourth edition, 

New York, 1954 (Chapter 15). 

Hart-Smith, L. J., Induced Peel Stresses in Adhesive-Bonded Joints, McDonnell Douglas Report 

MDC J9422A, August 1982. 

Barkey, D., Madan, R. C., and Sutton, J. O., Analytical Approach to Peel Stresses in Bonded 

Composite Stiffened Panels, Douglas Paper 7907, 1987. 

Dickson, J. N., Biggers, S. B., and Starnes, J. H., Stiffener Attachment Concepts for GraphiteEpoxy Panels Designed for Postbuckling Strength, presented at the Seventh DoD/NASA Conference on Fibrous Composites in Structural Design, Denver, 1985. 

Madan, R. C., Walker, K. A., Hanson, B. A., and Murphy, M. F., Impact Damage Analysis for 

Composite Multistringer Bonded Panels, Douglas Paper 7908, presented to 3rd Annual 

ASM/ESD Advanced Composites Conference/Exhibition, Detroit, Michigan, September 1987. 

Wang, J.T.S., and Biggers, S. B., SkinBtiffener Interface Stresses in Composite Stiffened Panels, 

NASA CR- 172261, January 1984. 

Lekhnitskii, S. G., Anisotropic Plates, translated from the second Russian Edition by S. W. Tsai 

and T. Cheron, Gordon and Breach, Science Publishers, Inc., New York, 1968. 

Barkey, D. A., and Madan, R. C., Manual for Program PSTRESS: Peel Stress Computation, 

NASA CR-178408, Contract NAS1-17970, December 1987. 

Nuismer, R. J., and Whitney, J. M., Uniaxial Failure of Composite Laminates Containing Stress 

Concentration, Fracture Mechanics of Composites, ASTM STP593, 1975, pp 117-142. 

Nuismer, R. J., and Labor, J. D., Application of the Average Stress Failure Criterion Part I1 - 

Compression, Journal of Composite Materials, Volume 13, 1979, pp 49-60. 

Whitney, J. M., and Nuismer, R. J., Stress Fracture Criteria for Laminated Composites Containing Stress Concentrations, Journal of Composite Materials, Vol. 8, July 1974, p 253. 

Madan, R. C., Crack-Arrestment in Composite Panels State-of-the-Art Review, Douglas Aircraft 

Company, Report MDC 53787, September 1985. 

Avery, J. G., Design Manual for Impact Damage-Tolerant Aircraft Structure, AGARD-AG-238 

North Atlantic Treaty Organization, October 1981. 

Avery, J. G., and Porter, T. R., Survivable Combat Aircraft Structure Design Criteria and 

Guidelines, Final Report, Contract AFFDL-TR-74-49, 1974. 
Nalml &mulcs and I- Sflacr Administralcm 

17. Key Words (Suggested by Author(s)) 

Damage tolerance, aeroelastic tailoring, 

residual strength, blade-stiffened 

bonded panel, delamination, peel stress 

Report Documentation Page 

18. Distribution Statement 

Unclassified - Unlimited 

1. Report No. 1 2. Government Accession No. 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 

NASA CR 178409 

4. Title and Subtitle 

Composite Transport Wing 

Technology Development 

21. No. of pages 22. Price 

7. Author(s1 

Dr. Ram C. Madan 

Unclassified 

9. Performing Organization Name and Address 

Douglas Aircraft Company 

McDonnell Douglas Corporation 

3855 Lakewood Blvd 

T.nno 

National Aeronautics and Space Administration 

Langley Research Center 

Hampton, Virginia 23665-5225 

12. Sponsorin< A-ss 

ia 

15. Supplementary Notes 

Langley Technical Monitor: Dr. Mark J. Shuart 

Final Report 

Unclassified 92 

3. Recipient’s Catalog No. 

5. Report Date 

February 1988 

6. Performing Organization Code 

8. Performing Organization Report No. 

10. Work Unit No. 

11. Contract or Grant No. 

NAS1-17970 

13. Type of Report and Period Covered 

14. Sponsoring Agency Code 

16. Abstract 

This report deals with design, fabrication, testing, and analysis of stiffened wing cover panels to assess 

damage tolerance criteria. The damage tolerance improvements were demonstrated in a test program 

using full-sized cover panel subcomponents. The panels utilized a “hard” skin concept with identical 

laminates of 44percent @degree, @percent f. 45-degree, and 12-percent 9Gdegree plies in the skins 

and stiffeners. The panel skins were impacted at midbay between the stiffeners, directly over the stiffener, and over the stiffener flange edge. The stiffener blades were impacted laterally. Impact energy 

levels of 100 ft-lb and 200 ft-lb were used. NASTRAN finite-element analyses were performed to 

simulate the nonvisible damage that was detected in the panels by nondestructive inspection. A closedform solution for generalized loading was developed to evaluate the peel stresses in the bonded structure. , 

Two-dimensional delamination growth analysis was developed using the principle of minimum potential energy in terms of closed-form solution for critical strain. An analysis was conducted to determine 

the residual compressive stress in the panels after impact damage, and the analytical predictions were 

verified by compression testing of the damaged panels.