AGARD-AG-238(Addendum)

NORTH ATLANTIC TREATY ORGANIZATION

ADVISORY GROUP FOR AEROSPACE RESEARCH AND DEVELOPMENT

(ORGANISATION DU TRAITE DE LATLANTIQUE NORD)

AGARDograph No.238

Addendum to

DESIGN MANUAL FOR IMPACT DAMAGE TOLERANT

AIRCRAFI TSTRUCFURE

by

MJJacobsonNorthrop Corporation

Aircraft DivisionHawthorne, CA 90250, USA

I

This Addendum has been prepared at the request of the USAF for presentation to the Structures andMaterials Panel of AGARD.

THE MISSION OF AGARD

According to its Charter, the mission of AGARD is to bring together the leading personalities of the NATO nations inthe fields of science and technology relating to aerospace for the following purposes:

- Recommending effective ways for the member nations to use their research and development capabilities for thecommon benefit of the NATO community;

- Providing scientific and technical advice and assistance to the Military Committee in the field of aerospace researchand development (with particular regard to its military application);

- Continuously stimulating advances in the aerospace sciences relevant to strengthening the common defence posture;

- Improving the co-operation among member nations in aerospace research and development;

- Exchange of scientific and technical information;

- Providing assistance to member nations for the purpose of increasing their scientific and technical potential;

- Rendering scientific and technical assistance, as requested, to other NATO bodies and to member nations inconnection with research and development problems in the aerospace field.

The highest authority within AGARD is the National Delegates Board consisting of officially appointed seniorrepresentatives from each member nation. The mission of AGARD is carried out through the Panels which are composed ofexperts appointed by the National Delegates, the Consultant and Exchange Programme and the Aerospace ApplicationsStudies Programme. The results of AGARD work are reported to the member nations and the NATO Authorities throughthe AGARD series of publications of which this is one.

Participation in AGARD activities is by invitation only and is normally limited to citizens of the NATO nations.

Published March 1988Copyright 0 AGARD 1988

AU Rights Reserved

ISBN 92-835-0443-7

&Sn dpil by S04hedhrws"e vs L*Im ed40 CtwelLow, LoNA00t Enss IGIO3"Z

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SUMMARY

Addendum to ")eign Manual for Impact Damage Tolerant Aircraft Structure

In 1981 the Structures and Maeril kPae of AGARD published a Design Manual for Impact Damage Tolerant AircraftStructures (AG 23S) Since thad date, theme have been aiguifcant advances in design to resit impact damage. The Panel hastherior considered it appropriate to record tdin information in an addendum to the AGARDograph.

Supplkment an "Guid de, conception de bs tolcnce des stnacicuravion l'enrdommagement dfi & I'impect.

En 1981 le Panel des stnactures et matdix doI'AGARD apubW no ide deconcepionWa to~irance des strcurd'avn Areaommlleentdi i rmpac (AG 238) Depuis cotte date, sucun progris important nWest h signaler dana le

domaine de In conception de Ia toldrance & I'endommagement dfi h I'impact. Le Panel eat done dc I'avis qu'il seat opportund'en hare itat sou forme de suppl~ment h I'AGAflDograplie en question

Aco!ssion For

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PREFACE

Aircraft mut maintain structural integrity despite variou sources of damage auch as. for exa mlfigue cracking ororrosio. Military aircraft mwtalo withstand, as far asis pr ctica e- duae hilicted by hostile military weap a. The

resistance of the structure to the impact of projectiles is an important patreter in considerationa of vulnerability. It isnecessary to determine the impact lur dracttic ofthe structure under oad, and its residual strength after damage. 71edetail design features of the structure are important in determining the spread of the damage. Where neighbouni system orfuel tank are vulnerable, the degree of penetration of the projectile into the structure is important Hydrodynamic ram, blastand fragmentation effects must be considered.

An essential feature of AGARD activities is the pooling of relevant knowledge with the NATO community, aided by thebringing together of specialists for informed discussions. AGARDograph No.238 was, therefore, produced by the Structuresand Matenals Panel (largely through the efforts of .GAvcry of the Boeing Company, Wichita, as coordinator, compiler andeditor of the manual) to aid the designer making assessments of the tolerance of the structure to various threats, and theprobability of the aircraft surviving the impact, completing the mission and returning safely to base. It describes methods whichexist to determine both the damage resulting from the impact of various t"es of projectilesand the resulting capabilities of thedamaged structure. It also embraces an analogous problem, arising mainly on transport aircraft, of the resistance of thestructure to impact of debris from engine disintegration.

Since AGARDograph No.238 was published in 1981, further work has been performed on the subjects which itaddressed. Therefore, plans for preparing and publishing an addendum to AGARDograph No.238 were initiated, withMJJacobson of Northrop Corporation, Hawthorne, acting as the U.S coordinator, and G.Kagerbauer of MBB acting as theEuropean coordinator.

This document is the result of the SMP's more recent activities in this field, and it contains considerable information that isin written reports of (1) L.Oefly and GJ.Czarnecki of the USAF Wright Aeronautical Laboratories, Ohio (2) J.GAvery,SJ.Bradley and M.R.Allen of the Boeing Military Airplane Co., Wichita, and (3) MJJacobson, R.M.Heitz and J.R.Yamane ofNorthrop Corporation, Aircraft Division, Hawthorne, California.

Thanks are due to all the above contributors.

Otto SensburgChairman, Sub-Committee onSurvivability of Battle DamagedComposiste Structures

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CONTENTS - -

SUMMARY l

PREFACE I

LEST OF FIGURESA

INTRODUC71ON I

SECTION I1- HYDRODYNAMIC RAM I~

IL-I OVERVIEW . --. - -,1.2 -'KIEY DESIGN PARAMETERS INFLUENCING DAMAGE SIZE AND TYPE: I

1.2.1 'A.Sdfenm Attachmentl, 21.2.2 PanelEdge Conditions 21.2.3 Panel Thicknesa, 41.2.4 Projectile Incidence oftIipact, 41.2.5 Test Fhald 51.2.6 Hybhri Desilgna- 71.2.7 Projectile Sheae 31.2.8 Comat Projectile Kinetic Enem3 and Momnentunm, 81.2.9 Fluld Levels, 11.2.10 Slack~ng Seiseuce (No Air Flow) >121.2.11 Temperature sdMoisture 121.2.12 Aillion - 121.2.13 SImdinlonoUnwe, 12

1.3 IHYDRODYNAMIC RAM ANALYSIS~ 131.4 CtONCLUDING COMMENTS 13

REFERENCES FOR SECTION 1 14

SECTION 2 - 6ESIGN GUIDELINES FOR SURVIVABLE STRUCTUJRE INCOMBAT is12.1 MATERIALS AND TEST METHODS! Is2.2 TEST RESULTS AND DISCUSSION: 15

2.2.1 Cliaetelatenof Basic Damaus Graph~e/EpoxyLaminatesL 162.2.2 Residual Strength olilsNstk Dunnged Graphte/Epoxy 19I2.2.2.1 Effective Flaw 19

2.2.2.2 Applied Lead Effects 192.2.3 '-Aslysis Verlicadeon - Warbead Area Teat 24

2.3 REPAIRAIILATY - "x24REFERENCES FOR SECTION 2 25

LJF OF FIGURES

Figure Page

I Aluminium and Carbon/Epoxy Material Comparison 2

2 Stiffener Attachment Techniques 3

3 Test Panel Design Details 3

4 Effects of Panel Boundary Support 4

5 Effect of Panel Thickness on Panel Delamination (C-Scan Measurements) 4

6 Effects of Projectile Incidence 5

7 Effect of Fluid Density 6

8 Northrop Tests 69 Residual Compression Strength 6

10 Relative Residual Compressive Strength 6

11 Hybrid Composite Residual Tensile Strength 8

12 Effect of Fragment Shape and Velocity on ASI/3501-6 Entry Panels 9

13 Effect of Fragment Shape and Velocity on AS 1/3501-6 Exit Panels 10

14 Effect of Simulated Liquid Level Following Simulated Detonations Above an Aircraft 11

15 Overview of Survivability Testing of Advanced Composites 15

16 Ply Orientations Included in Composite Survivability Test Program 16

17 Physical Aspects of Ballistic Damage in Graphite/Epoxy Laminates 17

18 Predictive Models for Physical Damage Resulting from Penetration of Graphite/Epoxy Laminates

by Non-Exploding Projectiles 17

19 Comparison of Ballistic Damage in Hybrid and All-Graphite Laminates 18

20 Damage from Superquick-Fused 23-mm HEI Projectiles, Graphite/Epoxy Laminates, 0.05 to 0.50-in

Thicknesses 19

21 Effect of Laminate and Threat Parameters on Effective Flaw Size 20

22 Effective Flaws for Cubical Fragments Impacting Graphite/Epoxy Laminates 21

23 Effective Flaws for Aligned and Tumbled Armour-Piercing Projectiles Impacting Graphite/Epoxy 21

24 Correlations of Test Data with Residual Strength Predictions Using Effective Flaw Concept 22

25 Correlation of Residual Tension Strength Prediction for 13 Laminates Using Effective Flaws 23

26 Effect of Stress at Impact on Residual Strength and Fracture at Impact (48 Ply 0/±45/90 GR/EP) 23

27 Effect of Load at Impact on Residual Strain Capability, 23-mm HEI Damage 24

28 Damage and Residual Strength Analysis Methods were Validated by Warhead Arena Test Results 24

_ a -i_

Since the publication of the -Design Manual for Impact Damage Tolerant Aircraft Stmucture"AGARDograph No.238 in1981, considerable activity in the NATO community has been performed to develop design concepts and analytical methodsfor composite aircraft structures (including fuel tanks) subjected to impacts from ballistic threats. Much of the activity had theprincipal objective of obtaining a better understanding of the damage modes and their effects on the survivability andrepairability of the damaged vehicles. Although some of the activity featured the use of composite materials of the secondgeneration (Le. bismakimides and toughened resins) and third generation (ie. thermoplastics). the experimental data that weredeveloped with these newer composites have generaly been deemed proprietary and, therefore, have not been published in theopen literature. On the other hand, information developed in programs featuring first generation composites and addressingtopics such as the effects on composite aircraft structure of ballistic impacts, hydrodynamic ram and repairability are availableand are presented in this document.

Effects of key struetural design parameters and threat/tank encounter conditions on hydrodynamic ram damage andresidual strength of composite fuel tanks impacted ballistically are presented in Section 1. Design guidelines for ballistically-survivable composite structure, in the absence of hydrodynamic ram effects, are presented in Section 2. In addition, briefcomments on battle damage repairability, as influenced by structural design and rapid repair needs, are also presented inSection 2.

SECTION I

HYDRODYNAMIC RAM

1.0 HYDRODYNAMIC RAMThe subject of hydrodynamic ram and its potential consequences (e.g. excessive fuel loss, structural failures, fire and

explosions, and aircraft kill) were addressed in Section 2.2.4 of AGARDograph No.238 (Reference 1-1 ). Hydrodynamic ramis the phenomenon that occurs when one or more solid objects, such as missile warhead fragments and armour-piercingprojectiles impact, penetrate, and traverse a fuel tank and generate intensive pressure waves that act on the fuel tank. SinceReference 1-I was published, several hydrodynamic ram research programs have been performed to obtain a betterunderstanding of different aspects of the technology and develop hardening concepts to reduce or eliminate the detrimentalconsequences of hydrodynamic ram.

A major portion of a pilot paper entitled "Damage Tolerant Composite Structure Development" addressedhydrodynamic ram and was presented by Mr Larry Kelly, USAF, to the Ad-hoc Group on Survivability of Battle-DamagedComposite Structures at the 60th Meeting of the AGARD Structures and Materials Panels in San Antonio, Texas, (April22-24, 1985). Much of the pilot paper is incorporatedherein. Other recent reports (e.g. References 1-2 and 1-3) addressinighydrodynamic ram have been made available by AGARD.

1.1 OVERVIEW

Composite aircraft structural response to exploding warhead fragments and penetrating projectiles can vary significantlydepending on the design approach employed, the amount and type of load being carried by the structure at time of impact, andwhether the structure serves as a fuel container. In this latter case, a shock wave is generated at the impact site resulting in veryhigh instantaneous dynamic pressures (> 2,000 psi). These pressures are commonly referred to as hydrodynamic ram and canresult in the delamination, brooming and spalation of large amounts of material from the fuel tank structure. In turn, these leadto significant fuel loss or even catastrophic reduction in the fuel tank strength and/or stiffness. The amount of damage caused byhydrodynamic ram effects is dependent not only on the fuel tank materials and type of construction, but also on the physicalfeatures of the fuel tank and many characteristics of the penetrating projectile. Several of the more important parameters havebeen evaluated through the design and ballistic testing of representative carbon/epoxy fuel tank structures. Results of thistesting are presented herein to provide researchers in the field with the knowledge gained to date, and to provide militaryairframe designers with design concepts and techniques that can be utilized in the conceptual phase of a new airframedevelopment, It is at the early design stage that more damage-tolerant designs or even hydrodynamic ram-tolerant fuel tankconstruction can be incorporated with minimal weight and overall system performance penalties.

1.2 KEY DESIGN PARAMETEl INFLUENCING DAMAGE SIZE AND TYPE

Current carbon/epoxy aircraft structure has been developed to achieve significant weight savings, especially whencompared to ceweaional al-stringer aluminium construction. Therefore, to obtain a feel for the response of carbon/epoxycomposite maserial to hydrodynamic run, 2024-T3 aluminim alloy and equivalent thickness carbon/epoxy (ASI/3501-5A)laminates were subjected to impact by 5/8-inch diameter steel spheres under simulated fue tank test conditions in USAF tests.The visible damage in the composite laminates was orders of magnitude greater than the visible damage in the aluminimpanes (Figure I ). These results were a stimulus for a series of programs evaluating the vulnerability of composite fuel tanks tohydrodynmi raw.

Pertinent test parameters for provi1ng realistic simulaton of actual operational conditions were identified and theirelfects were evaluated. The parameters were:

(1) Projectile shape, hardness and velocity(2) Projectle kinetic energy and momm(3) P!jectlle incidence of impact

2

MATERIAL COMPARISON

is

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2

07076 2024 24 37 57 37 37 20242024 46 46 2024696 6

PLY PLY PLY PLY PLY PtyI L L3600 3000 3500 2400 2600 2620 3072 3010 366 3015 355714324 4566U 4664Y

.250 .0-

Figure 1. Aluminium and Carbon Epoxy Mtaterial Comparison.

(4) Amount of Nuet tank ullage or fluid level(5) Type of test fluid(6) Tank geometry(7) Panel attachnut conditions(8) Type of panel construction and thickness(9) Temperature and moisture conditions of the composite

(10) Airflow

To evaluate the relative importance of these parameters numerous stiffened and unstiffened composite panelsrepresentative of current aircraft wing construction techniques and some baseline metallic panels were subjected to balliaticimpact and hydrodynamic ram forces. Relevant results are contained in the following presentation.

1.2.1 Stlffener Attachment

Stiffener attachment was a very important variable. Stiffeners were either bonded, bolted or stitched to flat panels (Figures2 and 3) in a manner representative of typcal wing cover to spar construction. These panels formed one side of a simulated fueltank teat fixture. All panels were impacted with 5/8 inch diameter steel spheres, weighing 250 grains, (43 7.5 grains - I ounce)at varying velocities. Panel response varied from almost complete stiffener separation to confinement of the total visibledamage between the spars. All panels experienced extensive internal delamination. The order of increasing survivability(measured by total visible damaged area) was:

Bonded (stiffeners disbonded at low projectile velocity)

Bolted (fastener pull-through and large area damage)

Through the (pull-throuigh of the tows)Skin Tows

Stitched (some damage under stiffener)

Double Stitched (visible damage contained between stiffeners and some internal damage under stiffener)

1.2.2 hueld EAI Casdfth

Three panel edg conditions were evaluated under a USAF contract with the Northrop Corp. (Reference 1-4): (1) smplysupported, (2) fully clamped and (3) clamped-sliding The contractor tested 48-ply (0o/4,/-5o9x ASI/3501-6carbon/epoxy unstiffened panels without mechanical preloeds. When attached to the tea tank, an 18 by 8 inch central zone ofeach -an wa in contact with the liquid. The -an boundary condons were applied at the edge of the 18 by 8 inch enutralzone. The panels were impacted as entry panels in hydrodynamic ram teats with 129 grain spheres and cubes at 4,300 and5,200 fps and normal incidence (0 degree obliquity). Edge conditions simulating sunspy supported conditions were achievedby attachig an edge flexibility franse between the teat panel and the adapter plate on the teat fixture. Following thehydrodynanice ram tests, a 7 inch wide hydrodynamic ram-damaged section of each panel was tested for residual tensionstrength. The rang of values obtained for each support condition is presented in Figure 4.

There was less cistry-panel damage with sumply-supported edge conditions. There was only minal tear-through of thepanels at the Joints with the supporting structure for simply-supported edge conditions in contrast with much greater paneltear-through at the panel boundaries with the damped and clamped-sliding edge conditions. The damped panels experienceddelaminaton failures over the entire wetted test section whereas the suiply-supported panels had intermittent delanmtons.

AMILS

3

BCOEDTH U MTO

DMUE STITHED BOLTEDWELINE

Figure 2. Stiffener Atteachant Techniques.

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130 EMG CONCIMow120I

RESZDUTENSILE 110

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CS- CLAMPED SLIDING00 SS- SIMPLY SUPPORTED

C&CS asSc

Figure 4. Effects of Panel Boundary Support.

PANEL THiKJE"100'

70'

at 40' 1030 I

1000 2000 3000 4000 SW0WACTV=f.CflY ffS)

Figure 5. Effect of Panel Thickness on Panel Delamination(C-Scan Measurements).

The greater strains to failure and lesser hydrodynamic ram damage with simply-supported edge conditions were attributed tothe greater in-plane flexibility of the support structure and therefore lesser membrane stresses during the hydrodynamic ramresponse. Analysis showed clamped conditions to be more representative of actual wing box structure.

1.2.3 Panel Thickness

Laminate thickness (i.e. number of plies) had a significant influence on damage size in USAF tests (Figure 5). Visibledamage size in thin composite laminates (<40 plies) that were entry panels in hydrodynamic ram tests tended to conform tothe shape of the projectile on the plane of the impacted laminates. Increased skin gauges can reduce the entry velocity of theprojectile and thus reduce the resulting fluid pressure. Increased skin gauges can also reduce panel deformation due to raminduced wall loadings. However, intermediate thickness laminates can be extensively damaged as a result of delamination andrear surface spallation. Excessive gauges are required to minimise damage. Therefore, increasing skin thickness to preventpenetration is not an effective means of improving structural survivability. A more efficient approach is to design the structureto accept penetration damage by using damage tolerant concepts (see Hybrid Designs, Section 1.2.6). Tests conducted underthe Northrop contract (Reference 1-4) showed no substantial difference in 0.25, 0.34, and 0.45 inch AS1/3501-6 paneldamage in that C-scan detected damage was observed throughout the wetted areas of all the panels in the tests with 129 grainthreats impacting the panels at velocities of 5,200 fps. There were no mechanical pre-loads applied to the test panels in any ofthe Northrop ballistic tests of Reference 1-4.

1.2.4 Prejeedie Ieldenee f lmpeet

Under the Northrop contract (Reference 1-4) the projectile impact tingle was varied from 0 to 45 degrees in a series oftests with 129 grain steel spheres representing an exploding warhead fragment. Impact velocity was 5,200 fps. The residualstrength (Figure 6) of the panels, measured by loading the panels to failure after impact, decreased substantially when theobliquity angle was increased from 0 to 45 degrees.

This indicates that entry panels of fuel tanks are more vulnerable to projectiles penetrating at high obliquity angles asopposed to normal incidence. High obliquity angles are known to cause far more delamination. The visible damage to an entrypanel when a projectile entered at non-zero obliquity was not symmetrical about the entry location, but was very much greaterin the direction of the projection on the panel of the projectile path through the liquid. This strongly implies that the concept ofan instantaneous release of kinetic energy of the projectile as it impacts and/or traverses the liquid is generally invalid for use inthe pmdicton of the structural response of the entry (or exit) panel.

In Reference 1-5 the exmnt of damnage is estimated to vay inversely with the cosine of the angle of obliquity. Thie higherobliquity angles increase the distance te projectile travels before exiting and the more the projectile may deform. The larger

5

LEGEN¢D:

0 5,200 fps shot

7,000 0 4,310 fps shot

0 DATA TREND IN SHOTS WITH

6o01 5,200-FPS NOMINAL VELOCITY

4.00000

5,000

3,000

40

02,000

1,000

0 10 20 30 40 50

Obliquity Angle (deg)

Figure 6. Effects of Projectile Incidence.

that the projectile cross-sectional area is, the greater the drag will be; therefore, there will be more kinetic energy transferred tothe fluid. However, more kinetic energy transferred to the fluid will not always result in greater damage to a panel For example.if a tank is deep enough to slow down a projectile sufficiently so that it does not perforate an exit wall, the damage to the exit wallwill be less severe than in cases when the exit panel is perforated and less projectile kinetic energy is transferred to the liquid

1.2.5 Test Fluid

The Air Force tests showed a distinct difference (Figure 7) in the amount of damage area incurred on entry panels whenwater was substituted for f, el. However, in terms of residual strength the Northrop tests (Figure 8 from reference 1-4) did notsubstantiate the Air Force findings.

The Northrop tests employed ASI/3501-6 entry and exit panels. Residual tensile strengths were obtained after impactwith spheres at a nominal velocity of 4,300 fps for both water and JP-4 test fluid. The lack of a discernible difference in damagewas considered consistent with the finding that strains to failure were not sensitive to fragment impact velocity and the kineticenergy loss of the projectiles in traversing the liquids was essentially insensitive to the type of liquid. There were a number ofdifferences in the Air Force and Northrop tests, including entry panel thickness, exit panel thickness and material type and testtank design. These variables are believed to cause the different test behaviour. Another reason could have beet the fact thatresidual strength does not generally correlate with damage area but rather the damage extent transverse to the applid load.Because doubt remains as to the effect of test fluid it is recommended that only parameter screening be done with water as thetest fluid for safety and expediency reasons. Final design verification testing of a hydrodynamic ram tolerant concept snould beaccomplished with the actual fuel as the test fluid. As an aside, it is noted that the USAF recommends that fuel simulanti shouldnot be substituted for JP-4 fuel in fuel leakage tests.

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6

30 FLUID DENSITY

20

(INCH)10P

0 1000 2000 3000 4000 O00 600

IMPACT ENERGY LB-FT

Figure 7. Effect of Fluid Density.

120 JP-4 & WATER100

RESIDUAL 80TENSILELOAD so

(K-LBS) W

0 J4 JP.4 JP.4 JP.4 WATEER WA ER WATERADimR EXIT amR EWf

Figure 8. Northrop Tests.

25 AVERAGE STRENGTH20

CMRSIN15

(KSI) 10

5

0 AS4/ -~ A1 CELION-STf AS41 .110 i~35014 3501 -A 52460 PEEK 3501-SA HI IO

Figure 9. Residual Compression Strength.

UNDRIAMAGED ULTIMATE STRENGOh1.0 - OMPRS",

0.8

RELATIVE 0.6RESIDUAL. RESIDUAL STRENGTHST~eH 04

02.0 AIW ASIJ CEUXNT 481 EAw ~I

25014 301A GMC PEEK S1dl4A NXl1I

Figure 10. Relative Residual Coxqpressive Strength.

7

1.2.6 Hyb DeignsMany damage-tolerant ehancement design concepts have been proposed for use in carbon/epoxy pri nay stncture.

These generally take the form of various interply, intraply or parasitic damage containment additives. The additions are toprovide crack arrestment or load path redundancy. In addition, the parasitic additions featuring foams are primarily intendedto reduce peak hydrodynamic ram pressures. Hybridizing the laminate with plies of a more compliant material such as wovenS-glass or Kevlar 29 is most common. In addition, higher strain to failure fibres and tougher resins have been employed.Interply hybrid laminates of S-glas and Kevlar 29 have not been very successful in containing damage kr dynamic impactforces, however, nor has material substitution alone been very useful (Reference 1-6). The newer materials were stronger incompression for a given damage size but sustained greater damage for a given threat impact. Under the Boeing contract(Reference 1-5), six different materials were subjected to multiple impacts of three 230 grain fragments at 2,540 fsindry shotsand the residual compression stress measured. The results are shown in Figure 9 and the materials were as follows:

ASl/3501-SA Current Carbon/Epoxy System

IM6/3501-SA High Modulus Carbon/Epoxy

Celion-ST/5245 High Strain Carbon/EpoxyIM6/HX1518 Carbon/Bismaleimide Resin

A4/PEEK CarbowTermoplastic Resin

AS4/3501-6 Current Carbon/Epoxy System

There was considerable variation in damage size obtained, but the compression residual stresses were relatively closeconsidering the wide variation in materials tested. The best approach is to incorporate a crack-arrestment region through theuse of buffer or crack-arrestment strips. This technique provides a region of lower stiffness/higher toughness than the carbon/epoxy primary laminate. As a crack approaches one of these regions, the stress intensity at the crack tip is reduced because ofthe lower stress in the buffer region which, combined with the higher toughness, prevents continued crack propagation.

Hydrodynamic rm damage to composites is too complex to model as a simple crack extension or damage zonepropagation. Composites are intolerant of large out-of-plane deflections resulting from hydrodynamic ram pressures.Extensive ply delamination, along with laminate peeling and fibre breakage, can occur. Such damage results in significantcompression strength loss. Figure 10 from Reference 1-5 shows that regardless of the material type, compression strength lossis dramatic under impact forces. In addition, the strength loss is augmented when the structure is already stressed byoperational loads at the time the impact and/or ram pressures forces are applied.

Tensile or compressive fracture can occur at impact even though the applied stress is below that static residual strengthcorresponding to the damage size induced. The applied stress level at which this type of fracture occurs is called the "impactfracture threshold". The impact fracture threshold for composite laminates containing O-deg plies impacted with non-exploding projectiles is 80% of the average residual strength of impacted panels that are unstressed (Rderence 1-6). This levelwas obtained from single ballistic impacts of tensile loaded panels and found to correlate with low velocity impact data.Recently (Reference 1-5), this 80% threshold was extended to compression-loaded panels subjected to multiple ballisticimpacts. If hydrodynamic ram pressure load damage mechanisms are coupled with already augmented strength loss, dramaticchanges in design concepts are required for damage containment to be realised. Using composites that are tougher and/or withhigher strains to failure or tailoring the laminate orientation has not proven to be adequate in increasing residual compressionstrength (Figure I0). Vulnerability reduction from hydrodynamic ram-induced threats requires the incorporation of damage-containment principles in the initial design process. Selectively employed hybrid materials can permit built-in break points for a"blow out panel" that prevents pressure damage to the primary load-carrying structure. Examples of some hybridizingmaterials that have shown promise from the Northrop contract (Reference 1-4) are woven Keviar fabric 49 and 29. In addition,parasitic rigidized reticulated foam and ballistic foam resulted in reduced hydrodynamic ram pressures and greater residualstrengths of the entry panels (Figure 11 ). The test results of Figure II were obtained in tension tests and may not be applicablequantitatively and/or qualitatively in compression tests. Further tests are needed to augment the meagre amount of availablecompreson data for hydrodymnic ram-damaged panels. The more promising approaches for reducing hydrodynamic rameffects have been those that provide significant energy absorption and/or controlled damage confinemenL The best designconcepts create a tolerable failure path within the stncture by directing the failure along defined boundaries selected so thatcritical adjacent structural elements remain undamaged. An example is the double-stitched concept in Figure 2. With thisapproach the induced damage is confined primarily to the skin between intermediate spas and ribs, so that failure ofhe spar/skin attachment is reduced. When hybrid materials are applied with skin pads at ribs and spars, the propagating fracturedamage is stopped, providing a battle-dmnage-tolerant skin.

Typically, exit pands experienced much greater hydrodynamic num damage than entry panels in the vicinity of theprojectile perfomaion site in o-degve obliquity shots. The difference in the typical hydrodynamic rm damage near thefragment perforation site of the entry versus exit composite panels is attributed to the.'iowing; factors: (1) Te entry panels inthe vicinity of the fragment perforation site were somewhat shielded from the hydrodynamic rmn pressures by the cavity,visible in high speed films, in the wake of frapnents that penetrated the pands and (2) the exit panels of W taks and soepartially-filled tanks were continuously subjected to extremely high pressures due to the liquid compression as the fragmentstraversed the last few inches of liquid immediately before perforating the exit panels.

limb

RIESIDUAL TNSSE STRENGTH120

sgo

Figure 11. Hybrid Composite Residual Tensile Strength.

1.2.7 Pr*ecl SbapeSoft steel cubes ad soft and hard steel spheres were used as the projectile types in Northrop tests (Reference 1-4). The

weight f each projectile was 129 grains. ASI/3501-6 graphite/epoxy and 2024-TS51 aluminium alloy entry and exit panels of0.25 inch thickness and with clamped edge conditions were used in these tests. These shots were all at normal incidence andprojectile velocities ranged from 4,000 to 6,000 fps. The following items were the principal results of this testing.

The scatter in strain-to-failure data and visible damage data of hydrodynamic ram-damaged AS 1/3501-6 entry and exitpanels was independent of the projectile type (i.e. spheres versus cubes). Strains to failure of AS1/3501-6 entry panelsimpacted by cubes were generally substantially less than when impacted by spheres (Figure 12). Strains to failure ofAS 1/3501-6 exit panels impacted by cubes were generally substantially greater than when impacted by spheres (Figure 13). Inshots wi.h AS1/3501-6 entry and exit panels the strain to failure of the entry panel impacted by cubes was generallysubstantially less than that of the entry panel. Only at the greatest fragment velocity (6,000-fps shots) were the strains to failureof AS 1/3501-6 exit panels substantially less than those of the AS 1/3501-6 entry panels. In shots with the 129 grain spheres inthe 4,009 to 5,200 fps range, the strain to failure of the AS 1/3501-6 entry and exit panels penetrated by spheres were in thesame range.

The failure modes of the ASI/3501-6 entry panels included panel tear-through at the edge clamps when the impactvelocity exceeded a threshold of 4,600 fps. As the fragment impact velocity increased, the tear-throughs at the entry paneledges became a prominent failure mode. The failure mode at the AS I/3501-6 exit panels included massive damage in the formof brooming ant delamination on the dry side of the panels in the vicinity (e.g. a 5 inch diameter zone) of the fragmentpenetration site and usually no evidence of even partial panel tear-throughs at the edge damps. Delaminations of theAS1/3501-4 entry and exit panels. detected by C-scans, were throughout the panel zones that were in contact with the liquidprior to the shots.

The visible damage of the 2024-T851 entry panels was confined to the fragment penetration site in normal incidenceshots anl was a relatively clean hole in all the shots. The spheres and cubes generally did not completely penetrate the 2024-T851 exit parels. Generally, the residual tensile strengths of the 2024-T851 entry and exit panels remained essentially constantas the fragment velocity at the impact with the entry panel increased from 4,000 to 6,000 fps.

The test results led to the following conclusions. It does not appear necessary to simulate cubes with spheres (note thatcubes general!y simulate missile warhead fragments better than spheres) in the hydrodynamic ram tests of panels of 0.25 inchthickness without mechanical preloads since the data scatter of strains to failure was approximately the same for both projectiletypes. An adjusted velocity of a sphere cannot generally be achieved that will simultaneously produce entry panel damage andexit patel damage that would have resulted from a shot with a cube. Based on the visible damage and strain-to-failure data, theAS I/3501-6 pnels of 0.25 inch thickness generally were substantially more vulnerableto the 129 grain threats than the 2024-T851 aluminum alloy panels if the fragment velocity at the entry panel exceeded approximately 4,000 fps. However, until thefragment wipac velocity at the entry panel reached approximately 6,000 fps, the strains to failure of the AS 1/3501-6 laminatesexceeded 4,000 nicro-in/m, which typically is the design limit strain for such laminates.

1.2.8 Cemt" Prisjete ie EAM and Momemam

Hydrodynamic ram tests with AS 1/3501-6 entry and exit panels of water-filled tanks were performed by Northrop(Refeutnce 1-4) with spheres having different combinations of projectile nas and velocity that would produce a constantkinetic energy of die sphere at its impact with the entry panel. In other hydrodynamic ram tests in this test series, the mass andvelocity combinations of the spheres were adjusted to produce a constant fragment momentum a the entry panel. Fragment

7000

0129-grain steal cubes

0• 129-gran stool pheres 0I V 311-inch diameter 06ooo I ari.led hole - 8 s

* SPHEE

5000

NdOTE: The strain to failure3000 at 2,800 Ila/ta occurred In a paenl in

r the grip sone at a locatie of panelRtear-through at an edge clsmp during

the hydrodynfmtc ram structuralresponse.

2000

1000

00 1000 2000 3000 00 5000 500

Fragment tet Velocity at Entry Panel (fps)

Figure 12. Effect of Fropet Shape and Velocity on MI13501-6Entry Panels.

'

10

0 129-grain steal tube

0 12"9-grai steel spbere6000 y 3/16-tech diater

drilled bole

30W'

20000

5 , 0

1 000

I t

0 3000

0

120000 J 005 OO6 0

ituesat ipact Velocity at Etry Pawl (fps)

iSure 13. ffeet of traegast 9Sape mad Velocity on A81/3501-6Exit POnels.

AA

mass aid veocity combinaions were 252 gramn at 3,070 fps, 129grins at 4,30 fs, and 87grains at 5,200 fps in constantkinetic usuall teass and were 129 grains at 4,000 fps and 87 grains at 5,930 fps in constant momentum tests in these tests anlprojectileimacts were at aormal incidence,

Mthe psneips tearesult wmsdo th taim to fire wene notoserdysensitive todiffiereae mm and vedocitycombiations,in the coats" fragment kinetic energy teat and in the constant fragment momentum tests. Sinc hydrodynamic ram tests withspheres at -n arbitry contain fragment kinetic energy resulted in average strans to failure within a range of approximatelyI0M it was concluded that strains to failure may be reported on the badis of fragmsent kinetic energy at the cntry panel. It isanticipated that the same conclusion would be reached if the spheres were replaced by cubes in hydrodynamic rain tests.However. this conjectre has not been verified by tests.

1.23 Flu Lasi

7Ue effects of fluid levels (Reference 1-4) were investigated in upward vertical shots, in horizontal shots simulating theupward vertical shots, and in horizontal shots simulating; downward vertical shots. Horizontal shots simulating upward anddownward shots were performed because a missile detonation may occur shove or below an aircraft in level flight. In all theshot, the distanice between the entvy and exit panels (iLe., the tank depth) was 7.6 inches and the projectile imspacts were at

nmdincidnce.

The stris to failure follow*n the upward vertical shots into the tank with ASI/3501-6 panels and the horizontal shotssimulat the vertical shots were withina 20% range and are not considered to be overly sensitive to the liquid levelsi verticalshots and simulated liquid levels in horizontal shots of 2.0, 3.8, 5.8 and 7.6 inches. In the horizontal showt with 129 grainspheres fired at 5,200 fps, -ae tear-througha at edge clamps that occurred in shots into liquid-filled tanks did not occur inshots into simulated partially-filled taks.

In horizontal shots at normal incidence and simulatingsa missile detonation above an aircraft, lkiid levels of 2.5, 3.5, and4.3 inches were simulated in the 7.6 inch deep test tankt with ASI/13501-6 eit panels. The hydrodynamic ram damage to theexit panels in the shots with the 2.5 and 3.5 inch simulated liquid level was substantially greater than that experienced by anyother 0.25 inch thick test panels in the Northrop test of Reference 1-4. The strains to failure of these panels dropped below2,000 miacro-inchi/inch (Figure 14) and were substantially less than all the other strains to failure obtained in the teats ofReference 1-4.

7.000

IO Eit an~e 06.000 EnrPae

05.000

r EXIT PANGEL DiATA TEND

-4.000

c3.000

2.000 NOTES:1. The vertical arrow as

that the panel's residualstrength ws " , to beat least of the aagtultude

1.000 A2. The distance between theI entry sod sxit penels was0 7.6 Inches.

0 1 2 3 4 5 6 7

lwilated Liquid Level (inch)

Figure 14. Effet of USated Litquid Level Following simulatedDetonations Above an Aircraft

12

Th re itts in this tt series led to the following conclusions: The hypothesis that an integral tank is moe vulnerable to amiie detomadion below the aerait tha above the aircraft and that the compression skin is always more vulnerable than thetension skin to warhead fagmet encounters may have to be reversed to account for the dum in tension slims of tanks withcertain liquid levels (eg., approximately 2 to 4 inches) if the detonation occurs above the aircraft. The residual strength of atension skin ofan integral iak an aircraft may not be increased substantiallyif dthe tank is penetraed in aparty-filled condition(se. >2 inches of liquid) ratha than a full condition.

1.2.10 Sli Sequons e (N. Air Flaw)

AS1/3501-6 entry panels of 0.25 inch thickness with two different stacking sequences were subjected to hydrodynamicram tests in Northrop tests without air flow followed by residual tensile stength tests (Reference 1-4). There was no effect of thestacking sequence on the residual tensile strength in the absence of air flow. However, stacking sequence may have a profoundeffect on residual strength when there is air flow, as noted subsequently in Section 1.2.12.

1.2.11 Tepesturwal MlshM

Hydrodynamic ram tests and residual strength tests of AS 1/3501-6 panels were conducted by Northrop (Reference 1-4)with water as the test fluid with the external surface of the entry panels heated to 180F in two of four tests. The other two testswere conducted at ambient temperature. The panels had been conditioned to a moisture content of 1.1% for the tests at 180T1.

The principal result of these tests was that the strains to failure of the AS 1/3501-6 panels were essentially insensitive tothe environmental conditions. The principal conclusion is that the ambient temperature test conditions of the other tests ofAS 1/35016 panels of Reference 1-4 produced strain-to-failure data applicable to the conditions of elevated temperature up toapproximately 180"F and moisture contents up to 1. 1%.

1.2.12 Msrmw

When a composite wing's fuel tank is impacted by a high-velocity projectile, hydrodynamic ram combines with projectiledamage causing bundles of delaminated fibres to protrude. Therefore, a sensitivity study was conducted by the FlightDynamics Laboratory (AFWAL/FIES) at Wright-Patterson AFB to determine the effects of high speed airflow (400 knots)over battle-damaged composite surfaces. Using an A-7 airfoil section as the test bed, effects of airlw were monitored as afunction of the following parameters: material type, threat, impact location, ply stacking sequence, fluid level, fluid type, panelthickness, static tank pressure, and tank volume.

Material types evaluated were graphite/epoxy (AS4/3502) similar to that used on a wing now in service, aluminium(2024 and 7075), and graphite/bismaleimide (T300/V378A). The composite skins sustained significantly more damage thantheir aluminium counterparts, with airflow increasing the original level of projectile/hydrodynamic ram damage in mostinstances. Protruding fibrous material was torn back toward the wing's trailing edge. Woven E-glass outer surface coatings,designed to attentuate delamination cause by low velocity impacts such as tool drops, proved detrimental to structuralsurvivability. Caught in airflow, the relatively ductile woven material tended to hold together and be torn off in one large piece.Along with the E-glass, many adhering subsurface plies were peeled back until reaching an edge, where they separated from thestructure.

Test panels were subjected to threats ranging in magnitude from 12.7 mm armou r-piercing incendiary (API) to the 30 mmhigh-explosive incendiary (HEI). Of all threats, only the 12.7 non API created damage so insignificant as to disallow theoccurrence of airflow effects. Thus, composite aircraft that have been designed to survive the 12.7 mm and lesser threats will beextremely vulnerable in a realistic hostile environment where the 23 mm and 30 mn threats abound.

The impact location significantly affected the amount of airflow damage received. Composite skin panels were foundvulnerable to airflow when impacted near their leading edges. Delamination spread easily along this discontinuous edge,allowing the airflow to wedge the plies apart and peel them back.

Stacking sequence changes affected the shape of original damage and the extent of subsequent damage. Damage on panelshaving a quasi-isotropic lay-up was somewhat circular, whereas on panels having an anisotropic lay-up, the damage spreadperpendicular to the airflow stream (increasing the potential for airflow effects).

The effects of hydrodynamic ram and airflow remained proportional to the fluid level when the tank was more than halffull. The extent of ram/airflow damage was not affected when the panel thickness was increased by one-third.

1.2.13 SihulsdaselUBg

When horizontal shots are performed in tests to simulate vertical shots, there is a need to simulate ullage (ie. airspaceabove the fuel of partially-filled fuel tanks). The motivation for conducting horizontal shots to simulate vertical shots is that atmany test facilities it is muds easier and less costly to conduct the horizontal shots.

Hydrodymic ram tests with vertical and horizontal shots were conducted by Northrop (Reference 1-4) with 7475-T61a, lumun alloy panels of 0.085 lich thickness and ASI/3501-6 graphite-epoxy panels of 0.25 inch thickness. Encapsulatednigidized reticulated fami, encapsulated non-tigidized reticulated foam, and styrofowm were used to simulate the air gapin thecases when a horimatl sot was perforned to simulate a vertical shot. The encapsulation was performed to keep thereticulated fan. dry prior to the shots. The horizontal shots simulated missile detonations below a aircraft in level flight. Allve. ical shom wore -par.

Vertial stohtwith 129grain spheresfired at 5,200 fps nominal velocityin thetestsofthe 7475-T61 patselsof0.085 inchthickness resubled nv Matropi exit panel Maie when there was I inch ol uagbutoot when therewere2inchesfullage.The so-aatohcansd catastrophic failuries of 7475-T6l a*i panels obtained in the vertical shots with allages of I and 2

13

inchs, rspectivly, wer replicated when heiontal shots were performed and the ulle was simulated by encapsulatedrildized reticulated fom. Howevr, when styrofam was used to replace the encepse ridized reticulated foam, therewe catastrophic failures of 7475-T61 exit panels in the shots with both I and 2 inches of ullage. The rigidized reticulatedfoam was ess sf in flatwise com ession, but hevier than the styrofoam.

The entry panel damage. horizontal shots sinulating vertical shots with a 1.8 inch ullage (i.e. a 5. inch liquid level) wasimsemtive to the use of encapsulated gridized reticulated foam versus styrfam in simulating the ul. However, in thevertical shots with 5.8 inch liquid level the entry panels expetienced tear-throughs along their long damped edges in contrast tothe absence of taur-throughs at clamped panel edges in all the horizontal shots with a 5.8 inch simulated liquid level.

The test results led to the following conlusions. A mimm amont of uule which is dependent on p geometry(especially on panel thickness) and mawial, tank depth, and threat encounter conditions is needed to Prevent a clatsrohicfailure of panels above the ulisge in upward shots. The cause of AS 1/3501-6 panel tear-through; at the edge damps thatoccurred in the vertical shots with a 5.8 inches of liquid is not clearly under-tood, but is being attributed principally to pre-loadsdue to the cantilevered tank arrangement and the somewhat different cantilevered edge conditions in de vertical versushorizontal shots. There were no entry panel tear-throughs at edge clamps in vertical shots with liquid levels of 2.0 and 3.8inches end in the horimntal shots simulating these vetical sots. The need for applying pre-lods to panels in horizontal shotsversus vertical shots should be further investigated. The air gaps in vertical shots can be simulated with encapsulated rigidizedreticulated foam or styrofoam in horizontal shots. The encapsulated rigidized reticulated foam will provide a somewhat bettersimulation of the air gaps, but the styrofoam is somewhat less costly and is easier to work with. Encapsulated nor-rigidizedreticulated foam should not be used to simulate the ullage since it crushes due to static liquid presures.

1.3 HYDRODYNAMIC RAM ANALYSIS

The HRSR finite element code is a promising analytic tool for predicting Hydrodynamic Ram Structural Response thatcouples a fluid pressure analysis method developed by E.Lundstrom at the Naval Weapons Center with the BR-IFC finiteelement code developed by Boeing for structural response to highly transient loading (see paragraphs 2.2.4.3.1 and 2.2.2.2.3.1of Reference 1 -1). The HRSR code was used satisfactorily in Reference 1-4 to predict the transient nature of hydrodynamicram structural response. However, the HRSR code has not resulted in accurate predictions of the extent of hydrodynamic ramdantage, which is needed for subsequent predictions of residual strength.

Furthermore, no analysis methods for predicting the extent of hydrodynamic ram structural failures for arbitraryproblems have been developed. Hence, general techniques of predicting residual strength of battle-damaged structure, such aspresented in Section 2, have not yet been developed for application to hydrodynamic ram-damaged structure.

1.4 CONCLUDING COMMENTS

Key conclusions that have been presented are summarised below.

(i) Current carbon/epoxy construction in aircraft can be intolerant to ballistic damage and hydrodynamic ram if notdesigned to survive such threats.

(2) Improvements in composite materials, both in terms of increased fibre strain capability and tougher resins, have notproduced a corresponding increase in ballistic damage tolerance as measured by the failure strain of impacted

compression-loaded panels.

(3) To survive the very high pressure forces created by hydrodynamic ram, damage containment concepts have to beincorporated into the structural design from the start.

(4) Stitching of stiffener elements to the panels provides transverse reinforcement and prevents local buckling prior tofailure and has been successful in resisting propagation of delamination damage.

(5) The shape (eg. sphere versus cube) of a simulated missile warhead fragment may have a substantial effect on thehydrodynamic ram structural response of a fuel tank panel because of the resulting effect of the fragment shape onfluid drag pressure.

(6) In hydrodynamic ram tests of liquid-filled tanks perforated near the centre of a tank wall by a simulated warheadfragment, the very severe structural damage of graphite/epoxy exit panels was concentrated in the vicinity of thefragment perforation sites, wereas the very severe damage of the entry panels was in the vicinity of both the paneljoints and the fragment perforation sites. Hydrodynamic ram damage increased and residual tensile strength of thedamaged panels decreased experimentally with increasing kinetic energy of the simulated warhead fragments.

(7) A missile detonation above an aircraft in level flight may result in massive damage to the exit panels of fuel tanks if.there is a critical depth of the fuel. This critical depth depends on factors such as panel geometry and missile fragmentweight and obNqtaty, and may be less tm three inches of fuel.

(8) In cases when missile detonations occur beneath an aircraft in level fPght, the hydrodynmic ram damage ofcomposite entry paneh wilnt vary greatlys afunctionof liquid level ifthe liquid Wevl exceedsacritical depth (e.&,appromately two inches, depending on panel geometry, the encounter conditions, etc.). Te hydrodynamic ramdamage will increase sbstantially with increasing fragment obliquity angles geat than approximately ten degreeswhen the liquid level is not a variable.

(9) In general, the stacking sequence (epecially of the surface plie) of conpoest panels will have a sbttial effect on.# " "the extent ofhydrodynmae ram damage when the panelsare subjected to airflow conditions, insmuchasthe airflow.tend to greatly enhance damage characterized by fibres with loose ends in te airstream.

14

(10) Elevated temperatures; of 180? and a 1.1% moisture content of AS1/3501-6 panela do nt appear to influencesbanaythe hydrodynamic ram damgeand reutn resal" strength of the damaged panel

(11) The effct of water vaseas JP-4 6We ms the fluid in hydrodynamic ranm tests is not completely understood. Untilfurther test are evaluated, water w recommended for safely reasons in hydrodynamic ram screenin tests, and JP-4tadl is recommended fom the final hydrodynamic ram testing. JPF4 fuel rather than it sulalats, are recommendedfor fued leakage tests.

(12) ff horizontal show ane to be performed to simulate vertical shoos in hydrodynamic rame tests, simulating the allegewith encapsulated rigidized retculte foam is geneiy prfeahle to simsulations with styrofoans. Simulations ofthe ullse with encapsulated non-riidized reticulated foam should be avoided.

(13) The research and development needed to establish dear design guidelines for hydrodynamic ram-tolerantcomposite fue tankls is still in its early staes and much more work needs to be done.

3E3ENC1A FOR SKCFON I1-1 J.GAvery, IDesip Manualfor Impact Damage Tolerant Aircraft Structure," AGARDograph No.238, October 198 1.

1-2 O~Gerbauer, D.Weisburger, and O.Seaaberg. Improvemeent of Battle Dasmage Toleranc for Composite Structures,"Imupact Dawu Tolerance to Composit Sinectuve AGARD Report No.729, Februy 1986.

1-3 T.Hess, -Balistic Survivability Considerations for Aircraft Structures," Impact Deange Tolerance to CompositeStructures, AGARD Report No.729, February 1986.

1-4 NfJJacobso, R.M.Kitz, and JR.Yamane, "Survivable Composite lntegral Fuel Tanks," Contract F33615-82-C-3212,AFWAL-TR-85-3085,iJanuary 1986.

1-5 "Survivable Characteristics of Composite Compression Structure," Contract F33613-83-C-3228, Boeing MiltaryAirplane Co.

1-6 SJ.Bradley and i.GAvery, "Design Guide for Survivable Structure in Combat Aircraft, Contract F33615-80-C-341 1,AFWAL-TR-84-30 15, April 1984.

SEACTON 2

DESiraGL NtSWMR AzAC3ATSThucnmE1

relonsiopredict daage and residual strengthwere very lintited. Subsequendly~the USAF sponsoredaprorm.(Rerence2-2) conduacted by wo ith a mqor objective of updating the semi-empirical methods of Relierasce 2-1 for damage andresidual atreadt The opens objectives oflthe progra were to extend desiw technology for achieving combat-surivalestructure ad so prepare an aecn ddp uid fr implemmnting thet nchology an new aircraft The program Primarilyinvestigated sebe compoutes subjected to conventional weapon threats. Research to the ara of metallic structure waa limitedto the incorporation of peemtig analysis oethods ad threat effects into the design guide. The program included anextensve weapo elets and residuial strengt test program eqiloling graphite/epoxty and hybrid structural elementsdamaged by machined flasm holistic perforations or bind overpressures. The test results were used to finalize analyticalmodels for predicting damage swe and structural deradation io nr composites impacted by small arms projectiles, AAA

hihepoieproectiles and misil warhead fragments. The teat program consisted of over 360 tests designed to refypreviously identife deficies. and included a warhead wan teat to investigte the dfects of warhead detonation oncomposite structural elements and components and to validat the frgent damage simulation methods used in thelaboratory.

2.o MATERIALS AND TEST METHODSFigure 15 shows an overview of the tea program. There were three phases involving non-exploding penetrators hig-

explosive AAA projectles and the fragmentation warhead.

AS4/3501-6 grade 145 tape was selected as the graphite/epoxy product for the test specimens, and grade 145 S-glass350146 and Kevlar 49/3501-6 were utilized for hybrid specimens. T7he nominal ply thicknesses of all the products was0.0055-in/ply.

Ply orentatimon of 0. +45 and 90 degrees were employed. Thirteen different graphite/epoxy lanainate configurations andthree hybrid laminate configurations were tested, as shown in Fgure 16. Specimen thicknesses ranged from 8 plies to 96 plies.

2.1 TEST ItESULTS AND DISCUSSION

The follow"n paragraphs summarize the significant conclusions drawn from analysis of test data obtained under thisprogram and from other sources.

5"5s in e a SLAS AU S ~

P.vDim X;muse s m-av 0 w

REIDUAL 1

WARHEAD AN1

* Figure IS. Overview of flurviabuity iteni of aence composites.

4No"

16

M w - mmiet

63A USIIS3 . 12.5 a5 MS IS U

03 IUS/ts311 GO a 25 U 0.11W1

038loss211 258 U w8 U Ue

65 14 2It6S ". I I Ik ONUS311

as3 14%4S/4% SA n Ila 9 0.01W

GOR *W4U6 U% I Ila a I @ mE

09m t01t4MW-4 2J, ns MA Is G AMn

@7 OSr.46011-UJ* 0* 2 67.8 13 DAM6

07M "U" ,~/4S VAO 33 DA 12 OAM6

%e 9 @AI 0 O* 40

pwim - 111. of I N

Gas lwadSLMOS as 0 w 46 x5 a5 QJ111

Uqi it is GJ1111aL

OKt P4og0am

2.2.1 a ofuleibide D~ gI Grapbe/1xy l(eminua

Deft m whudred bltfzwm fiapere amdlcfctehkdm beib dmmll n epolutlep esand sodevIo preictive mu These icluded I S m aling non-exloinh projeccilm incudivilowp frapfmen(4 so250 Vain cube.) eid snB-r proe*s (.30 mud .50cetibme Wiped and tumle4 There were 10 ifinp of 23 un HENprojecile si 21 ll enmd 26 static charge detonation ofts for Not response. The warhead arewst mnplayed 32

a) The "usl"wv"ledth e meet 'Sigfica pmImes hiuenciq, bdllicpeneraion dIa @sife are projectilepreea lengh bpectsangle (obliqutyN mid bmb k~ 1%anas ni rmssipertsnstoo. ctby non-ildingProjCtie at Ie ewd theo II tai ic dL

b) As a 1-~ xa) lhepmimuaIo Ium avelfrovmuw S~ frg t I asmidl-urnsprojectiles m beF ",dlftnlaimuatdamu moal in dowedloan.The m odelbe th m 1i wl d sin i m ae hite/epmyo w kbmmFlpml~mdISmidhbow

D, -(lS,coa)+2*t

17

WiTIAALOOL AII

Figure 17. Physical Aspcts of Bal lstic Damge? In Graphite/SpOxyLaminates.

t -- - - - . aaiu

2itL

7::*.

00 O.ZiIA 2A

Lawy 1jcm -m - lome~ an to -

(a) VISI OPA

K - -- - - -

I ~ ~ ~ ~ .I ---- - AIU,, U2A1s0

3 00

W MM-

IP4O10 rdc--eMdesfrPhscl,-aeReutnfrmPeertino GapieImyL intsbIl-

Uploti q~ -oectles

NI

I8

where: D, - maximma extent of visible damage including extended peeling of the surface ply;

Lp - projectile presented length;

0 - obliquity angle;

t - laminate thickness.

c) The extent of internal damage surrounding the visible damage (consisting of delamination revealed by ultrasonicscan), depends upon the same parameters as visible damage (defined in (a) above), but internal damage sizeincreases with laminate thickness at a faster rate than visible damage. The predictive model developed for internaldamage in graphite/epoxy is shown in Figures 17 and 18(b) and is given below

D, - (1.8 Lp/cos0)+ 3.1 t

where: D, - maximum extent of internal damage, consisting ofdeamination revealed by ultrasonic scan.

d) For graphite/epoxy laminates, the extent of significant damage is relatively independent of ply orientation exceptthat peeling of the surface ply in laminates fabricated from tape is always most extensive in the direction of thesurface ply.

e) Visible damage induced in graphite/epoxy laminates which are under an applied load at the time of impact does notdiffer significantly from damage induced in unloaded laminates if the applied load is below the threshold for fractureat impact (see Section 2.2.2. Residual Strength, for additional comments).

f) Ballistic damage due to multiple impacts does not differ significantly from single impact damage unless the spacingbetween impacts is such that the damage regions overlap. This was verified for simultaneous impact by HE projectilefragments and by simultaneous impact of warhead fragments.

g) Ballistic damage in thin Kevlar 49-graphite hybrids is similar in size and character to damage in all-graphite/epoxylaminates of the same thickness. However, damage induced in thin S-glass-graphite laminates is larger than all-graphite/epoxy laminates. Figure 19 illustrates these effects.

4.0

0.0 L.901 AP.

04" OmKKtY. RAIR.ID.- UO00 ff 0 OSiOultY.

- 2300ftFigure 19. Comparison of Balliatic Damage in Hybrid and

All-Graphite Laminates.

h) The damage size induced by contact-donating superquick fused 23 mm HEI projectiles is a function of projectileimpact velocity as shown in Figure 20. At 2.200 fps and 0-depee obliquity, a six-inch diameter hole is produced,with one to three additional inches of delamination around the hole. The hole size is of primary significance instrength degradation (as discussed subsequently under Section 2.2.2. Residual Strength).

i) The large base and fuse fragments from a 23 mm HE! projectile, which typically weight over 40 grains. have beenshown to penetrate laminates up to 0.5 inch thick. Most side-spray fragments, which typically weigh less than 40grains, can be stopped by aminates of 0.2 inches thick.

j) Threshold stand-off n for inducing dmage due to te blast effects of enemy 23 mm HE projectiles werestimated for 0/±45/90 and ±45 laminates. Dare charges consisting of 210 gains of Composition C-4 weredetonated in proximity to de lasinat. for this assessment. The ±45 laminates resist blast pressures better than0/±45/901 mtaofthesamedickms.b tovap urcaninducedaminationin thick0/+45/90laminateseven when shere bIs no visible trndm .Tp s nom-visftIh M can severely degrade comprcagion strength.

L . , 2.2.2 seaddSo qlhr3se n d ralis/gu

More dan 130 tsam and compression residual strenag ts were pe formed with specimens contaning -ac-neaws habti dase 23 m HEI damage, ad blat dmage. In addition, re l shear strength resilts wee analysed.

Caedolom draw. from thee tests are ummirised below.

AIL

19

.-

t- -S.T

IICI-

6 I I

:I]IL-IMrmaGs VILOC5? - PTIr

Figure 20. Damage from Superquick-Fuaed 23-m HEI Projectiles,Graphite/Epoxy Laminates, 0.05 to 0.50-in Thicknesses.

2.2.2.1 Effective Flaw

The fasibility of the effective flaw concept for predicting the residual strength of ballistic-damaged composite laminateswas established. This concept correlates ballistic damage with idealized flaws of a shape and size which produce the samestrength degradation as the ballistic damage. Once this correlation has been established, conventional analysis methods such aslinear elastic fracture mechanics can be used for residual strength prediction, employing the effective flaw size as thecharacteristic flaw dimension.

The results obtained from more than 100 residual strength tests with single penetrators and machined flaws verified thefollowing items. The appropriate effective flaw shape for ballistic damage in graphite/epoxy is a sharp-edged through-crack.The effective flaw size is a function of laminate thickness, projectile sie and impact angle, and is insensitive to ply orientation.

The predictive model developed for the size of the effective flaw resulting from non-explosive penetrators is given by the

relationship:

D,f - (0.8 L,/cos ) + 4.3t

where: D5 - effective flaw size, i.e. length of a sharp-edged through-crack causing the same strength degradation asthe ballistic flaw;

L -- projectile presented length;

0 - obliquity angle;

t - laninate thickness.

The data used to develop the model and applications of the model for common threats are shown in Figures 21, 22.and 23.

For 23 un HEI contact detonations, it was established by testing that the effective flaw is approximately the diameter ofthe perforation. Based on limited tests employing 23 min HEI standoff detonations, the effective flaw appeared to be the widthof the region intercepted by fragmnts spaced such that overlapping damages occur., ' heory, at certain critical combinationsof projectile velocity and stand-off distance, this region can be as wide as 20 inches. However, these conditions representextremes that have not been demonstrated.

Figure 24 shows application of the effective flaw concept in predicting the residual tension, compression and shearstrength of ballistic-damnged panels. Effective flaw sizes were computed using the model presented above and used with amodified version of the Whitmey-Nuismer point stress failure prediction method. These predictions are in good agreement withthe residual strength test results, as indicated in the figures. Figure 25 demonstrates applicability for a range of ply orientations.

2.2.2.2 Appld Leai Efwts

For non-exploding projectiles, unstable crack propagation can initiate at impact in certain laminates if the tension appliedloed i above 80% of the static resdual strength. This result applies to laminates containing plies parallel to the principal loaddireetion (0 degree pes in end-loaded panels, for example). If loaded below this threshold at impact, these laminates canexhibit reducel static residual strength when loaded to failure after impact. Reductions of up to 15% were observed, as shownin Figure 26.

A& A,

20

RESIDUALSTRCUOIHOF 0MD L6NGTH OFSHAAP

PANEL. CONTAINING &ALLWIC EDGID FLAW CAUSING Ill SAMIEDAMAGE STRENGTH ONRADATIONI

I- ut ll.Wt W-

0-

Figure 21. Effect of Laminate and Threat Parameters on EffectiveFlaw Size.

2!

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.OPINIO

- ' .-. UPPICIVV~2:. F&AW

LA~MINT TMUESSI - NICHE LAMMIATI TNICWgN - es

FF~IE6EFFECTIVE .J.I LAW F A

EFFECTIVE EFETV

M. I64

16 -IL U --IL?..API 15..IL

0Si TIKM-O AIAT HCW ISO

Figure 2. Effective Flaws for Cuignal Fagnds ImpdactorgPirigPoetlsIpcigGraphite/Epoxy.Lmnts

22

0 00@-EWL@SIR PU~ILESb23-m Wt. COTACT OTOUATION

to a

EPPECf VEffon SIZE 1IN) - - -

POECILES

U)EC~v PLAN~t SIZ (1

-vu-

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23

±10%

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Figure 25. Correlation of Residual Tenaton Strength Predictionfor 13 Laminates Using Zffective, Flaws.

IN

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Figure 26. af fect of Stress at r~act. on Residual Strngth andFracture at tapect. (40 Ply 0/±43/90 GM/ZP.

q ,

24

LAminates witieady ±45 degre plies loaded a tensi o o aeibit either of the showe degradations due to the loadwhea Impacted by non plod peciectiles, amid will survive aspact at loads up to the static residual strength level.

150 expluive 23 ow HEI projectiles cn initiate hlre at impact at applied testaih load between one-third anid one-hlf of the static residual strNgt d the laminate. This is shown in Figure 27 for contact detonation.

s$.saw

SYRIAN

1111- 3! GINaOe

2.2.3 Andysia Verifllaa - Warbmad Area Tedt

A warhead area teat was conducted to establis the rsne~ Of composite structural elements and componenits to afragmentation warhead, and to verify the damage. aepon d residual stengt predictive nmodels developed fromn labortrtesting, In the test arrangement, composite panels were arranged insa 12 foo radius around the wartead. Figure 28 shows acomparson of predicted and measured damage siues and residual strengths for two sections of 01±45190 40 ply panels. Thegood correlation is evident.

apm 5.5 I's

Opans am ast ,saw Los 0.0 at as I's-a" use &- at mam u Tam

a* assan as ammaro minusm

Figure 28. DOW90 and rMeideal Strength, Afelygia Methods wereValidated by Warbead Arena Test Umulta.

2.3 71M2RALTV

The establishment of repair concepts for composite structures is of great important sinc ill-conceived reparr conceptsmay result in the loss of both substantial weight savage and increased arcraft performsnce which tohe haow been thediving force fOr the development mid mse of composite aicraft sucre.Another factor tha affected the development ofrepair cosceptu the ONeed for rapid repairs of bett damage in oider toreturnaircraftto flihtad susain thesortie tteofthefleet in a short war.

* ~ ThM it lb1 mi inremlegly-F -parn that the design of repairal comsposite structure should be addressed aly inthe design age lofff ros. ie. Stuctural alyseiod oesdfreautn earcner ag rmlshtOWemN ploput lso lmetive computer propun. Thie fliieelemen piuurans are most advfageaouslymusd atmasrctid

baindevlopdfoeaelthe emaesthingavaltabe rpidrir ofabatte-damoled hetd

Memuchths~ee repir tht whe eldukfisui're by theiitial design. .tudesifh u h oaanrosn oueamovrarcefepslo repairk t od daipie stUctue h nuiodymhapo ovm Covaed putsaft

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25

am diflico ID repair than conventiona baked meufli skn-atringer contruction. Honeycomb construction poses particular

ven oeoeh suh-slrvcfsre as wedl as the types aoius, my have a suatial impact on battle damage propagation(uecisely lnaem of hydrodynamic -a and blals the e-a of the damage and the resulting repairs. Fsilures (eg.deleainalk and flilter pvll-lhoua) of composite sub-structure when hydrodynamic rain and blast pressures arm

tru5e bornva airrs skim to the sbatrucr, we generally of much am concern thin failures oftmetallic sub-structureNumerous composite repar cocepts bavn bee. oulnessgtd. For example, thick composit strucwrcs& Senrally

sevesaitae boaked repars (too amuch material would bane to be removed to obtain a satisfactory bonded repair), whereasboWde repairsawe eacrally suitable farthinner tructure. AN ftheoonceptsmusof e aitddress;the foiowngftors:(1) the eqfipment needed to asertain the damage extet, (2) the repmircnlsii (eg. aerodynamic smoothnessflutter, etc.) andthe repai concept (eag boiled repar external bonded pateh. flush scar bonded patch, etc., and (3) dhe site selection for tierair (eag depot vesu held repair

ADl of the aforementioned issues are be"ngaddressed (e-g. Reference 2-3) by aircraft repair specialists and designsers in thefurther development of the aircraft repai technology. The continual evolu of the repair techanology is proceeding to ensurethat the moat satisfactory use of composites, to reduce structural weight ad die part count of military aircraft will be achieved.

REFER11ENCES FMR SECTION 2

2-1 J.G.Avery, 'Deips Manual for Impact Damage Tolerant Aircaft Structure," AGAilDograph No.238, October 1981.2-2 SJI.Bradley and .G.Avwy, "esign Guide for Survivable Structure in Combat Aircraft," Contract F3361 5-80-C-341 1,

APWAL-TR-84-30 IS, April 1994.

2-3 LGOKely, Composite Structure Repair, AGARD Report No.7 16, February 1984.

,fA iaw,

mEPOUTr DOCUMENTATION PAGE1. Rgdpgat'u Ralriac 2. Od~.wrs rs5 3. FurhAc bhmc. 4. seawff autcih

AGARD-AG-238 ISBN 92-835-0443-7 UNCLASSIFIED

5. Ovar- Advisory Group for Aerospace Research and DevelopmentNorth Atlantic Treaty Organization7 rue Ancefle, 92200 Neuilly star Seine, France

6.1111.ADDENDUM TO DES1GN MANUAL FOR IMPACT DAMAGE TOLERANTAIRCRAFFSTRUCTURE"

7. Presaed

3. Auihuls)/Eidtes) 9. Date

MiJacobson March 1988

1&A~s/E~s~d~ess Northrop Corporation 1.Pr

Aircraft Division 34Hawthorne, CA 90250, USA

12. DitribuUOII Sttnth This document is distributed in accordance with AGARDpolicies and regulations, which are outlined on theOutside Back Covers of all AGARD publications.

LI. Keywor~Desak rs

Airframes Vulnerabilityimpact DesignDamage Manuals

14. Abstac

In 1981 the Structures and Materials Panel of AGARD published a Design Manual for ImpactDamage Tolerant Aircraft Structures (AG 238). Since that date, there have been significantadvances in design to resist impact damage. The Panel has therefore considered it appropriate torecord this information in an addendum to the AGARDograph.

This Addendum has been prepared at the reqest of the USAF for presentation to the Structuresand Materials Panel of AGARD.

192 0

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