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NASA Technical Memorandum 102679

(NASA-T_-IO?oTV) UNIQUE FAILURE BEHAVIOR OF

METAL/CnMPOSITb AIRCRAFT STRUCTU_&L

CnMPONENTS U_OER CRASH TYPE LnAO5 (NASA)

32 D CSCL 20K

G_/39

N_0-24o6o

Unclas

029075Z

Unique Failure Behavior of

Metal/Composite Aircraft

Structural Components Under

Crash Type Loads

Huey D. Carden

May 1990

N/LRA

National Aeronautics and

Space Administration

Langley Research Cenlef

Hampton, Virginia 23665- 5225

UNIQUE FAILURE BEHAVIOR OF METAL/COMPOSITE AIRCRAFT

STRUCTURAL COMPONENTS UNDER CRASH TYPE LOADS

Huey D. Carden

NASA Langley Research Center

Hampton, VA

SUMMARY

Unique failure behavior results are

presented on some of the crash dynamics

research conducted with concepts of

aircraft elements and substructure which

have not necessarily been designed or

optimized for energy absorption or crash

loading considerations. To achieve desired

new designs which incorporate improved

energy absorption capabilities often

requlres an understanding of how more

conventional designs behave under crash

type loadings. Experimental and analytical

data are presented which indicate some

general trends in the failure behavior of a

class of composite structures which include

individual fuselage frames, skeleton

subfloors with stringers and floor beams

but without skin covering, and subfloors

with skin added to the frame-stringer

arrangement. Although the behavior is

complex, a strong similarity in the

static/dynamic failure behavior among

these structures is illustrated through

photographs of the experimental results

and through analytical data of generic

composite structural models. It is believed

that the thread of similarity in behavior is

telling the designer and dynamists a great

deal about what to expect in the crash

behavior of these structures and can guide

designs for improving the energy

absorption and crash behavior of such

structures.

INTRODUCTION

The NASA Langley Research

Center has been involved in crash

dynamics research for several years dating

to the early 1970's. For about 10 years the

emphasis of the research was on metal

aircraft structures during the General

Aviation Crash Dynamics Program

(References 1 to 13) and a transport aircraft

program which culminated with the

controlled crash test of a B-720 aircraft in

1984 (References 14 to 16). Subsequent

to the transport work, the emphasis has

been on composite structures with efforts

directed at developing a data base of

understanding of the behavior, responses,

failure mechanisms, and general loads

associated with the composite material

systems under crash type loadings (See

figure 1). Considerable work has

been conducted into determining the

energy absorption characteristics

(References 17 to 20) which indicated that

composites can absorb as much if not

considerably more energy than comparable

aluminum structures. However, because

of the brittle nature of the composites,

attention must be given to proper geometry

and designs which will take advantage of

the good energy absorbing properties while

at the same time providing desired

structural integrity when the composites are

fabricated into aircraft type structural

elements and substructures. To achieve the

desired new designs often requires an

understanding of how more conventional

designs behave under crash type loadings.

The purpose of this paper is to

preSehfdata on the unique failure behavior

from some of the research conducted with

concepts of aircraft elements and

substructure which have not necessarily

been designed or optimized for energy

absorption or crash loading

considerations. Experimental and
analytical data are presented which

indicatessomegeneraltrendsin thefailure

behaviorof a classof compositestructures

which include individual fuselageframes,

skeletonsubfloorswith stringersandfloor

beams but without skin covering, and

subfloors with skin addedto the framestringer arrangement. Although the

behavioris complex,a strongsimilarity in

thestatic/dynamicfailurebehavioramong

these structures is illustrated through

photographsof the experimentalresults

and through analytical data of generic

compositestructuralmodels.It is believed

that thethreadof similarity in behavioris

telling thedesigneranddynamistsa great

deal about what to expect in the crash

behaviorof thesestructuresandcanguide

designs for improving the energy

absorption and crash behavior of such

structures.

IMPACT DYNAMICS RESEARCH

FACILITY

The information presentedin this

reportwasgeneratedandpublishedduring

the transport and composite aircraft

components phases of the impact

dynamics researchprograms at NASA

Langley ResearchCenter.The research

reported herein was conducted by

personnelat theLangleyImpactDynamics

ResearchFacility (shownin figure 2) and

with other testing equipmentassociated

with the installation. The facility is the

formerLunarLandingFacilityusedto train

astronautsfor moonlandings.Thefacility

is 220 feet high and400 feetlong. Three

setsof legs on the sidesand two on the

eastend supportthe upperlevels of the

gantry.Accessto thetoplevelsis provided

throughan elevator. In the early 1970's,

the structure was converted for crash

testing of full-scale general aviation

aircraft. Reference21 providescomplete

detailsof thefacilityandtesttechniquesfor

full-scale aircraft testing. Figure 3 is a

photographof a70foot highVerticalDrop

Test Apparatusoften usedfor full-scale

aircraft section,components,and/or seat

testing.Statictestingmachines,andother

apparatusarealsoaddedcapabilitiesat the

facility for metal and compositeaircraft

structuraldynamictesting.

ANALYSIS TOOLS

To gain understanding of

fundamentalphenomenaandthephysicsof

behavior,the experimentalresearchwith

structuresundercrashloadingsis generally

accompaniedby analytical prediction or

correlation studies whenever feasible.

Thus,variousfinite elementcodeswhich

have capabilities for handling dynamic,

large displacement, nonlinear response

problemsof metalandcompositestructures

areusedastoolsin theresearchefforts.

DYCAST ComputerCode

The analyticalresultspresentedin

this overview were generated with a

nonlinear finite element computer code

calledDYCAST(DYnamicCrashAnalysis

of STructures(Reference22)developedby

Grumman AerospaceCorporation with

principal supportfrom NASA andFAA.

The basic elementlibrary consistsof (1)

stringers or rod elements with axial

stiffnessonly; (2) three-dimensionalbeam

elements with 12 fixed cross-sectional

shapes typical of aircraft structures with

axial, two shear, torsional, and two

bending stiffnesses; (3) isotropic and

orthotropic membrane skin triangles with

membrane and out-of-plane bending

stiffnesses;(4) isotropic plate bending

triangles with membrane and out-of-plane

bending stiffnesses; and (5) nonlinear

translational or rotational spring elements

that provide stiffness with user-specified

force-displacement or moment-rotation

tables (piece-wise linear). The spring

element can be either elastic or dissipative.

The springs are useful to model crush

behavior of components for which data are

available and/or whose behavior may be

too complex or time consuming to model

otherwise. A contractual effort is

underway to add curved composite beams,

composite plate, and curved shell elements

to the DYCAST element library.

2
TESTSPECIMENSAND DESCRIPTION

Full-ScaleMetalAircraftStructures

NASA Langley ResearchCenter

hasconductedseveraltestsof metalaircraft

sections to support transport aircraft

researchefforts. Selecteddataonthecrash

behavior of full-scale metal transport

categoryaircraft sections(References23

and24)areincludedin thepresentpaperto

demonstratewhatappearsto beimportant

similaritiesin behaviornotednotonly with

themetalfuselagestructuresbutalsoin the

compositestructuresdiscussedherein.

Two 12-footlong fuselagesections

cut from an out-of-service Boeing 707

transport aircraft were drop tested to

measure structural, seat and occupant

responsesto vertical crashloads, and to

provide datafor nonlinearfinite element

modeling.Onesectionwasfromalocation

forward of theaircraftwingsandonewas

from aft of the wing location. Figure 4

presents a photograph of one of the

sectionssuspendedin the Vertical Drop

Test Apparatusat the Impact Dynamics

ResearchFacility. The sections were

loaded with seats, anthropomorphic

dummies,dataacquisition systempallet,

power pallet, and camerabatteriesto test

not only structural, seat, and occupant

responses but also to test the pallet

equipment to be used in the full-scale

transportcrashconductedlater.Thereader

should refer to the particular reports for

more complete descriptions of the test

articles since such information is not

repeatedin thisreport.

CompositeStructures

Single Composite Fram¢s.- Various crosssectional shapes for fuselage frames are

used in metal aircraft and are often

proposed for composite structures. Figure

5 presents sketches and photographs

illustrating four of the more common

geometries, , I-, J-, C- and Z-cross

sectional shapes of which several circular

frames were fabricated for testing to add to

the composite structures data base. To add

out-of-plane stability to the frame concepts

(with the exception of the Z-section

frames), 3 1/2 inch wide skin material was

added which enhanced the ease of testing

of both symmetrical and other

nonsymmetrical sections. The skin, a

1+45/0/90] 2s lay-up sixteen ply (.08

inches) thick, was cocured with the 6 foot

diameter frames which have the lay-ups as

indicated in Table I. The frames were

constructed in two different heights, 1 1/2

inches high and 3/4 inches high, to

investigate the effect of frame height on

behavior and responses.

One of the first geometries to be

studied under static and dynamic loadings

was the Z-cross section. Figure 5(b) is a

photograph of Z-cross section fuselage

frames used in the initial studies of the

behavior of composite structural elements

under impact loads. Figure 6 shows a Zframe suspended in the drop apparatus

prior to testing. The apparatus was

constructed with guide rails, a rear metal

backstop, and a front plexiglas sheet.

During free-fall the specimen was guided

and the front and rear backstops prevented

appreciable (but not all) out-of-plane

bending or twisting during impact, and

allowed photographs/motion picture

coverage through the front plexiglas plate.

The six-foot diameter frames were

constructed of 280-5HA/3502, a five

harness satin weave graphite fabric

composite material. The height of the

frame was 3 inches with a total width of

2.25 inches and about 0.08 inches thick.

Lay-up of the frames was quasi- isotropic.

Initial tests were with 360 degree frames

made from four 90 degree segments joined

with splice plates as shown in figure 5(b).

Additional tests were conducted with half

frames since the top half of the complete

frames were undamaged in the tests.

The approach of studying simple

structural elements and then moving to

combinations of these elements into more

complex substructures has been taken in

the development of a data base on the

dynamic response and behavior of

composite aircraft structures. The

approach parallels the one used during the

general aviation and transport aircraft
programs.Consequently,threecomposite

subfloor structures were fabricated

following theinitial investigationof theZframesdiscussedabove.

Subfloor Structures.- Figure 7 is a

photograph of the skeleton and skinned

subfloor specimens constructed with three

of the single Z-section frames similar to

those that were studied earlier. Pultruded

J-stringers attached the three frames

through metal clips and secondary bonding

methods. Aluminum floor beams tied the

top diameter of the frames together to form

the lower half of the subfloor. Notches in

the frames allowed the stringers to pass

through the frames. Two subfloors

without skin (called skeleton subfloors)

were fabricated. A third specimen (called

skinned subfloor) had a +45 lay-up skin

bonded and riveted to the frames to form

the lower fuselage type structure.

Full-Scale Composite Aircraft.- Two fullscale composite general aviation aircraft

structures, two complete wing sets, and

landing gears have been obtained for future

testing. Because of the scarcity and

expense associated with obtaining fullscale composite aircraft for impact testing,

plans include multiple usage of the

structures. Various acoustical, structural

mechanics, and impact dynamic tests are

contemplated for the aircraft to provide a

broader data base than just crash behavior

of the structures. These tests will be

integrated into the other research programs

of the Landing and Impact Dynamics

Branch at the NASA Langley Research

Center.

RESULTS AND DISCUSSION

Figures 8 to 18 present results from

the studies of full-scale aircraft structures,

composite fuselage frames, and subfloors

under static and/or dynamic loadings.

Analytical results are also included which

illustrate crash related behavior of some of

the structures. Photographs are included

which emphasize the failure behavior of the

composite and metal aircraft components

and show a strong similarity in their

behavior. The behavior is thought to be an

important aspect which must be considered

in the design of new structures for

improving the energy absorption and crash

behavior of these type components and

structural elements.

Full-Scale Metal Aircraft Structures

Experimental and analytical results

from studies with full-scale transport

category aircraft sections (from References

23 and 24) are presented in figure 8.

Dynamic tests.- Structural damage and

behavior of the transport aircraft structures

resulting for the 20 ft/s drop tests is shown

in figure 8(a) and (b). The damage to the

iransport sections was confined to the

lower fuselage below the floor level.

Under the vertical impact of 20 ft/s, all

seven of the frames ruptured near the

bottom impact point. Plastic hinges formed

in each frame along both sides of the

fuselage at about 50 degrees up the

circumference from the bottom contact

point (See figure 8(c)). The upward

movement of the lower fuselage was

approximately 22-23 inches at the forward

end and 18-19 inches at the rear for the

section taken from forward of the wing

location (figure 8(a)) whereas in the section

from aft of the wing location (figure 8(b))

the crushing was about 14 inches forward

and 18 inches in_t_e rear. Although the

aircraft structures are metal and the failures

discussed above involve plastic

deformations with some tearing of the

metal rather than brittle fractures, the

general observed failure pattern and

locations for the transport fuselage sections

will be shown to be quite similar to the

results of the composite frames and

subfloors discussed later.

Analytical studies.- A DYCAST model of

the section (from forward of the wing

location) was constructed with sufficient

detail the model the floor, two seats with

lumped mass occupants, and the fuselage

structure to determine if such a model

could predict the response of the complete

section with fidelity. The finite element

model is shown in figure 9. Stiff ground

springs simulated the concrete impact

4
surface.Eachframeof thefuselagebelow

the floor was modeled with eight beam

elementsandfloor andseatrails werealso

appropriate beam elements. Fuselage

structureabovethefloor (notexpectedto

fail) wasmodeledin lessdetail.

The predicteddeformationpattern

of thetwo framemodelis shownin figure

10. As may be noted, the overall

impression from the analytical model

deformationpatternis quite similar to the

visual behavior seen in the experiment

shown in figure 8. The full section

behaviorwasbasicallycontainedin thetwo

framemodel.

In the following sections the

compositeimpact dynamicsstudieshave

taken the building block approach of

utilizing a sequenceof testingandanalysis

which beginswith 'simpler'elementsand

moveto more'complicated'componentsor

substructures. As mention earlier, this

approach was used in the General Aviation

and Transport programs although the GA

data base was being concurrently

developed through full-scale testing.

Eventually it is desired to add to the

program the full-scale tests using currently

available composite aircraft specimens

and/or other full-scale structures that are

designed and constructed for that purpose.

Composite Single Frame Studies

Static tests.- Figure 11 presents static

results from reference 25. A photograph

of the static test apparatus in figure 1 l(a)

shows that the splice plate was at the load

point, thus, the frame failed just outside the

doubler splice plate area by a complete

fracture across the Z-section. Loaddeflection data and the location of failures

of the frame are shown in figure i l(b).

The load-deflection data show a sawtoothed behavior under the essentially

linear behavior of the frame up to initial

failure, and subsequent loading of the

frame after failure being at a new, reduced

stiffness for the frame as a result of that

failure(s). Photographic data in figure

l l(c) shows that the initial failure was

induced by a local buckling of the frame

which occurred at about 18 degrees from

the bottom loading point outside the splice

plate area. Second and third fractures

occurred up the side again at about 54 and

58 degrees under continued loading as may

be noted in the sketch at the right of figure

1 l(b). The sequence of events was: local

buckling near the splice plate induced out

of plane deformations which led to a

fracture of the section, the ends then

remained in contact (jammed together

under the compressive loads in the frame),

the point of fracture moved vertically

essentially in a guided manner until two

additional fractures occurred farther up the

frame.

Static analytical studies.- To demonstrate

analytically the apparent behavior of the

frames under load (exclusive of the local

buckling which actually initiated the failure

in the static case), a DYCAST finite

element model was constructed. For ease

of analysis, an I-section was modeled from

the specimens described in the "Single

Frame Studies" section. The frame was

loaded at the top and a simulated ground

plane (ground contact springs) resisted the

vertical movement of the frame during load

application. Boundary conditions were

imposed at the bottom node of the model to

account for the symmetrical situation thus

only half the frame had to be modeled. The

top node was constrained to allow only

vertical displacement thus simulating the

effect of a very stiff floor across the frame

diameter. The static load was slowly

increased until an input failure strain for the

material (0.0086) was exceeded at the point

of loading and failure was indicated. The

curve labeled unbroken frame in figure

12(a) is the load-deflection plot for this

case I.

An examination of the normalized

distribution of the bending moment on the

frame shown in figure 12(b) provides

insight and a better understanding of the

failure/behavior. Maximum moments are

indicated (just prior to failure) to be at the 0

degree location and between 50-60 degrees

from the bottom contact area. The

locations correlate well with the failure

5
locations in the experiment with the Zframe.

A secondDYCAST case,caseII,

wasalsorun wherein thebottompoint of

theframewasmodeledwith two shortskin

segments to represent the different

boundarycondition following the initial

failureof theframe. Thisconditionwill be

discussedfurtherrelativeto thecomposite

subfloor test with the skinned subfloor

specimen. The curve in figure 12(a)

labeledbroken frame is the load-deflection

response for this frame loading case.

Essentially, the frame load increases to the

point of initial failure at the bottom of the

frame. After the frame fractures the

boundary changes to one considerably

weakened--down to the bending stiffness

of the skin alone at that location. The load

then drops to the lower curve which

represents the stiffness of the section with

the weakened boundary on the bottom end

of the frame, as indicated by the dashed

vertical line between the case I and curve

and the case II curve. The load continues

to increase again along the broken frame

case until a failure at some other locations

on the frame circumference occurs.

Figure 12(c) presents the

normalized bending moment distribution

on the broken frame. As may be noted, the

distribution is quite similar to the initial

model results in figure 12(b). The failure

location is at the maximum bending

moment location predicted to be about 45

degrees which is somewhat lower than the

same location shown in the initial model of

the unbroken frame. The agreement

between the two models, however, is still

considered good. The effect of diameter

on the moment distribution of the frame

was assessed with a model having a 75

inch diameter (twice the initial diameter).

The distribution was identical to the smaller

diameter results, however, the loads

producing the moments differed between

the two models (as expected).

Furthermore, a comparison of the

analytical cases (figure 12(a)) with the

actual static load-deflection of the Z-section

(figure 1 l(b)) indicates very similar loaddeflection behavior patterns as discussed

above. Although the Z-frame had no skin,

if the ends jam together (as they did in

several cases), the boundary is effectively

between the skin stiffened case and a

guided boundary. Thus, the predicted

failure location in the simple beam-frame

model at about 50 degrees (See figure

12(b)) agrees well with the 54 and 58

degree failure locations in the experiment

with the Z-section frame.

Without a priori knowledge of the

manner of the failure noted and discussed

above, the initial formulation of a finite

element model would unlikely incorporate

the necessary failure mechanism/behavior

for the frames. However, knowing the

pattern of behavior can enable the analysts

to formulate adequate finite element models

to predict dynamic responses including the

failure/loads. Additionally, such

information is important to designers of

new structures to be able to design for

impact loads on such structural elements of

an aircraft fuselage.

_D_ynamic tests.- Figure 13 presents results

from reference 26 on the dynamic studies

of the response of composite frames. As

noted in figure 13(a) or (b), the splice

plates joining the segments of the frame are

45 degrees up the circumference from the

point of impact. As shown in figure 13(b)

complete failures (fractures) of the Zsection frames occurred at the bottom and

approximately 60 degrees from the bottom.

Potentially, it appears that the presence of

the splice plates may have influenced the

locations by moving the top failure points

up a few degrees to about the 60 degree

hx:ations.

Composite Subfloor Studies

For the three composite subfloor

specimens used for impact studies, _two

static and two dynamic tests were

conducted on the subfloors. With the

skeleton subfloor, a static and a dynamic

test to destruction was conducted. With the

skinned subfloor, a non-destructive static

test followed by a dynamic test to failure

was conducted.
Static test_.- Figure 14 presents

experimental results (Reference 27) of the

skeleton subfloor specimen following a

static test. As noted in figure 14(a) and

(b), failures on the three Z-section frames

occurred at 13 discrete locations. Unlike

the unnotched single Z-frame, the failures

in this specimen occurred at notches

(which served as stress risers) in the frame

through which the stringer passed.

However, as shown in figure 14(b) the

failures were still near the point of load

application (approximately 12-14 degrees)

and at other circumferential locations of

approximately 55 degrees. In the absence

of skin material, twisting and bending outof-plane occurred with the frames. The

stringers had minimal effect on the

subfloor response with the exception of

maintaining the lateral spacing of the three

Z-frames.

Dynamic test_.- Figure 15 shows the

skeleton subfloor after an impact test onto a

concrete surface at 20 feet per second. In

the dynamic test of the skeleton subfloor

fractures were produced at notches in the

frames (Figure 15(a)). The locations,

shown in figure 15(b), were also near the

point of impact (about 14 degrees because

of the splice plate) and at three other

locations up the circumference of the

frames (55 degrees and 85 degrees) and

involve all three frames for a total of 15

fractures. The impact energy exceeded the

energy absorbed by the local fractures and

the floor bottomed out in the impact.

Figure 15(c) is a normalized strain

distribution measured on the first (end)

frame during the dynamic test just before

first failure. A comparison of the

distribution to the moment distribution of

Figure 12(b) and (c) shows essentially

identical shape between the single frame

and skeleton frame distributions.

Maximums at 0 degrees and at

approximately 50 to 55 degrees agree well

with the analytically predicted locations on

the frame.

Figure 16 presents impact results

for the subfloor with skin after an impact

of 20 feet per second. Figure 16(a) is the

subfloor specimen after the test. Points of

failure of the frames in this specimen are

indicated in figure 16(b). Again the points

of failure are at/near the impact point

(within 12 degrees) and circumferentially at

about 56 degrees up both sides of the

frame on the middle and back frame and

45, 12 and 22.5 degrees on the front

frame. It was observed that the subfloor

impacted first on the front area which

possibly explains the 12 and 22.5 degree

fractures being different from the other

locations. Again all three frames were

involved in the failures. Some

delamination of the frames from the skin

was evident but the skin remained intact.

Figure 16(c) is a normalized strain

distribution (just prior to first failure)

measured on the first (end) frame during

the dynamic test. A comparison of the

distribution to the moment distribution of

Figure 12(b) and (c) and the strain

distribution in Figure 16(c) shows

essentially identical shape as the single

frame and skeleton frame distributions. As

was the case for the skeleton subfloor

maximums at 0 degrees and at

approximately 50 to 55 degrees agree well

with the analytically predicted locations on

the frame.

As mentioned previously in the

frame studies, once the frames fail at/near

the point of impact the broken ends of the

frame often jammed together and moved

upward in a guided manner. In the

subfloor structure, the frames may still fail

completely across the section but the skin

remains intact and serves as a much less

stiff boundary condition for the broken

frames as the deflection increases. Little

energy is involved in snapping the skin

through as the load increases on the

structure (See reference 30 on snapthrough of composite arches). In this

manner, the structural stiffness of the

frame/skin before fracture changes to the

skin only after frame fracture. The

analytical models discussed under the static

frame response/behavior simulated this

type of behavior.

Anldyti¢ol _tudies.-The contribution of the

skin to the stiffness of the section with the
nonsymmetrical frames is illustrated in

figure 17. Static load-deflectiondatafor

the unskinnedsubfloor and the skinned

subfloor along with the DYCAST

predictionsareshownin thefigure. It can

benoted that the subfloor stiffness(with

skin) is approximately three times the

stiffnessof the skeletonsubfloor,thusthe

skin's contribution to the structureis to

maintainin-planedeflectionsof the nonsymmetrical Z-section and prevent any

substantialtwisting of theframes.Out-ofplanebendingand twist were allowed in

the skeletonsubfloor predictions. As a

further noteof interest,if for theskeleton

subfloor, load in the load- deflection

results(with threeframes)is reducedby a

factorof three,goodcorrelationwith single

framedatais evident.

Full-ScaleCompositeAircraft

Other than two support tests for

Army composite aircraft programs (See

ref. 28 and 29), no testing of full-scale

composite aircraft has been conducted to

date at the Langley Research Center as part

of the composite impact dynamics

research. However, as mentioned earlier

composite aircraft fuselage specimens have

been obtained and multiple usage for the

structures is planned because of the

expense and scarcity of such specimens for

crash testing.

General Observations

The response behavior determined

during the studies of full-scale aircraft

sections, fuselage frames, and subfloors

are summarized in figure 18. Figure 18(a)

shows normalized moment distribution on

a representative frame of the various

specimens and Figure 18(b) shows the

failure locations which were noted from

static or d.ynamic tests. The visual

impression is quite striking among the

various specimens. It is suggested that

from the results of simpler frames to the

more complex subflo0rs and full-scale

sections, a strong similarity is evident in

the failure behavior of the structures. The

structures share in common the generally

circular or cylindrical shape, the normal

loading situations, and what appears to be

a similar pattern of failure behavior.

Analytical models of frame structures

under vertical loads have moment

distributions which have maximums at the

point of loading and at approximately 45 to

50 degrees (depending on boundary

conditions) around the circumference from

the ground contact point. Failures of the

structures were noted at these same

locations. Such observations can help

dynamists gain a better understanding of

what to expect from such structures in

crash loading situations, can guide

designers of new structures to better

account for the vertical crash loads, and

allow better energy absorption to be

included in the new designs. Additionally,

the observations can help analysts better

model the aircraft structures for predicting

the failure responses and behavior under

crash situations. The latter task is a

difficult and challenging one, not only for

composite structures but for metal

structures as well. Studies are currently

underway to improve the analysis

capabilities of code and to add composite

elements to finite element libraries such as

the DYCAST program. In addition, new

analysis approaches are being explored

through grants to universities as an

extension of NASA Langley Research

Center's efforts.

CONCLUDING REMARKS AND

OBSERVATIONS

Some unique failure behavior

results from the research with composite

full-scale aircraft sections, composite

structural elements, and subfloors have

been presented. Some observations on the

failure behavior of these structures have

been made and discussed and analytical

results have been included to help explain

some of the behavior noted.

From the observations made in the

overview the following conclusions are

made:

(1) Starting with simple representative

structural elements and moving to more
complex componentscan often provide

better understanding of the complex

local/global structural responses and

behavior.

(2) Uniquefailure behaviorpatternswere

found to be common among full-scale

aircraft sections,compositeframes,and

subfloorswith andwithout skin.

(3)Generallocationsof failuresappearto

occuratthe samestructuralregionsamong

the specimens as a result of similar

geometry (cylindrical shape), similar

loading (vertical), and similar moment

distributiononthestructuresundervertical

loads.

(4)Notedfailureswerelocatedin thesame

regionsas the maximumsin the moment

distributiononthestructures.

(5) The shapeof the distribution of the

moment was independent of the size

(diameter) of the frame/component.

Loads, however, which produced the

failuresvariedwiththestructuralsize.

Basedupontheconclusionsdrawn

from thevariousresearcheffortsdiscussed

in this paper,the following observations

arealsosummarized:

(1) The generalsimilarity of the failure

behavioramongtheaircraftstructurescan:

(a) assist the designer and

dynamists to better anticipate how the

structuresprobablywill fail,

(b) provide guidanceon how and

whereto incorporateand/oroptimizebetter

energy absorption into new aircraft

structuraldesigns,and

(c) aid analyststo bettermodelthe

structures for predicting failure/loads

behaviorundercrashsituations.

(2) To analyticallypredict,in a dynamic

loading situation, such complex failure

events and the loads which initiate the

failuresasnotedin thecompositestructural

elements and sub-components is a

challenge.

(3) Compositecurved beam,composite

plate and shell elements are being

developedand includedin finite element

codeto improvethecapability to analyze

compositetypestructures.

REFERENCES

1. Alfaro-Bou, Emilio; and Vaughan,

Victor L., Jr.:Light Airplane CrashTests

at ImpactVelocities of 13and 27 m/sec.

NASA TP 1042,November1977.

2. Castle,ClaudeB.; andAlfaro-Bou,

Emilio:LightAirplaneCrashTestsatThree

Flight-PathAngles. NASA TP 1210,June

1978.

3. Hayduk, Robert J.: Comparative

AnalysisofPA-31-350Chieftain(N44LV)

Accident and NASA Crash Test Data.

NASATM 80102,October1979.

4. Castle,ClaudeB.; and Alfaro-Bou,

Emilio:Light AirplaneCrashTestsatThree

Roll Angles. NASA TP 1477, October

1979.

5. VaughanlVictor L., Jr.; andAlfaroBou,Emilio: LightAirplaneCrashTestsat

Three Pitch Angles. NASA TP 1481,

November1979.

6. Vaughan,Victor L., Jr.; andHayduk,

Robert J.: CrashTests of Four Identical

Itigh-Wing Single-Engine Airplanes.

NASA TP 1699,August1980.

7. Carden, Huey D.; and Hayduk,

RobertJ.: Aircraft Subfloor Responseto

CrashLoadings.SAEPaper810614,April

1981.

8. Williams, M. Susan;and Fasanella,

EdwinL.: CrashTestsof FourLow- Wing

Twin-Engine Airplanes with TrussReinforcedFuselageStructure.NASA TP

2070,September1982.

9
9. Carden,Huey D.: Correlation and

Assessmentof StructuralAirplane Crash

Data with Flight Parametersat Impact.

NASA TP2083,November1982.

10. Carden,Huey D.: ImpulseAnalysis

of AirplaneCrashDatawith Consideration

Given to Human Tolerance.SAE Paper

830748,April 1983.

11. Castle,ClaudeB.; andAlfaro-Bou,

Emilio: Crash Tests of Three Identical

Low-Wing Single-Engine Airplanes.

NASA TP 2190,September1983.

12. Thomson,RobertG.; Carden,Huey

D.; and Hayduk, Robert J.: Survey of

NASA Research on Crash Dynamics.

NASA TP2298,April 1984.

13. Carden,Huey D.: Full-ScaleCrash

Test Evaluation of Two Load-Limiting

Subfloorsfor GeneralAviation Airframes.

NASA TP2380,December1984.

14. Hayduk,RobertJ.(Editor):Full-Scale

Transport Controlled Impact

Demonstration. NASA CP 2395, April

1985.

15. Fasanella,Edwin L.; Widmayer,E.;

and Robinson,M. P.: StructuralAnalysis

of theControlledImpactDemonstrationof

a JetTransportAirplane.AIAA Paper86-

0939-CP,May 1986.

16. Fasanella,Edwin L.; Alfaro-Bou,

Emilio; and Hayduk, Robert J.: Impact

Datafrom a TransportAircraft During a

Controlled Impact Demonstration. NASA

TP 2589, September 1986.

17. Farley, Gary L.: Energy Absorption

of Composite Materials. NASA TM

84638, AVRADCOM TR-83-B-2, 1983.

18. Bannerman, D.C.; and Kindervater,

C.M.: Crashworthiness Investigation of

Composite Aircraft Subfloor Beam

SectiOns. IB 435-84/3(1984), Deutsche

Forschungs-und Versuchasanstalt fur Luftund Raumfarht, February 1984.

19. Cronkhite, J.D.; Chung, Y.T.; and

Bark, L.W.: Crashworthy Composite

Structures. USAAVSCOM TR-87-D10,

U.S. Army, December 1987. (Available

from DTIC as AD B121 522.)

20 Jones, Lisa E.; and Carden, Huey D.:

Evaluation of Energy Absorption of New

Concepts of Aircraft Composite Subfloor

Intersections. NASA TP 2951, November

t989.

21. Vaughan, Victor L.; and Alfaro-Bou,

Emilio: Impact Dynamics Research Facility

for Full-Scale Aircraft Crash Testing.

NASA TN D-8179, April 1976.

22. Pifko, A.B.; Winter, R.; and Ogilvie,

P.L.: DYCAST A Finite Element

Program for the Crash Analysis of

Structures. NASA CR 4040, January

1987.

23. Williams, M. Susan; and Hayduk,

Robert J." Vertical Drop Test of a

Transport Fuselage Section Located

Forward of the Wing. NASA TM 85679,

August 1983.

24. Fasanella, Edwin, L.; and AlfaroBou, Emilio: Vertical Drop Test of a

Transport Fuselage Section Located Aft of

the Wing. NASA TM 89025, September

1986.

25. Boitnott, Richard L.; and

Kindervater, Cristof: Crashworthy Design

of Helicopter Composite Airframe

Structure. Fifteenth European Rotorcraft

Forum, Amsterdam, Holland, Paper Nr;

93, September 1989.

26. Boitnott, Richard L.; Fasanella,

Edwin L.; Calton, Lisa E.; and Carden,

ttuey D.: Impact Response of Composite

Fuselage Frames. SAE Paper 871009,

April 1987.

27. Boitnott, Richard L.; and Fasanella,

Edwin L.: Impact Evaluation of Composite

Floor Sections. SAE Paper 891018,

General Aviation Aircraft Meeting and

Exposition, Wichita, Kansas, April 1989.

10
28. Cronkhite, James D.; and Mazza,

L.T.: Bell ACAP Full-Scale Aircraft Crash

Test and KRASH Correlation. American

Helicopter Society, 44th Annual Forum

and Technology Display, Wash. D.C.,

June 1988.

29. Cronkhite, James D.; and Mazza,

L.T.: K_RASH Analysis Correlation with

the Bell ACAP Full-Scale Aircraft Crash

Test. American Helicopter Society National

Specialists' Meeting on Advanced

Rotorcraft Structures, Williamsburg, Va.,

October 1988.

30. Carper, Douglas M.; Hyer, Michael

W.; and Johnson, Eric R.: Large

Deformation Behavior of Long Shallow

Cylindrical Composite Panels. Virginia

Polytechnic Institute and State University,

VPI-E-83-37, September 1983.

ORIGINALS PAGE

BLACK AND WHITE PHO'rOGRAPvt

Figure 1.- Progression of research areas in crash dynamics at NASA Langley Research Center.

ORIGINAL PAGE IS

OF POOR QUALITY

11
ORIGINAl.:PAGE

BLACK A.I_. _HITK RHOEOGRAP_H

Figure 2. - Impact Dynamics Research Facility (IDRF) a NASA Langley Research Center.

Figure 3. - 70 Foot Vertical Drop Test Apparatus.

12
ORIGINA[ PAGE

BLACK AND WHIT£ PHOTOGRAPH

ORIGINAL PAGE IS

OF POOR QUALITY

_ Ir T _ _ _"

Figure 4. - Metal transport section suspended in Vertical Test Apparatus.

- 0.75 "--_

°;--11

C-SECTION

t

R:36"

±

ORIGINAl- PAGE IS

OF POOR QUALITY

0.75" 0.75"

,_C_ _L,

! !

I

I

""3.5" _k 3.5" "

I-SECTION J-SECTION

I

(a) Cross-sectional dimensions.

Figure 5. - Various cross-sectional shapes of composite fuselage frames.

13
ORIGINAL PAGE

BLACK AND WHIT£ PHOTOGRAPH

(b) Z-cross section fuselage frame.

Figure 5. - Concluded.

Backstop

Figure 6. - Composite Z-frame in drop apparatus.

14
ORIGINAE PAGE

BLACK AND WHITE PHOTOGRAPH

Figure 7. - Composite subfloor sections (skeleton/unskinned (top), skinned (bottom)).

ORIGINAL PAGE IS

OF POOR QUALITY

15
ORIGINA[ PAGE

BLACK AhID WHITE pHOTOGRAPH

(a) Sect/on frorn forward of wing2 _

OR!G;N_I. PAGE IS

OF POOR QUALITY

(b) Section from aft of wing.

Figure 8. - Structural damage to metal aircraft structures.

16
Z

90"_ - 1_ 90"

.51 • ..._52

0°

(c) Angular location of failures.

Figure 8. - Con,'luded.

17
UPANI-_ NONLINEAR SPRING

1 1 / /'-WALL-FLOOR

\ \ AT-/"- c-sECnON

_ouggco,__c,

Figure 9. - Finite element two-frame model of metal transport section.

Figure 10.-Computer graphics showing analytical failure of metal transport section.

18
ORIGIHAE PAGE

BLACK AND WHITE PHOTOGRAPH

o

Ple×igIasW _j_i

(a) Static test apparatus

Pl

pounds

Lood-deflection response

3000 -

Follure

2qO0

1800

1200

6OO

I

0 q

/

I

2

/

/

6 8 10

Pi

2

Deflection, in,

(b)Load-displacement and failure locations.

Figure 11. - Static results from tests of single composite Z-frame.

locQtlons

Pm

2

19
(c) Frame local instability.

Figure 11. - Concluded.

Vertical load,

Ibf

2800

Case I

2400

_g , W

I _ . CaseU

1200800lit / ] _"

r_ _ Case I unbroken frame/skin

400 II f

0 1 2 3 4 5 6 7 8 9 10

Displacement, inches

(a) Load-displacement.

Figure 12. - Typical analytical results for composite frame skin using I-section for ease of analysis.

20
Normalizedmoment

abouty-axis

1.0

0.5

0.0

-0.5

-1.0

-1.5 ' ' ' ' ' '

-100 -80 -60 -40

• I i I I , I a I t.. I •

-20 0 20 40 60 80

Angularlocafion, degrees

(b) Case I .... Normalized moment (Prior to failure).

Normalized moment

about y-axis

1.5

1.0

0.5

0.0 r

-0.5 J

• I • ! a I a I ,

-80 -60 -40 -20 0

• 1 , ! . ! • ! •

20 ,10 60 80

Angular location, degrees

(c) Case H---Normalized mement (Broken frame-intact skin).

100

100

Figure 12.- Concluded.

21
ORIGINAE PAGE

BLACK AND WHITE PHOTOGRAPH

's

r_

t_

8

'-a

22
0

0

0

!

0

!

E

I

{}

C:}

!

ORIGINAl: PAGE'

BLACK AND WHITE PHOTOGRAPH

9

¢)

r,j ,,

O

O

O')

I

¢:)

¢)

-=1

..-g

< {.aJE,

[

I

0

0

!

B

!

,4

23
iDRIGINAL PAGE

BLACK AND W_HIT£ PHOZOGRAPj-I

(a) Failed skeleton subfloor.

Frame A, B, and C

"90°-_ _ I--_--_ _'_-- 90 °

X_ 45° l\ 45_///

Splice plateJ _

(b) Failure locations,

Figure 15. - Behavior of skeleton composite subfloor under dynamic loading tests.

24
Normalized experimental

strain

1.6

1.2

0.8

0.4

0.0

-0.4

-0.8

-1.2

b,

-1.6

-100

I I I I | n

-80 -60 -40 -20 0

I • ! ! • I

20 40 60 80

Angular location, degrees

100

(c) Normalized strain distribution.

Figure 15.- Concluded.

25
O C}

ORIGINAL PAGE

BLACK A_D WHITE PHO_OGRAP.ti

o

0

8

o

!

1

0

0

i

26
Normalized experimental

strain

1.6

1.2

0.8

0.4

0.0

-0.4

-0.8

-1.2

-1.6 n I I I * I n

-100 -80 -60 -40 -20 0

I I I n i u

20 40 60 80 I(30

Angular locatioJl, degrees

(c) Normalized str',fin distribution.

Figure 16. - Concluded.

27
Verticalload,

lbf

1000

8OO

600

400

ection

_ _ Sk'mnedsection

200

0 ! I

O.0 0.2 0.4 0.6 0.8

Displacement, inches

Figure 17. - Comparison of experimental and analytical stiffness of skeleton and skinned composite subfloors.

28
!

_d

1.0

0.5

0.0

-0.5

-1.0

/

!

Observed fail ure

regions -

-1.5

- 100 -80 -60 -40 -20 0 20 40 60 80

Angular location, degrees

(a) Normalized moment distribution.

1(_

_62°

-45 ° 38 °

-12 ° 12

Figure 18.-

(b) Failure location,,,.

Comparison of nondimensional analytical moment distribution predicted by finite element frame

model and observed failure locations on metal and composite fuselage components and structures.

29
Report Documentation Page

_1_ ¢_rOr_OtlCS a_-_3

SC_e AOc_r.SttatO_

1. Report No.

NASA "EV[- 102679

2. Government Accession No.

4. Title and Subtitle

Unique Failure Behavior of Metal/Composite

Aircraft Components Under Crash Type Loads

7, Author(s)

Huey D. Carden

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, VA 23665-5225

12, Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546-0001

3. Recipient's Catalog No.

5. Report Date

May 1990

6 Performing Organization Code

8. Performing O"rganization Report No.

10. Work Unit No.

505-63-01- l I

11, Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14 Sponsoring ,_gency Code

15. Supplementary Notes

16. Abstract

Unique failure behavior results are presented on some of the crash dynamics research

conducted with concepts of aircraft elements and substructure which have not necessarily been designed

or optimized for energy absorption or crash loading considerations. To achieve desired new designs

which incorporate improved energy absorption capabilities often requires an understanding of how

more conventional designs behave under crash type loadings. Experimental and analytical data are

presented which indicate some general trends in the failure behavior of a class of composite structures

which include individual fuselage frames, skeleton subfloors with stringers and floor beams but

without skin covering, and subfloors with skin added to the frame-stringer arrangement. Although the

behavior is complex, a strong similarity in the static/dynamic failure behavior anaong these structures is

illustrated through photographs of the experimental results and through analytical data of generic

composite structural models. It is believed that the thread of similarity in behavior is telling the designer

and dynamists a great deal about what to expect i_t_sbrash behavior of these structures and can guide

designs for improving the energy absorption and eras- h behavior of such structures.

17. Key Words (Suggested by Author(s))

Aircraft crash dynamics

Nonlinear structural dynamics

Impact dynamics

Composite materials

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Unclassified- Unlimited

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