NASA Technical Memorandum 102681

USAAVSCOM Technical Report 90-B-003

(NASA-TM-I02681) _EflAVIOR nF

COMPOSITE/METAL AIKCRAFT STRUCTURAL ELEMENTS

AND COMPGNENTR UNDER CRASH TYPE LOADS: WHAT

ARE THLY TELLING US (NASA) Ib p CSCI 20K

N90-25368

Unclas

G_/39 0291540

Behavior of Composite/Metal

Aircraft Structural Elements and

Components Under Crash Type

Loads--What Are They Telling Us?

Huey D. Carden, Richard L. Boitnott, and

Edwin L. Fasanella

May 1990

N/LRA

Nalional Aeronaulics and

Space Administralion

Langley Research Cenler

Hamplon, Virginia 23665-5225

US ARMY_

AVIATION

SYSTEMS COMMAND

AVIATION R&T ACTIVITY
m
BEIIAVIOR OF COMPOSITFdMETAL AIRCRAFI" STRUCFIJRAL ELEMENTS AND

COMPONENTS UNDER CRASIt TYPE I.OADS--WIIAT ARE TtlEY TEl.LING US?

lluey D. Carden*

NASA Langley Research Center

tlampton, VA

Richard I.. Boitnott**

U'.S. Amly Aerostmctures Directorate, AVSCOM

I.angley Research Center

llampton, VA

Edwin 1.. Fasanclla_

l._:khecd F.ngineering and Sciences Company

1lampton, VA

,5_.b.gr_Ec_t

Failure behavior results are presented from crash

dynamics research using concepts of aircraft elements and

substructure not necessarily 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 Ioadings.

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 similarity in behavior is giving the designer and dynamists

much information about what to expect in the crash behavior

of these structures and can guide designs for improving the

energy absorption and crash behavior of stlch structtlres.

lntroduqliorl

The NASA l.angley Research Center has been

involved in crash dynamics research since the early 1970's.

For abot,t the first I(] years the emphasis of the research was

on metal aircraft structures during the General Aviation Crash

Dynamics Program I - 13 and a transport aircraft program,

the Controlled Impact Demonstration (CID), which culminated

with the remotely piloted crash test of a B-720 aircraft 14 - 16

in 1984. 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 loadings.

(See Figure 1._. Considerable work has been conducted into

determining the energy absorption characteristics 17 - 20 of

composites. These results indicated that composites can

absorb as mttch if not considerably more energy than

comparable aluminum structures, llowever, because of the

brittle nature of the materials, 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

+ AsSisii_,ii_afi, La,liiing-a-nd--l-_pactDynamics Branch

**Research Engineer

1" Staff Engineer

desired structural integrity during normal flight loading

conditions. To achieve the desired new designs often requires

an understanding of how more conventional designs behave

under crash Ioadings.

The purpose of this paper is to present an overview of

the research conducted using 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 indicates some general trends in the failure behavior of

a class of composite structures which include individual

fi,sclage frames, skeleton subfloors with stringers and floor

beams but without skin covering, and subfloors with skin

added to the frame-stringer arrangemertt. 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 stn_ctural models. It is believed that

the similarity in behavior is giving the designer and dynamists

much infomlation 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.

impact Dynamics Research Facility

The informatirn presented in this report is the result of

a research program to inve._tigate the impact response of

transporl and composite aircraft components conducted at the

NASA l.angley Research Center's Impact Dynamics Research

Facility (IDRF). The IDRF (shown in figure 2) is the former

Lunar Landing Facility used to train astronauts for moon

landings. The facility is 220 feet high and 400 feet long. In

the early 1970's, the structure was converted for crash testing

of fnll-scale general aviation aircraft. Reference 21 provides

complete details of the facility and test techniques for full-scale

aircraft testing. Additionally, a photograph of a 70 foot high

Vertical Drop Test Apparatus often used for full-scale aircraft

section, components, and/or seat testing is shown in Figure 3.

Static testing machines, and other apparatus are also available

at the facility for metal and composite aircraft structural

testing.

Analysis Tools

To gain an understanding of fundamental physical

behavior of complex stn_ctures, the experimental research with

structures under crash Ioadings is generally accompanied by

analytical prediction/correlation studies whenever feasible.

Thus, various finite element ccxles which have capabilities for

handling dynamic, large displacement, nonlinear response

problems of metal and composite structures are used as tools

in the various research efforts.
|)YCAS;F Computer Code

The analytical results presented in this overview were

generated with a nonlinear finite element computer code called

DYCAST (DYnamic Crash Analysis of STructures) 22

developed by Grnmman Aerospace Corporation with principal

support fiom NASA and FAA. The basic element library

consists of (1) stringers with axial stiffness only; (2) beam

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

stiffi_esses;(4) isotropic plate bending triangles with membrane

and out-of-plane bending stiffnesses; and (5) nonlinear

tnnslational 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 mc_|el crush

behavior of components for which experimental or analytical

data are available and/or whose behavior may he leO complex

or time consuming to model otherwise. An effort is underway

to add curved composite beams, composite plate, and curved

shell elements to the DYCAST element library.

_Speyimens ond Description

_Full:_S_- tale Melat Aircraft StnJctures

NASA Langley Research Center conducted three

vertical drop tests of metal aircraft sections to support

transport aircraft research efforts. Selected data on the crash

hchavior of fltll-scale metal transport aircraft sections 23 - 24

are included in the present paper to demonstrate what appears

to be important similarities m behavior noted for both the metal

fuselage structures and the composite structures discussed

herein.

Two 12-foot long fitselage sections cut from an out-ofservice Boeing 707 transport aircraft were drop tested to

measure structural, seat and occupant responses to vertical

crash loads, and to provide data for nonlinear finite element

modeling. One fuselage section was cut from forward and

one was cut from aft of the wing location of the aircraft. A

photograph of the forward section suspended in the Vertical

Drop Test Apparatus at the Impact Dynamics Research Facility

is shown in Figure 4. The aft section tests served two

purposes: 1) to test structural, seat, and anthropomorphic

dummies response, and 2) to lest the data acquisition system

pallet, power pallet, and camera batteries and instrumentation

later used in the full-scale transport crashprogram. The reader

should refer to the particular reports 14- I oand 23-24 for more

complete descriptions of the test articles since such

information is not repeated in this report.

C_'Q!!)pos_i!e Structur_es

i_,qi_gle Composite Frames.- Various cross-sectional

shapes for fuselage frames are used in metal aircraft and are

often proposed for composite structures. Sketches and

photographs illustrating four of the more common geometries,

I , J-, C- and Z-cross sectional shapes are shown in Figure 5

(a), (h) and (c). Several circular frames with these shapes

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), 2 1/4 or 3 1/2

inch wide skin material was added which enhanced the ease of

testing of both symmetrical and nonsymmetrical crosssections. The .08 inch thick, sixteen ply skin, with a

[-+45/0/9012s lay-up was co-cured with the 6 foot diameter

frames which have the lay_ups as indicated in Table I. The

frame cross-sections 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.

Table I.- Cc_mposite frame section lay-ups.

Fuselage Serial Configuration Lay-up Weight (Kg)

Frame Number

[_al',el

FR(_).ll 1 l-section (_+45/0/90) s 1.443

FRI'_'pSI I l-section (+_-45/0/90)2s 1.996

FR00IJ 2 J-seclion (_+4510/90)2s I.g53

FR(X)5C I C-section (+_.45/0/90)2s 1.229

FR005C 2 C-section (__.45/0/90)2s 1.229

One of the first geometries to be studied under static

and dynamic loadings was the Z-cross section. A photograph

of Z cross section fuselage frames uged in the initial studies of

the behavior of composite structural elements under impact

loads is shown in Figure 5(c). A Z-frame suspended in the

drop apparatus prior to testing is shown in Figure 6. 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 some (but not all) out-old-plane bending or twisting

during impact, and allowed photographs and motion picture

coverage through the front plexiglas plate . The six-foot

diameter frames were constructed using a quasi-isotropic layup of 280-5HA/3502, a five harness satin weave graphite

fabric composite material. The Z-cross section of the frame

was 3 inches high with a total width of 2.25 inches and about

0.08 inches thick. Initial tests were with 3(_0 degree frames

made from four 90 degree segments joined with splice plates

as shown in figure 5(c). 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, three composite subfloor structures were

fabricated following the initial investigation of the Z-frames

discussed above.

_,!bfloor Structures.- 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 is shown in Figure 7. Pultruded J-stringers connected

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 structure.

Full:_iLl¢ Composite Aircraft.- Two full-scale

composite general aviation aircraft structures, two complete

wing sets, and landing gears have been obtained for testing.

Because of the scarcity and expense associated with obtaining

full-scale composite aircraft for impact testing, plans include

multiple usage of the structures. Various structural, and impact

tests will provide additional information for the composites

data base.

2
Results and Discussion

Experimental and analytica.I results from the studies of

full-scale aircraft structures, composite fuselage frames, and

subfloors under static and/or dynamic loadings are presented

in Figures 8 to 18. 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 stn,ctures for

improving the energy absorption and crash behavior of these

lype ct_mponents and structural elements.

Experimental and analytical results from studies with

full-scale transport category aircraft sections 23 - 24 are

prescnted in figure 8.

L)v__n_a_n_.- Structural damage of the transport

_fircraft structures resulting for the 20 ft/s drop tests is shown

in figure 8(a) and (b). The damage to the transport 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 about 50

degrees up the circumference from the bottom contact point

(See figure 8(c)). The crush of the lower fuselage was

_,pproximately 22-23 inches at the front end and 18-19 inches

_,t the rear for the section taken from forward of the wing

location (figure 8(a)) whereas for the aft section (figure 8(b))

the crushing was about 14 inches in front and 18 inches in the

rear. Although the aircraft stn_ctures 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

resvlts of the composite frames and subfloors discussed later.

Analylica0LI _ji¢_.- A DYCAST model of the section

(from forward of the wing location) was constructed to model

the floor, two seats with lumped mass occupants, and the

fuselage structure to detcmfine if such a model could predict

the response of the complete section with fidelity. The finite

clement model is shown in figure 9. Stiff ground springs

simulated the concrete impact surface. Each frame of the

fuselage below the floor was modeled with eight beam

elements and floor and seat rails were also appropriate beam

elements. Fusela_,e structure above the floor (not expected to

fail) was modeled in less detail.

A comparison of the two frame analytical predictions

and the full section experimental responses are shown in

figure 1(}. The correlation of the vertical displa('en_eul,

wall/floor, and dummy pelvis accclerali_ms are considered

good. As may be noted in figure t0(d), the overall

impression from the analytical model deformation pattern is

quite similar to the visual behavior seen in the experiment

shown in figure 8. Thus, the full section behavior was

basically contained in the two frame model.

In the following sections the composite impact

dynamics studies have taken the building block approach of

utilizing a sequence of testing and analysis which begins with

'simpler' elements and move to more 'complicated'

components or substructures. As mention earlier, this

approach was used in the General Aviation (GA) and

Transport programs although the GA data base was being

concurrently developed through full-scale testing. Full-scale

tests using currently available composite aircraft specimens

and/or other full-scale structures are part of the on-going

research program.

Other than two support tests for Army composite

aircraft programs 25-26, no testing of full-scale composite

aircraft has been conducted at the Langley Research Center as

part of the composite impact dynamics research. However, as

indicated previously, composite aircraft fuselage specimens

have been obtained for crash testing.

_C,_Qmpositei_Si_Je Fr0m¢ Studies

Sty.- Results from the static test of a semicircular frame with a Z cross-secti,m 27 are presented in Figure

I I. A photograph of the static test apparatus in figure 1 I(a)

shows thal the splice plate was at the load point.

Consequently, the frame failed just outside the doubler splice

plate area by a cgmplete fracture across the Z-section. Loaddeflection data and the Iocatinn of failures of the frame are

shown in figure I l(b). The load-deflection data show a sawtoothed behavior of loading and unloading. The load

increased linearly until initial failure, then falls off to under

600 pounds force. Subsequent loading of the frame after

initial failure was at a new, reduced stiffness. Second and

third fractures occurred tip 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). Photographic data in figure lO(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.

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), two DYCAST finite-element models were

constructed. For ease of analysis, a typical 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. Only

half the frame was modeled using thirty-four (34) l-section

beam elements with boundary conditions imposed at the

bottom node of the model to account for the symmetrical

situation. 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, case I, in figure i 2(a) is

the load-deflection plot for this case I.

A second DYCAST model, broken frame, case H,

was also run wherein the bottom point of the frame was

modeled with two short skin segments to represent the

different I_'_undary condition following the initi:d failure of the

frame. This condition will be discussed further relative to the

composite subfloor test with the skinned subfloor specimen.

The curve in figure 12(a) labeled cc_,e H is the load-deflection

response for this frame loading case. Essentially, the frame

load increases along the case I curve 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, case II, which represents the

stiffness of the section with the weakened boundary on the

bottom end of the frame. The load continues to increase

along the case II curve until secondary failures at some other

location on the frame circumference occur.

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 failurefloehavior.

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 locations in the experiment with the Z-frame.
Normalizedbendingmomentdistributiononthe

brokenframeispresentedinFigure12(c).Asmaybenoted, thedistributionisquitesimilarto the initial model resuhs in

figure 12(b). The failure location is at the maximuin 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

prc(lictcd behavior by 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 (Lwice the initial diameter). The distribution

was identical to the smaller diameter results, however, the

loads producing the moments differed between the two n'_els

(as expected). Moment in terms of geometry and load for a

point loaded ring 2R shows that the maximum moment will

occur at the same angle for different vah,es of diameter.

Furtherrnore, a comparison of tile analytical cases

(figure 12(a)) with the actua/static load-deflection of the Zsection (figure If(b)) indicates very similar load-deflection

behavior patterns as discussed above. Although the Z-frame

had no skin, if tile ends jam together (as they did in several

cases), the boundary is effectively between the skin stiffened

ca_e and a guided _undary. Thus, the predicted maximum

moment 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

ira me.

Without a priori knowledge of the manner of the

failure noted and discussed above, the initial formulation of a

finite-element rnodel would unlikely incorporate the necessary

failure mechanism/behavior for the frames, ltowever,

knowing the pattern of behavior can enable the analyst to

formulate adequate finite:element models to predict dynamic

responses including the failure/loads. Additionally, such

infom_ation is important to designers of new stn_ctures to be

able to design for impact loads on such structural elements of

an aircraft fl|selage.

D_,'_s.!s_- Resnhs from reference 29 on the

dynamic studies of the response of composite frames are

shown in Figure 13. Photographs of the failed Z-frames in

the drop apparatus are shown in Figure 13(a). It should be

note(I, 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 faih|res (fractures) of the Zsection frames occurred at the bottom and approximately 60

degrees from the bottom. It appears that the presence of the

splice plates _nloved the top failure points up a few degrees (to

about the 60 degree locations).

__rnics_.__n_a_li_- Experimental and

an:|lytical m(xtel resuhs fi_r the composite Z-frame subjected to

a 2(1 ft/s impact are shown in Figure 13(c). The first 15

milliseconds of the experimental acceleration of the floor and

analytical predictions are shown. In one analysis, in-plane

deformations were constrained to tile plane of the frame

whereas in tile other, the frame was free to twist and bend outof plane. As noted in the figure, agreement between the "free"

model and the experiment is good for the initial peak load.

l,ater the agreement is less because the backstop and clear

fence in the experiment, which was not modeled, began to

provide support to the twisting and bending frame. The inplane model results are about 60% higher when compared to

the experiment and the free model resuhs. Such resuhs

emphasize the importance of knowing and modeling the

experimental boundary conditions to understand the dynamic

and static behavior of structures under crash loadings.

Unfortunately, mixed boundary conditions often exist or occur

in the experiments.

_nTposJte Suflb_O___r_Smdi_

Two static and two dynamic tests were conducted on

the three composite subfloor specimens used for impact

studies. Two skeleton subfloor were tested to destruction,

one statically and the other dynamically. The skinned subfloor

was subjected to a non-destructive static test followed by a

dynamic lesl to failure.

SI._.- Experimental results 30 of the skeleton

subfloor specimen following a static test are shown in Figure

14. As noted in figure 14(a),and (b), failures on the three Zsection frames occurred at 13 discrete locations. Unlike the

unnotched single Z-frame, the faiht,res in this specimen

occurred at notches (which served as stress risers) in the frame

through which the stringer passed. Ilowever, as shown in

figure 14(b) the failures were still near the point of load

application (approximately I1 degrees) and at other

circumferential locations of approximately 45 to 67 degrees.

In the absence of skin material, twisting and bending out-ofplane occurred with the frames. The stringers had only a

minimal effect on the subfloor response and that was to

maintain the lateral spacing of the three Z-frames.

__D_namic tests.- A photograph of a skeleton subfloor

after an impact test onto a concrete surface at 20 feet per

second is shown in Figure 15. In the dynamic test of the

skeleton subfloor, fractures were produced at notches in the

(rames (Figure 15(a)). The iocations, shown in figure 15(b),

were also near the point of impact (about I 1 degrees because

of the splice plate) and at three other locations up the

circumference of the frames (55 degrees and 78 degrees)

totalling I5 fractures for all three frames. The impact energy

exceeded the energy absorbed by the local fractures and the

floor bottomed out during the impact. The normalized

circumferential strain distribution measured on the flange of

the front frame dufing.th.e dynamic test just before first failure

is shown in Figure 15(c). A comparison of the measured

distribt|fion to the predicted moment distribution of Figure

12(b) and (c) shows essentially identical shape between the

single frame and skeleton frame distributions. Maximum

values at 0 degrees and at approximately 50 to 55 degrees for

the experiment agree well with the analytically predicted

locations on the frame.

Impact results for the subfioor with skin after an

impact of 20 feet per second are shown in Figure 16. A

photograph of the subfloor specimen after the test is shown in

Figure lr(a). Points of failure of the frames in this specimen

are indicated in figure 16(b). Again the points of failure are at

or near the impact point (within 12 degrees) and

circumferentially at about 56 degrees tip 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 subfl_r

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. Normalized strain distribution

(just prior to first failure) measured on the skin at the location

of the front frame during the dynamic test showed 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

for the experiment agree well with the analytically predicted

locations on the frame.

As mentioned previously in the frame studies, once the

frames fail at/ne,'u 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, l,ittle energy is involved in

4
skin snap-through as the load increases on the structure (,See

reference 31 on snap-through 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 herein simulated this type of

behavior.

3_o_,l_t!dies.-The contribution of the skin to the

stiffness of the section with the nonsymmetrical frames is

illustrated in figure 17. Static load-deflection data for the

unskinned subfloor and the skinned subfloor along with the

DYCAST predictions are shown in the figure. It can be noted

that the subfloor stiffness (with skin) is approximately three

times the stiffness of the skeleton subfloor, thus the skin's

contribution to the structure is to maintain in-plane deflections

of the non-symmetrical Z-section and prevent any substantial

twisting of the frames. Out-of-plane bending and twist were

allowed in the skeleton subfloor predictions.

The response behavior determined during the studies

of full-scale aircraft sections, filselage frames, and subfloors

are summarized in figure 18. Normalized moment distribution

on a representative frame of tile various specimens are shown

in Figure 18(a). Figure 18(b) shows the failure locations

which were noted from static and dynamic tests. The visual

impression is quite striking among the various specin'tens.

The structures all 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 maximnms 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 may help dynamlsts

gain a clearer understanding of what to expect from such

structt,res in crash loading situations, may guide designers of

new structures to account for the vertical crash loads, and

allow increased energy absorption to be included in the new

designs. Additionally, the observations may help analysts to

model the aircraft structures more adequately 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

current codes 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 in-house efforts.

Concluding Re_

Some important failure behavior results from the

research with composite full-scale aircraft sections, composite

slructural elements, and subfloors have been presented.

Observations of the failure behavior of these stta_ctures have

been made and discussed and analytical results have been

included to help explain some of the behavior noted.

From the research presented in the overview the

following conclusions are made:

(I) Comparing the results from simple representative structural

elements with more complex components provided insight

into the the local/global structural responses and behavior of

complex aircraft structures.

(2) Relatively simple analytical models provided generally

good correlation with experiments; however, guidance from

experimental data was required to allow adequate or improved

analytical models to be fommlated.

(2) ('Olllnlonaiity in the t'ailure behavior patterns among rt,llscale aircraft sections, composite frames, and subfloors with

and v,ithout skin was found.

(3) General locations of failureg appear to occur at the same

structural regions among the specimens as a result of similar

geometry, similar loading, and similar moment distribution on

the strt,ctures under vertical loads.

(4) Noted faih,res were located in the same regions as the

maximums in the moment distribution on the structures.

(5) The shape of tile distribution of the moment was

independent of the diameter of the frame/component. Loads,

however, which produced the failures varied with the

structural size.

.C.on_Qbservations

Based upon the conclusions drawn from the various

research efforts discussed in this paper, the following

observations are also summarized:

(1) The general similarity of the failure behavior among the

aircraft stn,ctures can:

(a) assist the designer and dynamists to anticipate how

the structures probably will fail,

(b) provide guidance on how and where to incorporate

and/or optimize improved energy absorption into new aircraft

structural designs, and

(c) aid analysts to more adequately model the

structnres for predicting faih_re/loads behavior under crash

situations.

(2) To analytically predict, in a dynamic loading situation, the

complex failure events and the loads which initiate the failures

as noted in the composite structural elements and subcomponents is a challenge, but possible; however, guidance is

often required from experiments.

(3) Composite curved beam, composite plate and shell

elements are being developed and included in finite-element

codes to improve the capability to analyze composite

structures.

Rci.e,rrm.¢_

1. Alfaro-Bou, Emilio; and Vanghan, Victor L., Jr.:

Light Airplane Crash Tests at Impact Velocities of 13 and 27

nVsec. NASA TP 1042, November 1977.

2. Castle, Claude B.; and Alfaro-Bou, Emilio: Light

Airplane Crash Tests at Three FlightPath Angles. NASA TP

1210, June 1978.

3. itayduk, Robert J.: Comparative Analysis of PA-31-

35(1 Chieftain (N44LV) Accident and NASA Crash Test Data.

NASA TM 80102, October 1979.

4.. Castle, Claude B.; and Alfaro-Bou, Emilio: Light

Airplane Crash Tests at Three Roll Angles. NASA TP 1477,

October 1979.

5. Vaughan, Victor I,., Jr.; and Alfaro-Bou, Emilio:

l,ight Airplane Crash Tests at Three Pitch Angles. NASA TP

148 l, November 1979.

6. Vat,ghan, Victor I,., Jr.; and ttayduk, Robert J.:

Crash Tests of Four Identical tligh-Wing Single-Englne

Airplanes. NASA TP 1699, August 1980.

5
7. Carden, lluey D.; and Hayduk, Robert J.: Aircraft

Subfloor Response to Crash Loadings. SAE Paper 810614,

April 1981.

8. Williams, M. Susan; and Fasanella, Edwin L.: Crash

Tests of Four Low- Wing Twin-Engine Airplanes with TrussReinforced Fuselage Structure. NASA TP 2070, September

1982.

9. Carden, lluey D.: Correlation and Assessment of

Structural Airplane Crash Data with Flight Parameters at

Impact. NASA TP 2083, November 1982.

10. Carden, IIuey D.: Impulse Analysis of Airplane Crash

Data with Consideration Given to Human Tolerance. SAE

Paper 830748, April 1983.

I I. Castle, Claude B.; and A[faro-Bou, Emilio: Crash

Tests of Three Identical Low-Wing Single-Engine Airplanes.

NASA TP 2190, September 1983.

12. Thomson, Robert G.; Carden, ttuey D.; and llaydnk,

Robert J.: Survey of NASA Research on Crash Dynamics.

NASA TP 2298, April 1984.

13. Carden, Iluey D.: Full-Scale Crash Test Evaluation of

Two Load-Limiting Subfloors for General Aviation

Airframes. NASA TP 2380, December 1984.

14. ttayduk, Robert J.(Editor): Full-Scale Transport

Controlled Impact Demonstration. NASA CP 2395, April

1985.

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

M. P.: Structural Analysis of the Controlled Impact

Demonstration of a Jet Transport Airplane. AIAA Paper 86-

0939-CP, May 1986.

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

Hayduk, Robert J,: Impact Data from a Transport Aircraft

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.

24. Fasanella, Edwin, L.; and Alfaro-Bou, Emilio: Vertical

Drop Test of a Transport Fuselage Section Located Aft of the

Wing. NASA TM 89025, September 1986.

25. Cronkhite, James D.; and Mazza, L.T.: Bell ACAP

Full-Scale Aircraft Crash Test and KRASH Correlation.

American ttelicopter Society, 44th Annual Forum and

Technology Display, Wash: D.C., June 1988.

26. Cronkhite, James D.; and Mazza, L.T.: KRASIt

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.

27. Boitnott, Richard L.; and Kindervater, Cristof:

Crashworthy Design of tlelicopter Composite Airframe

Structure. Fifteenth European Rotorcraft Forum, Amsterdam,

Holland, Paper Nr; 93, September 1989.

28. Van Den Brock, J.A.: Elastic Energy Theory, 2nd

Edition, John Wiley and Sons,Inc., New York, August 1946,

pp. 100-104.

29. Boitnott, Richard L.; Fasanella, Edwin L.; Cahon,

Lisa E.; and Carden, Huey D.: Impact Response of Composite

Fuselage Frames. SAE Paper 871009, April 1987.

30. Boitnott, Rich,'u-d L.; and Fasanella, Edwin L.: Impact

Evaluation of Composite Floor Sections. SAE Paper 891018,

General Aviation Aircraft Meeting and Exposition, Wichita,

Kansas, April 1989.

31. Carper, Douglas M.; Hyer, Michael W.; and Johnson,

Eric R.: Large Deformation Behavior of Long Shallow

Cylindrical Composite Panels. Virginia Polytechnic Institute

and State 1983.

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

Crashworthiness Investigation of Composite Aircraft Subfloor

Ream Sections. IB 435-84/3(1984), Deutsche Forschungsund Versttchasanstalt fur Luft-und Raumfarht, February 1984.

19. Cronkhite, J.D.; Chung, Y.T.; and Bark, L.W.:

Crashworthy Composite Structures. ASAAVSCOM TR-87-

Di0, U.S. Army, December 1987. (Available from DTIC as

AD BI21 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 1989.

2 I. 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 Finiie Element Program for the Crash Analysis of

Structures. NASA CR 4040, January 1987.

23. Williams, M. Susan; and ltayduk, Robert J.: Vertical

Drop Test of a Transport Fuselage Section Located Forward

of the Wing. NASA TM 85679, August 1983.

Figure 1.- Aircraft impact dynam!cs

been where we're

Figure 2. - Impact Dynamics Research Facility (IDRF) at

NASA Langle.y Research Center.

6

ORIGINAE PAGE'

BLACK AND WHITE PHOTOGRAPH
Figure 3. - Vertical Drop Test Apparatos used in testing of

aircraft sections and components.

Figure 4 - Metal transport section suspended in Vertical Test

Apparatus.

"T--- _F .TS"-_ ._._" t

1.5" and .75 I]

1 P-_._"-_ /

r ----2.25"-- _ :_ P. 1.16"

_c-seclion Z-_ection

_-- _.s"--_ p-.7s"-_

3.5" , 3.5"

I-Section J-Section

(a) Cross-sectional dimensions.

(b) I and C section frames.

8LACK ORI_NAE PAeE

AND WHITE PHOTOGRAPH

7

(c) Z-cross section fuselage frame.

Figure 5. - Various cross-sectional shapes of composite

fl]selage frames.
Figure 6. - Composite Z-frame in drop apparatus.

I:igtlre 7. Composite subgloor sections (skeleton/unskinned

(top), skinned (bottom)).

ORT(3iNAE PAGE

BLACK A,_D WHITE PHOTOGF_APH

(a) Section from forward of wing.

(b) Section from aft of wing.

0"

(c) Angular location of failures.

Figure g. - Structural damage I,3 metal aiwraft stn_ctures

resulting from impact lesls.
Groundconladsprlncls I

I:ignre 9, - Finite element tWo-frame model of metal

transport section.

intersedion

Displacemenl,

inches

28

24

2O

16

12

Experimental

-//_ ".. -_Analytical

,/

I t _1 I I I t I _

.04,0g .12 .16 .20 .24 .28 .32 .36

Time.sec.

(a) Vertical floor displacement.

Accel. G

15

10

5

0

-5

0

/,,,(- Analytical

•"_"x:, ,' 'J/,, \-E._lr_._l

_a__a_ _ J- t l _ a___ l ..... l___

•04 .08 .12 .16 .20 .24 .28 .32

Time,sec,

(c) Vertical pelvis acceleration.

(d) Deformed model.

Figure 10. - Comparison of experimental and analytical

fuselage crushing and accelerations of metal

transport section.

J

.36

• 1s

|o

5

Acce],G

0

5

|0

,

V ',-Experimental

.04 .08 .12 .16 .20 .24 .28 .32 .36

Time,sec.

b) Vertical accelerations at wall/floor interface.

ORIGINAL: PAGE

BLACK AND WHITE PHOTOGRAPH

9

(a) Static test apparatus.

Figure 11. - Static results from tests of single composite Zframe.
Load P,

Lbf Load-Deflection Response Failure Locations

3000

2400

1800]

12001

6001 I_I _131

0 2 4 6 8 i0

Deflection, in.

P

2

P

2

(b) Load-displacement and failure locations.

Figure I I. - Continued.

(c) Frame local instability.

Figure I I. - Concluded.

ORtO_NAEPAO£

BLACK AND WHITE PHOTOGRAPH

28OOr

[ Case I

20O0

,60o i!

1200fi800i i _/P i

400 [_ ---0-- Case I unbroken frmne/skin .

0 1 2 3 4 5 6 7 8 9

Displacement, inches

(a) Load-displacement.

Figure 12. - Typical analytical results for composite frame

skin using l-_ction for ease of analysis.

10

|

10
Normalized moment

al',.(nit y-a×is

1.0

o.5 f =

0.0 /

-0.5 }r

- 1.0

-1.5 . , . _ . . t • ,

-I(X) -80 -60 -40 -20 0 20 40 60 80 I00

Angular location, degrees

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

Normalized moment

about y-axis

].5

1.0

°,

-0,5 L

|

-I.0 ' '

-tO()-80 -60

• , I • • l • I I I

-40 -20 0 20 40 60 80

Angular location,degrees

00

(c) Case ll---Normalized moment (Broken frame-intact skin).

Figure 12. - Concluded.

i

20 fps

100 ibm. _V_

Initial im,

X_ 68°._

38°

(a) Photographs of failed frames.

20 fps

93 Ibm.

)act point Initial im

(b)Location of failures in frames.

Figure 13. - Behavior of composite Z-frame under impact loading.

_act point

ORIGINAl.: PAGE

BLACK AND WHITE PHOTOGRAPH

11
Accel, G

25'0 r D/.YCAST (In-Plane)

| /_" 20 fps Frame A (Front)

°0o ,0o

5,0 [ _-$plice plate

[ I I I O_

0 5,0 10,0 15,0 Frame B (Center)

Time, HSEC

(c) Experimental and analytical accelerations.

Figure 13. - Concluded.

ORICINAE PAGE

BLACK AND WHITE PHOTOGRAPH

-90°--[----_ _

_'78"30 _/

\X/' /

0 o

Frame C (Rear)

\X/-""/ ,o,

0

(a) Failed composite skeleton subfloor. (b) Failure locations.

Figure 14. - Behavior of composite skeleton subfloor under static loading tests.

12
(a) Failed composite skcleton subfloor:

Frame A, B, and C

Splice plate

0<'

(b) Failure locations.

90 °

Nonnalizexl experimental

strain

!.6

1.2

0.8

(1.4

0.0

-0.4

-i.2

-I.6 i

-60

I • • ! I •

-t00 -80 -40 -20 0 20 40

Angular location, degrees

I . l • I • I

60 8(1

(c) Nomlalized circumferential strain distribution.

Figure 15. - Behavior of composite skeleton subfloor under

dynamic loading tests.

I00

BLACK

ORIGINALS PAGE'

AND WHITE PHOTOGRAPH

13

(a) Failed composite skinned subfloor.

Frante A (Front)

00o 1 fT°°°

k,,.,_//

Splice plate--"

0 _,

• Frame B (Center)

"_ -- 90"

0 o

Frame C (Rear)

-90°-_ /'7 i'_-56 3" -

(b) Location of failures.

Figure 16. - Behavior of composite skinned subfloor under

dynamic loading tests.

90 _
Nonualized experimental

strain

1.6

!.2

0.g

0.4

olo

-0.4

-0._

-!.2

-1.6 ! • | • | • ! I • | • I • !

-!00 -80 -60 -40 -20 0 20 40 60 80 100

Angular location, degrees

(e) Nomlalized circumferential strain distribution.

Figure 16. - Concluded.

Vertical load,

Ibf

I(X)O

Nommli zed moment

about y-axis

1.0 i

' r-Observed Failure

ions ,..._

0,0

0.5

- 1.0 I _"

-t.5 ' " ' " ' " ' " " ' ' " ' " ' "

-t00 -gO -60 -40 -20 0 20 40 (10 80 100

Angular location, degrees

(a) Normalized moment distribution.

8(X)

6(X)

4(_)

2(_)

0

0.0

ion

on

i ": ........ ! , • I •

0.2 0.4 0.6

Displacement, inches

Figure 17. - Comparison of experimental and analytical

stiffness of composite skeleton and skinned

subfloors.

0.8

._j.-'J ':

-62 _ 68". J

i_. o '"'_

(b) Failure locations.

Figure 18. - Comparimn of non-dimensional analytical

moment distribution predicted by finite element

frame model and observed failure locations on

metal and composite fumlage components and

strllctures.

14
1. Report No.

NASA TM-102681

USAAVSCOM TR-90-B-003

4. Title and Subtitle

Report Documentation Page

i 2. Government Accession No.

I

I

Behavior of Composite/Metal Aircraft Structural

Elements and Components Under Crash Type Loads--

What Are They Telling Us ?

3, Recipient's Catalog No.

5. Report Date

7, Author(sl

Huey D. Carden, Richard L. Boitnott, and

Edwin L. Fasanella

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, VA 23665-5225 and

Aero6tructures Directorate, USAARTA-AVSCOM

Langley Research Center, Hampton, VA 23665-5225

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546-0001 and

U.S. Army Aviation Systems Command

St. Louis, MO 63120-1798

15.

May 1990

6. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

505-63-01-11

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring ,_gency Code

Supplementary Notes _ Carden • NASA Langley Research Center,

Hampton, VA

Richard L. Boitnott: US Army Aerostructures Directorate,

Hampton, VA

Edwin L. Fasanella: Lockheed Engineering and Sciences Co., Hampton, VA

16. Abstract

Failure behavior results are presented from crash dynamics research using concepts of aircraft

elements and substructure not necessarily 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

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

subfioors 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 and

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

similarity in behavior is giving the designer and dynamists much information about._ha_t to expect in

the crash behavior of these structures and can guide designs for improving the energffabsorption and

crash behavior of such structures.

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

Aircraft crash dynamics

Nonlinear structural dynamics

Impact dynamics

Composite materials

! 19. Security Classif. Iof this report)

Unclassified

18. Distribution Statement

Unclassified- Unlimited

Subject Category 39

20. Security Classif. (of this page)

Unclassified

21. No. of pages 22 Price

15 A03

NASA FORM 1626 OCT B6