impact damage formation on composite aircraft structures...hail ice impacts on sandwich panels...

UCSD FAA JAMS Paper 2014 1

Impact Damage Formation on Composite Aircraft Structures

Principal Investigator: Hyonny Kim*, Professor

Student Researchers: Zhi Ming Chen, Sara White, Monica Chan, Sean Luong

Dept. Structural Engineering, Univ. of California San Diego, La Jolla, CA 92093

Project Description Paper Supporting Presentation Given at Federal Aviation Administration JAMS 2014 Technical Review Meeting

25-26 March 2014, University of Washington, Seattle, WA

* corresponding author: [email protected]

Abstract

Ongoing research at UCSD is focused on understanding the key

phenomena governing the formation of damage to composite structures when

subjected to blunt impact sources. This includes establishing, physically-accurate

modeling capabilities that are developed based on, and validated by, impact

experiments conducted on specimens representing several levels of the building

block. Impact sources of interest are those acting over a wide area and

potentially affecting multiple structural elements, while leaving little/no exteriorly

visible signs of the damage. Specifically, impact by ground service equipment

(GSE) having soft/rubber-covered bumpers, hail ice impact, and large radius

metal tips are being investigated. Key results are: (i) modeling capability has

been established allowing prediction of the onset and propagation of interlaminar

failure modes, including inter-ply delamination and intra-ply fiber failure, (ii)

successful modeling methodology has been demonstrated for predicting blunt

impact damage formation in specimens ranging from simple small-sized


elements, to large-sized 5-frame panels; models are validated by experiments,

(iii) acoustic emission measurements have been shown to provide information

identifying the occurrence of failure events such as the fracture of shear ties

during a blunt impact event, (iv) four distinct internal core failure modes were

identified for blunt impact onto honeycomb core sandwich panels, ranging from

slight core wrinkling (evidence of cell wall buckling) to core wall fiber fracture, (v)

hail ice impacts on sandwich panels produce significant internal core damage

even with low dent visibility.

1.0 Introduction

1.1 Motivation

Impact damage remains a major source of damage to composite aircraft,

including the new generation of composite-intensive aircraft coming into service.

This is a major concern due to the resilience of the composite outer skins which

can sustain a high degree of deformation without developing cracks, even though

the internal sub-structure is damaged. Thus, the traditional reliance on visual

detection to find damage, which worked well for metal skins that dent easily, may

be inadequate for composite airframes, particularly those having skin-stringer

construction. Additionally, sandwich structure exhibits complex behavior when

subject to blunt impacts, with internal damage to the core affecting the core shear

and peel strength.

Investigations are ongoing that: (i) establish a clear understanding of the

development of damage from blunt impacts, and (ii) address the difficulties in

being able to visually detect and predict the corresponding damage. The


research conducted at University of California at San Diego (UCSD) has been

focused on blunt impact sources that are not well understood. As summarized in

Figure 1, these include ground service equipment (GSE) and hail ice.

Furthermore, impacts by widely-ranging radius metal tips are being used as a

generic impact source, representing GSE corners, railings, and unknown-

geometry foreign objects. UCSD’s activities in these areas are closely tied in with

industry and directly address aviation industry-driven needs for increased

knowledge in these areas, from both experimental and analytical/computational

viewpoints, ultimately aiding in the development of more efficient and safe aircraft

structures.

Figure 1. UCSD blunt impact focus: hail ice and ground service equipment.

Blunt impacts can be defined as impact sources that can affect large

areas or multiple structural elements, while potentially leaving little or no

externally visible detectable signs of damage. Blunt impacts come from a variety

Hail Ice Impact• upward & forward facing

surfaces• low mass, high velocity

Ground Vehicles & Service Equipment• side & lower facing

surfaces• high mass, low velocity• wide area contact• damage possible at

locations away from impact

Blunt Impacts• blunt impact

damage (BID) can exist with little or no exteriorvisibility

• sources of interest are those that affect wide area or multiple structural elements


of sources, as illustrated in Figure 1, and can involve a wide range of energy

levels. While the side and lower facing surfaces of the aircraft are subject to

contact with GSE, all exposed upper and vertical surfaces are subject to ground

hail impact (terminal velocity + wind gust) and forward-facing surfaces are

subject to in-flight high velocity hail impacts. These impact sources are described

in more detail:

i. Impact/contact by ground service equipment (GSE) includes ground

vehicles, cargo loaders, and any other equipment coming in close proximity

to a commercial aircraft. GSE is a source of great damage potential due to

the high mass of the GSE (typically 3,000 to 15,000 kg) and subsequently

high energies involved. Figure 2 shows the kinetic energy levels of GSE to

be up to the 103 J range (e.g., 10,000 kg cargo loader moving at 2 mph, or

0.894 m/s, has 4,000 J of kinetic energy). The length scale of contact for

GSE impact can range from ~20 cm for a belt loader’s single D-shaped

bumper, to over 2 m for a long cargo loader bumper (both shown in Figure

2). It should be noted that GSE historically accounts for a major percentage

of damage occurring to commercial transport aircraft [1], and is expected to

continue to be a major source regardless of whether the aircraft is made

from aluminum alloy or carbon/epoxy composite.

ii. Hail ice impacts involve high velocities and thus high energy levels,

exceeding 1,000 J (see Figure 2), mainly by virtue of typical aircraft in-flight

speeds (well over 200 m/s). Even higher velocities (energy) are possible for

rotating components like helicopter rotor and engine fan blades. Ground hail


occurs at lower energy levels, with relatively high velocity, ranging from 20 to

30 m/s, which excites a localized dynamic response in the impacted

structure. The ice projectile is complex in that it exhibits an initial elastic-type

response and then severely crushes during the course of impact. This

crushability results in a larger zone of contact, as shown in Figure 2, which

significantly reduces the propensity to impart visual damage. Meanwhile, ice

impact can produce significant internal damage such as large-area skin-

stringer separation or internal sandwich core crushing, with no external

visibility, unless penetration is achieved.

iii. Low velocity impacts by metal tips of large radius (up to 76 mm radius) are

being investigated as a representation of impacts by generic sources, such

as dropped equipment, or contact with GSE features that are not covered by

rubber bumpers, e.g., railings and rounded corners. The results of impacts

by large radius tips are contrasted with a 12.7 mm radius (1 in. diameter) tip

that is most commonly used for creating barely visible impact damage (BVID)

of a specified dent depth (e.g., 0.3 mm). The 12.7 mm radius tip damage

source, while commonly employed as part of an airframe component’s

damage tolerance program, may not be representative of actual blunt

sources (often of unknown geometry) of damage which can involve higher

energy levels. Larger radius tips allow higher contact forces to be applied

before a dent or other externally visible evidence of damage is left on the

impacted surface.


Figure 2. Blunt impact energy-damage spectrum.

Due to the quasi-brittle behavior of composites, and the large elastic strain

energy stored up until failures initiate, extensive sub-surface damage can

develop when impacts occur at levels just exceeding the amount needed to

initiate failure [2-4], i.e., when the failure threshold is exceeded. Additionally,

damage from blunt impacts to internal stiffeners can be extensive, existing in the

form of separation of stiffeners from the skin and fracture of the stiffeners. Of

critical concern is whether damage is extensive enough to result in the structure

losing ultimate (or even limit) load capability.

Among the three impact sources being studied (see Figure 2), both the

low velocity metal tips and high velocity ice impact can usually be associated with

a damage size limit. This damage limit reflects the ability for these impact

sources to induce localized failure at the point of impact/contact. Thus,


maximum-level cases of velocity and energy will typically produce penetration

damage as a result of this localized response. Penetration damage has a size

limit roughly equivalent in dimension to the impacting projectile. For this reason,

the worst-case scenario might not always produce the most critical mode(s) of

damage.

For the GSE impact, a non-local response has been observed to be

possible. Specifically, internal damage can occur at locations away from the point

of impact/contact, e.g., at joints or stress concentrations along the path of internal

reaction. Thus, the damage size limit is not clearly known and, depending on the

severity of the impact event, can possibly be larger than the length scale of the

contact between the projectile and structure. Furthermore, the GSE contact

speed can be quite varied, even very slow (quasi-static) level as indicated in

Figure 2, and so the energy-based description might not fully describe GSE-

aircraft interactions.

1.2 Objectives

The objectives of this research program focuses on impact damage

formation by a range of sources, including: (i) low velocity wide-area blunt impact

– vehicle/ground equipment contact with composite aircraft structure, (ii) high

velocity hail ice impact, and (iii) impact by low velocity metallic tips of large

radius. A common set of objectives exists for these three project focuses:

1. Characterize blunt impact threats and the locations where modes of

damage can occur.


2. Understand blunt impact damage (BID) formation and visual detectability,

specifically seeking to:

determine key phenomena and parameters governing BID formation,

understand how damage is affected by bluntness/contact-area,

identify and predict failure thresholds (useful for design),

establish what conditions relate to development of significant internal

damage with minimal or no exterior visual detectability, and

provide guidance on the inspection and detection of BID to internal

structural members.

3. Develop analysis methods, including new modeling capabilities validated

by tests. This focuses on failure modes and phenomena that are currently

difficult or even not possible to predict with conventional FEA approach.

Additionally, guidance will be established on defining how to analytically

predict whether damage is visually detectable or not.

4. Develop blunt impact testing methodologies, including defining appropriate

boundary conditions between full-scale panels vs. complete structure for

GSE impact studies.

1.3 Approach

While each of the project focus areas has unique challenges related to

their length scales and velocity regimes, a common general approach to achieve

these objectives is the following:

1. Experiments: impact representative structure/specimens, specifically


wide area high energy blunt impact – e.g., from ground service

equipment,

hail ice impacts – in-flight and ground hail conditions, effect of

internal stiffeners, internal damage to honeycomb sandwich core

with little/no dent depth,

low velocity impacts – non-deforming impactor which serves as a

generic impact source, understanding large radius effects on

damage visibility.

2. Modeling development capable of predicting the various impact damage

modes, including delamination and just-visible surface cracking –

nonlinear FEA, analytical simple models, energy balance.

3. Communication of results to industry and collaboration on relevant

problems/projects via workshops and meetings (at UCSD or at industry

site, by teleconf/WebEX meeting).

1.4 Expected Outcomes

The expected research outcomes will provide information and establish

methodologies that will aid engineers in assessing whether an incident could

have caused damage to a structure, and if so, what sort of damage mode, extent,

and location such damage would occur. This information assists with

understanding of what kind of inspection techniques should be applied to assess

the extent of damage. Furthermore, it is expected that design engineers can

make use of the research outcomes to: (i) improve the resistance of composite


aircraft structures to damage from blunt impacts sources, (ii) provide critical

information on the mode and extent of seeded damage for use in damage

tolerance considerations, and (iii) define relationship between blunt impact threat

and expected damage severity – e.g., specifying safe GSE operation speeds, or

at for what size hailstones should deeper inspection be necessary.

Modeling capabilities are being developed to establish clear guidance on

how to predict blunt impact damage modes and whether the blunt impact event

will leave visible evidence of any significant damage formation. This includes

providing the ability to predict the initiation, progression, and final failure modes

resulting from the three impact sources being studied.

1.5 Research Partners

UCSD’s research on impact of composites is of direct interest to industry,

as reflected by the research partners. Table 1 summarizes the research partners

that are involved in UCSD’s FAA-funded research activities. Large and small

aircraft manufacturers, a small composites-specialty engineering firm, and a

material supplier are represented.


Table 1. UCSD research partners on FAA-funded activities.

Company/ Organization

Description/ Expertise

Role in UCSD Project

Boeing OEM – Large Transport Aircraft

Provide guidance and input on blunt impact. Particular focus on blunt impacts onto panels of stiffened-skin construction. Possibly supply test panels. Participate in on-site meetings.

Bombardier OEM – Small/Regional Aircraft

Provide guidance and input on hail ice impact, particularly focused on sandwich panels. Supply test panels for ice impact.

San Diego Composites

Composites Design and Manufacturing

Provide technical advice on the direction of the project, guidance on the design of the large-scale blunt impact composite test panels, guidance on the design of tooling for manufacturing the test panels, and access to large autoclave for curing panels.

Cytec Engineered Materials

Materials Supplier

Provide technical advice on project directions. Provide guidance on use of materials. Supply carbon/epoxy prepreg materials to support fabrication of test specimens at UCSD for both blunt impact and hail ice studies.

United Airlines

Airline Provide guidance and feedback on project directions particularly with reference to operator view.

Delta Airlines Airline Provide guidance and feedback on project directions particularly with reference to operator view.

JCH Consultants

Consultant on Aircraft Safety and Composites

Advise on direction of project, provide guidance on tests, data reduction, results interpretation and dissemination to the public and senior-level individuals in industry as well as to military (Air Force).

Avanti Tech, LLC

Small Business – Nondestructive Evaluation

Coordinated activities focused on measuring ultrasonic guided wave signals originating from active damage sources, to be used for damage detection systems.

2.0 Wide Area Blunt Impact Damage 2.1 Background and Motivation

With the increased use of composites in airframe primary structural

components (e.g., fuselage and wing), there is a need to better understand the

damage mechanisms caused by accidental transverse impact loading,

particularly for high energy levels. The low resistance of composite laminates to

transverse impact damage is a major weakness of these materials.

The largest source of damage to a commercial aircraft is caused by

accidental contact with ground service equipment (GSE) [1]. GSE can weigh up


to 15,000 kg and travel at speeds up to 1 m/s while within close proximity of the

aircraft. The high mass of GSE can result in very high impact energy upon

collision with the aircraft, and can result in widespread internal damage. In

addition, rubber bumpers typically outfitted to GSE can reduce the local surface-

visible damage in the aircraft skin and mask the occurrence of an internally-

damaging impact event.

Non-destructive evaluation (NDE) and structural health monitoring (SHM)

methods are of increasing interest for aerospace structures. Aside from

understanding and predicting damage that occurs from GSE, it is of interest to

quickly determine what damage may be present after an impact has occurred by

using methods, such as NDE, that do not involve disassembling the aircraft for

tedious and costly interior visual-based inspections. Although accurately

predicting small damage occurring from wide area impacts can be very difficult

due to complexities in geometry and materials, identifying large-sized damage

modes is of key interest.

2.2 Summary of Previous Results

GSE Blunt impact experiments on large scale fuselage panels have been

completed up to Phase II (see Figure 3). The last 5-frame specimen tested,

Frame04, was geometrically identical to Frame03. This panel was tested twice,

once mounted with composite shear ties (ID: Frame04-1) and again with stout,

aluminum shear ties (ID: Frame04-2). Comparing the composite shear tie tests

to the aluminum shear tie test, the overall failure path has been altered. In the


composite shear tie panel tests, the initial failure mode would always be shear tie

delamination and fracturing. Then, the C-frames become free to rotate or make

contact with the stringers, resulting in C-frame failures from large deformation or

high contact stress (see Figure 4). In contrast, the Frame04-2 panel sustained

higher impact load without shear tie failure and with minimal deformation of the

shear ties. Subsequently, the C-frames were more directly loaded, resulting in

localized shear fracturing of the C-frames as both the initial and final failure mode

(see Figure 4).

Figure 3. Blunt impact program phases and modeling capability development.

Basic Elements- Excite Key Failure Modes- Model Correlation Data

- Understand Damage Formation &Relationship to Bluntness Parameters

Large Panel- e.g., 5 Bays

- Damage Excitation- Damage Thresholds- Model Correlation

OEMHardware- 1/4 to 1/2Barrel Size

- Vehicle Impacts

Scaling,B.C. EffectsDynamics

Scaling,B.C. Effects

Increasing LengthScale, Complexity,and Specificity

Phase III(Year 3)

Phase II(Year 2)

Phase I(Year 1)

Modeling CapabilityDevelopment & Correlationwith Test are Key Aspects

at Each Level

High Fidelity Explicit

Dynamic FEA- Complete Simulation- Damage

Progression- Final Damage

State

Elastic Static FEA

- Damage Initiation

Simple Models: Reduced Order 2DOF, Energy & Momentum

Balance

UCSD Developing Methodologies at All Levels

Increasingly Complex Phases of Activity

- gain fundamental understanding at bottom

- use modeling to generalize


Figure 4. Comparison of failure modes for panels having composite (Frame03) vs. aluminum (Frame04-2) shear ties.

Significant efforts were made toward FE modeling of the GSE blunt impact

tests. An Ogden material model was used to simulate the OEM bumper material

03

Frame03 – Non‐local failure of frame away from center due to load transfer between stringer‐frame contact & frame rotation. No visible cracks or dent.

Near BCs Near

BCs

04‐2

Frame04‐2 – Local failure of frame @ center. No shear tie failure, thus no major frame rotation. Minor skin cracks visible away from impact site.


to obtain accurate impact contact pressure distribution. The modeling

methodologies were refined to accurately model the deformation, failure modes,

and failure propagation in the StringerXX experiments, specifically for specimens

Stringer05 and Stringer06. Finally, Python scripts were developed to enable

parametric studies of the large-sized FrameXX panel impacts. Basic models

have been established for Frame03 and Frame04 up to C-frame failure. These

models provided some insight into damage evolution within the panels, but were

missing the initial delamination failure modes which subsequently affect the

overall panel response.

2.3 Recent Results

2.3.1 Lab-Scale Element-Level Experiments

An inverted testing pyramid approach [5] was taken to analyze the blunt

impact experiment results. The locations of key failures from the previously-

conducted experiments have first been identified, and then isolated small-scale

experiments were conducted on the individual elements of the large panel to

recreate these failure modes. Element and coupon size specimens were

manufactured, and were tested in loading conditions designed to simulate the

loading sustained by these components during the blunt impact experiments.

2.3.1.1 Stringer Element Specimen Tests

Stringer element specimens have been fabricated to examine the bumper-

to-skin interaction and initiation of skin surface damage (i.e., visible surface

cracking) during the StringerXX tests. The stringer element specimens are 76


mm wide with skin and stringer components cut from the undamaged part of

specimen Stringer00. Specimens are potted in epoxy resin and gripped in an

uniaxial test machine as shown in Figure 5. The stringer element is indented

with the same 76 mm wide D-shaped rubber bumper as the StringerXX panel

tests. One stringer element, SE01 has been tested so far. During the test, the

specimen was quasi-statically indented until initial skin-visible surface damage

was created. Due to the potting constraint of the specimen (see Figure 5), the

majority of the deflection occurred within the skin.

Figure 5. Test setup for stringer element indentation.


High bending stress concentration on the skin near the stringer-to-skin

joint location led to visible skin surface failure. Figure 6a shows the visibility of

surface cracks in specimen SE01, and the photomicrograph in Figure 6b shows

detail of the crack observed at the joint corner. It can be conclude from these

figures that at the point when barely visible external damage is created, the

external plies (near surface plies on both sides of the skin) will experience

through-thickness cracks, and some modest delamination. This focused

experiment provides critical data for establishing analysis-based criteria defining

the onset of damage visibility.

Figure 6. (a) SE01 post failure, visible skin crack at the joint location, and (b) photomicrograph of the SE01 showing the formation of cracks and delamination

at the skin-stringer joint location.

The SE01 test is simulated using continuum shell (SC8R in ABAQUS)

elements to establish the ability to predict the formation of visible cracks. The

quarter symmetry model shown in Figure 7 uses the Hashin-Rotem failure

criterion to represent in-plane ply failure. The force vs. actuator displacement

curve of the test and FE model is shown in Figure 8. As can be seen from the

figure, a nearly 1:1 matchup is found between these curves in the elastic region.

(a) (b)


In the experiment, failure initiation was measured at 21.55 kN, whereas the FE

model failed at 19.04 kN. The failure mode of the FE model matches the failures

found in the experiments, with cracks forming in the laminate outer plies, as

shown in Figure 6b and 7.

Figure 7. Quarter-symmetric model of the test.

Figure 8. Load vs. displacement plot of the SE01 test and FE model.

2.3.1.2 Shear Tie Compression Element Specimens

During Phase II large-panel blunt impact experiments on the specimen

Frame03 which was configured with composite shear ties, damage to the shear

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Contact F

orce (k

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Actuactor Displacement (mm)

SE01 Test

SE01 FEA Model


ties always occurred first, thereby absorbing much strain energy and allowing the

impact loading to distribute over a larger number of structural elements. The

sequence of failure modes in the shear ties are: delamination in the curved

region, followed by local crushing and fracturing under compression loading, and

finally buckling-induced bending failures resulting in complete fracture of the

shear ties. Subsequent panel deformations are then affected by these shear tie

failure modes. Hence, it is desirable to understand and accurately model the

complete damage sequence in the shear ties.

A focused experiment on shear tie behavior under direct compression has

been conducted to make observations on damage initiation and evolution which

will support model development. Figures 9 and 10 show the test photos and the

load vs. displacement plot of this experiment, respectively.

Figure 9a shows the shear tie compression experimental setup prior to

loading. A 66 mm wide shear tie section was bolted to a curved aluminum

support base plate, as shown. A top loading plate with V-groove was used to

apply compression loading on the top of the flange, while permitting free rotation.

As the load is applied, the shear tie undergoes bending and shearing at the

curved corner of the shear tie. In response, the shear tie corner flexes in an

opening-moment mode, until it delaminates, as seen in Figure 9b (opening

moment produces radial tension stress in curved laminate). The Initial

delamination, developing at an applied load of 7.0 kN, resulted in the first load

drop (see Figure 10 curve labeled Experimental Data). Additional delamination

at the radius then continued to accumulate and grow. In Figure 9c, outer ply


failures are shown to occur at a load level of 4.9 kN (see displacement of 5 mm

in Figure 10). This ply failure is driven by the highly localized transverse shear

stresses at the radius region. Crushing failure eventually softened and flattened

the radius region against the base plate (see Figure 9c). At this point, the shear

tie flange is subject to direct compression loading which induces buckling and

subsequent bending-failure at 10.5 kN (see Figure 9d and 13.9 mm displacement

in Figure 10).

Figure 9. (a) Pristine shear tie compression specimen, (b) initial delamination at shear tie corner, (c) crushing failure and flattening of the corner, and (d) buckling-

induced bending failure.

Figure 10. Load vs. displacement plot of the SE01 test and FE model.

02468

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Contact F

orce (k

N)

Crosshead Displacement (mm)

Experimental DataFE 4 LayersFE 6 Layers


The experiment was modeled in ABAQUS using cohesive surface

interactions for simulating inter-ply delamination. The Hashin-Rotem failure

criterion for composite ply failure was not used because it cannot account for

intra-ply transverse shear failure modes (i.e., fracturing of the fibers under

transverse shear stress). Instead, the Hill failure criterion for orthotropic

materials was used. This criterion allows for the modeling of shear tie crushing

and severing after the shear tie has buckled.

Figure 11 shows a series of FE modeling results for the composite shear

tie undergoing failure, analogous to the sequence shown in Figure 9. The shear

tie layup consisted of 12 woven carbon fiber plies. The analysis has been

conducted using four and six layers of elements (i.e., ply groups) stacked through

the thickness, with groups composed of three or two plies, respectively. The

ability of continuum shells to predict transverse stresses tends to improve with

increasing number of elements stacked through the thickness [6]. In addition,

dividing the plies into multiple layers that can deform separately allows directly

modeling of the interface layer separation with cohesive zone techniques. As

can be seen from Figure 9d, delamination occurred at multiple (approximately

four to five) interface layers at the shear tie corner, and each delamination

contributed to softening of the shear tie. Hence, this multi-layered modeling

approach was taken to ensure the correct residual stiffness was captured. The

shear tie residual stiffness would be over-predicted if fewer interface layers were

used in the model.


Figure 11. (a).Unloaded shear tie compression model with 4 layers of elements through the thickness, (b) initial delamination at shear tie corner, (c) crushing

failure and flattening of the corner, and (d) buckling-induced bending.

The FE analysis results are plotted in Figure 10 together with the

experimental data. As can be seen from this figure, modeling the shear ties with

four layers of elements is insufficient to capture the initial delamination load.

However, the overall deformation response is still correct, as can be seen in

Figure 11. Increasing the number of layers through the thickness to six resulted

in better response of the shear tie coupon for the initial delamination load, as well

as the shear tie crushing behavior up to 6 mm of crosshead displacement. After

the ply element failure had initiated, the loads oscillate due to stress singularity

experienced by elements at the corner. When these elements fail, they are

removed from the simulation which caused load drops in the model (a result not

seen in the physical experiment). Additional work needs to be done to overcome

this shortcoming and improve modeling fidelity. Note that in this model,

physically-accurate properties best representing these composite materials were

used. No adjustment and “tuning” of properties or parameters was done to attain

better correlation.


2.3.2 Large Scale Panel Modeling Development

The Frame03 (and Frame04-1) large-sized panel model has been updated

using techniques developed from the shear tie compression test specimens. The

previous 1st-generation Frame03 model yielded an approximate overall panel

response, but the exact deformation and pattern of failures were not captured

(see UCSD 2013 JAMS Paper Figure 10). Key failure modes in the shear ties

that would later influence subsequent panel failures were missing. Incorporating

the initial shear tie failure modes, namely delamination and subsequent

softening, improved the modeling accuracy. The updated half symmetric model

of Frame03 is shown in Figure 12.

Figure 12. Updated Frame03 model: (a) shear tie delamination occurs, and (b)

shear tie buckling-induced bending failure.

(a)

(b)


Figure 12a shows the Frame03 model soon after impact when shear tie

radius delamination initiates, while Figure 12b shows shear tie failure and large

rotation of the frames. Figure 13 shows the load vs. indentation plot of the

Frame03 test compared with the FE model results. In the FE model, the load

increased to 72.3 kN before delamination and crushing of the shear ties caused

an initial load drop, matching the response seen in the experimental loading

curve. The shear tie flange made contact with the panel skin, similar to the

behavior observed in the shear tie compression tests, thus allowing the load to

build up again to a load of 97.4 kN, at which point the center shear tie flange

buckled and fractured (see Figure 12b) and then load transfer is through direct

stringer-frame contacts. As shown in Figure 13, this model is accurate for up to

37 mm of indentation. Due to stress singularity issues and how the failure

criterion was implemented, the FE model cannot predict the failure pattern past

37 mm of indentation. In the model, the load rose to 113 kN before the 2nd row

of shear ties buckled and failed in bending.

Improvements to the modeling capability are ongoing, following the

approach of establishing accurate predictions for small element-sized geometry

with simple loading, before applying to predicting the response of a complex

built-up structure (e.g., compare shear tie compression in Figure 11 and full

panel model in Figure 12). The overall plan for modeling capability development

is shown in Figure 14. Clear definition of modeling methodologies to accurately

capture important failure modes, based on physically-accurate modeling of the

experiments, will subsequently be applied to investigating various new structural


configurations that have not been experimentally tested. This includes aspects

such as glancing impact, and the effects of additional internal structure such as

the floor beams and the joints between the frames and the floor.

Figure 13. Load vs. indentation plot of the Frame03 test and FE model.

Figure 14. Modeling capabilities plan.

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Force (kN)

Indentation (mm)

FE Hill Criterion with Cohesive Zone

Experimental (Combined Frame03 & 04)

Glancing ImpactSize, Complex Internal Structure, Geom., Joints

Various Impactors & Scenarios (vo)

Models of Generic Curved Panel Specimens- Static- DynamicExperimental Validation

Capture Key Failure Modes (Major Damage)

Damage Initiation Criteria

Damage Progression

Dynamic Effects

Externally Visible?

EstablishCapabilities

Define Methodologies

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How to predict this with FEA?

Apply to study and predict response for:


2.3.3 Frame-to-Floor Joint

In order to better represent the structural configuration of a fuselage,

future test specimens have been designed to incorporate floor beams and the

frame-to-floor joints. These components create a high stiffness region which can

greatly affect the stress distribution within the frame and the failure modes,

particularly cracking near/in the joint region.

During a meeting with the FAA and Boeing engineers on November 8,

2013 in Seattle, WA, future plans were made regarding plans to incorporate

frame-floor interactions in the GSE impact studies. Key impact regions along a

frame were identified and are shown in Figure 15:

Region 1 is the central (blue color) portion of the frame which is most

compliant and would develop higher deformations and bending stresses

during a GSE blunt impact.

Region 2 includes the regions near the floor joints (green zones within 3-

4X frame depths from floor joints) will be more stiff during impact, and will

develop high beam shear stress within the frames due to the proximity of

these regions to the highly stiff floor joints which act as stiff boundary

conditions to the frames.

Region 3 includes the floor joint region (red) which would be the most stiff

as an impact at these locations would be directly parallel with the floor

joint. Direct GSE hits to these locations are anticipated to easily damage

the frame and frame-to-floor joint.


Figure 15. Impact regions and floor structures.

The frame-to-floor bracket has been designed to fit tightly within the floor

beam, connecting to the web and flange. The current design is shown in Figure

16, where the aluminum alloy bracket will connect the floor beam flange and C-

frame via bolts. The frame-to-floor joint is first being modeled in FEA (currently

ongoing) prior to development of a detailed specimen and conducting

experiments to investigate joint-adjacent failure. The model geometry shown in

Figure 17 will include a portion of the skin and stringers, as well as shear ties and

frame. This model development is ongoing, and it includes boundary condition

studies to determine the appropriate rotational and translational stiffness that

should be applied as boundary conditions to the ends of the C-frames.

Region 1

Region 2

Region 2

Region 3

Region 3


Figure 16. Frame-to-floor bracket design.

Figure 17. FEA model geometry of floor structure incorporated in to frame.


Failure modes of interest include the frame cracks observed in previous

UCSD large-panel tests of specimens Frame01, Frame02, Frame03, and

Frame04. These cracks developed due to high bending and torsional

deformation of the frames, as well as high stresses developed when stringers

and frames made contact during blunt impact. Examples of these frame cracking

failure modes are shown in Figure 18. Additional failure modes of interest will be

cracks anticipated to develop at the fastener holes of the frame-to-floor bracket,

and under direct crushing of the frame during impact at Region 3.

Figure 18. Examples of frame flange cracks resulting from blunt loading.


2.3.4 Non-Destructive Evaluation Measurements

Supporting the large-scale panel blunt impact tests, non-destructive

evaluation (NDE) measurements were performed. Each panel was equipped with

macro fiber composite (MFC) acoustic sensors which were bonded to the inside

skin surface of the panels. The acoustic emissions resulting from impact and

proceeding panel failures were recorded. The MFC sensors were arranged in a

rosette pattern, as shown in Figure 19, with the aim of determining the incoming

wave direction by methods described by Matt and Lanza di Scalea [7].

Figure 19. MFC rosettes used to determine a wave source.

The general concept of the MFC rosettes is that the three sensors placed

next to each other at a known separation angle can be used to determine the

incoming direction angle of a Lamb wave. This is due to the varying voltage

response excited depending on the incoming angle with respect to the orientation


of the MFC sensor (sensitivity is directional). When a minimum of two rosettes

(six sensors) are used, each which will provide an incoming wave angle which

can be projected to intersect, thereby revealing the wave source location (a

damage event produces waves). This approach is not dependent on the velocity

of the wave source but instead relies on the directivity of voltage output of the

MFC sensors, which is a favorable for complex geometries, anisotropic

laminates, and dispersive materials where determining wave speeds is difficult.

Specimens Frame 03 and Frame 04 were both equipped with four rosettes

(12 MFC sensors total) located as shown in Figure 20.

Figure 20. Rosette locations and shear tie labeling on specimens Frame03 and Frame04 prior to C-frames being assembled to panels.


Unfortunately, the acoustic emissions recorded from Frame 04 were

saturated as a result of too small of a window chosen for the voltage, thus

making much of the data unable to be analyzed. For the Frame 03 test, the

window was large enough to capture all the data from the test and will be the

focus of discussion from here on. Using the methods described by Matt and

Lanza di Scalea [7], along with incorporating small time windows to analyze each

perceived damage event, the predicted damage locations are plotted in Figure

21.

Figure 21. Predicted damage locations for Frame 03 for entire test duration by methods of reference [7].

A closer look at the damage developed as a function of time was

documented using the high speed videos. The data were recorded with the

same time trigger as the high speed video. Much of the visible damage occurred

at the shear ties and in the C-frames. These events were then compared to the


responses of the sensors at a given location. Table 2, was constructed using the

high speed video with each row representing the shear ties which are labeled

according to Figure 20. The colored block indicates visible cracking and a

colored block with an 'x' denotes major failure of the shear tie.

Table 2. Time table of shear tie failure based on high speed video. Shear tie

label in rows with colored block denoting visible cracking and colored block with an 'x' denotes major failure.

Time (ms) 1

21

122

123

124

125

126

127

128

129

130

131

132

133

134

135

136

137

138

139

140

141

M x x x R

Time (ms) 1

42

143

144

145

146

147

148

149

150

151

152

153

154

155

156

157

158

159

160

161

162

H x x x x R

Time (ms) 1

63

164

165

166

167

168

169

170

171

172

173

174

175

176

177

178

179

180

181

182

183

R x x

no events detected between 184 ‐ 220 ms

Time (ms) 2

21

222

223

224

225

226

227

228

229

230

231

232

233

234

235

236

237

238

239

240

241

G x x x x x I x x x x x x x L x x x x N x x x x Q x x x x x x x x x S x x x x


The spectrogram from the signals, which visually shows the frequency

content as it varies with time along with a relative amplitude, is generated from a

selected event and compared with the timing of damage events visible in the high

speed videos. Shear tie R failure, which is located near rosettes 2 and 3 (as

shown in Figure 20), was of particular interest because there were no other

visible events occurring at that time. The spectrograms shown in Figure 22

compare the results from one sensor near shear tie R with one far from shear tie

R, to provide some insight on how much information from a shear tie failure event

is transferred into the skin and transmits along the skin (through stringers and

rows of shear ties along the way) before being detected by the MFC sensors.

Figure 22. Spectrogram results from two different sensors (near and far) during a shear tie R failure.


This comparison reveals a strong attenuation of the signals for increasing

distance, with a high concentration of the signal having ~160 kHz frequency

content. Delamination damage and minor-level cracking in shear tie R initiated at

121 ms and was constantly evolving and growing, and thus rosette 3 constantly

picked up a significant-strength signal. Finally, when major fiber-failure occurred

at 167 ms, a more intense broadband signal is abruptly visible in rosette 3 at 167

ms time, and a significant weaker narrow signal also appears in rosette 1 at this

time.

2.4 Discussion

Focused, small-scale experiments on element-level specimens were

conducted to gain insight into the key failure features of the large experiments.

Element-level FE models were then created to accurately predict these

experiments which help to validate choices made in defining the models and

demonstrate the capability to predict these failure modes. Once validated

models were established at the element level, they were incorporated back to the

larger-scale models.

These high fidelity physics-based models, created at both small and large

scales, can predict the key failures modes within a narrow error margin. The

stringer elements were able to predict skin cracking initiation within a 10% load

error. In the case of the shear tie compression coupons, the results were in

accurate agreement with the experiments after refining the mesh through the

thickness of the shear tie to six layers of elements. The modeling methodologies


were then transferred to the large StringerXX and FrameXX models. The results

for StringerXX modeling was shown in the 2013 JAMS report. It is now shown

that by using a refined shear ties model in the FrameXX models, the progression

of damage following failure of the shear ties is much better captured. However, it

is important to note that this work is ongoing and there is still much to be done to

improve the predictions (e.g., address singularities created during element

deletion, overcome the force oscillation created when elements fail).

The NDE results from the Frame 03 test using methods described by Matt

and Lanza di Scalea [7] show discrepancies between predicted damage

locations and those found visually on the specimen and observed via the high

speed video. This is likely due to the complexities of the test specimen along

with the inability of this method to differentiate between multiple acoustic

emission sources occurring simultaneously. Slight variations in the sensor

sensitivity could also strongly affect the location prediction results. Further

development is needed for reliable damage detect\ion, especially for wide area

impacts which cause a number of discrete damage events at various locations.

The spectrogram results along with cataloging the failure events are used to

show that the sensors near the shear tie failures exhibit a higher amplitude

response in comparison with the sensors further away from the failure event.

This indicates that acoustic waves are capable of being transferred from the

shear tie to the skin, even though the shear tie and skin are not bonded together.


2.5 GSE Blunt Impact Conclusions

To gain insight into the large scale blunt impact experiments, an inverted

building block pyramid approach to testing was used. Small-scale

experiments were conducted on element-level specimens extracted from

the large-sized panel specimens to directly study the individual failure

modes and to provide data for model development and validation.

The methodology presented allowed for the in-depth analysis of failure

initiation and propagation in the experiments, resulting in the development

of accurate finite element models that are based on actual physical

observations and test data.

Acoustic emissions from shear tie failures are able to be identified using

sensors bonded to the skin of a panel, even if the shear ties are only

bolted to the skin (i.e., not tightly acoustically coupled).

3.0 Blunt Impact Damage to Sandwich Panels 3.1 Background and Motivation

Composite sandwich panels are being used more widely in various

aerospace structures because of their lightweight yet stiff nature. However, their

impact characteristics are not well understood, particularly when impacted by

blunt sources creating internal damage with little/no exterior visibility (no dent, no

cracking). An in-service aircraft may encounter impacts from a number of

projectiles such as blunt impact from ground service vehicles and runway debris,


tool/equipment drops, or hail impact while the plane is on the ground or in flight.

The varying types of projectiles and range in speeds at which impacts may occur

can result in different damage modes such as fiber failure, matrix cracking,

delamination, facesheet disbonding, core crushing, and the interaction between

any of these damage modes. Furthermore, such damage may grow and

eventually lead to sudden large-scale failure due to repeated loading cycles,

such as pressure loading/unloading process during repetitive ground-air-ground

flight cycles. Often, such damage may initiate at a very small scale (mm) or

below the facesheet surface and cannot be visually detected from the exterior.

Therefore, it is important to understand what type of damage to expect after

impact has occurred to best determine what actions need to be taken.

3.2. Summary of Previous Results

This sandwich impact research activity started in early 2013. Since no

significant results were available by March 2013 to be reported in the last 2013

JAMS paper, no summary of previous sandwich impact results is included here.

3.3 Recent Results

3.3.1 Low Velocity Blunt Impact by Large Radius Metal Tips

Aircraft routinely experience impact damage from maintenance related

sources, from contact with GSE, or dropped tools and equipment. In the case of

composite sandwich panels, external damage is not always present or visible,

making internal damage difficult to detect. This impact study focuses on low

velocity impacts from a wide range of radius tips to determine the correlation


between internal and external damage as a function of impactor radius. Due to

the lightweight nature of the sandwich specimens, this investigation is particularly

interested in impacts of low energies (2 to 14 J) in which external damage is

barely visible.

The test specimens in this investigation were composite sandwich panels

with a honeycomb core obtained from an Airbus A320 rudder. Four impactor tips

of different radii, as shown in Figure 23, were mounted to the pendulum impactor

shown in Figure 24. The test panel was held in a 165 mm square window and

impacted with several tip radii and kinetic energy level combinations, as

summarized in Table 3.

Table 3. Number of tests conducted at each radius/energy combination.

Figure 23. Impact tips of radii 12.7, 25.4, 50.8, and 76.2 mm (left to right).


Figure 24. Pendulum impactor test set-up.


3.3.1.1 External Damage and Data

At each impact, the maximum impact force and impact energy are

recorded. The dent is measured within five minutes after impact to obtain the

immediate dent profile and maximum dent depth. Because sandwich panels tend

to experience relaxation over time, the dent is measured again after 24 hours to

document a relaxed dent profile. A representative example is shown in Figure 25.

Figure 25. Dent profile of test site impacted with R12.7 at 4 J.

The visibility of the dent is then characterized based on a parameter

defined as the maximum depth D divided by the dent span S. The span is

determined as the span of the dent at one-third of the depth below the original

surface plane. Because visual detectability of damage is dependent on both span

and depth, this “visibility” parameter is more indicative of its visual-detectability

than depth alone. The visibility, plotted in Figure 26 as a function of impacting

kinetic energy, shows increasing visibility with kinetic energy, but no differences


between tip radii is apparent – i.e., the different tip radii tend to produce visibility

(depth/span) values falling on a common curve over the range of energies tested.

Figure 26. Measured kinetic energy vs. visibility for lower energies.

A plot of the relationship between measured kinetic energies and

maximum dent depth can be seen in Figure 27. The correlation between dent

depth and kinetic energy is fairly linear up to 6J, and very similar for all four tip

radii in the lower energy range (up to 6J). At higher energies of 10 to 14 J, the

dent depths vary considerably for different radii, suggesting the depths increase

at different rates with increasing impact energy. In most cases, the panel

experienced only dent damage. However, the uppermost four outlying data

points are instances where the facesheet was also cracked, which only occurred

with tip radius 12.7 mm (blue) and 25.4 mm (red).


Figure 27. Measure kinetic energy vs. maximum dent depth; circled points indicate facesheet cracking.

3.3.1.2 Internal Damage and Data

Each test site was sectioned (cut through impact site) to observe the

damage to the honeycomb core. In general, the damage tends to exist at a depth

which plateaus to fixed level, which is visible in Figure 28. In this figure, the

damage progression can be observed for panels impacted with tip radius 50.8

mm at energy levels of 4, 6, 10, and 14 J. Although the depth reaches a

maximum level, the internal damage span, defined as the measureable width of

the damage parallel to the facesheet, increases with increasing kinetic energy.

Face Sheet Cracked


4 J

6 J

10 J

14 J

Figure 28. Sectioned view of damage progression for R50.8 mm tip impact with increasing impact energy: 4, 6, 10, and 14J from top to bottom.

The cross sectional inspections (e.g., in Figure 28) and additional optical

microscopy (Figure 29) were used to identify four distinct modes of internal

damage to the Nomex honeycomb core. These modes are identified as:

Mode A: Slight wrinkling of cell walls (see Figure 29a).

Mode B: Visible wrinkling of cell walls (see Figure 29b).

Mode C: Buckling of cell walls (see Figure 29c).

Mode D: Fracture/Bursting of cell walls (see Figure 29d).


(a) Mode A (b) Mode B

(c) Mode C (d) Mode D

Figure 29. Sandwich core damage modes; impact applied to facesheet located at

top of each photo.

Further correlation of damage modes with impact energy, peak forces,

and external damage visibility is currently still in progress.

3.3.2 Ground Hail Impact

Large-sized hail impact is a realistically-occurring event for aircraft. The

size of hail may reach up to 75 mm and the speed of impact can range from

terminal velocity (~25 m/s) to in-flight speeds (over 200 m/s). The test specimens


utilized for this research were composite sandwich panels obtained from an

Airbus A320 rudder that was removed from service. Due to the limited number of

test specimens available from this part, a selected range of testing conditions

were explored in order to focus on damage morphology for given threat levels

(velocity) and glancing incidence angles, when subject to impacts by 50.8 mm ice

spheres (representing hailstone).

The goal of these tests is to determine the condition at which there is little

to no visible surface damage, while still having internal core damage present.

While the results are specific to this specific lightweight core sandwich panel

configuration, the methodology and approach are generically applicable to a

broader range of the sandwich panel design space.

The panels were oriented at glancing angles of 90 degrees (normal

impact), 40 degrees, and 25 degrees relative to the path of the incoming ice

sphere. All impacts were conducted using 50.8 mm diameter ice spheres

projected at nominal velocities of 25 m/s and 50 m/s. An additional parameter of

interest was the facesheet thickness which was either ~1.19 mm or ~1.87 mm

(these will be referred to as thin specimens and thick specimens, respectively).

Recent data are summarized with all pertinent damage measurements and

observations in Table 4.


Table 4. Summary of recent results with damage observations.

After each specimen is impacted, a dent profile is measured within 15

minutes, and the peak dent is obtained from that profile. As Table 4 shows, most

of the thin facesheet specimens suffered visible surface damage as well as

significant internal damage. Figure 30 shows the relationship between impact

energy and peak dent depth of the thin facesheet specimens. The points that are

located on the x-axis represent the tests that fractured the skin of the specimen,

so no useful peak dent data were measured from those tests. Over the range of

energy tested, the data show that there is a linear relationship between impact

energy and peak dent depth for each glancing angle, until a certain threshold is

reached, at which point the skin fractures. As expected, the dent depth increases

as a function of impacting kinetic energy at a higher rate for 40 degrees than at


25 degrees. The 90 degree impacts caused such high damage with relatively low

energy levels (see Tests 001 and 002 in Table 4) that no more specimens were

devoted to that condition.

Figure 30. Impact energy vs. peak dent depth of thin specimens.

Figure 31 shows an example of the level of internal damage resulting from

a 25 m/s impact at 40 degrees (panel sectioned through impact site). Other

details about the impact can be found in Table 4. Even though the dent is not

clearly visible, there is still a large amount of wrinkling visible in the core near the

impact-side facesheet (upper side in photo).

Cratered


Figure 31. Sectioned view of thin panel impacted at 23.67 m/s and 40 deg (Test 005).

An example of severe core damage is shown in Figure 32. This panel was

impacted at 50 m/s (nominal) velocity at a 40 degree glancing angle. The energy

level was 59 J, resulting in a significant-sized and clearly-visible crater.

Figure 32. Sectioned view of thin panel impacted at 44.76 m/s and 40 deg (Test 006).

For comparison, the same testing conditions were replicated on the thicker

facesheet specimens. Many of the thick facesheet specimens did not have


readily visible dents, and the peak dents were essentially negligible. However,

these specimens have not yet been sectioned, so the internal damage state of

the core has not been revealed. More work is needed to draw qualitative

conclusions from these specimens. Figure 33 compares the relationship

between impact energy and peak dent depth for the thick facesheet specimens.

Figure 33. Impact energy vs. peak dent depth of thick specimens.

At similar energy levels, the thicker facesheet specimens show lower dent

levels than their thin facesheet counterparts, as expected. The two outlying (low

dent-level) points for the 25 degree impacts represent two tests in which the ice

crushed before impact or broke apart upon impact, and thus less energy was

absorbed by the panel. Similar situations occurred for the 25 m/s (nominal)

speed tests at 40 degrees, but the 50 m/s (nominal) speed test created

significant visible damage (crater) to the panel.

Cratered


3.4 Discussion

3.4.1 Low Velocity Blunt Impact

There exists a correlation between dent depth and impact kinetic energy

separated by tip radius. Most of the dent depths seem to overlap at lower energy

levels (i.e., no radius dependence), with the exception of cracked facesheet

cases. When higher energy levels are considered, it is clear that resulting dent

depths separate based on tip radii. In the case of a sharp tip of radius 12.7 mm,

cracking of the facesheet occurred at 4 J. Tip radius 25.4 mm did not crack the

facesheet until 14 J. All other tip radii only created a shallow dent at similar

energy levels, with a maximum dent of 1.58 mm.

The majority of dents had maximum depth below 1.0 mm. While visibility

of these dents increased with energy, there is no clear indication that it is

dependent on tip radii. It is therefore important to note that these shallow dents

are difficult to detect without careful visual inspection. Some dents are only

visible when carefully observed under oblique lighting. Dent depth relaxation may

further obscure the ability to visually detect damage.

In all cases, internal damage was present when the specimens were

sectioned. While the analysis of this damage is still in progress, relationships

between damage modes, damage extent, dent depths, tip radius, and kinetic

energy are being sought.


For ice sphere impacts, these results reflect that lightweight sandwich

panels (rudder) have low resistance to ice impacts. The presence of any visible


dent suggests there is going to be significant internal core damage as well. The

thick facesheet specimens do not show much visible surface damage but it is

anticipated that they will have significant internal damage. It is possible that since

these specimens are stiffer, the level of damage could be more variable due to

the potential dependence on the state of the hail ice on impact. Often, the hail

may break upon impact and impart less energy onto the panel than if it were to

stay intact and rebound from the surface, thus imparting the full impulse on the

panel. It is also important to note that the peak dents are not highly accurate

measurements due to the variability in the measurement process, particularly at

small depth levels (below 0.05 mm). The slightest bump or inconsistency of the

panel surface will offset the results, and thus the dent data are only significant

within a 0.02-0.05 mm range of measurement accuracy.

3.5 Sandwich Panel Impact Conclusions

3.5.1 Low Velocity Blunt Impact

Dent depth increases linearly with increasing energy levels for lower

energy levels up to 6 J. The rate of increase changes with different tip

radii, and is more evident in higher energy levels beyond 6 J.

Blunter impacts (larger radius) produce more shallow dents that exhibit

more relaxation over time. Percent relaxation tends to be low, however,

and reaches final state within 24 hours.

Composite sandwich panels with thin facesheets offer little resistance to

impact damage. Energy levels as low as 2 J produce both internal

damage with all impact tip radii investigated.



Any visible/measurable dent on the surface corresponds to significant core

damage, which ranges from core wrinkling to core fracture.

Barely-measureable dents have been found to produce significant core

damage.

With 50.8 mm diameter ice spheres, speeds as low as 50 m/s are capable

of fracturing the skin and leaving highly-visible facesheet cracking and

crater in the sandwich panel. Thicker skins allow higher impact energies

before fracturing the facesheet.

4.0 Benefits to Aviation UCSD’s research focused on impacts threats which can produce non-

visible damage have several key benefits to aviation. These are summarized

below:

4.1 GSE Blunt Impact

Understanding of damage produced from wide-area GSE impact events

through experimental data and finite element analysis. The results can

bring awareness to blunt impact damage modes and allow for prediction of

where damage will likely take place.

Demonstrated inverted building block pyramid approach to coordinated

tests that support development of accurate finite element modeling

capabilities. Large-scale experiments first identified failure locations and


modes. Then, small-sized element-level specimens were used to provide

in-depth data supporting analyses and modeling methodology

development.

Established FEA models that were capable of predicting:

o onset of cracks that lead to large-scale damage and degradation of

material/structural integrity, as well as correlating the visibility to the

internal damage produced, and

o large scale structure deformation, failure behavior, and loss of

global stiffness of substructure due to these failures.

Capabilities to model and predict damage of fuselage structure based on

impact location. GSE speed and mass are key parameters to define safe

GSE operations, and also provide guidance on what conditions would

merit further inspection for damage.

Established modeling capabilities can also be used for design of new

structures that may be more able to resist damage from commonly-

occurring impact sources

Using NDE methods to determine whether or not large damage has

occurred, to help in the decision on whether an aircraft should be removed

from service for further (invasive) inspection.

4.2 Blunt Impact Damage to Sandwich Panels

Increase understanding of the damage modes and governing mechanisms

in sandwich composite honeycomb structures when subjected to blunt

impact conditions.


Understanding of how internal damage develops provides insight into

improving the impact resistance of composite sandwich panels.

Establishing a correlation between the extent of core damage and visible

external damage allows for a prediction of damage based on external

observations.

5.0 Future Work and Follow-On Activity

The project activities summarized herein are ongoing. Future planned and

recommended activities are described below. In all areas, education/training and

dissemination of results via reports, workshops, and seminars is a key topic.

5.1 GSE Blunt Impact

• Develop FEA models that incorporate aircraft floor joints and floor beams to

better represent fuselage structure. Various impact locations will be explored

to determine how this affects development of damage.

• Studies to be performed incorporating different stringer and shear tie

geometries to establish how these components’ configurations affect blunt

impact damage formation and progression.

• Explore NDE methods for finding major damage, including frame cracks and

shear tie failures. Methods may involve using a series of sensors at various

locations where one sensor sends a signal and the others receive.

• Perform physical quarter-barrel or half-barrel fuselage test with specimen

having floor structures, to validate FEA modeling capabilities.


– full or ½ barrel – needs to include internal floors, joints, and other

structure

• relationship made to previous lab-tests – work out methodology to

account for effect of various threat types and contact locations, as

well as boundary condition and panel size effects

• internal joints affecting load paths – e.g., proximity of impact to

passenger and cargo floor or locations of high stiffness transition –

i.e., Regions 1, 2, and 3

– impact with actual GSE vehicle (or rolling-mass representative) –

incoming energy, allow to decelerate

– glancing impact effects

• investigate how contact zone moving across surface affects 1)

damage produced and 2) damage visibility (more surface shear,

scuffing marks?)

• modeling of the surface shear effects on damage formation – must

account for friction

• Blunt Impact on Other Structure Types

– metal-composite hybrid – could metal frames yield and “pull” skin in,

leaving more visible dent?

– all-metal construction (can serve as baseline particularly for dent

formation)

– aged metal structures from in-service fleet or retired aircraft – can blunt

impact increase widespread fatigue damage (WFD) state?


– sandwich construction

– non-fuselage locations – e.g., lower wing and empennage surfaces

Continued developments to establish high fidelity FEA modeling capability

– accurately predict damage initiation, progressive failure process,

damage extent, energy absorption

– accounting for interlaminar failures within context of laminated shell

element modeling approach to accommodate large-structure modeling,

also account for how interlaminar failures interact with in-plane failures

predicted by classical lamination

Define generally-applicable visibility metrics and failure criterion compatible

with FEA – focus on when crack formation is visible

Develop and refine reduced order models

– estimate damage onset for wide parameter range: GSE mass, velocity,

impact location

– relate test results to GSE field operations

Education/Training: dissemination of results, workshops.

5.2 Blunt Impact Damage to Sandwich Panels

Section test specimens and relate observations of internal core damage

depth and span to external visibility.

Relate extent of core damage to impact sources and external damage

formation.


Establish prediction capability within explicit FEA simulation framework to

predict blunt impact induced damage modes, size, and severity, with

relationship extended to post-impact residual strength reduction.

Detailed photography and microscopy of core damage modes and

morphology.

Conduct post-impact facesheet peel/fracture tests, and correlate results with

FEA. Relate to core damage modes and morphology.

Investigate effect of multi-hit and impact adjacency

Compression after impact testing of the panels tested - what is the residual

strength of panels that have experienced the types of damage that have been

observed?

Use photogrammetry to further characterize the surface dents and

deformations of the panels.

Investigate the effect of layup orientation on impact damage thresholds – how

does layup affect final delamination morphology?

Investigate impact onto stiffened skin panels and sandwich panels.

5.3 Low Velocity Blunt Impact of Monolithic Laminates

Residual strength of damaged panels. Is there a significant strength

reduction between different damage modes?

Expand experimental testing to modeling and prediction of damage.

Explore efficient and effective NDE methods to determine core damage.


7.0 References 1. International Air Transportation Association 2005, “Ground Damage

Prevention Programme Targets 10% Cost Reduction,” Industry Times, Edition 7, September, Article 4.

2. Kim, H. and Kedward, K. T., “Modeling Hail Ice Impacts and Predicting Impact Damage Initiation in Composite Structures,” AIAA Journal, Vol. 38, No. 7, 2000, pp. 1278-1288.

3. Kim, H., Kedward, K.T., and Welch, D.A., “Experimental Investigation of High Velocity Ice Impacts on Woven Carbon/Epoxy Composite Panels,” Composites Part A, Vol. 34, No. 1, 2003, pp. 25-41.

4. Rhymer, J., Kim, H., and Roach, D., “The Damage Resistance of Quasi-Isotropic Carbon/Epoxy Composite Tape Laminates Impacted by High Velocity Ice.” Composites Part A: Applied Science and Manufacturing, DOI: 10.1016/j.compositesa.2012.02.017. Available online 3 March 2012.

5. "Military Handbook 17 - MIL-HDBK-17-3F: Composite Materials Handbook, Volume 3 - Polymer Matrix Composites Materials Usage, Design, and Analysis, Chapter 7," U.S. Department of Defense. June 2002, Chapter 7.

6. Dassault Systèmes Simulia Corp., Providence, RI, USA, “ABAQUS v.6.11 Analysis User’s Manual,” Section 31.5, 2011.

7. H. M. Matt and F. Lanza Di Scalea, “Macro-fiber composite piezoelectric rosettes for acoustic source location in complex structures,” Smart Mater. Struct., vol. 16, no. 4, pp. 1489–1499, Aug. 2007.

impact damage formation on composite aircraft structures...hail ice impacts on sandwich panels...

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