Report 3762

[NASA-CR- 3762) LIGHTNI NG FHO_ECTION

GUIDELINES AND _EST D._TA fO_ ADH£SIVELY

BONDED AIRCRA_ S_RUC_URES Final Report

(iiqhtninq Technoloqies, Inc., Pittsfield)

175 p HC _08/MF A01 CSCL O|C H|/05

N86-21586

Unclas

06025
_m
NASA Contractor Report 3762

Lightning Protection Guidelines

and Test Data for Adhesively

Bonded Aircraft Structures

j. E. Pryzby and J. A. Plumer

Lightning Technologies, Inc.

Pittsfield, Massachusetts

Prepared for

Langley Research Center

under Contract NAS1-15884

N/LS/X

National Aeronautics

and Space Administration

Scientific and Technical

information Office

1984

TABLE OF CONTENTS

I 0 SU_LARY ........................

i I Introduction .....................

1 2 Lightning Protection Design Guidelines ........

1 2.1 General .......................

1 2.2 Basic Steps in Lightning Protection Design .....

2 0 EXPERIMENTAL PROGRAM ................

2 1 Purpose ........................

2 2 Test Methods .....................

2 2.1 Impulse voltage tests ...............

2 2.2 Impulse current tests ................

2 2.3 Test chamber ....................

2 2.4 Detection of electrical sparks ...........

2 3 Adhesively bonded aluminum ..............

2 3.1 Bonded aluminum lap joint specimens .........

2 3.2 Bonded aluminum fuel line brackets .........

2 3.4 Comparison of test results for various adhesives

2 3.5 Adhesively bonded aluminum honeycomb panel specimens.

2 4 Hardware Interfaces with Metals ............

2 4.1 Access doors riveted, fastened, or bonded and

fastened ......................

2 5 Adhesively Bonded Graphite/Epoxy (gr/E) ........

2 5.1 Bonded gr/E lap joint specimens ...........

2 5.2 Bonded gr/E stiffeners .............

2 6 Metal to gr/E Interfaces .............

2 6.1 Rivets in gr/E laminates .............

2 6.2 Access door dome nuts in gr/E laminates .......

2 6.3 Specimen description ...............

3 0 FULL SCALE STRUCTURES TEST PROGRAM ..........

3 1 Purpose ........................

3.2 Basic Test Methods ..................

3.2.1 Test setup .....................

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PRECEDING PAGE BLANK NOT FILMED"
TABLE OF CONTENTS- CONT'D.

3.2.2

3.2.3

3.2.4

3.2.5

32.6

Current distribution tests ..............

Bond line voltage measurements ...........

Ignition source tests ................

Induced voltage measurements in wing electrical

wiring .......................

Magnetic field measurements .............

33

33.1

33.2

33.3

33.4

Bonded aluminum structure ...............

Specimen description ................

Test setup .....................

Test procedures ...................

Test results ....................

3 4

34.1

34.2

34.3

34.4

gr/E Wing .......................

Specimen description ................

Test setup .....................

Test procedures ...................

Test results ....................

35

3.5.1

3.5.2

References

Discussion of Results .................

Comparison of the full scale structural environment

with the subelement test results ..........

Comparison of the bonded aluminum wing with the gr/E

composite structure .................

• • . ° . • • • • • • • • • • • . , • . • • • •

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iv
1.0 SUMMARY

The information presented in sections 2.0 and 3.0 of this

report were developed from electrical tests of groups of representative specimens of graphite/epoxy (gr/E) and adhesively

bonded aluminum. The purpose of the subelement program reported

in section 2.0 was to characterize the electrical performance of

typical structural bonds and joints and to determine the voltage

and/or current thresholds at which electrical sparking and loss

of physical strength occur. Then, complete, fully assembled

structures employing adhesive bonding techniques and other manufacturing concepts similar to those evaluated during the subelement test program were tested to see if spark or physical damage

threshold conditions could be reached in typical structures.

These data are reported in Section 3.0. The data from both of

these experimental programs are intended for use by aircraft

designers in avoiding hazardous lightning effects in other designs. Guidelines for use of this data and design of lightning

protection design are presented in Section 1.2.

Tests were performed on bonded aluminum and gr/E coupontype (sub-element) specimens to determine the dielectric voltage

breakdown and spark threshold levels of insulating adhesively

bonded specimens and the current spark threshold levels of electrically conductive, adhesively bonded specimens. The aluminum

specimens included lap joint, fuel line bracket, stiffener and

honeycomb sections, and fuel tank access doors, gr/E specimens

included lap joint, stiffener, fuel line feed-through elbow, and

access door dome nut specimens in gr/E laminates.

Results of the bonded aluminum subelement specimens showed

that many of the bonded specimens were electrically conductive but

conductivity was not related to bond line thickness. The electrically conductive specimens exhibited a wide scatter of current

spark threshold levels and many of the visible sparks were due to

burrs at the specimen edges and ends. Shear strength degradation

became significant at 2500 A/in 2 except for specimens bonded with

aluminum powder-filled adhesive which did not suffer appreciable

loss of strength at this current density. The addition of rivets

to the bonded specimens increased the spark threshold level by

400-900 percent above the level for non-riveted specimens.

For aluminum specimens with nonconductive bonds, a wide range

of voltage breakdown levels existed and most were due to edge

sparkovers. The highest sparkover voltages were exhibited by

specimens bonded with supported modified epoxies; the lowest readings were obtained on specimens bonded with aluminum powder-filled

adhesives. Controlled bond line specimens showed increasing bond

line voltage thresholds for increasing bond line thickness.
Contrary to expectation, no significant differences were

found in spark threshold levels for the several aluminum access

door and fastener configurations tested.

Results on the gr/E subelement specimens showed that some

of the lap joint specimens exhibited electrical conductivity

prior to test. Due to the small number of samples tested, no

conclusion could be made between adhesive used and bond line

breakdown voltage. On specimens with controlled bond lines, the

breakdown voltage increased with increasing bond line thickness.

Generally, the current spark threshold level was low for specimens which sparked at voltage breakdown and high for specimens

which had exhibited pretest conductivity. However, all of the

stiffener specimens exhibited pretest conductivity and had a low

spark threshold level.

Tests on the rivet, dome nut, and fuel line feed-through

elbows installed in gr/E laminates showed that the current spark

threshold level of these specimens was low. Tests of fuel tank

sealants as a means of containing such sparking showed that one

sealant resulted in somewhat higher spark threshold levels than

the other.

Tests on full scale bonded aluminum and gr/E wings included

measurements of current flow in individual components, bond line

voltages, induced voltages in wiring, and magnetic field levels.

Test results on the bonded aluminum wing showed that less

than 0.2% of the test current flowed through the fuel lines and

that 86% of the current was transferred from the skin to the main

spar. Low bond line voltages indicated that electrical continuity

existed through the bonding adhesive. Photographs of the fuel

tank interior during simulated strikes to the wing indicated that

no ignition sources existed within the fuel tank.

Induced voltage measurements in electrical wiring within the

bonded aluminum wing showed that circuits routed entirely within

the wing structure were relatively immune to severe induced voltages, but wiring to apparatus located in the plastic wing tip was

susceptible to induced voltages high enough to damage other electrical/electronic equipment powered from the same source. The

magnetic fields within the wing fuel tank were approximately 5%

of the level external to the fuel tank.

Current flow measurements in the gr/E wing showed that during the first few microseconds (_s) of current flow, the majority

of the current was carried by the gr/E skin and spars. After

about 20 _s had elapsed, a significant portion of the total current was carried by metal conductors, including a leading edge

conduit and the control cables, with smaller portions remaining

in the gr/E structure.
Bond line and structure voltage levels increased with increasing test current levels and increased distance between the

measurement location and the wing reference plane. However, the

gr/E skin resistivity was somewhat non-linear and decreased with

increasing current level. Voltage levels induced in wiring

routed through a metal conduit to equipment located within the

gr/E wing tip were generally low enough to be tolerated by electrical/electronic equipment associated with this wiring. One

circuit experienced levels that may require additional protection

against the effects of a severe strike.

The magnetic field amplitudes measured within the gr/E wing

interior were approximately 20% of the levels measured around

the exterior surfaces of the structure.

Measurements in the bonded aluminum wing showed that test

current levels of 7 kA to II kA resulted in bond line voltage

readings of 0.15 volts or less. Extrapolation of these measurements for a 200 kA severe lightning stroke results in a voltage

level of less than 5 volts. This level is below the 50 volt

spark threshold recorded during tests on the lap joint subelement

specimens. Fuel line currents were 160A or less during current

distribution measurements; this was higher than current spark

threshold levels measured on subelement fuel line bracket specimens (iI00A). Lack of visible fuel tank sparking during ignition

source tests, however, indicated that the amplitude of the current

flowing though any one individual fuel line bracket must have been

lower than the subelement spark threshold level.

Measurements in the gr/E wing showed that a test current of

190 kA resulted in a bond line voltage reading of 200 volts. This

is below the voltage spark threshold level of 1200 volts recorded

on the lap joint specimens during the subelement tests. Fuel line

currents, measured during application of 190 kA test currents,

reached a peak amplitude of 2 kA which was lower than the 5 kA

spark threshold level of the fuel line feed-through subelement

specimens.

I.I Introduction

The highly competitive marketplace and increasing cost of

energy is motivating manufacturers of general aviation aircraft

to achieve greater efficiency and economy through application of

advanced technologies in the design of new aircraft.

Among these are an increased use of composite materials to

obtain higher strength-to-weight performance and the use of

metal-to-metal bonding with adhesives in place of conventional

fasteners and rivets to obtain smoother outside surfaces and

reduce drag and to reduce costly hole drilling and fastening

operations. Other advantages include corrosion reduction and

extension of fatigue life.

3
Standing in the way of widespread use of some of these new

technologies, are potential problems posed by variability of

properties, impact resistance, environment effects, lightning

protection, and electromagnetic compatibility. The primary

emphasis of this program was protection from lightning effects.

The program began with a survey of the participating manufacturers to obtain information on advance-technology materials

and fabrication methods under consideration for future designs.

From the survey, typical structural designs were selected, and

these were fabricated as small "subelement" specimens which

were subjected to voltages and currents representative of the

lightning stroke. Measurements of bond line voltages, electrical

sparking, and mechanical strength degradation were made to comprise a data base of electrical properties for new technology

materials and basic structural configurations.

The second phase of the program involved tests upon full

scale wing structures which contained integral fuel tanks and

which were representative examples of new technology structures

and fuel systems. The purpose of these tests was to provide a

comparison between full scale structural measurements and those

obtained from the "subelement" specimens.

The new technologies dealt with included adhesive metal-tometal bonding; bonded honeycomb; kevlar, fiberglass and gr/E

structures; bonding and fastening techniques; fuel system internal components, and electrical systems located within the newtechnology structures.

The prime contractor was Lightning Technologies, Inc.,

Pittsfield, Massachusetts. Test specimens were provided by Beech

Aircraft Corporation, Wichita, Kansas; Lear Fan Corporation (US),

Reno, Nevada; and Cessna Aircraft Company, Wichita, Kansas. The

above-named participants, in addition to the Materials Division

and the Low-Speed Aerodynamics Division of NASA Langley Research

Center, provided engineering guidance and analyses through a

project steering group which met periodically to review test

results and plan further tests. The program period of performance

was I August 1979 through I July 1983.

This report is organized into three sections:

Section 1.0 contains the Summary, Introduction, and Design

Guidelines, which include a discussion of the basic steps involved

in lightning protection design, and some techniques for use in

anticipating possible lightning effects problem areas and protection design approaches;

Section 2.0 describes the "subelement" specimens and test

results:

Section 3.0 describes the full scale structure specimens,

presents the test results, and compares the results with those

4
obtained from the "subelement" tests. Test data from sections

2.0 and 3.0 are intended for use by designers in anticipating

lightning effects hazards in new designs.

Identification of commercial products in this report is to

adequately describe the materials and does not constitute official endorsement, expressed or implied, of such products or manufacturers by the National Aeronautics and Space Administration.

1.2 Lightning Protection Design Guidelines

1.2.1 General

In this section of the report, the basic procedural steps

in lightning protection design are reviewed and examples are

given of how to use the test data in sections 2.0 and 3.0 to

assess potential hazards or develop protection.

1.2.2 Basic Steps in Lightning Protection Design

In most cases, lightning protection designs are achieved

most efficiently if the process is begun early in the design

phase of a new aircraft and accomplished according to the following basic steps:

I. Determine the lightning strike zone(s) in which

the structure is located.

2. Establish the lightning environment.

3. Identify structural components, systems or subsystems that may be vulnerable to hazardous effects.

4. Establish the degree of protection required.

5. Proceed with protection design.

6. Verify adequacy of design by analysis or test.

The design and verification process outlined above is least

costly and results in the most efficient design if begun very

early in the preliminary design phase of a new aircraft or major

structure. Also, the design effort should be an interdisciplinary one, because lightning effects are both electrical and

physical in nature, and often interact with the structures as

well as with electrical and avionics, propulsion, fuel and other

systems contained within the aircraft. The following paragraphs

describe each of the above steps in greater detail.

1.2.2.1 Determine the lightning strike zone(s) in which the

structure is located

Lightning strikes attach initially to the extremities of an

aircraft, such as propeller tip, nose, or empennage tips. It is

extremely rare for lightning to initially strike a flat surface
such as the top or bottom surface of a wing or the center of a

fuselage; however, since most aircraft are moving forward when a

strike occurs and since most lightnin$ flashes have a lifetime

of approximately a second, the lightnlng arc channel may reattach

itself to aircraft surfaces aft of initial attachment points.

For example, a strike to the nose may reattach at spots along

the fuselage, and strikes initially reaching a propeller tip will

often reattach to the nacelle or a flap that lies behind it.

There are other surfaces of an aircraft which are not expected to

receive any strikes. These usually include the leading edges and

flat surfaces of wings and empennage structures that do not lie

behind initial attachment points such as the nose or a propeller.

These surfaces do not usually need to be designed to tolerate

direct strike effects. Nonetheless, they may lie between other

portions of the airframe which are subject to direct or swept

strokes. The lightning attachment process is described more

fully in Ref. I.

In recognition of this fact, lightning strike zones have been

defined for use by aircraft designers and regulatory authorities

in establishing the lightning environment for each major section

of an aircraft. These zones have been defined (Ref. 2) as follows:

Zone IA: Initial attachment point with low possibility

of lightning arc channel hang-on.

Zone IB: Initial attachment point with high possibility

of lightning arc channel hang-on.

Zone 2A : A swept stroke zone with low possibility of

lightning arc channel hang-on.

Zone 2B: A swept stroke zone with high possibility of

lightning arc channel hang-on.

Zone 3 : All of the vehicle areas other than those

covered by zone I and 2 regions. In zone 3,

there is a low possibility of any direct

attachment of the lightning arc channel.

Zone 3 areas may carry substantial amounts of

electrical current, but only by conduction

between some pair of direct or swept stroke

attachment points.

The location of each zone on a particular aircraft depends on

the geometry of the aircraft as well as other factors including its

intended flight envelope and airspeed. The zone locations for a

particular aircraft are usually established by comparison with

inflight experience of existing aircraft of similar shape or by

tests in which a scale model is subjected to simulated lightning

strikes applied from a variety of directions. Further guidance

on location of strike zones is available in Refs. 3 and 4.
Although some of the surfaces of an aircraft are not within

a direct strike zone (zone IA or 2A), nearly all internal structures are located within zone 3 because they lie between extremities. Thus, even though an aircraft skin may not need protection,

the complete structure which is comprised of the skins plus all

internal structural members, must be designed to safely conduct

zone IA currents (usually SAE current components A and C) from

one end to the other, or perhaps from one surface to another.

For riveted metal structures this requirement was easily met

with little or no specific changes in design to accommodate the

lightning current. However, adhesive bonds and composite structures may be susceptible to electrical sparking or loss of strength

when subjected to the same current levels unless specific measures

are employed in the design to counter these effects. Thus, it is

important that the lightning strike zones be located before assessment of potential lightning hazards or protection design methods

are begun. All parties involved inthe aircraft design and certification process should be aware of and understand the location(s)

of these zones.

1.2.2.2 Establish the lightning environment

In this section the actual amplitudes and other important

characteristics of lightning currents are established for the

surfaces and structures within each zone. If the criteria in

references 5 or 6 are followed, this step follows directly

from step one since the environments for each zone are defined.

The design process then becomes the determination of the

percentage of the total environment current that may be expected

to pass through a specific structural component or across ajoint

of interest. For example, if an integral fuel tank within a wing

is comprised of the upper and lower wing skins and three longitudinal spars as shown in Figure I.I, and if the skins are to be

adhesively bonded to the spars, the designer will have to estimate

the amount of current expected to flow across the adhesive bond

from the skin to each spar.

The calculation of electric current flow through structures

as complex as an airframe is extremely difficult to accomplish

without extensive computational facilities. In addition, considerable uncertainties surround the authenticity of even the most

sophisticated current flow models presently available, and this

aspect of lightning phenomenology remains the subject of research.

Therefore, the designer must resort to simple estimation methods

and considerable reliance upon test data (when available) to

determine the electrical environment. This task is made possible

by reliance on several "rules of thumb" as follows:

I. For metal structures, lightning currents flow initially

on the outermost conductors, which usually include outer skins

and leading edges, and diffuse more slowly into internal structural members such as spars and stringers. The current flow in

an aluminum spar within an aluminum wing might appear as shown in

Figure 1.2. 7
OF pOOR QUALITY

Wing Skin Current_

Entering Spar

through Adhesive

Bond

/

/

/

/

/

// 7

/////

/ /

/ //

/

/ //

//

//

St

0

ightning Current

Entering Wing Skin

Figure I.I - Current Flow from Wing Skin to

Spars through Adhesive Bonds.

Total

Current

I

I I

2ps 50_s

Current

in

Spar

I

I

// I

-- I

_30Vs

8

Figure 1.2 - Current Flow within an Aluminum Wing.

• Top Waveform - Total Current in Wing

• Bottom Waveform - Current in Spar
2. Electric currents diffuse much more rapidly throughout

gr/E structures so the lightning current may be expected to flow

into internal structural members (such as spars) at the same rate

as in external skins.

3. During the first few microseconds of current flow, current divides inductively. At later times, after a few tens of

microseconds have elapsed, currents divide in accordance with the

relative conductivity of structural members. For example, in a

gr/E structure containing some metallic items such as electrical

conduits or control cables, current flows through all paths in

accordance with their relative inductances (i.e. longer paths

have higher inductances and thus carry lower currents than shorter

paths) but at later times, the current will seek paths of lowest

resistance and thus begin to concentrate in the metal _conductors.

This characteristic is illustrated in the measurements of current

in the metal components within the gr/E wing, reported in paragraph 3.4.4.1. For estimating purposes, current division through

a composite structure may be considered inductive during the first

ten microseconds and resistive thereafter.

4. Current division through any structure is also influenced by the strike attachment point (ie., the location(s) where

currents enter and leave the structure). Since lightning currents

usually enter or leave a structure at single points, the currents

will become much more concentrated in the vicinity of these points

than elsewhere in the structure. Thus, nearby joints, bonds, and

materials usually must be designed to tolerate higher current densities than similar interfaces further away from current entry

or exit points. Current concentration factors between four and

ten should be applied in the vicinity of possible current entry

or exit points.

5. Smaller structures will always have to tolerate higher

current densities than larger airframes, since the total lightning

current is the same regardless of aircraft size. The goal of

spark or physical damage prevention within a general aviation

aircraft structure is, therefore, usually more difficult than for

structures of larger transport category aircraft.

With the above guidelines in mind, it is possible to make a

rough assessment of the currents that must be tolerated in a typical structure. An example of how this may be done is as follows:

A. Determine the possible direction(s) of lightning current

flow, based on the location of the structure within an aircraft.

B. Determine the cross-sectional areas of each of the conductive elements of the structure. This will included the upper

and lower skins, main spars, and any longerons or stringers, plus

any nonstructural components such as metal conduits or control

cables.

9
C. Determine the division of lightning current among the

elements. If the length of all of the above paths are approximately the same, the lightning current will initially divide

among the elements in proportion to their relative cross-sectional areas.

D. Determine the electrical resistivities of each of the

conducting paths, and calculate the resistances of each path or

conductor by the following expression:

pL

R -

A

where:

R = the total resistance of the conductor - ohms

p = the resistivity of the conductor material -

ohm-in. (ohm-cm)

L = length of the conductor - in. (cm)

A = conductor cross-sectional area - in_ (cm 2)

Electrical resistivities of typical airframe structural

materials are presented in Table I-I (Ref. 7).

Table I-I - Electrical Resistivities of Typical

Airframe Structural Materials

Material

Resistivity

a-in. xl0 -_ _-cmx 10 -6

Aluminum I. I (2.8)

Copper 0.67 (I. 7)

Titanium 17. (42.)

Stainless 28. (72.)

Steel (304)

Magnesium I. 8 (4.5)

gr/E 2000. (5000.)

The total resistance of any structural path, of course, is comprised of joint and bond resistances as well as the resistances

of the bulk structural materials themselves. In some cases, joint

or bond resistance can be higher than the resistance of the

structural member itself. Bulk structural resistances are readily

calculated, but joint or bond resistances are very difficult to

estimate and must usually be determined by test. These latter

resistances are also dependent on the amplitude of current flowing

I0
through them. Tests on candidate joint and bond configurations

should be performed to obtain these resistances. If such data is

available it should be included in the resistance calculation.

If not available, these resistances will have to be neglected

during the initial estimates.

Examples of the magnitude of structural resistance (including joint and bond resistance) which may be expected in bonded

aluminum and gr/E wings are recorded in paragraphs 3.3.4.4.2 and

3.5.2.1. For the bonded aluminum wing this value was 0.3 m_;

for the gr/E wing the value was 25 m_.

E. Estimate the lightning current in each structural element

at later times (after a few tens of microseconds), when current

flow is inversely proportional to the resistance of each structural member or other conductor. The following expression should

be used for this purpose:

iTR T

il -

RI

where:

i I = Current in a structural element or other

conductor of interest (amperes)

iT = Total lightning current in entire structure

(amperes, usually 200 kA)

R T = Electrical resistance of entire structure

along direction of current flow (ohms)

R I = Electrical resistance of structural element

or other conductor of interest (amperes)

In most cases, it will be found that current distribution based

on element resistances is different than distribution determined

by comparative inductances. This would seem to necessitate current

flow from one conducting element to another and this is actually

what happens. The result may be a concentration of currents at

interfaces between gr/E and metal conductors. Such current concentrations may produce sparking or loss of strength at gr/E-tometal interfaces, presenting difficult protection problems.

These probems can sometimes be avoided by breaking or eliminating

the current path through the metal conductor. Within a fuel tank,

for example, metal fuel or vent lines can be replaced with tubes

made of nonmetallic materials, or provided with nonmetallic inserts

to prevent current flow.

ii
F. Design structural materials, joints, bonds and other

critical interfaces to tolerate the highest current estimated for

each structural element. This may be the current determined by

either inductive or resistive distribution.

G. Determine if concentration factors are needed. At locations remote from lightning entry or exit points, the structural

element currents for design purposes may be assumed to be as

estimated by the foregoing procedures. However, in the vicinity

of the lightning entry or exit points, the estimated current

densities should be increased by a "concentration factor". For

structures directly beneath attachment points, such as a skin-tospar bond, the estimated current density should be increased by

a multiple of ten. In other cases, the multiple may be as low

as four. In cases where a higher degree of doubt exists, the

highest multiple should be applied.

H. Determine if spark or physical damage thresholds will

be exceeded. Refer to the spark or physical damage threshold

data presented in section 2.0 of this report to determine if

these thresholds are likely to be approached or exceeded at similar interfaces within the structure being designed. For example,

if adhesively bonded gr/E spar-to-skin joints are planned, the

data of paragraph 2.5.1.2.3. may be applicable. These data indicate that sparking occurs at current densities of 248 A/in. 2 or

greater. An example of how these data may be used in aircraft

design is shown in paragraph 3.5.1.2. In this example, the

average current density through the spar caps of the tested gr/E

wing was calculated to be 75 A/in. 2 T_hen current distributions,

as shown in Figure 3.36, were factored into the calculations, it

was found that the average current density through the spar caps

was approximately 55 A/in. 2 or 25% of the level which may cause

sparking as indicated by the subelement test results. If the

calculations had indicated that the current densities would

exceed 248 A/in. e, then additional protection methods would need

to be considered.

If the data of section 2.0 are not applicable to particular

structural designs, similar tests of appropriate subelement

specimens representative of the structure should be performed

in the same manner as described in section 2.0.

1.2.2.3 Identify structural components, systems or subsystems

that may be vulnerable to hazardous effects

The next step in the lightning protection design process is

to identify structural components and other associated systems,

subsystems or components that might be vulnerable to damage,

electrical sparking, or other adverse effects from the lightning

environment determined in the second step, paragraph 1.2.2.2.

12
A study of section 2.0 and 3.0 will illustrate particular areas

of concern. Some of these areas are:

Airframe and Structures

• Puncture of skins

• Delamination of composite skins and other structural members

• Adhesive debonding

• Crimping due to magnetic force effects

Fuel Systems

Electrical sparking at rivets, removable fasteners, nut plates,

and other types of fasteners. This hazard will be found most

often in gr/E structures where current concentrations near

metal fasteners are high; however, sparking has also been

found to occur at rivets and fasteners within all-metal fuel

tank structures. The reason is that most metal components

are treated with primers or other corrosion-resistance surfaces

prior to assembly, thus forcing lightning currents to flow via

rivets or fasteners.

Electrical sparking at fuel tank plumbing components such as

flexible couplings, bulkhead fittings, support clamps and at

other joints not specifically designed to tolerate lightninglike currents. Here again, this hazard will be most prevalent

in plumbing contained within a composite structure; however,

it also may occur within all metal structures.

Electrical sparking at adhesively bonded joints. This hazard

will appear most often in locations where no rivets or fasteners are used in either composite or all metal structures. The

ability of a rivet to increase the spark threshold level of an

adhesively bonded joint was shown by the test results on aluminum lap joint specimens reported in paragraph 2.3.1.2.3. The

results showed that the addition of a rivet to the lap joint

increased the spark threshold level from I kA to 5 kA.

Incendiary Sparks. The minimum electrical energy required to

produce an incendiary spark has been found to be 0.2 millijoules

(Ref. 8). Although the amount of current expected to flow in

an individual structural element can be estimated by the

methods described in paragraph 1.2.2, the amount of energy that

might be released in a spark between two structural elements

would be extremely difficult to estimate. The minimum spark

energy is, however, equivalent to the energy released by a current of several amperes flowing through a spark of several

microseconds duration. Since most structural elements will

experience currents of several hundred or several thousand

13
amperes, the possibility of at least one or two of those

amperes flowing in any spark that might occur must be considered a certainty.

Hot Spots. In addition to sparking, hot spots of sufficient

temperature and duration to ignite flammable vapors may be

formed. These hot spots would be formed by direct lightning

strikes to integral tank skins. It has generally been found

that the melting temperature of aluminum (1200°F) will be

reached before the short-time ignition temperature (2400°F)

of fuel air vapor is reached. This, however, may not be true

for composites. No work on this hazard was conducted during

this program.

Sparking at electrical apparatus such as fuel quantity probes

or float switches due to lightning-induced voltages in associated electrical wiring.

Electrical and Avionics

Surge voltages due to structural IR voltages that arise during

lightning current flow, appearing in electrical circuits

referenced to the airframe at two or more locations. This

hazard is most prevalent in composite structures whose resistances are comparatively high. gr/E wing resistance, calculated from data obtained during the full scale structures tests,

was found to be approximately 25 m_. Thus, a full threat

200 kA lightning strike could result in a voltage potential

as high as 5 kV near the wing tip. Electrical circuits routed

along these structures will "see" these IR voltages, which are

capable of driving substantial currents through electronic

equipment located at the ends of the circuits.

Surge voltages induced in electrical circuits by changing magnetic fields associated with lightning currents. This hazard

occurs most often in circuits that pass adjacent to apertures

in the airframe, such as windshields or along wing trailing

edges. Electric wiring that is completely contained within

metallic structures is least susceptible to this hazard especially if the structure is assembled with rivets. Circuits

within gr/E structures are more susceptible to magnetically

induced effects, but still less so than those exposed to large

apertures. A comparison of the shielding effectiveness of aluminum and gr/E was demonstrated during magnetic field intensity

tests on the bonded aluminum and gr/E wings. The ratio of the

magnetic field intensity measured within the bonded aluminum

wing to that of the gr/E wing was calculated in paragraph

3.5.2.2 and found to be approximately 4:1. Additional data

that shows the relative effectiveness of metallic and nonmetallic shielding is presented in paragraph 3.3.4.4.3 and

Table 3-5. These data show the voltages induced in electrical

14
wiring, located within the bonded aluminum wing, as a result

of a simulated lightning strike to the wing tip. Voltages

induced in wiring located entirely within the aluminum skin

ranged from 15V to 60V for test currents of 7 kA. Voltages

induced in navigational light wiring, which was partially

routed through the plastic wing tip, ranged from 2.9 kV to

4.1 kA for test currents of 7 kA. These higher voltage levels

resulted from a lack of electromagnetic shielding provided

by the plastic wing tip, thus allowing greater magnetic flux

interaction with these wires than for those located entirely

within the shielded environment of the aluminum skin.

Transient voltages "induced" in aircraft electrical wiring.

These voltages are a combination of magnetically induced and

structural IR voltages. In gr/E structures, the IR voltages

usually predominate; in metal airframes, the magnetically

induced voltages are usually highest.

Electrical wiring installed within nonconducting composite

structures such as those made of fiberglass or kevlar reinforced composites, may itself become the sole lightning conductor, resulting in puncture of the structure, explosion and

severe electrical damage to associated electrical or electronic

equipment. The effects of a lightning strike to a navigation

light mounted in a fiberglass wlng tip can be very destructive,

for example, unless protective measures are applied

Whether the above hazards, or other potential hazards actually

exist depends on the details of the specific design. The data of

sections 2.0 and 3.0 are intended as useful references against

which to compare new designs; however, development testing of more

representative specimens specimens is often advisable. Factors

such as adhesive type, bond line thickness, surface preparation,

edge treatments, and type and locations of fasteners have significant impact on spark and physical damage thresholds. Thus,

especially when new materials, bonding of fastening techniques

are being considered for critical structural or fuel tank applications, development testing should be conducted. The types of

tests require are similar to those described in section 2.0.

1.2.2.4 Establish the degrees of protection required

Once the potentially hazardous effects have been defined for

particular structures or systems, the extent of the desired protection can be determined. If the consequences of the lightning

effect would result in a flight safety hazard (i.e. ignition of

fuel vapors), protection design will be required. However, other

situations often exist, in which the lightning effect would not be

15
a hazard to flight safety but could result in an expensive repair

or partial impairment of aircraft performance. In these cases, a

decision might be made to forego protection, expecially if such

protection would incur a weight or cost penalty. Therefore,

proper completion of this step usually requires that alternative

methods be assessed, and costs, weight, and performance penalties

be determined. On more complex systems, trade-off studies are

also performed.

1.2.2.5 Proceed with protection design

Once the need for protection has been established, work may

progress on designing specific protection measures. Here again,

development testing is often appropriate to evaluate the effectiveness of various alternatives. Several alternatives may be appropriate for each hazard. The total number of possibilities is very

large and limited only by the imagination of the designer. Brief

descriptions of some common approaches are:

Improvements in electrical bonding at joints and other interfaces by conductive additives in adhesives or sealants, use

of rivets, conductive inserts within adhesive bonds or external

bond straps.

Elimination of electrical sparking by breaking or altering

current flow paths through structures, improvement of electrical

bonding, or design of highly conductive, spark-free interfaces.

Containment of sparking, by encapsulating or covering spark

sources with fuel tank sealant or other material capable of

withstanding spark pressures.

• Provision of metallic ground planes or electromagnetic shielding for wiring harnesses.

• Provision of surge suppression devices at the terminals of

electrical/electronic equipment.

Additional information on lightning protection devices is contained in Ref. 9.

1.2.2.6 Verify adequacy of design by analysis or test

The final step in the lightning protection design process is

verification of design adequacy. This can be accomplished by

simulated lightning tests of production-like structures, systems,

or components; or by analytical procedures which may be combined

with development tests. Qualification or certification testing is

most often used because it provides the most direct assessment of

performance. Typically, voltage or current waveforms of the stroke

components appropriate to the zone(s) in which the structure or

component is located are applied, and the condition of the

16
structure is carefully assesed afterwards. The test object may

range from a flight critical component such as an access door

installation or a wing tip assembly to a complete structure such

as a wing containing an integral fuel tank. Due to the complex

nature of many structures, it is not possible to verify a sparkfree design without subjecting the entire structure to full

scale lightning currents. Similarly, the response of internal

electrical wiring to magnetic fields and structural IR voltages

is dependent on characteristics of the complete structure (or

possibly the complete airframe); therefore the complete structure

must be represented in the tests. Tests similar to those described in section 3.0 are often applied.

There are other cases in which testing is not practical, or

perhaps not necessary. These include situations where the design

is sufficiently similar to previous designs formerly qualified by

test, or structures that are so large as to make testing of them

very difficult or impossible. In such cases, verification is

achieved by combining the results of development tests on critical

component specimens with analysis of current flow patterns or

other effects in order to achieve a high confidence in design

adequacy. An example of this latter procedure might be a wing

structure in which current densities have been estimated by analysis, followed by individual or component tests of representative

adhesively bonded joints and other potential spark sources. To

account for possible errors in the current flow analysis, test

current levels are usually set higher by a factor of approximately

100%.

Whichever approach is used, it should be carefully planned

in advance to address, as a minimum, all flight-critical lightning

protection designs. It is often appropriate to review the intended

verification plan with airworthiness regulatory authorities prior

to commencement. The verification plan should include definition

of the pass/fail criteria applicable to each test. For fuel tanks,

passage will be evidenced by lack of sparks or other ignition

sources. For other structures or systems, passage of the verification test will be established by some tolerable degree of damage.

Frequently, data from development tests conducted during

earlier phases of the lightning protection design program can be

used for verification purposes, if the development test specimen(s)

is sufficiently representative of the final design(s), and if the

development tests have been sufficiently documented. This is an

excellent way to reduce the cost of design verification.

17
2.0 EXPERIMENTALPROGRAM

2.1 Purpose

The purpose of the experimental program was to determine

the effects of lightning-like currents on typical bonded aluminum and gr/E aircraft structures, and acquire sufficient electrical effects data from which to develop guidelines for

protection of these structures against physical damage or

electrical sparking that could ignite flammable vapors.

The first phase of the experimental program included electrical tests of small subelement specimens representative of

adhesively bonded metal and gr/E interfaces that occur throughout adhesively bonded aluminum or gr/E structures. Emphasis

was on interfaces found in integral fuel tank construction,

where small electrical sparks may be a hazard. Where possible,

the subelement specimens selected were identical to mechanical

strength test specimens already in use so that adhesive bond

strength degradation due to electric current flow could be compared with adhesive strength data bases already in existence.

Other subelement specimens included typical hardware interfaces

with aluminum and gr/E skins, because these interfaces are often

in the paths of lightning currents flowing in fuel tank plumbing,

control cables or other electrically conductive parts. The types

of subelement specimens that were tested are listed in Table 2-1.

Table 2-1 - Subelement Specimens

Adhesively Bonded Aluminum:

Bonded aluminum lap joint specimens:

with varied adhesives

with varied bond line thicknesses

Bonded aluminum lap joint specimens:

with one rivet

Bonded aluminum fuel line brackets

Bonded aluminum stiffeners

Bonded aluminum honeycomb panels

Hardware Interfaces with Metals:

Access doors riveted, fastened, or bonded

and fastened

18
Table 2-1 - Subelement Specimens (cont'd)

Adhesively bonded gr/E:

Bonded gr/E lap joint specimens:

with varied adhesives

with varied bond line thicknesses

Bonded gr/E lap joint specimens:

with one rivet

Bonded gr/E stiffeners

Metal-to-gr/E interfaces:

Rivets in gr/E laminates:

covered with fuel tank sealant

uncovered

Access door dome nuts in gr/E laminates

Fuel line feed-through elbows in gr/E laminates

Basically, the subelement test program included impulse

voltage tests to determine the voltage required to break down

nonconductive adhesive bond lines or cause visible sparks to

appear, and impulse current tests to establish physical damage

and spark thresholds due to current flow across adhesive bonds

or metal-to-gr/E interfaces. The test methods utilized are

described in the following paragraph. Subsequent paragraphs

describe each type of specimen and the test results in detail.

2.2 Test Methods

2.2.1 Impulse voltage tests

Voltage breakdown tests were made on all of the subelement

specimens that were adhesively bonded with electrically nonconductive epoxies. The tests were conducted by applying a rapidly

rising voltage between both sides of the bonded specimen and

recording the voltage level reached when sparkover or breakdown

occurred. The test circuit and a typical voltage oscillogram

are shown on Figure 2.1

The test voltage was measured by a i00:i resistance voltage

divider and recorded by a cathode ray oscilloscope (CRO). The

generator charging voltage was set to provide a constant rate of

rise to peak of several thousand volts, well above the region

where breakdowns were expected. For most of the tests, the rate

of voltage rise was set at 600 volts per microsecond, which is

typical of the rate of voltage rise that may be produced throughout a gr/E structure by rapidly rising lightning stroke currents.

19
3mH

i000_

0.05 uF

0.0047 uF

T 1-1'

/

/

Subelement

specimen

ORIGINAL PACE ;_

OF POOR QUALITY

Voltage

divider

oscilloscope

O_

Test Volta e

dv/dt = 600V/_s

for 2.5 kV charge

400V/div 1.0_s/div

Figure 2.1 - Test Circuit for Applying a Controlled Voltage

Rate-of-Rise to Specimen for Determining Adhesive Breakdown Levels.

2.2.2 Impulse current tests

Impulse current tests were made on all of the subelement

specimens that were adhesively bonded with either conductive or

nonconductive epoxies, the access door specimens, and the metalto-gr/E interfaces. The purpose of the tests was to determine

visible spark and physical damage thresholds. Most of the bonded

specimens were tested at currents of 20 kA or less, whereas

riveted specimens and the complete fuel tank access door and skin

panel specimens were tested at currents of up to 200 kA. The

test circuit utilized to produce the impulse currents is shown

in Figure 2.2 together with a typical current oscillogram. The

circuit parameters shown produced the 3kA, 2 x50 Us current

pictured in the oscillograms. Different parameters were utilized

to produce currents higher than this level.

The 2 x50 _s (2 Hs to crest, 50 _s to 50% of crest on decay)

waveform is most representative of a natural lightning stroke

current, but could not be produced at peak currents higher than

about 15 kA because of the circuit resistance required. Therefore, for testing beyond this level, a unidirectional waveform

20
ORIGINAL '_'"_' ?J

OF POOR QUALI't'(

of shorter duration (8 x 20 _s) was utilized. To produce currents

above 50 kA, more resistance had to be removed from the test circuit, necessitating a damped sine waveform of approximately 25

kilohertz (kHz) frequency. Oscillograms of the three test waveforms utilized are shown in Figure 2.3.

/_ 3. O_

switch

cu

20 uF

i0 kV /_ ,

_ Transformer Current t///

to current /

measurement

oscilloscope

Subelement

specimen

4 turns

on

2 indlamet er

Voltage

divider

"_"Dk _' , _ tomeasurement voltage

Front Tail

i = IkA for

3kV of charge

I kA/div 1.0_s/div ikA/div 10_s/div

Test Current through Specimen

Figure 2.2 - Test Circuit Parameters for Generating the

2 x 50 _s Simulated Lightning Current Pulse.

During some of the impulse current tests, the voltage across

the bond line was measured, utilizing the same voltage divider

and CRO that was used for the impulse voltage tests.

2.2.3 Test chamber

Most of the subelement specimens were tested with a 2 ft. x

2 ft. x 6 ft. wooden, light-tight test chamber constructed for

this program. The chamber is pictured in Figure 2.4.

The chamber was fitted with a removable end panel to enable

arc entry tests of fuel tank access door specimens, and threaded

steel rods through opposite side walls to provide support and

electrical connections to smaller specimens suspended inside.

21
20 _s/div

2x 50_s

5 kA/div

i0 _s/div 5 kA/div

8 x 20 _s

20 _s/div 50 kA/div

0cillatory

25 kHZ

Figure 2.3- Test Current Waveforms.

22
ORIGINAL. PAGE PJ

OF POOR QUALITY

Figure 2.4 - Light-Tight Chamber in which Subelement

Specimens were Tested.

23
These test voltages or currents were conducted into and out of

these specimens via the steel rods. The light-tight chamber was

also fitted with a fiber-optic light source to serve as a reference light during photographs of electrical sparking and hinged

access doors to provide access to cameras and test specimens.

2.2.4 Detection of electrical sparks

Photography was selected for detection of electrical arcs or

sparks on the test specimens. ASA 2000 speed polaroid-type film

exposed through an F 4.7 lens opening has been shown to be capable

of detecting a spark of 0.2 millijoule which has been established

as the lowest spark energy capable of igniting a flammable vapor

(Ref. I0). At the beginning of the subelement test program,

cameras with polaroid-type 3000 speed film were utilized. Later

in the subelement test program, a 35 mm camera with 400 speed

film was added, to achieve improved definition of the spark locations and background. When exposed at a lens opening of F 3.5

and developed to 1600 speed with special developing procedures,

the film sensitivity is equivalent to polaroid-type 3000 speed

film exposed at F 4.7.

To assure that incendiary sparks could be successfully photographed, a series of tests were performed to determine the minimum

energy necessary to ignite a flammable fuel-air mixture, with a

spark produced between electrodes typical of the component interfaces found within an aircraft structure, and to verify that a

spark of this magnitude could be identified on exposed film.

Tests were performed with a set of flat electrodes cut from 0.5 in.

wide, 0.020 in. thick aluminum bar stock as shown in Figure 2.5.

Figure 2.5 - Spark Electrode Configuration.

In this respect, the tests differed from those reported in Ref. I0

in which point-tipped electrodes were utilized.

To create a spark, the electrodes were moved toward each

other until the gap breakdown voltage was exceeded.

24
Since a spark produced by a capacitor charged to 0.2 x10 -3

joules of energy has been previously established as the minimum

energy necessary to ignite a flammable vapor, a test circuit

capable of storing and discharging this much energy was arranged.

The circuit is shown in Figure 2.6.

DC

power

supply

_/_

0-10kV 0. 001 _F

-]-

Figure 2.6 - Test Circuit to Produce Small

Electrical Sparks.

A charging level of 632.5 volts results in a stored energy of

0.2 x10 -3 joules (0.2 millijoule).

For the ignition tests, a stochiometric mixture of I00

octane aviation gasolene and air at room temperature was placed

in a sealed explosion chamber. 0.2 millijoule electrical sparks

were produced by a 0.001 _F capacitor charged to 632.5 volts.

Sparks of higher intensity were produced by charging the capacitor to higher voltage.

Tests were begun with the capacitor charged to produce 0.2

millijoules. It was not possible to ignite the fuel with 0.2

millijoule sparks, so the tests were repeated at successively

higher spark intensities until ignition occurred. Three tests

were conducted at each energy level prior to moving to the next

higher level. The capacitor charging voltages stored energies,

and test results are presented in Table 2.2.

Table 2-2 - Spark Ignition Test Results

Test

No.

I

2

3

4

5

6

Charging

Voltage

632

1250

2000

3000

5000

7500

Stored

Energy

(millij oules)

0 2

0 8

2 0

4 5

12 5

28 1

Test Results

No ignition on three sparks

No ignition on three sparks

No ignition on three sparks

No ignition on three sparks

No ignition on three sparks

Ignition on third spark

25
ORIGINAL PAGE |_

OF POOR QUALITY

The failure of the 0.2 millijoule spark to ignite the fuelair mixture is probably due to the shape of the electrodes,

which were flat and thus presented a comparatively large heat

sink to the spark and the forming kernal of flame, thus quenching

it. Laboratory tests in which 0.2 millijoule was capable of

igniting a flammable vapor have always been performed with thin,

needle electrodes of low thermal mass. The flat-electrodes selected for investigation were more representative of the thermal

capacity that might exist among structural parts.

Spark tests were continued to establish that light from a

small spark could be identified on photographic film. The light

from a 2.0 x10 -3 joule spark at a distance of 4 ft from the

camera is shown on Figure 2.7. The camera was placed 4 ft away

from the specimen in the light-tight box during all of the subelement tests in Phase I.

upper: direct

lower: reflected

from mirror

behind spark

Figure 2.7 - Light from 2.0 x 10 -3 Joule

Spark. Camera is 4 ft from

light source. Polaroid-type

3000 film photographed via

F8.8 lens.

All of the subelement tests in which sparking thresholds were

to be determined were conducted within the light-tight chamber

pictured in Figure 2.4. Detailed descriptions of each type of

test speciemn and the results obtained are presented in the following paragraphs.

2.3 Adhesively bonded aluminum

Adhesively bonded aluminum specimens included aluminum lap

joint specimens as are commonly used to evaluate adhesive shear

26
strength; "T" section stiffeners and aluminum fuel line brackets

adhesively bonded to aluminum skins; sections of typical adhesively bonded aluminum honeycomb panels; and specimens of

riveted, fastened, or bonded and fastened, aluminum skin panels

and access doors. The latter specimens included typical rivet

and fastener installations coated with fuel tank sealant. All

of these specimens are representative of configurations being

utilized in new aircraft design.

Groups of the adhesively bonded specimens were tested with

each of eight different commercially available structural adhesives so that possible relationship(s) between adhesive properties and electric current transfer characteristics could be

determined. Two of the adhesives were partially electrically

conductive due to incorporation of aluminum powder. The rest

were electrically nonconductive and no attempt was made to

improve their electrical conductivity. Both supported and unsupported type adhesives were used. Supported adhesives utilize

a carrier material, such as a nylon or polyester woven scrim,

for bond line control or enhancement of bond line strength in

thick bond lines. Unsupported adhesives lack a carrier material

and may be used when low moisture absorption or good filleting

action are required.

2.3.1 Bonded aluminum lap joint specimens

2.3.1.1 Specimen description

The bonded aluminum lap joint specimens were fabricated of

0.060 in. aluminum with an overlap of I in. as shown in Figure 2.8.

The aluminum adherends were primed and bonded together with a

single layer of adhesive in accordance with the manufacturers'

recommendations. Most of the adhesives met the requirement of

US Military Standard MIL-A-25463 (Ref. ii) and US Federal Specification MMM-A-132 (Ref. 12).

Groups of seven specimens plus an untested control specimen

were cut from single lap joint panels. For each adhesive, four

such panels were prepared so the number of specimens available

for electrical tests was 28. The aluminum adherends were vapor

or alkaline degreased, acid etched, rinsed and dried prior to

priming and bonding. The characteristics of each adhesive and

primer are described in Table 2-3.

All but two of the adhesives were electrically nonconductive;

two of the adhesives exhibited partial electrical conductivity

due to aluminum powder incorporated in the epoxy. The purpose

of the aluminum powder was to maintain adhesive properties at

elevated temperatures rather than to provide electrical conductivity.

27
ORIGIt'_,_-,L PAGE 19

OF POOR QUALITY

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29
Bond line thicknesses ranged from a minimum of 0.001-

0.002 in. to 0.008-0.010 in. for the unsupported adhesives.

Supported adhesives had bond line thicknesses ranging between

0.002-0.006 in.

2.3.1.2 Test results

Prior to test, each lap joint specimen was given an electrical continuity check with a battery-operated ohmmeter to see

if physical contact existed between aluminum adherends. The numbers of specimens in each group of 28 which exhibited conducttivity prior to test are presented in Table 2-4.

Table 2-4 - Electrical Continuity of Aluminum

Lap Joint Specimens Prior to Test

Adhesive

Designation

AF-126-2

FM i000

EA 9602.3

R 7114

MB 1113

FM 61

HT 424

FM 400

Nominal

Bond Line

Thickness

(in.)

0.002

0 001-0.002

0 002-0 004

0 002-0 004

0 003-0 005

0 005-0 007

0 006-0 008

0 008-0 010

No. of

Conductive

Specimens/

Group Type

11/28 supported

1/28 unsupported

2/28 supported

1/28 supported

18/28 supported

2/28 supported

10/28 supported

5/28 supported

There was no apparent relationship between bond line thickness and occurrence of pre-test conductive paths, since even a

thin layer of adhesive may electrically isolate the adherends.

Even the aluminum-loaded adhesives appeared non-conductive in

some specimens at the low voltage (several volts) applied by the

battery-operated meter. The conductivity exhibited by some of

the specimens was undoubtedly due to burrs left on the edges of

the adherends by cutting operations.

30
The specimens that were nonconductive prior to the start of

tests were given voltage breakdown tests as described in Para.

2.2.1 to determine the voltage necessary to cause sparkover

across or through the adhesives. Those that exhibited continuity

would, of course, require no voltage to initiate conductivity.

Hence, these specimens were tested initially with impulse currents as described in Para. 2.2.2 to determine the amounts of

such currents that may produce visible sparks.

2.3.1.2.1 Bond line breakdown voltages. - Breakdown voltages of

bonds made with non-conducting adhesives ranged between 50 and

8 000 volts. In general, the thicker bond lines produced the

highest breakdown voltage, but the relationship was not well

defined due to the presence of burrs or other imperfections at

the edges. For this reason, most of the breakdowns were visible

as sparks at the edges of the specimens. Only 10% of the specimens indicated breakdown within the bond line, and in these

cases sparks were not visible.

Following these tests, a second set of specimens was prepared with controlled bond line thicknesses. The burrs were

also removed from these specimens to eliminate this source of

sparking. This group of specimens was identical to the previous

group except that only AF-126-2adhesive was used. None of the

specimens exhibited electrical conductivity prior to the tests.

The results showed that breakdown voltage increases with bond

line thickness, as shown in Table 2-5.

Table 2-5 - Breakdown Voltage vs. Bond Line Thickness

(AF-126-2 adhesive only)

Bond Line Thickness

in.

Breakdown Range

0.0035 1200 - 2400 volts

0.007 2400 - 2800 volts

0.010 3600 - 4700 volts

0.015 4000 - 4800 volts

0.020 4000 - 4800 volts

For thin bond lines, the breakdown voltage is about 500 volts per

mil (0.001 in.) thickness. For thicker bonds, the voltage required to cause sparkover diminishes to about 200 volts per mil.

All breakdowns occurred at the edges of the specimens, indicating

that the voltage withstand capability of the adhesive is greater

than that of air.

31
2.3.1.2.2 Spark threshold due to current flow.- Following the

breakdown tests, all lap joint specimens were subjected to

impulse currents as described in Para. 2.2.2 to establish the

minimum currents which would cause visible sparking. Since

many of the lap joint test specimens were nonconductive to

start with and breakdown occurred at the edges, visible sparks

occurred upon initial application of generator voltage and no

appreciable current flow was necessary to produce sparks.

These specimens were therefore eliminated from the current flow

tests. Thus, specimens given current flow tests included those

with partially conductive adhesive (FM 400 and HT 424) and

those with nonconductive adhesives in which breakdown has occurred within the bond.

Spark threshold ranged from less than 100A to greater than

I0 000 A as shown in Table 2-6 (the highest level applied). The

wide range of spark thresholds was undoubtedly due more to the

random locations of breakdowns within the bond line than to

characteristics of the adhesive.

Table 2-6 - Spark Threshold Currents for

Bonded Aluminum Lap Joint

Specimens (All specimens

exhibited electrical conductivity prior to test. Bonded

surfaces were I in2).

Adhesive

Designation

Range of Currents to

Produce Visible Spark (kA)

Adhesive

Conductivity

AF-126-2 <0.10-0.50 none

EA 9602.3 0.20 none

FM 61 <0.i0 none

FM 400 0.20-5.0 partial

FM i000 <0.i0->I0 none

HT 424 >0-I.0 partial

MB 1113 <0.I0-I0 none

R 7114 <0.I0 none

It should be noted that inclusion of the aluminum particles in the

FM 400 and HT 424 adhesives did not increase the spark thresholds.

Whereas the particles make the adhesive partially conductive,

current flow among the particles undoubtedly causes minute sparks,

32
ORIGINAL PAGE iS

OF POOR QUALITY

and if these should occur at or near the edges of such a bond,

the sparks are likely to be visible and may therefore constitute

a source of ignition of fuel vapors.

The data presented in Table 2-6 were derived from tests to

specimens including a i in. 2 bond. Due to the wide variations

in spark thresholds, the data should not be extrapolated to

predict the current that can be allowed to pass through a larger

adhesively bonded surface. Instead, the data indicate that

breakdown and visible sparks may occur under widely varying conditions, and means other than the adhesive must be provided to

allow spark-free transfer of current among adhesively bonded

metal parts.

2.3.1.2.3 Spark thresholds of sealed and riveted aluminum lap

joints. - One means of improving electrical continuity across an

adhesively bonded or sealed aluminum lap joint is to install rivets through the bonded joint. To determine the spark threshold

with a rivet installed, additional lap joint specimens were prepared, with a single rivet installed through the center of the

bonded surfaces, as pictured in Figure 2.9. Groups of five

specimens each were bonded with either of two commercially available fuel tank sealants, in a manner typical of sealed and riveted

fuel tank construction. The rivets were installed "wet" with

sealant but their heads were not sealed over after installation.

Similar groups of five non-riveted specimens with the same

tank sealants as used for the riveted specimens were prepared and

tested to establish a baseline for comparison with the riveted

specimens. Photographs of the riveted and non-riveted specimens

are shown on Figure 2.9.

.... ;

Figure 2.9 - Sealed Aluminum Lap Joint Specimens

with and without Rivet.

33
OF pOOR QUALITY

The specimens were bonded with nonconductive sealants, so

that all current would be forced (initially, at least) to flow

through the rivet. At moderate current levels all current is

conducted through the rivet and no sparking is visible. At

higher currents, of 5-10 kA per rivet, the voltage rise through

the rivet path is sufficient to establish other paths, which

result in visible sparks. A typical test is shown on Figure 2.10.

34

Figure 2.10 -Typical Test of Riveted Aluminum Lap Joint

Specimen. Spark at 20 kA occurs at corner

of bond. Reverse side of specimen is visible in mirror.
oRIGINAL p_GE ,_

Of POOR QUALITY

With the rivet installed, the impulse current spark threshold was raised to 5 kA when the bond was filled with one type of

tank sealant and I0 kA for the other type. Sparks occurred at

the edges or corners of the adherends in all cases. Comparisons

of test results with and without the rivet installed are presented on Figure 2.11.

20

"-"15

o

I

%

,- 10

I

, l

I

I

I

I J

I i

I

I i

I i

I I

I I

I

II ,

EC 3445

(RIVETED)

-- PRO-SEAL

------ E_ 3445

"---'P" THRESHOLD NOT REACHED

I

!

Illll II :;!IIll

PRO-SEAL EC 3445

(NON-RIVETED) (NON-RIVETED)

Figure 2.11 - Spark Threshold of Riveted and Non-Riveted

Sealed Aluminum Lap Joints.

2.3.1.2.4 Current density vs. shear strength. - Following the

spark threshold level tests, three aluminum lap joint specimens

with each of the adhesives described in Table 2-3 were tested

with impulse currents of i kA, 5 kA, I0 kA, and 50 kA peak

amplitudes. The shear strength of each specimen was determined

following these tests and is listed in Table 2-7. All of the

specimen bonds were destroyed at the 50 kA impulse current

amplitude.

35
ORIGINAL F,;_GE ;?:,,..

OF POOR QUALITY

Table 2-7 - Adhesive Shear Strength of Bonded

Aluminum Lap Joint Specimens following

Conduction of Current

Adhesive Shear Strength (Ibs/in 2)

Adhesive Before After After After

Desisnation Test IkA 5kA 10kA

FM 400 2820 2860 2507 897

FM 61 2898 2813 2813 2153

FM I000 -- 5507 6047 5690

HT 424 -- 3580 3869 3600

MB 1113 5600 6180 5973 5760

R 7114 4300 4420 2436 2407

AF-126-2 5270 5413 3580 5333

EA 9602.3 4940 4873 4500 4243

The data in Table 2-7 are plotted graphically on Figure 2.12.

NORMALIZED

SHEAR

STRENGTH

1.0

.8

.6

.4

.2

O

0 n4400 o

O F_ Io00

i_ HT 424

MBR71141113

AF.126.2 EA 9602.3

/

/

I

I

I

I

I

I

I

/

I I I I I I I

O 4 8 12 16 20 24

CURRENTDENSITY,kA/IN.2

36

Figure 2.12 - Effect of Current Density on Adhesive Shear

Strength of Bonded Aluminum Lap Joint

Specimens.
2.3.2 Bonded aluminum fuel line brackets

ORIGII";_L =:-,_ ,,_,_.",T: _

OF POOR QUALITY

2.3.2.1 Specimen description

The bonded aluminum fuel line brackets are typical of

brackets utilized to support small fuel or vent lines within

fuel tanks, as well as conduits or hydraulic lines in other

locations. The brackets were fabricated from 0.032 in. aluminum 0.5 in. wide. Groups of 12 brackets were bonded to single

sheets of primed aluminum with each of the adhesives described

on Table 2-3. Thus, twelve specimens with each type of adhesive

were available. A typical set of brackets is pictured on

Figure 2.13.

Figure 2.13 - Adhesively Bonded Aluminum

Fuel Line Bracket Specimens.

The aluminum brackets and panels were vapor degreased, acid

etched, rinsed and dried prior to priming and bonding. The characteristics of each adhesive and primer are described in Table 2-3.

These are the same adhesives as used for the lap joint specimens.

The bond line thicknesses were not controlled or measured, but

probably ranged between 0.005 and 0.020 in.

2.3.2.2 Test Results

Prior to test, each bracket was given a continuity check with

a battery-operated ohmmeter to see if electrical contact existed

37
between the brackets and the aluminum panel. Except for the

partially conductive adhesives, the presence of conductivity

means that physical contact existed between the bracket and the

panel. The numbers of brackets in each group that exhibited

conductivity prior to test are presented in Table 2-8.

Table 2-8 - Electrical Continuity of Bonded Aluminum

Lap Joint Specimens Prior to Tests

No. of

Conductive

Adhesive Description Specimens

Designation Type Conductivity Group

AF-126-2 supported none 8/12

FM I000 unsupported none 10/12

EA 9602.3 supported none 3/12

R 7114 supported none 2/12

MB 1113 supported none 5/12

FM 61 supported none 1/12

HT 424 supported partial 11/12

FM 400 supported partial 11/12

The partially conductive and unsupported adhesive exhibited

the highest incidence of conductivity. Electrical resistances

of "conductive" specimens ranged from a fraction of one ohm to

several thousand ohms.

2.3.2.2.1 Spark thresholds due to voltage stress.- Specimens

originally nonconductive exhibited bond line sparkover at the

voltages shown on Table 2-9 when tested as described in Para.

2.2.1. Most of the sparkovers were visible.

Since the range of sparkover voltages was so wide, no relationship could be derived between type of adhesive and sparkover

voltages. Some of the specimens with partially conductive

adhesives also exhibited high sparkover voltages. Others showed

some degree of electrical conductivity to start with, as indicated in Table 2-8.

38
Table 2-9 - Breakdown Voltage of Fuel Line

Bracket Specimens

Range of

Breakdown

Adhesive Description Voltages

Des ignation Type Conductivity (kV)

AF-126-2 supported none i. 8- 3.5

EA 9602.3 supported none <0.1-5.5

FM 61 supported none 1.0

FM 400 supported partial 0.5-10.0

FM I000 unsupported none 0.2-10.0

HT 424 supported partial 0. i-I0.0

MB 1113 supported none 0.2-5.0

R 7114 supported none 0.5-2.0

2.3.2.2.1 Spark thresholds due to current flow.- Following

the sparkover tests, all bonded 'fuel line brackets which had

indicated some conductivity prior to test were subjected to

impulse currents as described in Para. 2.2.2 to establish the

minimum currents which would cause sparking. The ranges of currents which produced visible sparking for each type of adhesive

are shown on Table 2-10.

Spark thresholds ranged from less that 0.I0 kA to greater

than I0 kA (the highest level applied) as shown in Table 2.10.

The wide range of spark thresholds was undoubtedly due more

to the random locations of breakdowns within the bond line than

to characteristics of the adhesive.

39
Table 2-10 - Spark Threshold Currents for Bonded

Aluminum Fuel Line Brackets

(All specimens 'exhibited electrical

conductivity prior to test. Bonded

surfaces were 0.25 in.m)

Adhesive Description

Designation Type Conductivity

Range of Currents

to Produce Visible

Sparks (kA)

AF-126-2 supported none 0.2-2.0

EA 6902.3 supported none <0.I-I0.0

FM 61 supported none 1.0

FM 400 supported partial 0.5-10.0

FM i000 unsupported none 0.2-10.0

HT 424 supported partial 0. I-i0.0

MB 1113 supported none 0.2-5.0

R 7114 supported none 0.5-2.0

2.3.3 Bonded aluminum stiffeners

2.3.3.1 Specimen description

The bonded aluminum stiffeners are typical of "T" section

stringers or stiffeners used in wing and integral fuel tank construction. Three stiffeners were fabricated from 0.120 in.

aluminum and adhesively bonded to sheets of primed aluminum

with each of the adhesives described on Table 2-3. "T" sections

were 1.5 in. wide and I0 in. long, so the bonded surface area

was 15 in. 2 Thus, the stiffeners represented the largest

bonded areas tested. Typical specimens are shown in Figure 2.14.

2.3.3.2 Test results

Prior to test, each lap joint specimen was given an electrical continuity check with a battery-operated ohmmeter to see if

physical contact existed between aluminum adherends. The numbers

of specimens in each group of 28 which exhibited conductivity

prior to test are presented in Table 2-11.

40
(_

E

c4

(}.)

,,-4

0

!

o_

41
Table 2-11 - Electrical Continuity Tests of Bonded Aluminum

Stiffener Specimens Prior to Test

Adhesive Description

Designation Type Conductivity

AF-126-2 supported none

FM i000 unsupported none

EA 9602.3 supported none

R 7114 supported none

MB 1113 supported none

FM 61 supported none

HT 924 supported partial

FM 400 supported partial

Nominal

Bondline

Thickness

(in.)

0.002

0 001-0.002

0 002-0 004

0 002-0 004

0 003-0 005

0 005-0 007

0 006-0 008

0 008-0 I0

No. of

Conductive

Specimens

Group

3/3

3/3

2/3

3/3

3/3

0/3

3/3

0/3

The specimens that were nonconductive prior to the start

of tests were given voltage breakdown tests as described in

Para. 2.2.1 to determine the voltage necessary to cause sparkover

across or through the adhesives. Those that exhibited continuity

would, of course, require no voltage to initiate conductivity.

Hence, those specimens were tested initially with impulse currents

as described in Para. 2.2.2 to establish the visible spark thresholds due to current flow.

2.3.3.2.1Spark thresholds due to voltage stress.- Sparkover voltages of the seven specimens that exhibited no conductivity prior

to test ranged between I00 volts and 7100 volts. Visible sparks

occurred at the edges or corners of the specimens, and the low

sparkover voltage and pre-test conductivity was due to metal burrs

or other cutting tool marks along the edges or ends of the specimens.

42
2.3.3.2.2 Spark thresholds due to current flow.- All bonded

aluminum stiffener specimens that exhibited some degree of electrical conductivity were subjected to impulse currents as described

in para. 2.2.2 to establish the minimum currents which would

cause visible sparking. The ranges of currents that caused visible sparking are shown in Table 2-12.

Table 2-12 - Spark Threshold Currents for Bonded

Aluminum Stiffener Specimens

(All specimens exhibited electrical

conductivity prior to test. Bonded

surfaces were 15 in. 2)

Adhesive Description

Designation Type Conductivity

AF-126-2 supported none

FM I000 unsupported none

EA 9602.3 supported none

R 7114 supported none

HB 1113 supported none

FM 61 supported none

HT 424 supported partial

FM 400 supported partial

Nominal

Bondline

Thickness

(in.)

0.002

0 001-0 002

0 002-0 004

0 002-0 004

0 003-0 005

0 005-0 007

0 006-0 008

0 008-0 010

Range of

Sparkover

Cur rent s

(kA)

0.1-0.5

5.0-I0.0

2.0-I0.0

I .0-5.0

0.2-10.6

0.0-2.0

1.0-5.0

0.2-5.0

2.3.4 Comparison of test results for various adhesives

Bond line sparkover voltages for aluminum lap joints, fuel

line brackets and stiffener specimens bonded with each of the eight

adhesives are presented on Figure 2.15.

The sparkover voltages were highest (up to 8.6 kV) for the

supported modified epoxies, probably due to the presence of the

nylon carrier which maintains a positive separation between

43
v

6-

o0

_J

,-4

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@

supported supported

modified modified

epoxy epoxy film

film 0.002-

1 O-- O. 002 in. O. 004 in.

I

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ORIGINAL PAGE Ig

OF POOR QUALITY

AF-126-2 _ 9602.3 _ 61

support_

composite

film

O. 005-

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I

I

i

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I i

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t' 1

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Bonded

2-

HT 424

supported

FM 400 alumlnum

supported FM I000 filled

aluminum unsupported modified

filled polyimlde epoxy

modified modified phenolic

epoxy film epoxy film resin

0.008- 0.001- 0.006-

0.010 in, O,O021n. 0.008 in.

I

I_\\"

Aluminum Stiffeners

Aluminum Fuel Line Brackets

Bonded Aluminum Lap Joint Specimens

MB 1113 R 7114

supported supported

modified modified

epoxy film epoxy film

0.003- 0.002-

0.005 in. 0.0041n.

Figure 2.15 - Ranges of Sparkover Voltages for Bonded Aluminum

Specimens.

bonded surfaces. Within this group, the adhesives with the

thickest bond lines allowed the highest voltage. The low sparkover voltages that sometimes occurred were due to burrs which

penetrated the adhesive at the edges of the specimens, or to

sparkovers across the adhesive at the edges of the specimens.

The highest sparkover voltage (1200 V) for the unsupported

adhesive was somewhat less than that recorded for the supported

adhesives, due to the absence of the carrier.

The sparkover voltages of the two aluminum filled adhesives

did not exceed i00 volts due to the aluminum particles. In

contrast to the other adhesives, which usually sparked over at

the specimen edges, the aluminum filled adhesives allowed conduction to take place within the bond, and visible sparks were not

usually visible until higher currents of sufficient magnitude

were applied to blow arc products out of the bond.

44
Loss of adhesive shear strength due to current flow is due

to arcing and pressure buildup within the bond. Thus, shear

strength was affected only when the sparkover took place within

the bond. Sparkovers that occurred at an edge of one of the

adherends or between the sides of the adherends did not result

in pressure buildup within the bond line, nor affect the bond

in any other way. Thus, loss of bond strength depends greatly

on whether sparkovers may occur within the bond. The adhesives

that permitted this to occur most often were the aluminum powder

filled modified epoxies or unsupported modified epoxies. Those

for which sparkovers occurred most often at the edges were the

supported modified epoxies with no aluminum powder.

Normalized shear strengths following exposure to lightninglike currents are presented as a function of current density for

each of the adhesives on Figure 2.16.

The amount of current necessary to degrade or debond a lap

joint depends on the number of sparkover paths that exist through

the adhesive. If a large number of paths exist, as is possible

when aluminum powder is present throughout the adhesive, arcing

of a large number of partially conducting paths may occur so the

current and pressure buildup associated with each path would be

small. Beneath some level it would be insufficient to degrade

the bond. On the other hand, if the adhesive bond is a comparatively good electrical insulator, only a few sparkovers may

occur within the bond, but these will conduct much higher amounts

of current and the associated pressure buildups can be sufficient

to cause loss of bond strength. For this reason, data that relate loss of shear strength to current density (amperes per unit

of bonded surface) give a very general indication only. Since

the number and location(s) of internal sparkovers are random

occurrences, data produced from small coupon-type test specimens

give only a general indication of the capability of much larger

bonded surfaces. From Figure 2.16, it may be concluded that:

° The shear strength of all adhesives degrade a

minimum of 20% when sparkovers occur inside the

bond, regardless of current level.

, For nonconductive adhesives, strength degradation

becomes significant at 2500 A/in. 2 , at which level

very little shear strength may remain.

• Aluminum-filled adhesives do not suffer appreciable

loss of shear strength at 2500A.

45
OR;GI,NAL _'_"

OF POOR QUALITY

1.0[_

0

0.9 0 03 02

03

08o o3 ,._ O [-3 O

44

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_ 0.6 -,

,=0.5

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'-"(_ LEGEND

O. 3 ADHESIVE

0

z 0 FM 400

[] FM 61

0.2<_ FM I000 []

ix, HT 424 Ch

[_ MB 1113

0.I N R 7114

0 AF-I 26-2

EA 9602.3

0 I 2 3 4

Current Density, kA/in}

Figure 2.16 - Normalized Shear Strength of Bonded

Aluminum Lap Joint Specimens Following Current Flow (data shown is for

internal sparking only).

Note: Numbers next to symbols indicate replicated data.

2.3.5 Adhesively bonded aluminum honeycomb panel specimens

2.3.5.1 Specimen description

The bonded aluminum honeycomb panel specimens are representative of structural members used for rib or spar applications. They

are comprised of aluminum face sheets adhesively bonded to an

aluminum trussgrid core. Each specimen was i in. wide and 15 in.

long. Eight specimens were provided with each of the eight adhesives described on Table 2-3. A typical specimen is shown in

Figure 2.17.

46
ORIGINALPACE I_ii

OF POOR QUALITY

Face Sheet

t

Face Sheet

Figure 2.17 - Edge View of Adhesively Bonded Aluminum

Honeycomb Panel Specimen.

2.3.5.2 Test results

Since the adhesive bond between the aluminum core and face

sheets is not normally exposed to fuel vapors, and since much

of the bonded interfaces are out of view, no attempt was made

to determine sparking thresholds. Instead, tests were conducted to determine the current levels at which adhesive bond

strength becomes degraded, or other physical damage effects

occur.

For this purpose, test currents were conducted from one

face sheet to the other, so that currents would have to flow

across adhesive bonds on both sides of the core. The test setup

is pictured on Figure 2.18.

f

generator I

/

---7 .... ..;-" >:"-:_-_)-=-A _

face sheet

core

face sheet

Figure 2.18 - Circuit for Aluminum Honeycomb Panel

Specimen Tests.

47
ORIGINAL Pf.C_, _ ;_.;

OF POOR QUALI'_, -

Test currents of between 20 kA and i00 kA were applied to

the specimens. Following the current applications, each specimen

was given a ppel test in which the force necessary to peel a face

sheet from the core was measured. This also permitted inspection

of the bonded surfaces for evidence of arcing or other damage.

In general, currents of 50 kA or less (equivalent to

3.33 kA/in. 2) produced up to 30% loss of peel strength and some

evidence of sparking at the bonds was evident after the face

sheets had been peeled away. Currents of from 50 to I00 kA (3.33-

6.67 kA/in. 2) caused considerable loss of bond strength and the

pressure buildup due to internal arcing caused some of the core

to be blown out of the edges of the specimens. A typical example

is shown on Figure 2.19.

Face Sheet

Figure 2.19 - Bonded Aluminum Honeycomb Panel

Specimen following i00 kA Test.

Current was applied between Face

Sheets.

The test results for each adhesive are plotted on Figure

2.20, which shows normalized peel strength versus current density.

This is the same format in which the lap joint shear strength

data are presented in Figure 2.16. Due to the scatter of peel

test results, it is not possible to discern a difference that is

attributable to type of adhesive, except that the three specimens

which sustained the greatest loss of peel strength contained

aluminum powder and were partially conductive. This may simply

be an extension of data scatter and not related to adhesive

properties. No other explanation is apparent.

48
OF POOR QUALFfY

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49
2.4 Hardware Interfaces with Metals

The specimens in this category included access doors which

were riveted, fastened, or bonded and fastened. They represented

typical access door configurations utilized in aluminum aircraft

structure.

Each type of specimen was subjected to impulse currents as

described in Para. 2.2.2 to establish the minimum current which

would cause visible sparking. The test specimens were constructed

to evaluate the effect of various access door bonding and fastening configurations on spark threshold level.

2.4.1 Access doors riveted, fastened, or bonded and fastened

2.4.1.1 Specimen description

The entire access door assemblies were fabricated in 17.5 in.

by 22.5 in. aluminum panels which were 0.040 in. thick. The access

door was also of aluminum construction and 0.040. in, thickness.

A 0.050 in. thick aluminum splice plate was fastened to the aluminum "skin" and access door using rivets, fasteners, and adhesive

in various configurations which included the following:

Configuration 3

Configuration 4

Configuration 5

Configuration 6

Configuration I & 2 - Single rows of LZ4 rivets

- Stagger rows of LZ4 rivets

- Single rows of LZ5 rivets

- Single row of LZ4 rivets

attaching the "skin" to

splice plate. #I0 screws

attaching the access door-tosplice plate

- Adhesive bonding between "skin"

and splice plate. #i0 screws

attaching the access door-tosplice plate

All rivets were "wet" installed. Fuel tank sealant was applied to all fasteners on the wet side and to the skin-to-splice

and access door-to-splice interfaces. In addition, panel configurations 5 and 6 included T-section stringers representative of

typical aircraft construction. The test specimens are further

described in Figures 2.21 through 2.24.

Photographs of typical panel configurations with and without the T-section stringers are shown in Figure 2.25.

50
+

ORIGINAL Pf%C;_;::-_

OF POOR QUALITY

LZ'_N

10.7in.

.Aluminum

Access Door

l

!

/

/

I

I

Aluminum Splice

Skin

Centerllne for 27

access door-to-spllce

rlve_s

Centerline for 27

skin-co-spl_ce rivets

Fuel Side

Figure 2.21 - Access Door Configurations I and 2.

51
ORIGINAL PAGE IS

OF POOR QUALITY

\

\

\

k

11.1 in.

-L--Aluminum

Access

Door

Aluminum

-- Skin

Aluminum

Splice Plate

-Centerlines for

Stagger Row of

34 Skin-to-Spllee

Rivets

for

Stagger Row of

34 Access Doorto-Spllce Rivets

_Access

Door

0.040 in.

Fuel Side

Figure 2.22 - Access Door Configuration 3.

52
ORIGINAL PAGE IS

OF POOR QUALITY

Aluminum Splice

Plate

Aluminum

Skin

Lne for 27

ACcess Door-to_

Splice Rivets

Centerline for 27

Skin-to-Splice

Rivets

Fuel Side

Figure 2.23 Access Door Configuration 4.

53
oRiGiNAL PAG_ _5

OF pOOR QUAL_I_

--Aluminum Skin 0.040 in.

Alumlnum i

• _ 38 required Skin-to-Splice Fasteners

_---------------_(Configuration 5 only>

__ Aluminum Splice

kin - 0.193 in. diameter

/ r I _ / 0.025 in. max' _l _ _ _ _ / hole/NASI473A3 Nutplate

,, I l Gap between Skln and _ Iil I/ ,_i_ 28 required

7.,ln. [ I Accees Door / j I L_l" _ access door-to-

_x_ / I _ _4 I ',i splice fasteners

+

_, L+

_--3.25 in. H Splice Plate and

Stringers are on

_ Fuel Side.

A luminum

Stringer

Figure 2.24 - Access Door Configurations 5 and 6.

Note: Configuration 6 is the same

as above except splice plate is

bonded to aluminum skin.

54
ORIGINAL PAGE _'

OF POOR QUALI"t"_

Figure 2.25 - Photographs Showing Typical Access Door Panels.

Top photo - with T-section stringers

Bottom photo - without stringers

55
2.4.1.2 Test results

Direct lightning strikes to the skin-to-splice and the access

door-to-splice fasteners and to the center of the access door were

simulated during the tests. Each fastener or access door was

struck only once. During the fastener tests, the applied current

level was increased in the sequence of i0, 20, 50, I00, 150 and

200 kA until a fastener sparked. The sequence was repeated several times as allowed by the number of fasteners available.

Each access door was struck only once at the highest current

level available (150-200 kA), thus a spark or no spark result

occurred indicating the ability of the access door configuration

to suppress a fuel ignition spark. The test results are shown

in Figures 2.26 to 2.28.

The test results for a direct strike to the center of the

access door are shown in Figure 2.26. A no spark result is shown

by an upward pointing arrow indicating that the spark threshold

was higher than the tested level. The lowest spark threshold

level was 150 kA. Due to the limited number of tests and specimens, no definite relationship between spark level and fastener

configuration could be concluded.

A typical example of visible sparks resulting from a strike

to the access door is shown in Figure 2.29. As this photo shows,

the pressure built up by the transfer of current from the access

door to the splice plate caused the arc products to blow past the

door "0" ring causing visible sparking within the fuel tank area.

Figure 2.27 shows the spark threshold level for strike to

the skin-to-splice fastener. The minimum spark threshold level

was 180 kA for the series of tests. The graph indicates that

there was no configuration that was significantly better or worse

than the others - all appeared to provide about the same level of

protection as the others. Figure 2.30 shows the results of a

200 kA strike to a skin-to-splice rivet. Pressure buildup beneath the fuel tank sealant due to current transfer from the

rivet to the skin tore through the sealant at the rivet/splice

plate interface causing a potential fuel ignition spark in the

fuel tank interior.

Figure 2.28 shows the results of strike to the access doorto- splice plate fasteners. The minimum spark threshold level

was 50 kA. In general, fastener configuration did not seem to

affect spark threshold level except for configuration No. 5. In

this configuration, the simulated strikes were to #I0 screws and

the current return paths to the skin were through LZ4 rivets. Due

to their larger size, the #I0 screws may have provided a lower

current density path from the access door to the splice plate than

either the LZ4 or LZ5 rivets of the other configurations. The

56
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ORIGINAL PAGE iS

OF POOR QUALITY

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61
ORIGINAL PAGE !_

OF POOR QUALITY

lower current concentration would have prevented excessive pressure from building up to a level sufficiently high to tear through

the fuel tank sealant. Configuration six also utilized #i0 screws

to fasten the access door to the splice plate; however, the current return path to the aluminum skin provided by that specimen

configuration was through a bonding adhesive which was electrically nonconductive. Thus, the adhesive would have tended to aid

in pressure buildup due to its nonconductive properties.

Figure 2.31 shows internal sparking as a result of a 200 kA

strike to an access door-to-splice rivet.

Figure 2.31 - Internal Fuel Tank Sparking as a Result of a 200 kA

Strike to an Access Door-to-splice Rivet.

Current transfer from the rivet to the splice plate resulted

in pressure buildup beneath the sealant, eventually exceeding the

strength of the fuel tank sealant and tearing a hole in it at the

rivet/splice plate interface.

2.5 Adhesively Bonded gr/E

Adhesively bonded gr/E specimens included single lap joints

used to evaluate adhesive shear strength and bonded "T" section

62
ORIGINAL FA_ _3

OF POOR QUALITY

stiffeners. The lap joint specimens were evaluated using varied

adhesives to determine if a possible relationship between adhesive properties and electrical transfer characteristics could be

determined. All the adhesives tested were electrically nonconductive. Additional tests on controlled bond line lap joint

specimens were evaluated for a relationship between bond line

thickness and dielectric voltage breakdown. All of the specimens were representative of configurations being utilized in new

aircraft design.

2.5.1 Bonded gr/E lap joint specimens

2.5.1.1 Specimen description

The bonded gr/E lap joint specimens were fabricated of 4ply

0.056 in. gr/E, and were 2 in. in width, 12 in. in length and

had an overlap of 0.5 in. as shown in Figure 2.32. The gr/E adherends were bonded together with a single layer of adhesive in

accordance with the manufacturers' recommendations. The characteristics of each adhesive are described in Table 2-13.

The controlled bond line specimens had bond line thicknesses

ranging from 0.003 in. to 0.028 in.

4 ply gr/E Laminate - 0.056 in.

I

1

I

I

L I

12.0 in. ,-

6.25 in.

l

250 Film Adhesive EA962_ZW HYSOL

350 Film Adhesive EA9649 HYSOL

Room Temp Paste ADX3130 HYSOL

I

6,25in _--_

Figure 2.32 - gr/E Lap Joint Specimen.

63
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2.5.1.2 Test results

Prior to test, each lap joint specimen was given an electrical continuity check with a battery-operated ohmmeter to see

if physical contact existed between the gr/E adherends. The

number of specimens in each adhesive group which exhibited conductivity prior to test is presented in Table 2-14.

Table 2-14 - Electrical Continuity of gr/E Lap Joint

Specimens Prior to Test.

Nominal

Adhesive Bond Line

Type Thi ckne s s

Number of Conductive

Specimens/Group

EA9 628NW 0.005-0.007 in. 0/12

EA9679 0.005-0.007 in. 3/3

ADX3130 0.005-0.007 in. 0/3

The three EA9649 adhesive samples were conductive prior to

test, apparently due to incidental penetration of yarns through

the adhesive. Due to the small number of samples involved, no

relationship could be determined between electrical conductivity

and type of adhesive or bond line thickness.

The specimens that were nonconductive prior to the start

of tests were given voltage breakdown tests, as described in

Para. 2.2.1, to determine the voltage necessary to cause sparkover across or through the adhesives. Those that exhibited

continuity would require no voltage to initiate continuity,

hence, those specimens were tested initially with impulse currents

that may produce visible sparks.

2.5.1.2.1 Bond line breakdown voltages.- Table 2-15 shows the

breakdown voltage levels for the two nonconductive adhesives

tested. The lap joints bonded with EA9649 were not subjected

to voltage breakdown tests since they exhibited continuity prior

to test. Breakdown voltages for specimens bonded with EA9628NW

ranged from 1200V to 8000V for II of the 12 specimens tested, one

specimen broke down at less that i00 volts. The breakdown voltage range for the ADX3130 adhesive was 1600V to 2000V.

Visible sparking at breakdown, which was due to edge flashovers, occurred in twenty-five percent of the specimens bonded

with EA9628NW and all of the specimens bonded with ADX3130.

Thus, the majority of the specimens bonded with EA9628NW (seventy

five percent)and none of the ADX3130 specimens broke down within

the bond line.

65
Table 2-15 - Breakdown Voltage vs. Bond Line Adhesive of

Nonconductive $r/E Lap Joint Specimens.

Sample Adhesive Breakdown Voltage

I EA9628NW 4600

2 EA9628NW 3700

3 EA9628NW 6000 (a)

4 EA9628NW 6000 (a)

5 EA9628NW 4000

6 EA9628NW <I00

7 EA9628NW 3600

8 EA9628NW 3200

9 EA9628NW 1200

I0 EA9628NW 6800

II EA9628NW 8000 (a)

12 EA9628NW 3400

I ADX3130 1600 (a)

2 ADX3130 2000 (a)

3 ADX3130 2000 (a)

(a) visible spark

An additional set of specimens was tested to determine the

relationship between bond line thickness and voltage breakdown

level. The test results are summarized in Table 2-16. This

group of specimens was identical to the previous group except

that only EA9628NW adhesive was used. Prior to the tests, continuity measurements were made between the adherends to determine if electrical contact existed between them. None of the

specimens exhibited continuity.

The results showed that the overall breakdown voltage

level increased with increasing bond line thickness. The voltage "gradient", however, decreased from 350 volts per mil for

"thin" bond lines to about 150 volts per mil for "thick" bond

lines. All breakdowns occurred at the edges of the specimens

indicating that the voltage withstand capability of the adhesive was greater than that of air.

66
Table 2-16 - Breakdown Voltage vs. Bond Line Thickness for gr/E Lap Joint Specimens.

(EA9628NW Adhesive only)

Bondline Thickness

(in.)

0.003

0.008

0.010-0.013

0.015

0.027-0.028

Breakdown Range

(volts)

II00

1700-1800

2100

2200-2400

4200-4400

2.5.1.2.2 Spark threshold due to current flow.- Following the

voltage breakdown tests, all lap joint specimens were subjected

to impulse currents as described in Para. 2.2.2 to establish

the minimum currents which would cause visible sparking. Since

some of the specimens showed visible sparking during the breakdown voltage tests, they conducted no appreciable current before

sparking occurred. The test results are summarized in Table

2-17 and indicate that the specimens which were bonded with

EA9649 and which exhibited pretest conductivity were able to

conduct up to 5000 amperes before producing visible sparks.

This was more than five times the current conducted by the

EZ9628NW specimens before they exhibited visible sparking.

Table 2-17 - Spark Threshold Currents for

gr/E Lap Joint Specimens.

Bonded surfaces were 1 inJ

Adhesive

Range of Currents to

Produce Visible Sparks

(amperes)

EA9628NW <100-900

EA9649 5000

ADX3130 <I00

2.5.1.2.3 Current density vs. shear strength.- Following the

current flow spark tests, the gr/E lap joint specimens were

tested for shear strength. A comparison of the post test shear

strength with untested control samples is given in Table 2-18.

Loss of shear strength is due to arcing and pressure

buildup within the bond due to current flow. Thus, shear

strength was affected only when the sparkover took place within

the bond. Sparkovers that occurred at an edge or between the

67
Table 2-18

Specimen No.

Control #I

Control #2

Control #3

i

2

3 (a)

4 (a)

5

6

7

8

9

I0

II (a)

12

Control #I

Control #2

Control #3

i

2

3

Control #I

Contro I #2

Control #3

I (a)

2 (a)

3 (a)

(a) visible

Shear Strength of

Following Current

Bonded gr/E Lap Joint Specimens

Spark Threshold Tests.

Current

Adhesive Density

Type (A/in. 2)

EA9628NW --

EA9628NW --

EA 9628NW - -

EA9628NW 796

EA9628NW 874

EA9628NW <i00

EA9628NW <i00

EA9628NW 345

EA9628NW 259

EA9628NW 305

EA9628NW 259

EA9628NW 248

EA9628NW 2 75

EA9628NW <i00

EA9628NW 819

EA9649 --

EA9649 --

EA9649 --

EA9649 _5000

EA9649 _5000

EA9649 5176

ADX3130 --

ADX3130 --

ADX3130 --

ADX3130 <I00

ADX3130 <I00

ADX3130 <I00

sparking during voltage

Test Failure

Load

(Ibs./in. 2)

3190

3158

3693

2043

2022

3075

3919

1318

1711

2715

2362

1908

2741

2545

1761

1716

2119

1784

590

1641

2385

1660

1541

2011

2220

b r eak down tests

Failure Mode

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

_nterlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

interlaminate

bond destroyed

bond destroyed

interlaminate

a dhes ive

adhesive

adhesive

adhesive

adhesive

adhesive

68
sides of the adherends did not result in pressure buildup within

the bond line or affect the bond in any other way.

Those specimens which exhibited visible sparking during

voltage breakdown subsequently produced very low current spark

threshold levels. Since those specimens sparked over externally

(rather than internally within the bond), it was expected that

no loss of shear strength would result. Table 2-19 summarizes

the shear strength of the test specimens compared to the test

sparking mechanism. Those specimens which exhibited external

sparking showed little, if any, loss of shear strength when

compared to the control specimen values. Those specimens which

sparked within the bond showed loss of shear strength when compared to the control specimens.

2.5.2 Bonded gr/E stiffeners

2.5.2.1 Specimen description

The bonded gr/E stiffeners were typical of "T" section

stringers or stiffeners used in wing and integral fuel tank

construction. The "T" sections were of 4 ply 0.056 in. gr/E

construction and were 1.5 in. wide and 6 in. long, so the

bonded surface area was 9 in. 2 . They were bonded to 5.5 in. by

I0 in., 4 ply gr/E "skins" with EA9628NW adhesive. The specimen

is shown in Figure 2.33.

2.5.2.2 Test results

Prior to test, each gr/E stiffener specimen was given an

electrical continuity check with a battery-operated ohmmeter to

see if physical contact existed between the gr/E adherends. All

9 specimens showed electrical continuity. Since no voltage would

be required to initiate conductivity, the specimens were tested

only with impulse currents as described in Para. 2.2.2 to establish the visible spark thresholds due to current flow.

2.5.2.2.1 Spark thresholds due to current flow.- All nine gr/E

stiffener specimens produced visible sparks at applied current

levels of less than I00 amperes. Apparently, incidental penetration of yarns through the adhesive at the specimen edges resuited in visible sparks.

2.6 Metal to gr/E Interfaces

The metal to gr/E interfaces included rivets, access door

dome nuts and fuel line feed-through elbows in gr/E laminates.

These subelement specimens represented typical interfaces utilized in the production of gr/E integral fuel tanks and which

could be a source of fuel ignition sparks due to lightning cirrents flowing through them.

69
ORI(_INAL pAGE _.S

OF PoOR QUALITY

Table 2-19 - Summary of Shear Strength Levels Compared

to Current Threshold Sparking Mechanism

for Bonded gr/E Lap Joint Specimens

Adhesive

Designation

Control

Specimens

Shear Strength (ibs./in.2)

Specimens Sparking Specimens Sparking

within Bond Externally

EA9628NW

EA9649

ADX3130

3158-3693 1711-2741 2545-3919

1716-2119 590 ---

1641-2385 --- 1541-2220

/x_--gr/E Laminate 0.056 in.

CD

r-q

I >

6.0in.

2.0 in.

l

70

0.75 in.

2.75 in. _ _0.12 in.

S _bl 0.75in.

Figure 2.33 - Bonded gr/E Stiffener.
Each type of specimen was subjected to impulse currents as

described in Para. 2.2.2 to establish the minimum currents which

would cause visible sparking. Test specimens were fabricated

to evaluate the effect of fuel tank sealant, to compare different sealants, and to compare results with and without the use of

an adhesive carrier cloth.

2.6.1 Rivets in gr/E laminates

2.6.1.1 Specimen description

The rivet in gr/E laminate specimens represented metal-tographite interfaces utilized in gr/E integral fuel tank construction. Lightning currents conducted across this interface could

cause sparking which would be a fuel ignition hazard.

The specimens consisted of a 0.04 in. thick, 2.5 in. by

0.75 in. aluminum tab which was riveted to a gr/E skin with a

0.125 in. diameter titanium rivet. The gr/E skin was 3 in. by

6 in. and was constructed of four plies of gr/E laminate 0.056 in.

thick. In addition, some of the specimens also contained a

0.005 in. one ply layer of fiberglass between the aluminum tab

and gr/E skin to simulate an adhesive carrier cloth. The test

specimen is shown in Figure 2.34.

2.6.1.2 Test results

Currents were conducted through the rivet to gr/E interface

by connecting the generator "high" side to the aluminum tab and

returning the current to generator "ground" through a connection

to the gr/E "skin". Initial specimen tests were to determine

the effect of the carrier cloth and fuel tank sealant on the

spark threshold level. The applied test current levels were increased in a I, 2, 5 sequence on each specimen until a spark was

detected. The test results are shown in Figures 2.35 and 2.36.

The test results indicate that, in general, the spark levels

were higher for specimens without the carrier cloth than those

containing the carrier cloth. Apparently, the carrier cloth

forced all of the current to pass through the rivet; thus, the

carrier cloth increased the current density at the rivet-gr/E

interface resulting in visible sparks at lower applied current

levels. Comparison of specimens with the adhesive cloth carrier

to those without the cloth carrier indicates that the application

of fuel tank sealant had the effect of increasing the spark

threshold level. For the specimens with the adhesive carrier

cloth and sealant, the spark threshold level was 4.3 kA; for

specimens without the adhesive cloth but with sealant, the spark

threshold level was 7.5 kA. Without sealant, the threshold

71
ORIGINAL PAGE _

OF POOR QUALITY

3.0 in.

gr/E Laminate

\

6.0 in.

0. 056 in.

//

//

0. 125 in.

-- titanium rivet

_--I.I0 in._

2.1 in.-_

0.75 in.

-- fiberglass

72

[.-

aluminum

I ply

fiberglass

0. 005 in.

Figure 2.34 - Rivet in gr/E Laminate

Test Specimen.

I

0.04 in.
2O

v

•o 15

.=

10

with

sealant

Test Samples

without

sealant

Figure 2.35 - Spark Threshold Levels for Rivet in gr/E

Laminate with Adhesive Carrier Cloth.

20

---15

•'_ 10

I1with

eealan_

Test Samples

without

sealant

Figure 2.36 - Spark Threshold Levels for Rivet in gr/E

Laminate without Adhesive Carrier Cloth.

73
level for the carrier cloth specimens was 2.1 kA; those without

the carrier cloth were able to tolerate 7.5 kA before sparking.

A second series of tests utilized two types of fuel tank

sealant to determine if sealant characteristics had an effect

on the spark threshold level. Specimens with and without an

adhesive carrier cloth were tested. Table 2-20 lists the characteristics of the sealants tested.

Table 2-20 - Characteristics of Fuel Tank Sealants

Utilized in gr/E Test Specimens

Sealant

Manufacturer

Manufacturer

Part No.

Tensile

Type of Strength

Sealant (Ibs/in. 2)

Essex Chemical

Corp.

PRO-Seal Two part 300(a)

890 polysulfide

integral fuel

tank and

fuselage sealant

Products Research

and Chemical Corp.

Notes:

(a) 14 day cure

(b) 28 day cure

SEMKIT Two part, 269(b)

654 polysulfide

integral fuel

tank and

fuselage sealant

Applied current levels to each specimen were initiated at

5 kA and increased in 5 kA steps until a visible spark was detected. The test results are shown in Figures 2.37 and 2.38.

The specimens containing the simulated adhesive carrier

cloth, in general, sparked at lower current levels than those

without a carrier cloth. This result was the same as that obtained

during tests on the first set of rivet in gr/E laminate specimens.

The test results also indicated that the Semkit 654 specimens had, on the average, higher spark threshold levels than did

the specimens coated with Pro-Seal 890 sealant. In some cases

the Semkit sealant prevented visible sparking up to and including 17 kA which was the limit of the generator. The difference

in spark threshold level was less well defined between specimens

coated with a thin coat of Pro-Seal and those coated with a

normal coat of Pro-Seal. Most likely, this was the result of

difficulty inapplying a thin coat of Pro-Seal.

74
ORIGINAL PA;._ _;3

OF POOR QUALITY

2O

v

15

0

¢o

PRO-SEAL

,!11

PRO-SEAL

(THIN COAT)

II

i l

I I

I i

I

I

i

il

PRO-Seal 890

------ SEMKIT 654

I lel

I III

ill i

i ill

I

,EMKIT

Figure 2.37 - Comparison of Spark Threshold Levels of

Two Fuel Tank Sealants Applied to Rivet

in gr/E Laminate Specimens with an

Adhesive Carrier Cloth.

2O

15

v

i0 "r

u'J

LU

I-

_ 5

PRO-SEAL

(THIN COAT)

-- PRO-SEAL 890

.... SEMKIT 654

------_Threehold not reached

I

II

II iI

SEMKIT

Figure 2.38 - Comparison of Spark Threshold Levels of

Fuel Tank Sealants Applied to Rivet in

gr/E Laminate Specimens without an

Adhesive Carrier Cloth.

75
ORIGINAL PAGE lg

OF POOR QUALITY

The spark threshold level of the Pro-Seal specimens, with

and without the carrier cloth was 5 kA; the spark threshold

level of the Semkit specimens, with and without the carrier

cloth was i0 kA. In general, the Semkit sealed specimens conducted greater amplitude current levels before sparking than

did the Pro-Seal specimens. Figure 2.39 shows a typical spark

produced during these tests. The transfer of current from the

gr/E skin to the rivet resulted in pressure buildup beneath the

sealant and subsequent release through a tear in the sealant at

the gr/E skin surface.

Figure 2.39 - Spark at Rivet to gr/E Interface Due to

17 kA Simulated Lightning Strike.

2.6.2 Access door dome nuts in gr/E laminates

2.6.2.1 Specimen description

The access door dome nut specimen consisted of a 1.5 in. by

2.88 in. 7 ply, 0/I0 in. gr/E skin section to which a nutplate

was fastened by two rivets. A second 7 ply gr/E skin section of

identical dimensions was fastened to the first section by a

countersunk machine screw which threaded into the nutplate. A synthetic rubber gasket between the gr/E skin sections was used to

insure that all of the impulse current flowed through the dome

nut. The specimen is shown in Figure 2.40.

76
ORIGINAL PAGE IS

OF POOR QUALITY

7 ply gr/E

laminate

I 1.50 in.

;I __ L

_- 0.218 in. dta. hoIe

NAS1473C3 nutplate (cap)

MS20427F3-4 rivet (2)

2.88 in.

5.00 in.

---Synthetic Rubber Gasket

' 2.88 In. il

Figure 2.40 - Access Door Dome Nut in gr/E Laminate.

2.6.2.2 Test results

Generator current levels of 5, II and 17 kA were applied to

one gr/E skin section, conducted through the dome nut assembly,

and returned to generator "ground" from the second gr/E section.

Tests on a specimen were concluded when a visible spark occurred.

Test specimens were prepared with either Pro-Seal or Semkit fuel

tank sealant applied to the fuel side of the dome nut assembly.

The test results are shown in Figure 2.41.

The tests indicated that, in general, the Semkit sealed

specimens sparked at higher current levels than did the Pro-Seal

specimens. The minimum spark threshold level of the Semkit

sealant was, however, at 5 kA which was lower than the I0 kA

level of the normally applied Pro-Seal specimens and the same

as the specimens with the thinly applied coat of Pro-Seal.

Thus, the minimum spark threshold level for either the Pro-Seal

or the Semkit sealant was 5 kA.

77
ORIGINAL pAGE I_

OF pOOR QUALITY

20

,-_15

,._

.J

o

"1-

w

,- 10

v

PRO-SEAl

1

1

PRO-SEAL.

(THIN COAT)

t

I

I

J

i

i

I

I

I

Ill I

SEMKIT

--- PRO-Seal 890

.... SEMKIT 654

--_Threshold not reached

i

i

I

i

I'I

i

fi

Figure 2.41 - Spark Threshold Levels of Access Door

Dome Nuts in gr/E Laminates.

Figure 2.42 shows a photograph of a 17 kA spark through

Pro-Seal sealant. The spark was the result of current transfer

from the first gr/E skin section, through the dome nut assembly

and into the second gr/E skin section. The internal pressure

buildup, which resulted from current flow through the specimen,

tore through the sealant at the gr/E skin surface as shown by

the photograph. This spark was typical of those observed

during these tests.

2.6.3 Fuel line feed-through elbows in gr/E laminates

2.6.3.1 Specimen description

The fuel line feed-through elbow specimens were metal-tographite interfaces which represented interfaces between fuel

lines and internal fuel tank structural configurations such as

ribs.

Five feed-through elbows were mounted on a 3.0 in. x 18 in.

4 ply, gr/E laminate. Each surface of the laminate made contact

with the feed-through elbow by a 0.036 in. thick, 2.06 in. diameter washer. A drawing and a photo of the test specimen are

shown in Figure 2.43.

78
OF FCOR QbALITY

,_

4.)

.1...I

r,..)

i_.

0

.I-I

.I-I

z

0

O0

O_

_._

u_

4.,1 ,_

I

Qt

79
AN960C916L washer

__

3.0 in.

4 ply gr/E Laminate

18.0 in.

OF POOR QUALFi'Y

0. 59 in. dia. hole (5 req'd)

AN833-6J elbow

AN924-6J nut

1 each req'd)

AN960C916L washer

"Wet" Side

"Dry" Side

Figure 2.43 - "Wet" Side View of Fuel Line Feed-Through Elbow

in gr/E Laminate.

80
2.6.3.2 - Test Results

ORIGINAL PAGE ;_.

OF POOR QUALITY

The generator "high" side was connected to the "dry" side

of the feed-through elbow and generator return was connected to

the gr/E "skin". Thus, current flow was into the metallic feedthrough elbow, across the washer-gr/E interface, into the gr/E

skin and then returned to generator "ground". This circuit,

then, represented possible current flow through internal fuel

lines as a result of a lightning strike to the wing tip.

Each feed-through elbow was tested with a series of increasing current levels (6, i0, 14, 17 kA) using a 4 x50 ps wave

and 20, 25, 30 and 40 kA using an 8 x 20 Us current wave. Testing on an elbow ceased when a visible spark was detected. Both

Pro-Seal and Semkit fuel tank sealant was applied over the

washer - gr/E interface on both the "wet" and "dry" skin surfaces. A mirror was positioned to provide a view of the specimen

"dry" side during test.

The test results are shown in Figure 2.44.

4O

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= 20

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Position Number

--PRO-Seal 890

.... SEHKIT 654

--'--'_Threshold not reached

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Figure 2.44 - Spark Threshold Levels of Feed-through

Elbows in gr/E Laminates.

81
URIGINAL PAGE

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Examination of the test results indicates that, in general,

both Pro-Seal and Semkit sealants provided equal protection

against visible sparks. This is in contrast to the results on

the rivet and dome nut gr/E interfaces in which the Semkit sealant appesred to provide higher spark threshold levels. The

metal to gr/E contact surface provided by the feed-through

elbow washers was greater than the contact surfaces provided

by the rivet or dome nut specimens; hence, the current density

at the interfaces was lower. The lower current density level

may have tended to "equalize" the results between the two

sealants due to lower internal pressure buildups and could also

explain the higher current threshold levels obtained with these

specimens. Both sealants were able to prevent visible sparks

at levels of 40 kA or higher in some specimens. The minimum

threshold level was I0 kA for the Semkit specimens and 14 kA

for the Pro-Seal specimens. No difference in spark threshold

levels was noted between the "fuel" side and the "dry" side.

Figure 2.45 shows a spark at a feed-through elbow resulting from a current of 15 kA applied to the feed-through elbow.

The spark was the result of pressure buildup beneath the

sealant due to current flew across the metal washer to gr/Einterface. The pressure caused the sealant to tear at the metalgraphite interface allowing release of the arc products and

producing a visible spark.

Figure 2.45 - Spark at Washer to gr/E Interface Resulting from

a 15 kA Current Strike to the Fuel Feed-through

Elbow. Wet side view shown.

82
3.0 FULL SCALE STRUCTURES TEST PROGRAM

3.1 purpose

Measurements of the electrical behavior of adhesively bonded

aluminum and gr/E coupon-type subelement specimens were conducted

during phase one of the program. Whether the voltage and current

levels measured during these tests actually existed in full scale

structures was the subject of investigation during phase two.

Specifically, measurements and tests conducted included the

following:

i. Current distribution

2. Bond line voltages

3. Ignition source locations

4. Induced voltages in wing electrical wiring

5. Magnetic field levels

The structures tested were full scale wings containing integral fuel tanks and were typical of general aviation aircraft construction.

The first of the two structures tested was a right hand

aluminum wing which contained an adhesively bonded integral fuel

tank and outboard leading edge assembly. The remainder of the

structure was constructed with conventional rivets and fasteners.

With the exception of the leading edge, tip, and some control

surfaces, the second structure was constructed entirely of gr/E

composites. The wing was fabricated as one unit (left and right)

but only the right wing was tested.

3.2 Basic Test Methods

A brief description of the test setup and procedures is presented in paragraphs 3.2.1 through 3.2.6. A more detailed description of the test specimens and test procedures is presented

together with the test results in paragraphs 3.3 and 3.4.

3.2.1 Test setup

The full scale structure was positioned adjacent to the high

current generator. Connections between the generator and wing

simulated a strike to the wing tip with currents flowing through

the wing structure to an exit point at the wing root. Measurement signals were conducted from the wing structure through

shielded cables and electrical conduit to a nearby RF shielded

room where they were recorded by oscilloscopes.

83
3.2.2 Current distribution tests

Currents in fuel lines, conduits, control cables and other

structural members were measured using a wide-band current transformer which was placed around the conductor to be measured. The

measurement signals were transferred through shielded cables and

recorded by the oscilloscope.

3.2.3 Bond line voltage measurements

Voltages across bond lines were measured in the fuel tank

interiors. One measurement lead was connected to the wing skin

and the second was connected to a rib, spar, or stringer. The

signals were transmitted by shielded cable to the differential

inputs of an oscilloscope and the measurement waveform recorded

on an oscillogram.

3.2.4 Ignition source tests

Possible fuel vapor ignition sources were investigated by

the placement of cameras within the sealed fuel tank during a

simulated strike to the structure. The appearance (or absence)

of light ("sparks") on the film negative indicated the presence

(or absence) of an ignition source.

3.2.5 Induced voltage measurements in wing electrical wiring

The voltages induced in the wing circuits by the simulated

lightning currents were measured using the differential measurement technique. The circuit conductor under test was connected

to one input of the oscilloscope differential preamplifier;

circuit "ground" was connected to the second input. The measurement cable and instrumentation were shielded to minimize measurement errors due to external fields and electrical "noise".

3.2.6 Magnetic field measurements

Both internal and external magnetic fields were measured

using a search coil positioned in each of three orthogonal directions. The signal induced in the coil was transmitted through

shielded cable, integrated and recorded by an oscilloscope. The

magnetic field intensity was calculated using the search coil

calibration factor.

3.3 Bonded aluminum structure

3.3.1 Specimen description

The wing extended from inboard of the right engine nacelle

out to and including the wing tip. Except for the wing tip which

was made of acrylonltrile-butadiene-styrene (ABS) plastic, the

84
wing was of aluminum construction. The wing contained one integral

fuel tank and outboard leading edge assembly of adhesive bonded

construction spanning from nacelle to tip. This was the only part

of the wing which was adhesively bonded; elsewhere conventional

rivets and fasteners were utilized. Sketches of the wing are shown

in Figures 3.1 and 3.2.

The wing test locations hereafter referred to in this report

will be referenced in inches from fuselage center line by a wing

station (WS) number. For example, the wing box closeout rib which

is located 220 in. outboard of the fuselage center line, would be

designated WS 220.

3.3.2 Test setup

Figure 3.3 is a photograph showing the laboratory set-up for

tests on the bonded aluminum wing.

The wing was arranged to simulate a strike to the wing tip

with the lightning currents flowing inboard to an exit point elsewhere on the plane.

The wing was suspended from the ceiling with the leading

edge up and the inboard end located at the lightning current

generator. Twelve in. width aluminum flashing, connected to the

generator output, encircled the wing with provision for attachment

to the taxi and navigation lights as shown in Figure 3.4. This

arrangement allowed the current to be injected into the wing tip,

flow through the wing skin, and return to generator "ground"

through the front main, and rear spars at the wing root.

Currents flowing in the aluminum flashing conductors produced nearly equal and opposite magnetic fields between them;

thus, the influence of the test circuit magnetic fields on the

wing was minimized. In addition, the use of wide conductors minimized the test circuit inductance which was necessary for generation of peak currents of I00 kA or greater. The conductors were

positioned as far as practical from the wing surface to enable

access to the fuel filler cap and access doors for instrumentation

purposes and to minimize the influence of their magnetic fields on

the wing structure.

The simulated lightning currents were generated by a 17.25_F

70 kV high current capacitor bank in conjunction with waveshaping

elements. The currents were measured by a I00:I ferrite core,

impulse current transformer (C.T.) and were recorded by a

Tektronix 535Aoscilloscope. The simulated lightning current waveform oscillograms are shown in Figure 3.5.

85
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Test 2146

2kA/div 4x50_s l_s/div

Test 2145

2kA/div 4x50_s 20_s/div

Test 2207

5kA/div Oscillatory 20_s/div

(17kHz)

Figure 3.5 - Simulated Lightning Current Waveforms.

90
3.3.3 Test procedures

3.3.3.1 Current distribution

Measurement of currents in internal structural members and

fuel lines was limited to those members which could be physically

fitted with the current transformer. The current transformer was

a Pearson Model IIOA which had a current times time (IT) product

of 0.5 A.s and a frequency response of 1 Hz to 35 MHz. RG-58/U

coaxial cable connected to the current transformer, routed the

measurement signal from the wing to a Tektronix 7704A oscilloscope located in the RF shielded room. The cable was routed as

far as practicable within the wing structure to shield the signal

from the influence of external fields. Aluminum foil and electrical conduit provided shielding between the wing and the RF shielded

room.

The cable was terminated in 50_ at the Tektronix 7704A oscilloscope.

The simulated lightning current, which was injected into

the taxi light housing, was a 16 kHz oscillatory wave with peak

amplitudes ranging from 12 to 88 kA.

Measurements were made of currents in fuel lines (outside

the fuel tank) and the forward and rear spars.

3.3.3.2 Bond line voltages

Lightning currents flowing through the aluminum wing skin

can raise its voltage potential with respect to structural or

fuel system components which may be electrically isolated from the

skins. If the voltage potential difference between the skin and

these components exceeds the withstand capability of the insulating

medium (adhesive or air), breakdown will occur and the resulting

spark could be a possible fuel ignition source.

Measurements were recorded across bond lines with the fuel

tank interior between wing locations WS 153.75 and WS 164.65.

Both spar-to-skin and rib-to-skin bond line voltages were measured.

A "ground" measurement reference plane was established at

the fuel tank skin by attaching 14 in. wide aluminum flashing between the skin and a grounded electrical junction box. Two RG-58/U

coaxial cables, shielded with aluminum foil, conducted the measurement signals from the tank interior, to the electrical box, and

through electrical conduit to the measurement oscilloscope located

in the RF shielded froom. The influence of external magnetic

fields was minimized by maintaining contact between the shielded

measurement cables and the aluminum flashing ground plane as shown

in Figure 3.6.

91
ORIGINAL PAGE !_

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Figure 3.6 - Lab Setup for Measurement of Bond Line Voltages

92 on Bonded Aluminum Wing.

_c---
One test lead was connected to the wing lower skin with a

self-tapping screw; the second lead was connected to the rib or

stringer under test.

The simulated lightning current unipolar and oscillatory

waves were injected into the wing at the taxi light housing. The

voltage potential which developed across the bond lines was measured by the shielded coaxial cables and recorded by a Tektronix

7704A oscilloscope with a type 7A13 preamplifier. The test circuit schematic is shown in Figure 3.7.

3.3.3.3 Ignition source tests

Lightning Currents flowing through an aircraft structure,

such as a wing, can produce sparking across joints, especially

those which are bonded with a nonconductive adhesive. In the

aluminum wing, bonded joints were located within the fuel tank

area; thus, a spark across one of these joints would present an

ignition source hazard.

The method employed to detect possible sparks in the fuel

tank was similar to that described in paragraph 2.2.4 for spark

detection during the subelement tests. The size of the fuel

filler cap opening prevented insertion of a Polaroid camera into

the fuel tank; the spark detection was limited to the use of a

more compact 35 mm camera equipped with an equivalent lens/film

speed combination.

The simulated lightning strikes of 85 kA - I00 kA peak amplitude were conducted into the taxi light housing and returned to

the generator through the front, main, and rear spars at the wing

root end to represent the effects of stroke currents being conducted through the structure between the wing tip and wing root.

It was recognized that direct strikes to the fuel tank skin were

also possible since portions of this tank lie aft of the propeller.

However, as noted in paragraph 1.2.2.3 the effects of direct strikes

to integral fuel tank skins were not a part of this investigation.

The 85 kA - I00 kA peak current, which represents a moderately

severe stroke, represented the limits of the test facility generator.

The camera lens was positioned in each of the four orthogonal

wing directions during the simulated strikes to the taxi light

housing. These directions included views toward the upper and

lower wing skin surfaces and toward the wing tip and root ends.

The fuel tank was sealed for each tests and a test frame was recorded prior to each test to verify that the fuel tank was lighttight. The test circuit schematic is shown in Figure 3.8.

93
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3.3.3.4 Induced voltages in wing electrical wiring

The electrical circuits contained within the bonded aluminum

wing included the following:

Right hand taxi light

Right hand navigation light

Strobe lamp and power supply

Inboard & outboard fuel quantity transmitters

Auxiliary fuel pump

Alternator

Overvoltage relay

Fuel primer solenoid

Heater fuel solenoid

Right hand oil pressure transducer

Right hand cylinder head temperature transmitter

Right hand oil temperature sensor

Voltage regulator

Right hand stall switch

Fuel pressure transducer

A wire from each of these circuits was accessible for test

connections at the wing root..RG-58/U coaxial cable connected

each circuit under test to channel "A" of a Tektronix 7704A type

7A13 differential preamplifier. A second coaxial cable (circuit

return) connected a wing inboard rib to the preamplifier channel

"B" input. By differentially subtracting the readings of the

measurement and return lines, errors due to electrical "noise"

were minimized. Additional noise reduction techniques included

shielding of the coaxial cable measurement lines with aluminum

foil and electrical conduit, and locating the measurement instruments within an RF shielded room.

A I0:I resistive voltage divider, inserted into the measurement circuit, protected the oscilloscope preamplifier circuitry

by limiting the voltage level appearing at its terminals. The

divider consisted of two sets of five I00_ resistors which fed

the coaxial cables terminated in 50_ at the oscilloscope preamplifier. Simulated lightning currents were injected into the

taxi or navigation lights and were removed from the wing at its

root end through connections between the forward, main, and rear

spars and generator "ground"

A photograph of the instrument cable and test current connections is shown in Figure 3.9. The test circuit schematic is

shown in Figure 3.10.

96
ORIGINAL P_,_..,_,_

0_. POOR QUALI'T'Y

Figure 3.9 - Root End of Bonded Aluminum Wing Showing Shielded

Instrument Cables and Test Current Connections

for Measuring Induced Voltages in Wing Electrical

Wiring.

97
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3.3.3.5 Magnetic field measurements

The magnetic field intensity within the wing fuel tank and

along its exterior surface was measured with the aid of search

coils. The changing magnetic flux (d@/dt) produced by these fields

induced a voltage signal in the search coil. This signal was

then integrated by a passive RC integrator and recorded by a

Tektronix 7704A oscilloscope with Type 7A13 preamplifier. The

magnetic field intensity (H) at the coil measurement location was

then calculated from the measurement oscillogram and the calibration factor of the search coil.

Three different search coils and two passive integrators were

utilized. The specifications are given in Table 3-1.

Table 3-1 - Specifications for Search Coils and

Passive Integrators Utilized in

Magnetic Field Measurements

Search Coil

SingleLayer Multi-Layer Single Layer

Number of Turns 20 20 i00

Effective Diameter 2.4 inches 2.4 inches 2.4 inches

Inductance _30_H _23_H _885_H

Resistance _0_ _0_ _30_

Passive Integrator

R C

i0 k_ 0.I0 _F

Time Constant

1 ms

I k_ 0.01 _F i0 us

R

o---4A4 o

e(t) i V(t)

from coil C to

T oscilloscope

O O

Figure 3.11 - Passive Integrator Schematic.

99
The magnetic field intensity can be related to the integrated coil signal as follows (refer to schematic diagram of

passive integrator in Figure 3.11).

e(t) - C dV(t) -,,(R>>IXci

R dt

(3.1)

V(t) = R_I t

O

e(t) dt (3.2)

d_ = NA dB _

e(t) = N_-_ dt Po NA d__HH

dt

(3.3)

V(t) PoNA I t IdH_dt

RC o

(3.4)

H(t) - RC V(t) (3 5)

poNA

where:

e(t) = search coil voltage (volts)

V(t) = integrator output voltage (volts)

= magnetic flux through coil (webers)

B = magnetic flux density through coil (webers/meter 2)

H = magnetic field intensity (amperes/meter)

Po = permeability of free space (4_xl0-_weber/amper e-meter)

N = no. of turns of search coil

A = area of search coil (meter 2)

I00
R = resistance of integrator coil (ohm, _)

C = capacitance of integrator (farads, F)

t = time (seconds)

Equation (3.5) gives the theoretical relationship between

the magnetic field intensity at a given point and the integrated

signal of a search coil responding to the changing magnetic field

at that point.

The coil calibration factors was determined experimentally

using the test setup shown in Figure 3.12.

The calibration test fixture was approximately i0 ft long

with a center wire conductor which was connected to the simulated

lightning current generator "high" side. The conductor was connected to 16 return wires at the far end of the test fixture

(Figure 3.12 shows only four return wires to retain clarity). The

return wires were strung back to the near end of the fixture in a

radial configuration with each return wire equidistant from the

center conductor. The return wires were fastened together and

connected to the ground side of the generator. The search coil

being calibrated was located at a known distance, r, from the

center conductor of the test fixture with its axis perpendicular

to the conductor.

A current impulse which was measured by a current transformer

and recorded by a Tektronix 535A oscilloscope, was applied to the

center conductor, and the voltage induced in the search coil was

measured by the integrator and recorded on a Tektronix 7704A

oscilloscope. The magnetic field intensity (H) at the search coil

location was calculated from the relationship

I

H - (3.6) 2_d

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RI

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where the field intensity (H) is in A/m, the current, I, is in

amperes and the distance is in meters. The coil calibration

factor, then, is H/V or the magnetic field intensity per volt of

the integrated coil signal and has the units of amps per meter

per volt.

The test circuit used for the measurement of magnetic field

intensity within the wing structure is shown in Figure 3.13.

The RG-58/U coaxial cable measurement lines, which conducted

the search coil signals to the measurement and recording instruments, were routed as far as practical within the wing structure

to utilize its shielding capabilities. Externally, the cables

were shielded by aluminum foil and metal conduit until their entrance into the RF shielded room. The coil signals were measured

by a Tektronix 7704A oscilloscope with a Type 7A13 preamplifier.

Measurements of the magnetic field intensities were made in

the three orthogonal axes as shown in Figure 3.13.

The magnetic fields were generated by simulated lightning

currents as they flowed through the wing structure from the taxi

light housing at the wing tip to their exit point from the forward, main and rear spars at the root end of the wing. Unipolar

and oscillatory current waves were utilized for these tests.

3.3.4 Test results

3.3.4.1 Current distribution

A summary of test results is shown in Table 3-2 and typical

current oscillograms are shown in Figure 3.14. The locations of

the test measurements are given in Figure 3.15.

The noise level measured during test 2409 was 2 amperes.

Currents in the fuel lines ranged from a maximum of 16 amperes to

a minimum of less than 1 ampere for a test current of 12 kA. Increasing the test current to 88 kA resulted in a fuel line current

level of 160 amperes and a noise level of 30 amperes. At either

level, the currents measured in the fuel lines were less that 0.2%

of the test current.

The distribution of currents in the spar was recorded during

tests 2423-2426. During a 19 kA test current strike, a current

of 2.5 kA was measured in the forward spar and 190A was measured in

rear spar. The data indicate that the main spar carried 86% of

the stroke current, the forward spar 13% and the rear spar 1%.

103
ORIGINAL PAGE 19

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Table 3-2 - Current Distribution in Fuel Lines and Spars

of the Bonded Alumin'um Wing

Test

No.

2409

2410

2411

2412

2413

2414

2421

2422

2423

2424

2425

2426

Test Measured

Measurement Current Current

Location kA A

Fuel line

Fuel line

Fuel line

Fuel line

Fuel line

Fuel line

Fuel line

Fuel line

Forward

Spar

Forward

Spar

Rear Spar

Rear Spar

12 2

12 2

12 16

12 14

12 <I

12 <i

88 160

88 30

19 2500

19 800

19 60

19 190

Remarks

Noise measurement

Noise measurement

Noise measurement

Noise measurement

Test 2411

5kA/div

Test Current

20_s/div 4V/div 20_s/div

Measured Current

Figure 3.14 - Typical Current Oscillograms Obtained during

Current Distribution Measurements in Bonded

Aluminum Wing.

105
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3.3.4.2 Bond line voltages

Measurements of bond line voltages were made in the fuel

tank interior between WS 153.75 and WS 164.65 as shown in Figure

3.16. The test results are summarized in Table 3-3.

Table 3-3 - Bond Line Voltages Inside Bonded Aluminum Wing

Fuel Tank. Simulated Strike to Taxi Light

Housing.

Measured

Test Input Measurement Voltage

No. Current Location (a) (V) Remarks

2236 llkA,17kHz -- 0.04

2237 llkA,17kHz

2238 llkA,17kHz

2239 llkA,17kHz

2240 llkA,17kHz

2241 llkA,17kHz

2242 llkA,17kHz

2243 llkA,17kHz

2244 7kA,4x50_s

2245 7kA,4x50_s

2246 7kA,4x50Ds

2247 7kA,4x50_s

2248 7kA,4x50_s

2249 7kA 4x50_s

2250 7kA 4x50_s

2251 7kA 4x50_s

(a) See Figure 3.16

A

C

D

E

B

G

F

F

G

B

E

D

C

A

0.I0

0 14

0 i0

0 I0

0 Ii

0 O5

0 06

0 O5

0 03

0 I0

0 04

0 07

0 I0

0 08

0.02

Noise check - Test leads

shorted to each other

inside tank

Noise check - Test leads

shorted to each other

inside tank

Readings obtained with each of the input current waves were

0.14 volts or less. The noise level, which was 0.04 volts or

less, was greater than 10% of the measured values. The low test

readings seemed to indicate that the bond line voltage readings

were primarily due to noise pickup. Apparently the adhesive

joints were not isolated but had metal-to-metal contact at one

or more locations.

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3.3.4.3 Ignition source tests

Table 3-4 is a summary of results. The test photographs

showed that no light sources existed during the simulated lightning strikes to the taxi light housing indicating the lack of a

fuel vapor ignition source. Figures 3.17 through 3.20 are

photographs of the fuel tank interior during these tests.

Table 3-4 - Summary of Ignition Source Test Results

on ' Bonded Aluminum Wing

Test

Test Current Camera Lens Photograph

No. (kA) Direction Results Figure No.

2417 I00 Inboard No light 3.17

2418 85 Lower wing No light 3.18

surface

2419 90 Upper wing No light 3.19

surface

2420 90 Outboard No light 3.20

3.3.4.4 Induced voltage in wing wiring

Table 3-5 is a complete record of the induced voltage measurement test results. In addition to the measured values, the test

results have been extrapolated for an average stroke of 20 kA and

a severe stroke of 200 kA.

Table 3-6 summarizes the induced voltage measurement test

results which have been grouped by the circuit return path configuration. These paths include an independent return within the wing,

airframe return within the wing, and airframe return from the plastic wing tip. Representative circuits for each of these configurations are presented in Figures 3.21 to 3.24. The figures include

wire numbers, connectors, and pertinent interfacing electronics.

The accompanying oscillograms, which show the induced voltage waveforms, are referenced by arrows to the measurement point. Each

oscillogram includes the amplitude, sweep scale settings and test

number.

3.3.4.4.1 Independent return within wing= Figure 3.21 shows voltage induced in the righthand stall switch to battery line W4C22

by the 4x50_s unipolar and the 17 kHz oscillatory simulated

lightning current waves. Since the wing structure is not utilized

for circuit return, the induced voltage is proportional only to

the rate of change of the simulated lightning current (i.e. there

is no structural IR component). This concept is illustrated in

Table 3-7. The initial rates-of-rise of the unipolar and

109
ORIGINAL PAGE |_

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Figure 3.17 - Interior of Fuel Tank during Ignition Source Tests.

(camera view - inboard) No Light Sources.

Figure 3.18 - Interior of Fuel Tank during Ignition Source Tests.

(camera view-lower wing surface) No Light Sources.

ii0
ORIGINAL PAGE lg

OF POOR QUALITY

Figure 3.19 - Interior of Fuel Tank During Ignition Source

Tests. (camera view-upper wing surface)

No Light Sources.

Figure 3.20 - Interior of Fuel Tank During Ignition Source

Tests. (camera view-outboard) No Light Sources.

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Range of

Peak Voltages

(v)

20-40

Table 3-6 - Summffry of Induced Voltage Measurements

in Bonded Aluminum Wing

Traveling

Wave form S t ructural Wave

and Strike IR Frequencies Figure

Location (_) (MHz) No's.

Independent Return within Wing

7kA, 4x50_s --

Taxi Light

3.21,3.22

15-50

Airframe Return within Wing

7kA, 4x50Ds 0.3(10 -3 ) 8

Taxi Light

Airframe Return from Plastic Wing Tip

3.23

20 12kA, 17kHz

Taxi Light

2900 7kA,4x50_s 14(10 -3 ) 8

NAV Light

3.24

II,000 85kA, 17kHz

NAV Light

Table 3-7 - Induced Voltage Measurements Rate-of-Rise Comparison

in Bonded Aluminum Wing.

Rate-of-Rise Induced Voltage

Circuit Waveform (A/s) (V)

Stall Switch 7kA, 4x50Ds _6.0x109 40

to Battery

12kA, 17 kHz _l.4x109 I0

Stall Switch 7kA, 4x50_s _6.0x109 40

to Horn

12kA, 17 kHz _l.4x109 8

116
ORIGINAL PAGE I11

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oscillatory currents were approximated by dividing the peak test

voltage by the circuit inductance. The calculations indicated

that the initial rate-of-rise of the unipolar wave was approximately four times that of the oscillatory wave. The peak values

of the voltages induced by these currents also approximated a

4:1 ratio.

The 8 MHz oscillations which appear on the induced voltage

oscillograms were due to traveling waves being reflected at each

end of the wing.

A second example illustrating the relationship of peak induced voltage to current rate-of-rise is shown in Figure 3.22.

Induced voltages were measured on the right hand stall switch to

horn circuit (line W5A22) during unipolar and oscillatory current

strikes to the wing taxi light housing. The measurement oscillograms indicated induced voltage levels of 35 to 40 volts for the

unipolar current wave and approximately 8 volts for the oscillatory wave. As before, the ratio of the two induced voltage

levels was in the same proportion as the initial rates-of-rise

of the test current waves as shown in Table 3-7.

3.3.4.4.2 Airframe return within wing.- The electrical circuits

in this group utilize the wing airframe for measurement return.

Figure 3.23 illustrates a representative circuit. The measurement oscillogram indicated that the voltage on line L8AI8 was the

sum of two voltage components. One component was proportional

to the test current rate-of-rise as previously illustrated in

Table 3-7. The second component was the result of structural

voltage rise (IR) due to the test current flowing in the wing skin.

The equation for the measured voltage can be written as follows:

V = IR + L di

d-_ (3.7)

where IR represents the component due to the structural voltage

rise and L di/dt represents the component proportional to the

rate-of-rise of the test current. The wing resistance (between

the taxi light and wing root) can be calculated from equation 3.7

by measuring the voltage and current amplitudes at a time, t, at

which the rate-of-change of the test current is zero (di/dt = 0).

Equation 3.7 then becomes

V = IR

R-v I

I di/dt --0 (3.8)

For the unipolar current wave, di/di=0 at t = 5 _s (see

Figure 3.5). At 5 us, the test current amplitude was 7 kA and

the voltage amplitude was 2 volts (Figure 3.23, test 2150).

121
The wing resistance, from equation 3.8, is then

V R -

I di/dt = 0

2V R- 7kA

R = 0.3 (10-3)f_

3.3.4.4.3 Airframe return from plastic wing tip.- Figure 3.24

illustrates the right hand navigation light circuit and measurement oscillograms. Because the navigation light was located at

the end of the plastic wing tip, a separate wire was necessary to

connect the navigation light return to airframe ground.

Figure 3.24, test 2151, shows that the induced voltage was

the result of a component proportional to the test current-rateof-rise and a component due to the wing voltage rise (IR).

The wing resistance can be calculated as before using

equation 3.8:

R = _Idi/dt=0

As before, di/dt = 0 at time t= 5 _s. At t= 5 _s the test

current amplitude was 7 kA and the voltage amplitude was 175 volts,

(see Figure 3.24, test 2151). Substituting these values into

equation 3.8 yields the wing resistance as follows:

R= Vl di/dt= 0

175V

R -

7kA

R = 25(I0-3)_

This calculation results in a higher wing resistance than that

previously calculated for the taxi light circuit. The additional

circuit return wires and connections required in the navigation

light circuit may account for this difference.

It should also be noted that the peak induced voltage magnitude of the navigational light circuit was greater than that of

the electrical circuits located entirely within the metallic

wing. This is because the plastic wing tip did not provide

electromagnetic shielding for the navigation light circuit wires,

thus allowing greater magnetic flux interaction with these

122
ORIGINALPAGE I_

OF POOR QUALITY

wires than for those located entirely within the shielded environment of the metallic wing structure.

Test 2152 of Figure 3.24 shows that the electrical "noise" in

the measurement lines was less than 10% of the voltage measurement

signal.

3.3.4.5 Magnetic field measurements

Table 3-8 summarizes the results of the search coil calibration tests. Voltage and current oscillograms recorded during

calibration of the single layer 20-turn coil are shown in Figure

3.25

Table 3-8 Summary of Results Obtained during

Search Coil Calibration Tests

Coil

Description

Magnetic

Field

Intensity(H)

Integrator A/ft _/mx 103)

Integrated Calibration

Voltage Factor

V A/ft/V A/m/V

Multi-layer I0 _s 935 (3.07) II.50 0.81 (0.28)

20 Turn i ms 935 (3.07) 0.32 2.90 (9.60)

Single layer I0 _s 970 (3.18) 15.50 0.063 (0.21)

20 Turn I ms 935 (3.07) 0.26 3.60 (12.00)

Single layer I0 _s 935 (3.07)

I00 Turn i ms 902 (2.96)

45.00 0.021 (0.068)

0.63 1.40 (4.60)

2kA/div 20_s/div

Test Current

0.1V/div 20us/div

Integrated Coil Signal

Figure 3.25 - Oscillograms Recorded during Calibration of a Single

Layer 20-turn Search Coil with I ms Integrator.

123
The peak amplitude of the test current, as shown in

Figure 3.25, was 5.4 kA, and the distance, d, of the coil from

the current-carrying conductor was 0.92 ft. Substitution of

these values into equation 3.9 yields the magnetic field intensity, H, at the position of the coil.

I H -

2_d

H = 5.4kA

2_(0.92ft)

H = 935 A/ft (3.07(103 ) A/m)

The field intensity value of 935 A/ft (3.07(103 ) A/m) obtained above, divided by the integrated coil signal of 0.26 volts

(obtained from Figure 3.25), results in a calibration factor for

this coil-integrator combination of 3.6 (103 ) A/ft/V (11.8(103 )

A/m/V).

Initial measurements of the field intensity within the fuel

tank are summarized in Table 3-9. A typical measurement oscillogram is shown in Figure 3.26.

The search coil was oriented in each of the three orthogonal

axes for the tests as defined in Figure 3.13. Measurements were

initiated with the coil positioned midway between the upper and

lower skin surfaces in the vicinity of the fuel filler cap. The

fuel filler cap was in place during the tests to simulate the

actual structural configuration that exists during a lightning

strike. Table 3-9 indicates that the magnetic intensity amplitude

readings ranged from 4.6 to 7.9 A/ft (15 to 26 A/m). Electrical

noise in the measurement cables was equivalent to 4.6 A/ft

(15 A/m).

For comparison purposes, additional tests were made with

the coil positioned midway between the fuel filler cap and the

wing leading edge, at the wing leading edge, and at the wing

trailing edge. Table 3-9 shows that the field intensity measurement amplitudes were higher at the leading and trailing edges than

at the fuel filler cap location, but not significantly so. A

noise measurement made with the search coil located at the wing

trailing edge, yielded an equivalent field intensity level of

6.4 A/ft (21 A/m).

These measurements indicated that the amplitude of the magnetic field intensity within the wing tank was not significantly

higher than the amplitude of the electrical "noise" within the

measurement cables.

124
ORIGINAL PAGE lg

OF POOR QUALITY

Table 3-9 - Summary of Magnetic Field Intensity Measurements within a Bonded Aluminum Wing Fuel Tank, 20-turn Search

Coil, 1 ms Passive Integrator. 2.4 kA 4x50_s Conducted

Current Strike t0 Taxi Light Housing.

Coil Axis

Orientation (a)

Search Magnetic

Coil Intensity(H)

Location A/ft (A/m)

X

Y

Z

midway between skin

surfaces at fuel

cap location

11

7.3 (24)

7.9 (26)

4.6 (15)

4.6 (15)

X

Y

Z

midway between skin

surfaces and midway

between fuel cap

and leading edge

I!

9.5 (31)

7.9 (26)

7.9 (26)

X

Y

Z

midway between skin

surfaces at trailing

edge

7.9 (26)

9.5 (31)

7.3 (24)

6.4 (21)

Z midway between skin

surfaces at leading

edge

(a) See Figure 3.13.

9.2 (30)

Test 2265

Notes

Noise measurement

Noise measurement

7.2A/ft/div 20_s/div

(23.6 A/m/div)

Figure 3.26 - Typical Magnetic Field Intensity Oscillogram

Obtained during Measurements within Fuel Tank

of Bonded Aluminum Wing. 125
A search coil wound with I00 turns of wire was utilized during additional tests of the magnetic fields within the wing fuel

tank. Since the output signal or the search coil is proportional

to the number of turns of wire in its construction, as shown by

equation 3.3, theory would predict that the signal amplitude from

this coil would be approximately 5 times that of the 20-turn coil

used during the previous tests.

e(t)100T coil

e(t) 20T coil

NI00T d_/dt

N20 T d_/dt

(3.10)

e(t) 100T coil NI00T I00

e(t) 20T coil N 20T 20

- 5 (3.11)

Two sets of tests were conducted with this coil. The first

set of tests, with the search coil located in the vicinity of the

fuel filler cap, was performed with a 2.4 kA, 4x50 _s, conducted

entry current to the taxi light housing. The second set of tests

was performed with a 16 kHz, II kA oscillatory current wave. For

comparison purposes, each set of tests consisted of measurements

taken with the fuel cap in place and with it removed. Figure 3.27

shows oscillograms of the test current waveforms and typical integrated coil response oscillograms. Tables 3-10 and 3-11 summarize

the measurement results of these tests.

Table 3-10 - Summary of Magnetic Field Intensity Tests within an

Aluminum Wins Fuel Tank near the Fuel Filler Cap.

i00 Turn Search Coil, I ms Passive Integrator.

214 kA, 4 x50 Us Conducted Current Strike to Taxi

Lisht Housing.

(a) See Figure 3.13

Magnetic

Coil Axis Intensity (H)

Orientation (a) A/ft (A/m)

x 5.5 (18)

" 7.0 (23)

Y 8.5 (28)

" 5.5 (18)

Z 20.0 (64)

" 20.0 (.64)

1.5 (5)

Notes

Fuel cap ON

Fuel cap OFF

Fuel cap ON

Fuel cap OFF

Fuel cap ON

Fuel cap OFF

Fuel cap ON - Noise check

126
ORIGINAL PAGE Ig

OF POOR QUALITY

TEST CURRENT

INTEGRATED COIL RESPONSE

(H)

Test 2256

Test 2292

2.4kA

8.5A/ft

(28A/n0

IkA/div 4x50_s 20_s/div 7.6A/ft/div 20_s/div

(25A/m/div)

Test 2305

Test 2306

llkA

5kA/div 16kHz 20_s/div 25.9A/ft/div 20_s/div

(85A/m/div)

Figure 3.27 - Typical Test Current and Integrated Coil

Response Oscillograms. I00 turn search

coil; I ms passive integrator. Conducted

entry strike to taxi light housing of

bonded aluminum wing. Measurement inside

fuel tank with filler cap on. See Table 3-12.

127
Table 3-11 - Summary of Magnetic Field Intensity Tests

within a Bonded Aluminum Wing Fuel Tank

near the Fuel Filler Cap. i00 Turn Search

Coil, I ms Passive Integrator. llkA, 16 kHz

Conducted Current Strike to Taxi Light Housing

Orientation (a)

Magnetic

Intensity (H)

A/ft (A/m) Notes

X 14.0 (46)

" 14.0 (46)

Y 43.0 (140)

" 43.0 (140)

Z 85.0 (280)

" 98.0 (320)

-- 1.5 (5)

-- 2.8 (9)

Fuel cap ON

Fuel cap OFF

Fuel cap ON

Fuel cap OFF

Fuel cap ON

Fuel cap OFF

Fuel cap ON - Noise check

Fuel cap OFF- Noise check

(a) See Figure 3.13

The test results shown in Tables 3-10 and 3-11 showed that

the presence or absence of the fuel filler cap made little difference in the test results. This indicated that the internal

fields were the result of diffusion flux rather than aperture

flux penetration. Tables 3-10 and 3-11 also indicate that stronger

magnetic fields existed in the Z axis orientation than in either

the X or Y axis orientations. The Z axis field, as defined by

Figure 3.13, can be represented by field vectors normal to the

upper and lower wing skin surfaces. This field is the result of

wing skin penetration by the external magnetic field lines of

flux. The Y-axis field intensity measurements were approximately

one half the magnitude of the Z axis amplitude levels recorded

during the II kA oscillatory tests (Table 3-11) and were in the

noise level range for the 2.4 kA unipolar current tests (Table 3-10).

The X axis orientation field amplitude levels were less than I0

times the amplitude level of the electrical noise in the measurement cable for both sets of tests.

A comparison of the fuel tank internal field measurements is

shown in Table 3-12.

Magnetic field intensity measurements along the external surface of the wing were recorded utilizing a 20 turn multi-layer

search coil and a I ms passive integrator. The wing circumference

at WS 160.75 was divided into eight equidistant sections. Figure

3.28 is a schematic representation of the wing outline showing the

128
Table 3-12 - Comparison of Magnetic Field Intensity

Measurements in A/ft (A/M) within a -

Bonded Aluminum Wing Fuel Tank near

the Fuel Filler Cap fgr Conducted Current Strikes to the Taxi Light Housing.

Coil

Axis

Orientation (a)

20 Turn Coil

Filler Cap ON

I00 Turn Coil

Filler Cap ON

2.4 kA, 4x50 _s Test Current

X Y Z

7.3 (24) 7.9 (26) 4.6 (15)

5.5 (18) 8.5 (28) 20.0 (64)

Electrical

Noise

Reading

4.6 (15)

1.5 (5)

i00 Turn Coil

Filler Cap OFF

7.0 (23) 5.5 (18) 20.0 (64)

II kA, 16 kHz Test Current

I00 Turn Coil

Filler Cap ON

14 (46) 43 (i00) 85 (280) 1.5 (5)

i00 Turn Coil

Filler Cap OFF

14 (46) 43 (140) 93 (320) 2.8 (9)

(a) See Figure 3.13

129
oRIGINAL PAGE IS

OF POOR QUALITY

field measurement locations and the magnetic intensity amplitudes.

The measurement amplitudes were greater than eight times the noise

level.

Upper Wing Surface

3 2

4

5 <__i i Leading Edge

Control 7 8

Surface Lower Wing Surface

Measurement

Location

Magnetic Field Intensity(H)

A/ft x 103 (A/m x 103 )

i 0.61

2 2.10

3 0.89

4 0.28

5 0.28

6 0.18

7 1.00

8 I. 70

(2 0)

(6 9)

(2 9)

(0 9)

(0 9)

(0 6)

(3 4)

(5 5)

Figure 3.28 - External Magnetic Field Intensity Measurements

and Locations for a Bonded Aluminum Wing.

Distance between measurement locations was

approximately 1.2 ft. Test current was 12 kA.

The average magnetic field intensity around a current-carrying conductor is given by the equation

I (3.12)

HAV - p

where I is the conducted current and P is the circumference. For

a conductor with an irregular shape, such as an aircraft wing,

the field intensity will be greater than average where the radius

of curvature is smaller than average (such as the leading and

trailing edges) and less than average where the radius of curvature is greater than average (as midway between the leading and

trailing edge).

130
For the tested wing, the circumference was 9.5 ft at

WS 160.75 and the test current level was 12 kA. From equation

(3.12), then, the average magnetic field intensity along the

external wing surface was

I 12kA

HAV - p -

HAV = 1.3(103)A/ft (4.3(103 ) A/m) (3.13)

Magnetic field intensity measurements were made at locations

I through 8 as shown in Figure 3.28. They indicated that the

tested wing did not follow the theoretical intensity distribution.

The maximum field intensity amplitudes occurred at locations 3

and 7. A sharp drop in amplitude is recorded at locations 4 and

6 with an equally low reading at the wing trailing edge. Location I at the wing leading edge also recorded an intensity level

lower than average but was greater than twice the reading recorded

at the trailing edge.

Comparison of the field distribution readings shown in Figure

3.28 with the photograph of the wing laboratory setup reveals an

explanation for the divergence of the wing field pattern from that

predicted by theory. The aluminum flashing, which conducted the

test currents from the generator to the taxi light housing, was

positioned along the wing skin approximately 1/3 to 1/2 of the

distance between the leading and trailing edges. This corresponds

to the region of highest field intensity levels. Apparently, the

close proximity of the test current-carrying conductor to the skin

surface strongly influenced the magnetic field distribution in

their immediate vicinity.

3.4 gr/E Wing

3.4.1 Specimen description

The gr/E wing was fabricated as a single structure. The wing

box extended from the fuselage centerline to the close-out rib

226 in. outboard. With the exception of the leading edge and some

control surfaces, the wing was constructed of gr/E. The skin

thicknesses ranged from 0.05 in. at the wing tip to 0.125 in.

inboard. The wing had three full length spars with an integral

fuel tank between the front and the rear spars and extending

throughout much of the wing. The leading edge, which was to be

fabricated of nonconducting composite beneath a de-ice boot, was

not fastened to the wing structure during the test program. The

wing tip was fabricated of gr/E containing interwoven wires for

lightning protection. The wing was complete with ailerons, flaps,

internal sealants, wiring and components.

131
Electrical circuits contained within the tested wing included navigation light, position light, strobe light, trim tab

actuator motor, and capacitance-type fuel quantity probes. The

quantity probe wiring was located within the fuel tanks; the wiring for all other circuits was routed through an aluminum conduit

located within the leading edge of the wing.

The gr/E skins were adhesively bonded to the spars and ribs.

Some mechanical fasteners were utilized to provide additional

strength and to fasten hardware such as control surfaces, hinges,

etc. to the wing.

Rib, access door, connector, and test locations hereafter

referred to in this report will be referenced in inches from the

fuselage center line by a wing station (WS) number. For example,

the wing box closeout rib which is located 226 in. outboard of the

fuselage center line, would be designated WS 226.

A drawing of the wing is shown in Figure 3.29.

3.4.2 Test setup

A photograph of the test setup is shown in Figure 3.30. The

wing was positioned on a specially constructed wooden framework

with the tip of the right wing positioned at the simulated lightning current generator. The wing was positioned with its lower

skin surface facing up to allow access to the fuel tank access

doors. The framework structure provided for the positioning of

eight equally-spaced radial lightning current return lines. These

lines originated at a metallic ground plane, which simulated the

conductive cabin floor located in the aircraft fuselage, and were

terminated at generator "ground". By spacing the return lines in

an equidistant, radial pattern around the wing, the magnetic fields

created by these wires would have little or minimal effect on

wing measurements; this is because the magnetic fields surrounding

each wire would cancel in the areas between them. This is shown

schematically in Figure 3.31.

The simulated ground plane is shown in the photo of Figure

3.32.

The output of the generator "high" side was connected to the

static discharger which was located on the trailing edge of the

wing tip. The lab setup simulated a strike to the wing tip with

current flow through the wing skin to an exit point elsewhere on

the aircraft.

The simulated lightning currents were generated by a 20 _F,

i00 kV high current capacitor bank in conjunction with waveshaping

elements. The currents were measured by a I00:i current transformer and recorded by an oscilloscope located within an RF

shielded room.

132
ORIGINAL PAGE I_

POOR QUALITY

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ORIGINAL PAGE 19

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135
ORiGiNAL FAGE Ig

OE POOR QUALITY

©

Re turn

Wire

O

O

Wing

Magnetic

© Flux Line

Figure 3.31 - Cancellation of Return Wire Magnetic

Fields in Vicinity of gr/E Wing.

136

Figure 3.32 - Photo Showing Connections to

Simulated Ground Plane in gr/E Wing.
3.4.3 Test procedures

ORFGINAL _E 1&

OF POOR QUALrrY

3.4.3.1 Current distribution

Measurements were made of current through the following

structural elements:

Conduit in wing leading edge (aluminum)

Conduit braid (copper)

Wing attach fittings (titanium)

Drag angles (gr/E)

Hydraulic lines (aluminum)

Fuel lines (steel reinforced)

Each of the above elements was connected to the ground plane;

for the drag angles, four equally spaced copper braids connected

each drag angle to the ground plane. The wing attach fittings

(dagger fittings) were connected to the ground plane through

0.087 in. gr/E skins to simulate actual conditions in the aircraft,

as shown on Figure 3.33.

Figure 3.33 - Connections of Drag Angle and Wing

Attach Fittings (dagger fittings)

to Ground Plane in gr/E Wing.

137
The current measurements were made by a Pearson Model IIOA

current transformer which had an IT product of 0.5 A.s and a

frequency response of i Hz to 35 MHz. RG 58 coaxial conductor

cable routed the measurement signal from the current transformer

to the RF shielded room where the measurement was recorded by a

Tektronix 535A oscilloscope with a type IAI preamplifier.

Test currents of 900 A and 4450 A peak amplitude and 4x50_s

waveform were conducted into the static discharger located at

the trailing edge of the right wing tip and flowed through the

wing to the ground plane where they were conducted by the radial

return wires to the lightning current generator "ground". The

test current peak amplitudes were kept low to avoid "conditioning" the wing structure.

3.4.3.2 Bond line and fuel probe voltages

Lightning currents flowing through the resistive gr/E wing

skin can raise its voltage potential with respect to structural

or fuel system components which may be electrically isolated

from the skins. If the voltage potential difference between

the skin and these components exceeds the withstand capability

of the insulating medium (adhesive or air), breakdown will occur

and the resulting spark could be a possible ignition source.

Measurements were recorded across bond lines located within

the wing structure between locations WS 28 and WS 212. The bond

lines included upper skin to front, center, and rear spars and

stringers.

The number of measurements which could be made at each wing

location was determined by the physical accessibility of the

locations and the ability to insure a low resistance electrical

contact between the skin and measurement leads. Since the epoxy

resin electrically insulated the wing skin surface, modified test

procedures were employed to insure low contact resistance between

the skin and measurement leads. Experimentation with several

test methods resulted in the following procedure. The insulating

epoxy resin surface was removed by sanding the surface area with

medium grade sandpaper. Copper braid, approximately 0.75 in.

in length, was taped to the sanded surface with aluminum foil

tape, with a sufficient length of braid extending beyond the

aluminum tape to allow connection of the voltage measurement

probe. A wooden dowel was wedged between the aluminum foil/

copper braid and the opposite skin surface to provide the pressure

needed to insure good electrical contact between the braid and

the skin surface. This procedure was repeated at the adjacent

spar or stringer measurement point. The contact resistance of

each copper braid location was measured with an ohmmeter and

the whole procedure repeated if the contact resistance was too

high.

138
ORIGINAL r *_'_ =_

OF POOR QUALITY

The measurement signals were conducted by twinaxial cable to

the recording oscilloscope (Tektronix 535) located within the RF

shielded room.

Measurements of fuel probe to skin voltages were made using

a test procedure similar to that employed during bond line voltage

measurements. One measurement lead was connected to a fuel probe

terminal and the second lead was connected to the adjacent wing

skin surface using the technique described above. The fuel probe

circuits were electrically tied to the ground plane through a

dummy fuel quantity meter circuit.

Initial measurements were made using test currents of 900A -

1800A conducted into the static discharger on the wing tip. During the latter part of the test series, test current levels of

190 kA were injected into the aluminum rib at wing location

WS 226, the wing tip attachment point. A photograph of the test

setup is shown in Figure 3.34.

Figure 3.34 - Lab Setup for Bond Line and Fuel Probe

Voltage Measurements in gr/E Wing.

3.4.3.3 Induced voltages in wing electrical wiring

Measurements of voltages induced in the leading edge conduit

harness wires were recorded during 36 kA oscillatory test currents

conducted into the navigation light housing. In addition, one

139
OF POOR QU_,LIrY

170 kA oscillatory test current strike was applied across a I in.

air gap to the navigation light and the voltage in the navigation

light circuit conductor measured with respect to the conduit

braid. A schematic of the measurement circuit is shown in

Figure 3.35 for the 170 kA test current strike to the navigation

light. The conduit braid was connected to the simulated ground

plane (not shown). The circuit wire under test was pulled through

the conduit braid at the connector located at the inboard end of

the conduit and measured with respect to the conduit braid by a

I00:i voltage measurement divider. The divider consisted of two

measurement leads each of which had 5000 _ of resistance at the

measurement end and 50 _ to ground at the recording (oscilloscope)

end. The signals measured by the divider were conducted through

twinaxial cable to a Tektronix 535 oscilloscope located in an

RF shielded room where they were recorded differentially by the

oscilloscope equipped with a type IA5 preamplifier.

to 50_ terminations

and Tektronix 535A

pe

Electrode_ I

Braid connected,

to ground plane

(not shown)

Wires under test

ulled through

)raid at

:onnector

Figure 3.35 - Circuit of Induced Voltage Measurement Circuit.

Arc Entry Strike to Navigation Light Shown on

gr/E Wing. (Ground plane not shown.)

140
3.4.3.4 Magnetic field measurements

Magnetic field intensity readings were recorded for six wing

interior locations and two wing exterior locations utilizing the

20 turn multilayer search coil and i ms passive integrator employed during field measurement tests on the bonded aluminum

structure described in paragraphs 3.3.3.5 and 3.3.4.5.

The measurement coil was attached to shielded twinaxial cable

which conducted the signals to the I ms passive integrator and

Tektronix 535A oscilloscope located within an RF shielded room.

The internal field measurements were at the location of six wing

access doors. Aluminum flashing was used to shield the access

door opening during the test measurements with the exception of

the door at WS 40, since this door was constructed of kevlar and

would not normally have provided any shielding. The coil was

oriented in each of the three orthogonal axes as defined in

Figure 3.13, i.e., the X-axis direction is inboard to outboard,

the Y-axis direction is leading edge to trailing edge, and the

Z-axis direction is upper to lower skin surfaces.

Measurements of external magnetic fields were made at wing

location WS 182 and WS 80 with the axis of the coil positioned in

the axis of highest field intensity. Measurements at WS 182 were

made at I0 equidistant location points around the wing perimeter.

Measurements at WS 80 were made at five locations along the wing

lower skin which were also located at equidistant points along

the wing surface.

Test currents were conducted into the wing through the static

discharger located on the wing tip trailing edge and were removed

from the wing at the simulated ground plane and returned to

generator "ground" through the radial return line configuration.

3.4.4 Test Results

3.4.4.1 Current distribution

The test results are shown on the schematic drawings of

Figures 3.36 through 3.38. Typical oscillograms of total current

and currents measured in a metallic conductor are shown in Figure

3.39.

Figure 3.36 shows the initial current distribution at time

t = 4_s the time at which the applied current is at its maximum

value. At this time, 74% of the current was conducted by the

graphite skin and the remaining 26% by the metal conduit, control

cables, and hydraulic lines. The gr/E skin current was returned

to the ground plane through the two drag angles and the six dagger

fittings. It is interesting to note that the skin current returning to the ground plane through the dagger fittings did not divide

equally among them. Almost twice as much current returned through

141
ORIGINAL PAGE lg

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Total

- II7A

(14o7%)

99A(12.5%)

(rises to

100A at32us)

84A

"(i0.6%)

42A

C

(i0.7%)

62A

(7.8%)

II2A

(rises to

244A at

©

©

31A

(3.9%)

Total

- 130A

(16,4%)

590A (74%)

of current is

in gr/E

795A

(ioo%)

IN

Total Current

OUT - 792A

Notes:

lq----_ll paths go to ground plane

(not shown)

2. Test Current: 4 x 50Us, 900A peak

3. 0.087ingr/E skins between each

dagger and ground plane

4. Fuel lines are not electrically

connected to front spar due to

sealant.

Figure 3.36 - Current Distribution in gr/E Wing at

T = 4 _s for a 900 A Current during

a Simulated Strike to Static Discharger.

142
ORIGINAL PAW_ W

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8A

151A

(5 _7.)

7A 13_

IIA

(47.)

200A

17A

67.

(29%)

of current remaining in gr/E

structure at

T - 80us

/

Total Current

OUT - 285A

at80us

(input)

Notes:

_-'-_II paths to go to ground plane

(not shown)

2, Test Current: 4 x 50_s, 900A peak

3. 0,087_.gr/E skins between each

dagger and ground plane

4. Fuel lines are not electrically

connected to front spar due to

sealant

Figure 3.37 - Current Distribution in gr/E wing at

T = 80 _s for 900 A Current during

a Simulated Strike to Static Discharger.

143
ORIGINAL PAGE IW

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\

480 at 4_e (I0.

780 at 80_s(48.

430 at 4us (9.7%)

//44.bat 80_s (2.87.)

I. 6kA

at 80 ps

4.&5 kA

(100%)

IN

Notes:

I?'-'--All paths to go to ground plane

(not shown)

2. Test Current: 4 x 50ps,4500A peak

3. 0.087in. gr/E skins

between each dagger and ground

plane

4, Fuel lines are not electrically

connected to front spar due to

sealant

Figure 3.38 - Current Distribution in gr/E wing at

T = 4 Us and T = 80 Us for a

4500 A Current during a Simulated

Strike to Static Dischager.

144
ORIGINAL PAGE 19

OF POOR QUALITY

the center fittings as through either the forward or aft fittings.

It is reasonable to assume that the conduit braid at the leading

edge and the hydraulic lines and control cables at the trailing

edge bled off some of the current that might normally have been

carried by the forward and aft dagger fittings.

Figure 3.37 shows that during the latter part of the strike,

at t = 80 _s, the percentage of current conducted by the skin decreased from 75% to 29% as the metal conduit, control cables, and

hydraulic lines began to conduct a greater percentage of the

input current. Most of the current previously conducted by the

graphite skin was transferred to the metal conduit; a much smaller

percentage was transferred to the control cable and hydraulic

lines since trailing edge currents were also conducted by the wire

mesh along the rear spar.

Total Current

970 A/div

Conduit Current

20 ps/div

400 A/div 20 _s/div

Figure 3.39 - Typical Oscillograms of Total Test Current

and Current in Conduit of gr/E wing.

Additional current measurements were made in the leading edge

conduit braid, the right center dagger fitting, and the hydraulic

lines during a limited number of 4500A 4 x50_s conducted test

145
oRIGINAL PAGE

OF POOR QUALITY

current strikes. The purpose of these additional tests was to

determine if the current distribution varied as a function of peak

current amplitude. The results are shown in Figure 3.38. A comparison of current distributions at 4 _s and 80 _s for the two

levels of test current showed that the current distribution percentages for the 4500 ampere test currents were within 5% of the

readings obtained for the 900 A test current.

Measurements of current amplitudes in the fuel lines were

recorded during 190 kA conducted entry test currents to the aluminum closeout rib at WS 226. The current amplitudes measured in

the large and small diameter fuel lines, which exited at the front

spar, were 200 A and 1.96 kA, respectively.

3.4.4.2 Bond line and fuel probe voltages

A summary of bond line voltage levels is shown in Table 3-13.

Test currents, conducted into the static discharger, included a 900A18 kA, 4x50 _s unipolar wave and a 45kA-190kA, 16 kHz oscillatory

wave (waveforms are shown in Figure 3.5). Positive voltage readings

indicate that the stringer or spar under test was at a lower voltage

potential than the adjacent wing skin; negative readings indicate

that the stringer or spar under test was at a higher voltage potential than the adjacent wing skin. Typical bond line and fuel probe

voltage measurement oscillograms are shown in Figure 3.40.

Test 736 Test 742

IV/div 10_s/div 5V/div 10_s/div

Center Stringer to Skin Fuel Probe to Lower

Bond Line Voltage Skin Voltage

Figure 3.40 - Typical Bond Line and Fuel Probe Voltage

Oscillograms at gr/E Wing Location WS 128

for a 900 A Test Current.

Bond line readings obtained during the 960 A conducted entry

test currents ranged from 0.07 volts to 4.4 volts. Readings were

also obtained at selected locations during application of 4.5 kA

test currents. The bond line voltages recorded at this level were

approximately five times those recorded at 900 A; thus the bond

line voltages increased in the same proportion as the test current

levels for this test current range.

Two 45 kA test current strikes were applied to the wing following the 900 A and 4500 A test current series. The rear stringerto-skin bond line voltage was 63 volts.

146
Table 3-13 - Bond Line Voltages in gr/E Wing

Total Current Level (kA)

0.9 4.5 18.0

REAR SPAR TO UPPER SKIN

45.0

WS 88 0.25 1.50

after 45 kA 0.40 --

WS 128 0.80 4.00 ....

WS 183 1.80 II.00 ....

WS 206 1.90 ......

REAR STRINGER TO UPPER SKIN

WS 88 4.40

after 45 kA 1.80 -- 36 63.0

WS 128 2.50 ......

CENTER SPAR TO UPPER SKIN

WS 28 2 .50 ......

WS 69 0.07 ......

WS 88 I. 40 ......

WS 128 -0.70 ......

WS 212 I. 60 ......

CENTER STRINGER TO UPPER SKIN

WS 88 -0.70 ....

WS 128 2.5 ....

FRONT STRINGER TO UPPER SKIN

WS 28 2.90 ....

WS 69 -1.20 ....

FRONT SPAR TO UPPER SKIN

WS 28 2.20 --

WS 69 1.50 --

WS 88 ....

WS 128 ....

WS 168 ....

WS 183 ....

WS 212 1.60 7.00

190.0

200.0

147
A retest of WS 88 bond lines at 900A indicated that the

rear stringer-to-skin voltage which was previously 4.4 volts,

was 1.8 volts. This second measurement showed better correlation with bond line voltages of other locations indicating that

the initial reading may have resulted from a high resistance

measurement contact. The rear spar-to-skin bond line voltage

was 0.4 volts, an increase of 0.15 volts from the first measurement at this location. Variation in the measurement probe contact resistance may account for the difference in reading.

At the conclusion of an intervening series of tests at

45 kA and 190 kA, voltages at WS 88 were remeasured at test currents of 900 A. A comparison of the remeasured voltages with

the initial voltage readings at 900 A is shown in Table 3-14.

Examination of the data shows that there was no significant

difference in the fuel probe to lower skin voltage. The rear

spar to upper skin bond line voltage increased by 20%, but the

remaining spar and stringer-to-skin bond line voltages decreased

by 20%- 57%.

Table 3-14 - Voltages in gr/E Wing at WS 88 for

900 A Test Currents

Initial Voltage After Voltage After

Voltage 2 Tests at 45 kA 20 Tests at 198 kA

(V) (V) (V)

Fuel Probe to

Lower Skin

19.00 20.00 20.00

BOND LINE

Rear Spar to 0.25 0.40 0.30

Upper Skin

Rear Stringer

to Upper Skin

4.40 i. 80 3.00

Center Stringer

to Upper Skin

-0.70 -- -0.50

Center Spar 1.40 -- 0.60

to Upper Skin

Rib to 1.50 -- 1.20

Upper Skin

Front and rear spar bond line voltages are plotted in

Figure 3.41. The graph indicates that the bond lines in the

vicinity of the wing tip developed higher potentials than those

further inboard. This is consistant with the fact that the

148
ORIGINAL PAGE i_

OF. pOOR QUALITY,

[3- [] Front Spar

0-- ................... <D Rear Spar

i0-

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2 _WS 281yj

WS 69 ........ -

0

0.4 0.8 1.2 1.6 2.0 2.4 2.8 3.2 3.6 4.0

Total Current in Wing (kA)

Figure 3.41 - Bond Line Voltages between Front and Rear

Spars and Upper gr/E Wing Skin.

149
voltage potential developed along a wing, as the result of currents flowing through the wing skin, is highest at the point of

entry (tip) and decreases as the measurement location approaches

the current exit point (wing root reference plane).

Fuel probe to skin voltage measurement levels resulting from

900 A and 190 kA test current levels are plotted in Figure 3.42.

Fuel probe and bond line voltages measured at location WS 88 are

plotted in Figures 3.43 and 3.44. During these tests, the fuel

probes were electrically connected to the reference ground plane

at the wing root but were electrically isolated from the wing

skin at their points of attachment; thus, the fuel probe voltage

readings were actually a measurement of the wing skin potential

with respect to the reference ground plane. It should be noted

that the fuel probe voltage amplitude levels varied in a manner

similar to the bond line voltages discussed above, i.e. the amplitude level was the highest at the point of current entry (tip) and

decreased as the measurement location approached the point of current exit (wing root reference plane). Measurements of the peak

amplitude voltage between the fuel probes and each wing skin

surface indicated that an insignificant difference in voltage

potential existed between the two surfaces.

Figures 3.43 and 3.44 are graphs of fuel probe and bond line

voltages at wing location WS 88 as a function of test current

amplitude. The graphs indicate that the wing skin resistance was

non-linear, i.e., the resistance decreased as the peak current

amplitude increased.

Table 3-15 shows the fuel probe-to-skin voltages measured

during application of 190 kA conducted simulated lightning

strikes to the aluminum closeout rib at WS 226. Calculations of

the predicted voltage level at each fuel probe location for a

200 kA strike current were made based on wing resistance measurements recorded during previous laboratory tests at low current

levels (_3A) and were based on the assumption that WS 28 was at

zero potential.

Table 3-15 shows that the actual voltages which existed in

the wing were 20-40% lower than predicted (with the exception of

WS 29.03). As mentioned above, the resistance of the wing structure material decreased as the conducted current level increased;

thus voltage measurements recorded during the 190 kA strike predictably would be lower than voltages calculated based on resistance data obtained at 3 amperes.

150
ORIGINAL PAGE _r

OF POOR QUALITY

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30

1(

[] Voltages at 0.9kA

O Voltages at 190kA

©

3000

O

O_

r-q

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2000

r-q

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I000

2'8 !

69 J8 1128 16_

Wing Station (WS, inches from Center Line)

Figure 3.42 - Voltages between Fuel Probes

and Top or Bottom Skins at

Each Probe for gr/E Wing.

151
oRIGiNAL F._,.=_ Y_

OF poOR QUALITY

/

/

/

/

/

/

/

/

/

/

40 80 120 160

Current in Wing (Amps x 103 )

Bond Line

Probe

2000

Q

t

7

1500¢

,r

C

pc

500

0

200

Figure 3.43 - Fuel Probe and Bond Line Voltages at WS 88

of the gr/E Wing.

152
ORIGINAL PAGE I_

OF POOR QUALII"Y

60

45

_0

.r_

3O

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0

>

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0

CQ

/

/

/

/

/

/

/

/

j/

jJ

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/

/"

1/

//

/

//

i,, ̧

Jjl/// //

I0 20 30 40

Bond Line

Probe

Current in Wing (Amps x 103 )

/

/

51

Figure 3.44 - Fuel Probe and Bond Line Voltages at WS 88

of the gr/E Wing.

6O0

OD

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0

0

300_

0

150

153
Table 3-15 - Voltage between Fuel Probe and Adjacent

Skin for Simulated Strike to WS 226 Rib

in gr/E Wing

Predicted Measured

Probe Wing at 200kA(a) at 190kA

No. Station (V) (V)

I WS 29.03 40 1200

2 WS 70.53 1720 --

3 WS 89.03 2470 1950

4 WS129.03 4080 --

5 WS169.03 5700 3200

(a) from structural resistance measurements made at 3A

3.4.4.3 Induced voltages in wing electrical wiring

The results of the induced voltage test measurements are

summarized in Table 3-16 and typical test current and measurement

voltage oscillograms are shown in Figure 3.45.

The 36 kA conducted entry strikes resulted in measured voltages of 80 to 300 volts. If a linear relationship between the

measured voltage and the test current amplitude is assumed, then

the range of voltages expected to appear on the conduit wire

harness circuits for a full threat 200 kA strike would be 450 V

to 1.8 kV. The 170 kA arc entry test to the navigation light

resulted in a peak voltage of 1350 volts. Again, assuming a

linear relationship, the voltage reading for a full threat level

strike current of 200 kA would be 1.6 kV.

3.4.4.4 Magnetic field measurements

Prior to initiation of the magnetic field measurements, the

integrated output response of the 20 turn multi-layer coil was

compared with that of the test current. The result, which is

shown in the oscillogram of Figure 3.46, indicated close correlation between the integrated coil response and the test current

waveform.

154
Table 3-16 - Summary of Voltages Measured in Leading Edge Conduit Wire Bundle Circuits during Simulated Lightn_ng Strikes to Navigation Light in gr/E Wing

Peak Test

Circuit Current Entry Current

Measured Configuration (kA)

Peak Measurement

Voltage

(v)

Navigation

and Position

Light

(Pin 8)

Conducted

Entry to

NAV Light

Housing

36 330

Pot Control

(Pin I)

fv 120

Pot Control

(Pin 2)

w!

120

Power to

Motor

(Pin 6)

I! t! 230

Power to

Motor

(Pin 7)

11 230

Power to

Power Supply

(Pin II)

I! 320

Power to

Power Supply

(Pin 12)

T! 320

Pot Control

(Pin 3)

260

Chassis Ground

(Pin 4)

11 80

Noise Check

Navigation

and Position

Light (Pin 8)

It

Arc Entry to

NAV Light

t!

170

4

1350

155
ORIGINAL PAGE

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Test 128

50kA/div 10_s/div

Test Current

1000V/div 10_s/div

Voltage on Navigation

Light Wire

Figure 3.45 - Test Current and Voltage Measurement Oscillograms

Obtained during 170 kA Arc Entry Strike to Navigation Light in gr/E Wing.

Test 817

--integrated coil

response

--test current

Integrated Coil Response=0.05V/div

Test Current=200 A/div

i0 _s/div

Figure 3.46 - Comparison of Search Coil Integrated Output

Response with Test Current.

156
Figure 3.47 is a schematic representation of the wing showing the measurement location with the magnitude and direction of

the magnetic field readings obtained during 900 A simulated strikes

conducted into the wing tip static discharger. The oscillograms

show the search coil response waveforms for the Z axis (skin-toskin) magnetic field for each of the measurement locations.

Examination of the data presented in Figure 3.47 reveals

that the internal magnetic field intensity was highest in the Zaxis direction (upper to lower) and that the highest Z-axis readings occurred at WS 182 and WS 216 in the vicinity of the wing

tip and WS 40 located near the wing root.

Since an inverse proportionality exists between the circumference of a current-carrying conductor (wing skin) and its external magnetic field intensity, it follows that the average

external field intensity at the wing tip (which is the smallest

circumference) should be greater than the average intensity inboard near the wing root (largest circumference). The internal

field intensity, which results from diffusion and aperture penetration of the external field, would follow a similar pattern.

The access door at WS 40 was constructed of kevlar rather

than gr/E as in the other door locations. The tests at this

location were performed without the use of aluminum flashing

shield to simulate the electromagnetic environment that would

exist with the kevlar door in place. The test results show a

high Z-axis field intensity at this location comparable in magnitude to the measurement recorded at the wing tip locations but

higher than the reading obtained at the adjacent access door location which used a gr/E door. The high field readings may also be

due to other fields in the vicinity such as the field produced by

current flowing in the leading edge conduit. The influence of the

conduit field may also be the reason for the pronounced dip in

the leading edge of WS 40 to WS 138 waveforms of Figure 3.47.

The second highest readings occurred during tests in which

the coil axis was oriented to measure the leading to trailing

edge magnetic field (Y-axis). The waveforms from these measurements, as shown on Figure 3.48 for a typical example, rise to peak

more rapidly than do the Z-axis waveforms; thus, these fields

appear to be more dependent on aperture-type field coupling than

the Z-axis field measurements which appear to be more dependent

on diffusion of the external fields through the skin material.

The lowest readings occurred for fields in the X-axis direction (wing inboard to outboard). Measurements ranged from 2A/ft

(7 A/m) to 6A/ft (21 A/m), this compares to electrical "noise"

readings of I A/ft (4 A/m) to 3 A/ft (I0 A/m).

157
158

X

o

_q _ X

o

i

,I

ii

i

i

/

/

_oo_ _u3_

ORiGiNAL PAGE

OF POOR QUALI"rY

_t

-,1" e.l

r-4

I-i

(11

v _
ORIGINAL PAGE

OF POOR QUALITY

18A//m/div 50_s/div

Figure 3.48 - Typical Oscillogram of

Integrated Coil Response

to Y-Axis Field at WS 138

in gr/E wing.

The results of the external field measurements are shown in

Figures 3.49 and 3.50. Figure 3.51 shows typical oscillograms

of the test current and the integrated coil response.

As discussed in paragraph 3.3.4.5, the average magnetic

field intensity surrounding a current carrying conductor is

calculated by dividing the conducted current by the circumference of the conductor. The circumference of the wing at location

WS 182 (not including the control surfaces) was approximately

52 in. For a test current of 900 A, the theoretical average

field intensity at WS 182 should have been 208 A/ft (692 A/m).

At WS 80, with a circumference of approximately 84 in. (not including the control surfaces) the average field intensity should

have been 129 A/ft (423 A/m). Comparison of the calculated average magnetic field intensity magnitudes with the recorded values

obtained at WS 182 and WS 80 shows that in both cases the measured values are lower than expected.

Variation of the field intensity magnitude with location

along the wing perimeter agrees with the theoretical prediction;

i.e., the magnetic field intensity amplitude at locations with

a smaller than average radius of curvature will be greater than

average. The readings obtained at locations 1 and I0 (leading

edge), where the radius of curvature was smaller than average,

were higher than at any other locations. Readings obtained at

locations where the radius of curvature was larger than average,

as midway between the leading and trailing edges, were lower

than average. The test reading amplitudes again increased at

locations 4 and 7 where the radius of curvature again became

smaller than average. Measurement points 5 and 6 were located

at the trailing edge of the flaps approximately eight inches

from the current-carrying wing, thus the field readings were

understandably low.

159
ORIGINAL PAGE 13

OF POOR QUALITY

Control Surface---- 7

Upper Wing Skin

-- 122(400)

/161(527) _]D

Trailing 5_(182) /__ _/ -- _ --

141(463) 14_(472 )

Lower Wing Skin

145(476)

_$_) 199(654)

_ 227(745)

199(654)

Leading

Edge

Figure 3.49 - Peak External Magnetic Field Intensity Levels at

WS 182 in gr/E Wing. Test current was 900A.

Measurements are in A/ft (A/m). Distance

between measurement locations (circled numbers)

is approximately 8 in.

Control Surface

Upper Wing Skin

66(218) 89(291)

100(327) C__ _9 130(427)

Trai i ing _ '_

Leading

Edge Edge

2

105(345)

64(209) 94(309)

Lower Wing Skin

Figure 3.50 - Peak External Magnetic Field Intensity Levels at

WS 80 in gr/E Wing. Test current was 900A.

Measurements are in A/ft (A/m). Distance

between measurement locations (circled numbers)

is approximately 14 in.

160
ORIGINAL. PAGE T_

OF POOR QUALITY

200A/div 10_s / div

Test Current

55A/ft (180A/m) 10_s/div

Integrated Coil Response

Figure 3.51 - Typical Oscillograms of Test Current and

Integrated Coil Response Obtained during

External Magnetic Field Measurements at

WS 182 on gr/E Wing.

3.5 Discussion of Results

3.5.1 Comparison of the full scale structural environment with

the subelement test results

3.5. I. 1 Bonded aluminum wing

3.5.1.1.1 Current distribution.- Currents in the wing fuel lines

(external to the fuel tank) were recorded during applied current

strikes of up to 88 kA peak amplitude to the wing tip. The results are shown in Table 3-2. The maximum fuel line current

amplitude recorded was 160A for the 88 kA strike; when extrapolated for a full threat strike of 200 kA, the expected fuel line

current amplitude would be 364 A. Table 2-10 lists the spark

threshold levels of bonded aluminum fuel line bracket subelement

specimens and shows that visible sparks occurred at current levels

of 0.i0 kA - I0 kA for specimens which exhibited electrical conductivity prior to test. Thus the current levels measured in the

full scale structure indicated that the possibility for ignition

source sparking would exist if the following conditions were met:

a) One hundred amperes (or higher) of fuel line current

flowed through a fuel line bracket to the aluminum

wing skin. Since no evidence of internal fuel tank

161
sparking was recorded during the ignition source

tests (Figures 3.17 through 3.20), either the current flowing through individual fuel line brackets

was less than the spark threshold level, or the

bracket adhesive was electrically nonconductive.

b) A voltage potential of I00 volts or higher existed

between a fuel line bracket and the wing skin (see

the fuel line bracket sublement test results shown

in Table 2-9). Again, the lack of internal fuel

tank sparking during the ignition source tests indicated that the fuel line bracket-to-skin voltage

potential did not exceed the breakdown voltage level

threshold indicated in Table 2-9.

Interpretation of possible fuel tank sparking based on fuel

line current measurements must be made with care. Currents flowing between fuel lines and wing skin return may follow several

paths and may distribute unequally through several fuel line

brackets. Thus fuel line current levels which exceed the subelement spark threshold levels may not result in fuel ignition

sparks at the bracket/wing skin interface since any one bracket

may carry only a fraction of the total current.

3.5.1.I.2 Bond line voltages.- Bond line voltages recorded

within the fuel tank for test current levels of 7 kA and II kA

were 0.14 volts or less as shown in Table 3-3. Extrapolation of

this reading for a 200 kA, full threat lightning strike results

in a projected bond line voltage difference of 2.5 volts. This

low voltage level indicates that continuity existed between the

stringers, ribs, and aluminum skin. A check with a battery operated ohmmeter confirmed that continuity did exist between the

adherends and, therefore, it would not be possible to develop

a voltage potential between them. Thus, any sparks occurring in

the fuel tank would be the result of current amplitudes which

exceeded the current spark threshold level of the bonds.

Table 2-6 shows the spark threshold levels of several adhesives

tested on the subelement lap joint specimens. In particular,

EA 9602.3 adhesive, which was utilized on the full scale structure,

had a current spark threshold level of 200 amperes when current

was conducted through a bonded surface area of I in.2

Photographs taken during the ignition source tests (Figures

3.17 through 3.20) indicated that no internal fuel tank sparking

existed during 85 -I00 kA wing tip strikes. Thus, currents flowing between the adherends (wing skin and ribs or stringers) apparently had not exceeded the current density spark threshold

level of 200 A/in}.

162
3.5.1.2 gr/E wing

3.5.1.2.1 Bond line voltage and current thresholds.- The full

scale structure bond line voltage test results are tabulated in

Table 3-13. For test current of 900 A, the bond line voltages

ranged from 0.07 volts to 4.4 volts. If a linear relationship

is assumed and the bond line voltages obtained during the 900 A

current strikes are extrapolated for a full threat current strike

level of 200 kA, the range of anticipated bond line voltages

would be 16 to 980 volts. The rear stringer-to-skin bond line

at WS 88, which recorded the higher bond line voltage at 900A,

was measured during a 45 kA strike to the wing tip and found

to be 63 volts. This value extrapolates to 280 volts for a

200 kA strike. A retest of this bond line at the 900 A strike

level showed the bond line voltages to be 1.8 volts, down from

its original level of 4.4 volts. This result would seem to indicate that either the first measurement circuit contained a high

resistance connection causing a higher than normal voltage reading or the conduction of the 45 kA current through the wing

resulted in the formation of additional conductive paths in the

gr/E skin which reduced the voltage potential rise within the

skin.

Table 3-14 shows a comparison between the original 900 A

test current bond line readings at WS 88 and those recorded

following twenty 190 kA current strikes to the wing. With one

exception (rear spar to upper skin), the latter bond line voltage measurements were lower than the initial ones. This result

would tend to corroborate the conclusion that currents through

the structure form additional conductive paths thus reducing

the voltage potential rise in the structure skin.

A graph of bond line voltage at WS 88 (Fig. 3.43) versus

test current amplitude indicates that the resistivity of the

gr/E skin was not linear but decreased with increasing current

amplitude. Thus, if bond line voltage levels obtained with

900 A test currents, are used to predict bond line voltages

resulting from 200 kA test currents, the predicted results will

be higher than the actual measured voltages. Examination of the

test data presented in Table 3-13 shows this to be true. At

WS 88, the initial 900 A rear stringer-to-skin bond line voltage

was 4.4 volts. If the resistivity of the gr/E skin were linear,

the expected bond line voltage for a 200 kA strike would be 980

volts. The table indicates that the actual bond line voltage

for a 190 kA strike was 200 volts, approximately 20% of the

predicted value.

Table 2-15 tabulates the results of voltage breakdown tests

on nonconductive subelement lap joint specimens. With the exception of one test specimen, the voltage breakdown levels ranged

163
from 1200V to 8000V. (The voltage breakdown level of the one

specimen was less than I00 V. Pretest continuity measurements

with a battery-operated ohmmeter had indicated that the bond

line of this specimen was electrically nonconductive. Due to

the difficulty of probing through epoxy resin with test probes,

it is possible that good electrical contact with the graphite

fiber was not made and that the specimen may have actually been

conductive prior to the voltage breakdown tests. A conductive

specimen would, of course, require no voltage to initiate conductivity).

The bond line voltages obtained on the full scale gr/E

structure for a 200 kA current strike indicated the worst case

bond line voltage level expected within a wing is approximately

1/6 of the minimum voltage breakdown level determined by the

subelement test results. Thus, sparking at the bond lines would

not be expected as a result of voltage breakdown.

If a bond line was electrically conductive, voltage potential

could not be developed across it; hence, a fuel ignition spark

would not result from bond line voltage breakdown. An ignition

spark could result, however, if the current flowing though the

conductive bond line were to exceed the current density spark

threshold level. Table 2-18 shows the current density spark

threshold level resulting from tests on the gr/E subelement lap

joint specimens. Six of the specimens exhibited current spark

threshold levels of less that I00 A/in. 2 These specimens had

previously shown edge sparking during the voltage breakdown tests

and thus would conduct no appreciable current before sparking.

The results of tests on the remaining specimens indicated

that the current density spark threshold level ranged between

250 A/in. 2 and 5000 A/in. 2. An estimate of the current density

level occurring within the full scale gr/E wing can be approximated from its physical dimensions. A worst case current density

level would result by assuming that all of the skin current is

transferred to the full length spars through the spar caps. If

it is assumed that the average width of the spar caps is 2 in.

and the length of each spar is 226 in., then the current density

is given by:

average current density =

through spar caps

strike current peak amplitude

spar cap widthxspar lengthxno.of

spars x no. of skin surfaces

= 200 kA

2 in. x 226 in. x 3 x 2

Average current density through spar caps = 75 A/in. 2

164
The above analysis indicates that the current density in

gr/E wing spar caps is approximately 30% of the level required

to produce visible sparks as indicated by the subelement lap

joint test results. This result is on the conservative side.

Figure 3.36 indicates that at the time of the test current peak

amplitude (approximately 4 ps), the wing skin conducted only

74% of the test current; hence, the current density through the

spar caps is approximately 75% of 75 A/in. 2 or 55 A/in. 2 which

represents about 25% of the current density required to cause

sparking.

Peak current amplitudes in the fuel lines measured 2 kA

during a 190 kA simulated strike - this level is below the

spark threshold level of I0 kA determined from the subelement

fuel line feed-through test specimens, as shown in Figure 2.44.

3.5.2 Comparison of the bonded aluminum wing with the gr/E wing

3.5.2.1 Bond line voltage and wing resistance

The bonded aluminum wing exhibited extremely low bond

line voltage readings indicating that electrical contact probably existed between the structural members and the wing skin;

thus, projected bond line voltages for a severe 200 kA lightning

strike were approximately 2.5 volts. This contrasts with the

gr/E wing results where a maximum bond line voltage of 200 volts

was measured for a 190 kA test current, as shown on Table 3-13.

The resistance of the bonded aluminum wing skin was determined to be 0.3 m_ and calculated from the test oscillogram of

voltage induced in the taxi light wiring, as shown in Figure

3.23 and on Table 3-6. The resistance was calculated by dividing the voltage by the current at the time, t, at which the rate

of current change was zero. A schematic representation of this

measurement is shown in Figure 3.52.

Calculation of the gr/E wing resistance can be made with a

knowledge of the fuel probe-to-skin voltage measurements. As

described in Para. 3.4.4.2, the fuel probes were electrically

tied to ground through a dummy fuel quantity meter circuit and

thus the probe-to-skin voltages actually represented the wing

skin voltage potential with respect to the reference ground

plane, as shown in Figure 3.53.

Fuel probe voltage readings are plotted in Figures 3.43

and 3.44 as a function of wing current for the probe at wing

station WS 88. The structural resistance is the slope of the

line which best fits the data points on the graphs and can be

determined using the statistical method of linear regression.

165
oR_G_,_ . _ _x:_

OF t:_O_ Q_ALF_'

rw rw

Taxi

% g_" _ Light

Taxi Light _ i L

rw V = Voltage of Taxi Light Wlre

iL = Simulated Lightning Current

rw = Wing Skin Resistance

Figure 3.52 - Bonded Aluminum Wing Skin Resistance

Measurement (di/dt = 0).

r W

--___ .__._ Wing Tip

Quantity Meter

Circuit _ Prr_°5e _ / /

B

. V = Voltage between Fuel Probe

rw rw and adjacent skin

= Simulated Lightning Current

rw = Wing Skin Resistance

166

Figure 3.53 - gr/E Wing Skin Resistance Measurement.
This calculation indicated that the resistance between the reference ground plane and location WS 88 was i0.I xl0-3_/in. The

structural resistance to the closeout rib at WS 226 was then

226 in. x 115 xl0-_/in. = 25.8 x10-3_.

The test results and calculations above show that the

resistances of the two wings were 0.3 m_ for the bonded aluminum

wing and 25.8 m_ for the gr/E wing. The _-a-tio of wing resistances

was then :

gr/E wing resistance

bonded aluminum wing resistance

25.8m_ _

- I00:I

0.3m_

3.5.2.2 Magnetic fields

Some interesting comparisons can be derived from the magnetic

field data obtained on the bonded aluminum and gr/E wings. The

data were previously presented in Table 3-12 and Figure 3.28 for

the bonded aluminum wing and Figures 3.47--and 3.49 for the gr/E

wing.

The ratio of the external to internal magnetic field intensities for the two wings can be calculated by normalizing the

field intensity measurements with respect to test current levels.

For the bonded aluminum wing, this ratio can be determined

as follows (measurements are at the fuel filler cap location):

exterior to interior magnetic peak exterior field intensity/test current

field intensity ratio at WS 160 =

(Bonded Aluminum Wing) peak interior f_l_id intensity/test current

2.1(103 ) A/ft (6.6(103)A/m)/12kA

20 A/ft (66A/m)/2.4kA

175 A/ft/kA (570A/m/kA)

8.3 A/ft/kA (27A/m/kA)

exterior to interior magnetic

field intensity ratio at WS 160 _ 20:1

(Bonded Aluminum Wing)

In a similar manner, the external to-internal-_eld ratio

for the gr/E wing can be calculated.

167
exterior to interior peak exterior field intensity/test current

magnetic field intensity = peak interior field intensity/test current

at WS 182

(gr/E wing)

145 A/ft (480 A/m)/0.9kA

28 A/ft (92A/m)/l.9kA

145 A/ft/kA (480 A/m)

28 A/ft/kA (92 A/m)

Exterior to interior

magnetic field intensity = 5:1

at WS 182

(gr/E wing)

A comparison can also be made between the three orthogonal

field intensity measurements within the bonded aluminum wing

with those in the gr/E wing. Again, data from bonded aluminum

wing location WS 160 will be compared to gr/E wing location

WS 182, as shown in Table 3-17.

Comparison of exterior to interior magnetic field ratios

between the two wings shows that the shielding provided by the

aluminum skin was approximately four times more effective than

that provided by the gr/E skin. This correlates closely with

the internal field ratio comparison of 3.7.1 between the wings

for the Z-axis field orientation.

168
Table 3-17 - Data Tabulation and Results for

Internal Magnetic Field Comparison

between Bonded Aluminum and gr/E

Coil Orientation (a)

Peak Magnetic Intensity (H)

A/ft (A/m)

Bonded aluminum wing WS 160

Peak Test Current (kA)

Bonded aluminum wing WS 160

Normalized Magnetic

Intensity. A/ft/kA (A/m/kA)

Peak Magnetic Intensity (H)

A/ft (A/m)

gr/E wing WS 182

Peak Test Current (kA)

gr/E wing

Normalized Magnetic

Intensity. A/ft/kA (A/m/kA)

gr/E wing WS 182

X y Z

5.5(18) 8.5(28) 20(64)

2.4 2.4 2.4

2.3(7.5) 3.5(12) 8.3(27)

6 (21) II (36) 28 (91)

0.9 0.9 0.9

6.7(23) 12 (40) 31 (I00)

gr/E Wing Internal Field

Bonded Aluminum Wing Internal Field

_3:1 _3.5:1 _3.7:1

(a) See Figure 3.13

169
References

.

.

,

,

Fisher, F.A. and Plumer, J.A.: "Lightning Protection of

Aircraft" NASA RP-1008, 1977, pp. 35-55.

"Lightning Test Waveforms and Techniques for Aerospace

Vehicles and Hardware" Report of SAE Committee AE4L,

June 19 78.

Plumer, J.A.: "Further Thoughts on Location of Lightning

Strike Zones on Aircraft" Lightning Technology, Suppl.

NASA CP-2128, FAA-RD-80-30, 1981, pp. 81-98.

"Lightning Test Waveforms and Techniques for Aerospace

Vehicles and Hardware". Report of SAE Committee AE4L,

June 19 78, Para. 2.0.

.

.

"Lightning Test Waveforms and Techniques for Aerospace

Vehicles and Hardware". Report of SAE Committee AE4L,

June 1978, Table i.

"Lightning Qualification Test Techniques for Aerospace

Vehicles and Hardware". MIL-STD-1757, June 1980.

.

.

,

I0.

ii.

12.

Fisher, F.A° and Plumer, J.A.: "Lightning Protection of

Aircraft" NASA RP-1008, 1977, p. 190.

Gerstein, M. and Allen, Robert D.: "Fire Protection

Research Program for Supersonic Transport" Report of

Dynamic Science Corporation, SN-4400, 1964, p. 112.

Fisher, F.A. and Plumer, J.A.: "Lightning Protection of

Aircraft". NASA RP-1008, 1977, pp. 441-446.

Gerstein, M. and Allen, Robert D.: "Fire Protection

Research Program for Supersonic Transport". Report of

Dynamic Science Corporation, SN-4400, 1964, p. 113.

U.S. MIL-STD MIL A-25465, Rev. B. "Adhesive, Film Form,

Metallic Structural Sandwich Construction", March 1982.

U.S. Federal Specification MMM-A-132. "Adhesives, Heat

Resistant, Airframe Structure, Metal to Metal", October

1964.

170
Qb
I. Report No. 2. Government Accession No. 3. Recipient'% Catalog No.

NASA CR- 3762

4. Title end Subtitle

LIGHTNING PROTECTION GUIDELINES AND TEST DATA

FOR ADHESIVELY BONDED AIRCRAFT STRUCTURES

7. Author(s)

J.E. Pryzby and J.A. Plumer

g. Perfuming Organization Name and Addrm

Lightning Technologies, Inc.

I0 Downing Parkway

Pittsfield, MA 01201

2. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546

5, Report Date

J_nu_r¥ 1984

6. Performing Organization Code

8. Performing Organ;zation Report No.

LT-83-155

10. Work Unit No.

11. Contract or Grant No.

NASI-15884

13. Type of Report and Period Covered

Contractor Report

14. Sponsoring Agency Code E

505 -45-03-01

15. Supplementary Notes

Langley Technical Monitors - Bruce D. Fisher and William E. Howell

16. Abstract

The highly competitive marketplace and increasing cost of energy has

motivated manufacturers of general aviation aircraft to utilize composite materials and metal-to-metal bonding in place of conventional

fasteners and rivets to reduce weight, obtain smoother outside surfaces

and reduce drag. The purpose of this program was protection of these

new structures from hazardous lightning effects.

The program began with a survey of advance-technology materials and

fabrication methods under consideration for future designs. Sub-element

specimens were subjected to simulated lightning voltages and currents.

Measurements of bond line voltages, electrical sparking, and mechanical

strength degradation were made to comprise a data base of electrical

properties for new technology materials and basic structural configurations. The second phase of the program involved tests upon full scale

wing structures which contained integral fuel tanks and which were

representative of examples of new technology structures and fuel systems.

The purpose of these tests was to provide a comparison between full

scale structural measurements and those obtained from the "sub-element"

specimens.

17. Key Words (Suggested by Author($])

Lightning protection

Lightning effects

Composites

Adhesive bonding

Structures

19 Security Oa_if. (of this report)

Unclassified

Fuel systems

General aviation

20. Security Unclassified Classif. (of this _ge)

18 Distribution Statement

21. No. 173 of Pages [

22. Price

NASA-Langley, 1984