NASA Reference Publication 1008 

Lightning Protection 

of Aircraft 

Franklin A. Fisher and J. Anderson Plurner 

General Electric Company 

Pittsfield, Massachusetts 

Prepared for 

Aerospace Safety Research and Data Institute 

NASA Lewis Research Center 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Office 
For sale by the Superintendent of Documents 

US. Government Printing Office, Washington, D.C. 20402 
PREFACE 

This book is an attempt to present under one cover the current state of 

knowledge concerning the potential lightning effects on aircraft and the means 

that are available to designers and operators to protect against these effects. The 

impetus for writing this book springs from two sources--the increased use of 

nonmetallic materials in the structure of aircraft and the constant trend toward 

using electronic equipment to handle flight-critical control and navigation 

functions. Nonmetallic structures are inherently more likely to be damaged by a 

lightning strike than are metallic structures. Nonmetallic structures also provide 

less shielding against the intense electromagnetic fields of lightning than do 

metallic structures. These fields have demonstrated an ability to damage or cause 

upset of electronic equipment. 

Such concerns, when added to the continuing apprehension regarding 

the vulnerability of fuel systems to lightning, have led in the past decade to 

increased research into lightning effects on aircraft. The results of this research 

are contained in the technical reports and literature published by ourselves and 

by researchers in other laboratories who are also working on these problems. 

Conferences and symposiums have been held so that researchers could exchange 

ideas and information; there is a high degree of cooperation among all of those 

working towards the goal of complete safety-of-flight in the lightning environmen t. 

The persons who can best use information on aircraft protection from 

lightning are the aircraft designers and operators, but generally they are not 

among those who produced this information. Moreover, they are often unaware 

of its existence, and they seldom have the background to distill from it the 

important facts that can and should be applied to achieve safer designs. The 

purpose of this book is to present the most important parts of this body of 

knowledge in a manner most useful to the designer and the operator. 

This book is organized into seventeen chapters. In the first of these we 

review what lightning is and how it originates. The second chapter describes how 

the aircraft becomes involved with the lightning Slash and why it is that aircraft 

do not produce their own lightning flashes, but may, we think, sometimes trigger 

natural ones. Chapter 3 considers how often and under what conditions aircraft 

have been struck, reviews avoidance procedures now in use by operators, and 

reviews their degree of success. We also take up the question of whether or not 

strikes could be totally avoided. The fourth chapter summarizes the various 

effects which have occurred when lightning has struck aircraft, giving the 

operator an idea of the direct and indirect effects which he may expect when his 

aircraft is "zapped ." 

Since our main purpose is to help the designer protect against those effects 

that may be hazardous, the remainder of the book is devoted to this purpose. 

Chapters 5, 6, and 7 deal with protection against the direct physical damage 

effects. Chapter 5 sets forth three philosopl~ical steps which guide us in the 

design work that follows. Attention is also called to government standards or 
certification regulations which deal with aircraft lightning protection. 

Chapter 6 reviews in some depth what is known about lightning effects on 

aircraft fuel systems and tells how to design protection against these effects. We 

give considerable attention to this subject and urge the designer to do likewise 

because of the serious consequences in the past of lightning effects on these 

systems. Chapter 7 deals with the protection of the aircraft from structural 

damage resulting from lightning, with emphasis on the nonmetallic materials-- 

materials that may be more vulnerable than the metal structures they are 

beginning to replace. 

The remainder of the book deals with indirect effects. Chapter 8 

introduces the reader to the basic mechanisms by which induced voltages occur 

in aircraft electrical circuits, and Chapters 9 through 14 treat these mechanisms 

in greater detail. Chapters 15 and 16 then consider the impact of induced 

voltages upon solid state electronic devices and tell how these devices may be 

protected. Here, concepts such as the transient control level philosophy are 

presented. Concepts of this kind may form the basis for future specifications 

that define the roles to be played by both the aircraft designer and the 

electronics equipment designer. Finally, in Chapter 17 we show how aircraft can 

be tested to determine their actual susceptibility to indirect effects and how 

equipment can be tested to determine its vulnerability or prove that protection 

design goals have been met. 

To some extent, each chapter stands by itself and can be utilized without 

knowledge of the others. Dependencies often exist, however, among the 

lightning effects on structural, electrical, and fuel systems; for the most 

thorough understanding of lightning effects on any one of these systems, the 

reader is urged to read the entire book. 

Part of the research upon which this book is based was conducted by 

ourselves, but a significant amount was conducted by researchers at other 

laboratories. We have referred to or incorporated that work frequently. Without 

it our present understanding of lightning effects--as well as this book--would 

not have been possible. Not all of the work conducted in this field could be 

referenced, of course, but we have carefully studied most of it. The work we 

reference is that which we consider to be definitive, and we have taken care to 

provide complete source details so that the reader can refer to them for 

additional information. 

Even though much has now been learned about lightning effects on 

aircraft and how to design protection, there are still some lightning effects which 

are not fully understood. Examples of these are (1) the mechanisms by which 

lightning currents diffuse into interior structural members and conducting parts 

together with the extent to which this happens and (2) the effects of 

electromagnetic radiation from the lightning arc upon aircraft electrical and 

electronic systems. We have tried to identify these areas as they are encountered, 

and we caution the reader to remain alert for developments in these areas in the 

future. 

While much of this book may appear oriented to the designer, there is 
much here of benefit to the operator as well. Familiarity with its contents will 

enable him to know what to expect, what not to expect, and why particular 

thing happen when lightning strikes his aircraft. The book will show him where 

to look for damage after a strike, and, we hope, help him to better identify 

potential problem areas and to communicate them to researchers and designers 

in time to avert future problems. 

The preparation of this material was supported by the Aerospace Safety 

Research and Data Institute, Lewis Research Center, National Aeronautics and 

Space Administration under Contract NAS3-19080. We wish to acknowledge the 

support of that organization. We also wish to acknowledge the support of H. V. 

Bankaitis, Solomon Weiss and, in particular, of Paul T. Hacker of that 

organization. It was Mr. Hacker who first suggested that a book of this nature be 

written. 

Also, we deeply appreciate the help of Beryl I. Hourihan in digging 

through our fiIes for materials and in preparing the rough draft; and the 

diligence, skill, and long hours applied by Catharine L. Fisher, who edited this 

book, bringing order and clarity to a sometimes confusing array of facts and 

figures. 

F. A. Fisher 

J. A. Plumer 

April 1977 

TABLE OF CONTENTS 

Preface ................................................. iii 

Chapter 1 . The Lightning Environment ........................... 1 

1.1 Introduction ..................................... 1 

1.2 Generation of the Lightning Flash ....................... 2 

1.3 Further Development of the Lightning Flash ................ 16 

1.4 Lightning Polarity and Direction ........................ 18 

1.5 Intracloud Flashes ................................. 19 

1.6 Measured Characteristics of the Lightning Flash .............. 20 

1.7 Thunderstorm Frequency and Lightning-Flash Density ......... 27 

1.8 Engineering Models of Lightning Flashes ................... 31 

References .......................................... 34 

Chapter 2 . Aircraft Lightning Attachment Phenomena ................ 37 

2.1 Introduction ..................................... 37 

2.2 Aircraft Influence on Flash Formation .................... 37 

2.3 Swept Strokes .................................... 49 

2.4 Aircraft Lightning Attachment Zones ..................... 51 

References .......................................... 54 

Chapter 3 Lightning . Strike Experience .......................... 57 

3.1 Introduction ..................................... 57 

3.2 Altitude and Flight Path ............................. 57 

3.3 Synoptic Meteorological Conditions ...................... 59 

3.4 Immediate Environment at Time of Strike .................. 62 

3.5 Thunderstorm Avoidance ............................. 65 

3.6 Frequency of Occurrence ............................. 71 

References .......................................... 73 

Chapter 4 . Lightning Effects on Aircraft .......................... 75 

4.1 Introduction ..................................... 75 

4.2 Direct Effects on Metal Structures ....................... 78 

4.3 Nonmetallic Structures .............................. 85 

4.4 Fuel Systems ..................................... 89 

4.5 Electrical Systems .................................. 91 

4.6 Engines ......................................... 96 

4.7 Indirect Effects ................................... 98 

References .......................................... 102 

Chapter 5 . Direct Effects Protection ............................ 105 

5.1 Introduction ..................................... 105 

5.2 Basic Steps in Protection Design ........................ 105 

5.3 The Lightning Environment as Related to Attachment Zones ..... 110 

vii 
TABLE OF CONTENTS 

5.4 Applicable Airworthiness Regulations ..................... 111 

References .......................................... 114 

Chapter 6 Fuel System . Protection ............................. 115 

6.1 Intrndtlction ..................................... 115 

6.2 Fuel Flamn~ability ................................. 115 

6.3 Sources of Ignition ................................. 170 

6.4 Vent Outlets ..................................... 121 

6.5 Fuel Jettison and Drain Pipes .......................... 134 

6.6 Hot Spot and Hole Formation in Integral Tank Skins .......... 136 

6.7 Integral Fuel Tanks with Nonmetallic Sk~ns ................. 154 

6.8 Current Flow Through Structures and Components ............ 157 

6.9 Fuel System Checklist ............................... 178 

References .......................................... 181 

Chapter 7 . Structures Protection ............................... 185 

7.1 introduction ..................................... 185 

7.2 Metallic Structures ................................. 186 

7.3 Nonmetallic Structures .............................. 106 

References .......................................... 150 

Chapter 8 -Voltages and Currents Induced By Lightning ............... 253 

8.1 Definition and Elementary Consideiations .................. 353 

8 . 2 The Exterior Response of the Aircraft Stiucture .............. 754 

8.3 The Internal Response of tlie Aircraft Structure .............. 259 

References .......................................... 367 

Chapter 9 The External Magnetic Field Environment ................. 269 

9.1 Elementary Conside~ations ............................ 169 

9.2 Determination of the External Current and 

Magnetic Flux Density ........................... 271 

9.3 Examples of External Magnetic Fields .................... 185 

References .......................................... 292 

Chapter 10 -The Internal Magnetic Fields Created By Diffusion ........... 293 

10.1 Circular Cylinders .................................. 193 

10.2 Elliptical Cylinders ................................. 303 

10.3 Fields Within Cavities ............................... 320 

References .......................................... 327 

Chapter 11 . The Internal Fields Coupled Through Apertutes ............ 329 

1 1 . 1 Magnetic Fields ................................... 319 

1 1.2 Electric Fields .................................... 342 
TABLEOFCONTENTS 

11.3 Magnetic Field Zones ............................... 344 

References .......................................... 348 

Chapter 12 - Voltages and Currents Induced on Wiring ................. 349 

12.1 Elementary Considerations ............................ 349 

12.2 Transmission Line Effects ............................ 351 

12.3 Aids for Calculation of Voltage and Current ................ 357 

12.4 Voltage and Currents in the Space Shuttle .................. 359 

References .......................................... 360 

Chapter 13 . Effects of Shields on Cables ..................... 361 

13.1 Introduction ..................................... 361 

13.3 Some Elementary Illustrations ......................... 362 

13.3 Transfer Quantities ................................. 369 

13.4 Multiple Conductors and Shields ........................ 385 

13.5 Transfer Characteristics of Actual Cables ................... 393 

13.6 Transfer Characteristics of Connectors .................... 394 

13.7 Grounding of Shields ................................ 394 

References .......................................... 405 

Chapter 14 - Exaltiples of Induced Voltages Measured on Aircraft ......... 407 

14.1 Introduction ..................................... 407 

14.2 Wing from F-89J Aircraft ............................. 407 

14.3 F-8 Digital-Fly-By-Wire (DFBW) ........................ 413 

References .......................................... 427 

Chapter 15 . Design to Minimize Indirect Effects ..................... 429 

15.1 Some Premises and Goals for Design ...................... 429 

15.2 Improvement Through Location of Electronic Equipment ....... 429 

15.3 Improvement Through Location of Wiring .................. 430 

15.4 Improvement Through Shielding ........................ 433 

15.5 Improvement Through Circuit Design ..................... 438 

15.6 Circuit Protection Through Use of Protective Devices .......... 441 

15.7 Improvement Through Transient Coordination ............... 446 

References .......................................... 450 

Chapter 16 . Component Damage Analysis ......................... 451 

16.1 Introduction ..................................... 451 

16.2 Theoretical Models ................................. 451 

16.3 Empirical Models .................................. 451 

16.4 Limitations ...................................... 452 

16.5 Failure Mechanisms-- Semiconductors .................... 453 

16.6 Damage Constants ................................. 455 
TABLE OF CONTENTS 

16.7 Failure Mechanisms.. Capacitors ....................... 463 

16.8 Failure Mechanisms.. Other Components .................. 464 

16.9 Examples of Use of Damage Constants .................... 465 

References .......................................... 492 

Chapter 17 . Test Techniques For Evaluation of Indirect Effects .......... 495 

17.1 Introduction ..................................... 495 

17.2 Waveform Definitions ............................... 496 

17.3 Tests on Complete Vehicles ........................... 498 

17.4 Tests on External Electrical Hardware .................... 504 

17.5 Tests on Internal Equipment ........................... 505 

17.6 Design Notes for Transformer Pulse-Injection Equipment ........ 520 

References .......................................... 530 

Appendix 1 . SR-52 Calculations . Conductor ...................... 533 

Appendix 2 . SR-52 Calculations . Tube .......................... 535 

Appendix 3 . Program MACFLD ............................... 539 

Index .................................................. 545 
CHAPTER 1 

THE LIGHTNING ENVIRONMENT 

1.1 Introduction 

The lightning flash originates with the formation of electrical charge in the 

air or, more commonly, clouds. The most common producer of lightning is the 

cumulonimbus thundercloud. Lightning, however, can also occur during sandstorms, snowstorms, and in the clouds over erupting volcanos. Lightning has 

even been reported to occur in clear air, though this phenomenon is rare and is 

possibly a result of lightning originating in conventional clouds beyond the 

observer's field of vision. Lightning originating in sandstorms and volcanic 

eruptions is not of serious concern to aircraft, but lightning associated with 

snowstorms occurs sufficiently often as to present a problem, not because its 

nature is different from lightning associated with thunderstorms but because it is 

apt to occur when it is unexpected. 

The most common types of lightning are those involving the cloud and 

ground, called cloud-to-grou~zd lightning, and lightning between charge centers 

within a cloud, called intracloud lightning. This latter is sonletimes erroneously 

called intercloud or cloud-ro-cloud lightning. True cloud-to-cloud I~glltning 

between isolated cloud centers is possible; however, what appears to be cloud-tocloud lightning is often a spectacular manifestation of intracloud discharges. 

Most research on lightning has centered on cloud-to-ground l~ghtning; 

despite its importance to aircraft operation, much less information on the 

characteristics of intracloud lightning than on those of cloud-to-ground lightning 

exists, for a variefy of reasons. The first is simply that intracloud lightning, 

unlike cloud-to-ground lightning, is largely hidden from direct observation. The 

second is that conducting research on the characteristics of lightning is often a 

labor of love requiring both extensive apparatus and extreme patience. Observing 

lightning from a fixed ground station is rnuch easier and cheaper than observing 

lightning from a moving aircraft. The third is that most of the funding for research on lightning has come, directly or indirectly, from those who are concerned with the effects of lightning on electric power transmission and distribution lines, which are affected only by cloud-to-ground strokes. 

The characteristics of lightning are discussed in the sections which follow. 

However, much of the material relative to the physics of lightning will be discussed primarily in terms of cloud-to-ground strokes for the reasons cited 

above and because frequently it is difficult to say for any given flash which type 

was involved. While aircraft may be involved with any of the three types of 

lightning, cloud-to-ground and intracloud lightning flashes are the most common 

types. But where aircraft design and operation are of concern, the type of flash 

makes little difference. 

Figure 1.1 shows a generalized waveshape for the current flowing to 

ground from a typical negative cloud-to-ground flash and presents terms for the 

five main regions shown: the leader, the initial return stroke, an intermediate 
current, a continuing current, and one or more restrikes. Each of these aspects 

will be discussed. 

INTERMEDIATE RESTRIKE 

Figure 1.1 Generalized waveshape of current in negative cloud-to-ground 

lightning. (Note that the drawing is not to scale.) 

1.2 Generation of the Lightning Flash 

1.2.1 Generation of the Charge 

The energy that produces lightning is assumed to be provided by warm air 

rising upwards into a developing cloud. As the air rises it becomes cooler, and at 

the dew point, the excess water vapor condenses into water droplets, forming a 

cloud. When the air has risen high enough for the temperature to drop to -40 OC, 

the water vapor will have frozen to ice. At lower elevations there will be many 

supercooled water drops that are not frozen, even though the temperature is 

lower than the freezing point. In this supercooled region, ice crystals and 

hailstones form. 

According to one theory, the cloud becomes electrically charged by the 

following process (Reference 1.1). Some of the ice crystals which have formed 

coalesce into hailstones. These hailstones fall through the cloud gathering 

additional supercooled water droplets. As droplets freeze onto a hallstone, small 

splinters of ice chip off. Apparently, these splinters carry away a positive 

electrical charge, leaving the hailstone with a net negative charge. The vertical 

wind currents in the cloud carry the ice splinters into the upper part of the 

cloud, while the hailstone, being heavier, falls until it reaches warmer air, where 

some portion of it melts and the remainder continues to earth. Thus, the upper 

part of the cloud takes on a positive charge while the lower region takes on a 

negative charge. In some other manner, another smaller pocket of positive charge 

may be formed near the front of the base of the cloud and below the main body 

of negative charge. 

Other theories have been proposed to account for the electrification of the 

cloud (References 1.2 to 1.7). All of them are based on experimentally observed 

evidence that the charge in the top of the cloud is positive. There may also be a 

body of positive charge near the front of the base of the cloud. The charge in the 
rest of the cloud is negative. Figure 1.2 shows a typical cloud with the charge 

distributed as previously described. The cloud is moving to the left. The 

unbroken lines represent stream lines of air. 

NEGATIVE RAIN 

Figure 1.2 Generalized diagram showing distribution of air currents and 

electrical charge distribution in a typical cumulonimbus cloud. 

The air currents and the electrical charges tend to be contained in localized 

cells and the cloud as a whole to be composed of a number of cells. A typical 

cloud might have the cell structure shown in Figure 1.3 (Reference 1.8). The 

electrical charge contained within a cell might appear as shown in Figure 1.4 

(Reference 1.9). The temperature at the main negative-charge center will be 

about -5 OC and at the auxiliary pocket of positive charge below it, about 0 "C. 

The main positive-charge center in the upper cloud will be about 15 "C colder 

than its negative counterpart. 

The lifetime of a typical cell is about 30 minutes. At its mature stage the 

cell as a whole will have a potential, with respect to the earth, of 10' to lo9 

volts (V). It will have a total stored charge of several hundred coulombs (C) with 

potential differences between positive- and negative-charge pockets again on the 

order of lo8 to lo9 V. The cell as a whole will have a negative charge. 

1.2.2 Conditions on the Ground 

As the cloud passes over a point on the ground, an electrical charge is 

attracted into the ground under the cloud. The average electric field at the 

surface of the ground will change from its fair-weather value of about 300 volts 

per meter (V/m) positive (air positive with respect to the earth) to as high as 

several thousand volts per meter. Generally, when a cloud is overhead, the field 

from cloud to earth will be negative, but when a localized positive region is 

overhead, the field may be positive. The potential gradient will be concentrated 

around sharp protruding points on the ground and can exceed the breakdown 

strength of the air, which has a nominal value of 30 000 V/cm. When the 
rain 

snow 

- ~ce crystals w~nd speed scal$'%~ 

Figure 1.3 An idealized cross section through a thunderstorm cell in its mature 

stage. Key: 0, rain; *, snow; +, ice crystals. 

breakdown strength of the air is exceeded, current into the air increases sharply 

and a bluish electrical discharge called corona forms around a point. This 

discharge is the St. Elmo's fire that appears on such places as masts of ships, 

aircraft wing and tail tips, and from trees and grass on high ground. The 

magnitude of the current from a single discharge point may range from 1 or 2 

microamperes (PA) to as high as 400 PA. 
0 1 I 

Figure 1.4 Estimated charge distribution in a mature thundercloud (after 

Phillips). 

1.2.3 Development of the Leader 

At some state in the electrification of the cloud, a discharge towards the 

earth takes place. It starts as a slow-moving column of ionized air called the pilot 

streamer. After the pilot streamer has moved perhaps 30 to 50 m, a more intense 

discharge called the stepleader takes place. This discharge lowers additional 

negative charge into the region around the pilot streamer, recharges it, so to 

speak, and allows it to continue for another 30 to 50 m, after which the cycle 

repeats. 

A discharge propagating in this manner is called a streamer discharge; its 

development is illustrated in Figure 1.5. The streamer is initiated when a free 

electron is accelerated in a sufficiently high electrical field (a). An electron (b) so 

accelerated collides with neutral molecules of air (c), ionizes them, leaving them 

with a positive charge, and creates new electrons at a rate of a per unit length. 

The electrons, being much less massive than the positive ions, move under the 

influence of the electric field, leaving the positive ions behind. If the discharge 

continues to develop over the length a, there will be at the head of the discharge 

(d) a number of electrons given by 

N = exp (J: (a-n)) dx 

where a is the Townsend ionization coefficient and g is the attachment 

coefficient. In air at atnlospheric pressure, electron multipIication can exist only 

where the field is higher than 25 kV/cm because only in this case is a greater 

than 7. 
POSITIVE 

REGION 

- 

0 00 ++tt+ 

t + tt+t 

tttt + ++-- 

0 0 +tt t f + t++ 

000 +it+ tGT + - 

0 0 tit+ 

ttt 

000 

i-t 

++ 

ttt 

NEGATIVE ttt 

REGION ttt ttt (h) 4 t 

Figure 1.5 Stages in the development of a leader. 

Left in the wake of these electrons will be positively charged ions (e). If 

the electric field is high enough, the initial avalanche will reach a critical size 

(approximately lo8 electrons) for another avalanche of electrons (f) to be 

initiated by photoionization (g) from the initial discharge. The electric field that 

accelerates this secondary discharge is the sum of the initial electl-ic field and 

that produced by the positive space charge left behind by the initial avalanche. 

Under the action of the total field, these successive avalanches reach the positive 

space charge (h), neutralize it, and leave a new positive charge a little farther on. 

With such a mechanism a positive charge moves step by step into the un-ionized 

air leaving behind it a partially ionized filament (i). This filament is a conductor, 

though at this stage of its development perhaps only a poor conductor. 

The processes just described relate to a positive electric field. In a negative 

electric field a similar, though more complicated, phenonienon occurs. The 

initial avalanches seem to develop in the air farther ahead of the leader and to 

propagate both ways: into virgin air and back toward the more fully developed 

leader. The end result is much the same; behind the advancing head of the leader 

is an ionized column in this ease with a predominance of negative charge and 

' having at its center a more heavily conducting filament. 

If the initial development of the leader takes place in the charged cloud, 

the developing streamer branches and begins to collect charge from its 

surroundings. Because it collecrs charge in this way, the streamer may be viewed 

as connected to the cloud and at the same potential as the cloud. As the head of 

the leader moves farther into the un-ionized air, charge flows down from the 

charged regions of the cloud, along the partially conducting filament and toward 

the head of the leader, thus tending to keep all parts of the leader at a very high 

6 
potential. The amount of charge, qo, lowered into the leader will be on the order 

of 2 to 20 x loe4 C/m of length. A leader 5 km long would then have stored 

within it a charge of 1 to 10 C. 

Since the potential of the leader is very high, there will be a high radial 

electric field along the leader. This field will be high enough to exceed the 

breakdown strength of the air, and secondary streamers will branch out radially 

away from the central filament. The filaments will branch out radially until the 

field strength at the edge of the ionized region falls to about 30 kV/cm. 

It can be shown that the electric field strength at the edge of a cylinder 

containing a charge, qo, per unit length is 

Froni this and the above breakdown strength of air, it can be deduced 

that the radius of the leader will be 1.2 to 12 m. At higher elevations the 

breakdown strength of air is less; hence the leader radius may be more. 

Figure 1.6 The lightning leader as postulated by Wagner 

7 
As postulated by Wagner (Figure 1.6 [Reference 1.101 ), the head of the 

leader may have a larger diameter than that of the rest of the leader, though this 

is difficult to prove by photographs. The head of the leader, nevertheless, is 

generally visible because of the optical radiation associated with the extension of 

the electron avalanches, But once the growth ceases, the radiation stops; 

consequently, the corona sheath surrounding the central conducting filament is 

not visible. 

The process can be studied in the laboratory, albeit on a smaller scale than 

that of natural lightning. Figure 1.7 shows typical phenomena observed during 

the breakdown of the air between electrodes about 10 m apart. The sketch is 

based on a series of short-duration (about 1 microsecond [ps]) photographs 

taken at intervals of about 25 ps with an image-converter camera. As the 

electrons are attracted out of the initially ionized region at the head of the 

leader, the conducting filament lengthens and the corona discharge at the head 

of the leader occurs farther on in previously un-ionized air. 

Figure 1.7 The development of a leader in the laboratory. 

WAVESHAPE OF 

TO 

VOLTAGE 

ELECTRODE 

APPLIED - 

ELECTRODE A 

0 1 

VI 

f5 A- 

+ 

Photographs of actual lightning leaders may be taken with a Boys camera, 

a camera in which the film moves relative to the camera lens. An example of 

such a photograph is shown in Figure 1.8 (Reference 1.1 1 ). The leader is seen 

originating at the top left-hand corner of the picture and lengthening as time 

8 

2 

8 

GROUND 10 

6- 

- 

//////////////////////// /////////////////////////I f 
increases. The bright line at the right of the picture is called the return stroke 

and will be discussed shortly. 

Figure 1.8 Boys camera photograph of a lightning leader. 

From such hotographs it has been learned that the leader advances at tF about 1 to 2 x 10 m/s, or 0.03 to 0.06% of the speed of light (Reference 1.12). 

In order that a charge of 2 to 20 x 1 o-~ C be deposited by a leader advancing at 

the rate of 1 x lo5 m/s requires the average current in the leader, ip, to be 20 to 

200 A. A current of this magnitude could be carried only in a highly conducting 

arc discharge, the assumed central conducting filament of the leader. Such an arc 

would have a diameter on the order of a few millimeters and an axial voltage 

gradient, gp, of about 5 x lo3 V/m. A leader 4 km long would then have a 

voltage drop along its length of 2 x lo7 V. The longitudinal resistance, RQ, of 

the conducting filament gy/iQ would then be in the range of 40 to 400 ohms per 

meter (Rim). 

While of less importance as regards aircraft, it might be noted that leaders 

sometimes start at the ground and work their way toward the sky. This happens 

most frequently from tall buildings or towers, or from buildings or towers 

located atop hills. Generally, one can tell from the direction of the lightning 

flash branching whether the leader started at the cloud or at the ground: if the 

branching is downward (Figure 1.9), the leader originated at the cloud; if the 

branching is upward, the leader originated at the ground. 

1.2.4 The Return Stroke 

As the negatively charged step-leader approaches the ground, positive 

charge accumulates in the ground underneath it or, more accurately, negative 

charge is repelled away from the region under the leader. At some point the 

electric field strength around objects on the ground becomes sufficiently high 

that a streamer starts at the ground and works its way toward the downwardapproaching leader. When the streamers meet, the conducting filament in the 

center of each streamer provides a low-impedance path so that the charge stored 

in the head of the leader can flow easily to ground. As the current in the central 
CLOUD 

Figure 1.9 Leader direction as determined from direction of branching. 

(a) Downward-branching leader starts at cloud. 

(b) Upward-branching leader starts at ground. 

filament rises from its initial current of a few tens of amperes to higher values, it 

gets hotter, its diameter expands, its longitudinal gradient decreases, and it 

becomes an even better conductor, which in turn allows even more current to 

flow in the arc. As the charge in the lower part of the leader is neutralized, the 

heavily conducting arc reaches higher into the charged leader channel. The head 

of the region in which this neutralization takes place moves upwards at a rate of 

roughly 100 000 km/s (or one-third the velocity of light) until it reaches the 

cloud. This heavily conducting region, called the return stroke, produces the 

intense flash normally associated with the lightning stroke. Some stages in the 

development of the return stroke are shown in Figure I. 10 (Reference 1.13). 

The velocity with which this return stroke propagates, together with the 

amount of charge deposited in the leader channel, determines the amount of 

current developed in the return stroke. Let v be the velocity of the return stroke 

and q be the amount of charge deposited per unit length, dQ, along the leader 

channel. Since 

and 

then 
As a numerical example let 

The velocity of the return stroke is not constant from one stroke to the 

next. It seems to vary with the magnitude of current that is ultimately 

developed. The relationship between current and velocity may be deduced either 

from theoretical concepts or experimentally. The relationship derived by Wagner 

(Reference 1.14) is shown in Figure 1.1 I. Considerations of the return stroke 

velocity are primarily of importance in studying the time history of the electric 

field produced by the lightning flash. The velocity, however, may affect the 

surge impedance of the lightning stroke, and thus the way that the stroke 

interacts with a metallic conductor like an aircraft. 

TIME 

INAf C D E F G H 

p SEC -5 0 4 8 13 15 17 19 

Figure 1.10 Stages in the development of the return stroke. 

a 

z 8 1.2~10~ - w V) 

E w 0 

V) O.9x1O8' oz YI 

I 

0 50 100 150 200 

STROKE CURRENT IN KILOAMPERES 

Figure 1.11 Relation between stroke current and velocity of return stroke. 
The velocity of propagation of the return stroke is less than that of the 

speed of light for two basic reasons. The first reason involves the longitudinal 

resistance of the return stroke channel. 

Some of the factors associated with this longitudinal resistance are shown 

in Figure 1.12. Central to the phenomenon is the fact that the current in the 

lightning channel must increase fairly rapidly from the 200 A (approximately) 

(4 

Figure 1.12 Phenomena associated with passage of the return stroke. 

(a) Current 

(b) Longitudinal voltage 

(c) Longitudinal resistance 
current associated with the initial development of the leader to a current of 

perhaps 100 kA as the return stroke becomes fully developed. It is a 

characteristic of an arc channel discharge that if the current through the arc is 

increased, the arc channel expands in diameter, keeping a fairly constant current 

density across the channel. This channel cannot expand instantaneously, since 

energy must be put into the channel to cause the channel to heat up sufficiently 

to force it to expand. Accordingly, if the current through the arc channel is 

increased suddenly by a large magnitude, as in Figure 1.12(a), the longitudinal 

voltage gradient of the channel must suddenly increase. Since the rate at which 

energy is injected into the channel is the product of the current and the 

longitudinal voltage gradient, the increased longitudinal voltage gradient may be 

taken as the mechanism forcing the arc channel to get hot enough to expand to 

the diameter required to carry the high currents. It is not known what the 

maximum longitudinal voltage gradient would be in a lightning channel, but it is 

known from studies of arcs in laboratories that the gradient will fall to values on 

the order of 100 kV/m in a fraction of a microsecond. Presumably in a few 

microseconds, the channel dianreter will have expanded to its final value, and the 

longitudinal voltage gradient will have decayed back toward values on the order 

of 5 to 10 kV/m. The longitudinal resistance, then, would fall from values on the 

order of 40 n/m to values on the order of a small fraction of an ohm per meter, 

in times on the order of a few microseconds. 

This collapse of longitudinal resistance, however, is far from instantaneous. 

The initial resistance of the leader is sufficiently high to retard the development 

of the upward-going return stroke and hence reduce its velocity of propagation 

below that of the speed of light. Presumably, leaders which lead to the 

formation of high-amplitude lightning currents either have a sufficiently low 

longitudinal resistance to begin with or the longitudinal resistance is reduced to 

low values sufficiently fast by the high-amplitude return strokes that the 

longitudinal resistance presents less of an obstacle for the upward-going return 

stroke than it does for the flashes which involve lower peak currents, 

An additional factor that affects the velocity of the return stroke and is 

the second reason that the velocity is less than the speed of light is shown in 

Figure 1.13. As explained earlier, the leader deposits in its wake a column of 

electrical charge with diameters on the order of several meters. At the center is a 

highly conducting core, which has a diameter of a few millimeters for the leader 

and which expands to a few centimeters during the passage of the return stroke. 

The inductance of this return stroke is determined by the diameter of the highly 

conductive central core, and the capacitance by the diameter of the column of 

electrical charge. The lightning stroke may then be modeled as shown in Figure 

1.13(b), in which a highly conductive central conductor is fastened onto a series 

of projecting splines, much like the backbone of a fish. A better analogy might 

be to view the lightning flash as a piece of tinsel rope for decorating a Christmas 

tree: a central piece of string is surrounded by a tube of fine filaments projecting 

radially away from the central core. In either case, the radial filaments can carry 

a radial current, i,, but cannot carry an axial ~urrent, ia. Accordingly, the 
lightning return stroke has both a high capacitance and a high inductance per 

unit of length. In this respect it differs from a solid conductor of large diameter 

which, while possessing a high capacitance per unit length, simultaneously 

possesses a low inductance per unit length. It follows that the surge impedance, 

governed by the ratio of inductance to capacitance, is high while the velocity of 

propagation, governed by the product of inductance and capacitance, is less than 

that of the speed of light. 

(a) 

(b) 

Figure 1.13 Effect of corona cloud on velocity of propagation. 

(a) Distributed charge surrounding a highly conductive central 

core 

(b) Highly conductive central conductor fastened onto a series 

of projecting splines 

Wagner (Reference 1.15) concludes that the surge impedance of the 

lightning flash is of the order of 3 000 52 for return strokes of large amplitude, 

100 kA. This value is large compared to the surge impedance (~500 a) of a 

simple conductor in air and remote from a ground plane or other current return 

path. 

The waveshapes of lightning-flash currents measured at ground level are 

reasonably well known, principally from the work of Berger (Reference 1.16). 

Typical waveshapes detailing the front of the initial return stroke are shown in 

Figures 1.14 and 1.1 5 (References 1.17 and 1 .I 8). In all cases, the current is 
seen to have a concave front, the current initially rising slowly but then 

increasing to a maximum current rate of change just before crest amplitude is 

reached. It may be speculated that the initial slowly changing portion of these 

current oscillograms (which, of course, were measured at ground level) 

represents the growth of an upward-going leader from the lightning tower 

reaching upwards to contact the downward-approaching lightning leader. It can 

also be speculated that the maximum rate of change of current, which occurs 

just before crest, is most representative of the rate at which the current can 

increase in the lightning channel as the return stroke passes one particular point 

in space. This is supported by the observation that subsequent strokes in a - lightning flash, even measured at ground level, exhibit front times considerably 

faster than the rise time of the initial stroke in the flash. 

The true front time of the leading edge of the return stroke as it passes a 

point remote from ground has probably never been measured. It, however, seems 

appropriate to assume that it will be faster than the leading edge of currents 

measured at ground level. 

Figure 1. 14 Front waveshapes of lightning currents as measured by Berger. 

15 
Figure I. I5 Current oscillograms from single strokes or first downward strokes. 

1) Fast time-scale t 1 ; 2) Slow time-scale t2. In osc. No. 6 1 19 TI, 

chopping may be caused by a flashover in the measuring equipment. 

1.3 Further Development of the Lightning Flash 

After the charge has been drained from the leader by the upwardly moving 

return stroke, the current measured at the ground decays, though at a rate 

slower than that at which the current rose to its peak. Oscillograms showing 

typical decay times are shown in Figure 1.15 (Reference 1.18). The figure 
displays the current on two different time scales, emphasizing the front and the 

tail. Some of the oscillograms showing the front are the same as those shown in 

Figure 1.14 (Reference 1.17). 

As the return stroke approaches the cloud, it may encounter other 

branches of the leader, as shown in Figure 1.16. As it passes these branches, the 

charge stored in them will feed into the developing lightning stroke and 

momentarily increase the current. Eventually, the return stroke will reach the 

cloud. Our understanding of the phenomena occurring within the cloud is 

hindered by our not being able to see the phenomena, but we can infer some of 

the phenomena from measurements of the electrical radiation produced by the 

developing flash and from the usual behavior of the flash after the initial return 

stroke has passed. As the return stroke reaches into the cloud, it appears to 

encounter a much more heavily branched leader than it did in the air below the 

cloud. The return stroke can thus tap the charge diffused through a large volume 

of the cloud, rather than only the charge in the more localized leader. It would 

CHARGE ON BRANCHES 

UPWARD-MOVING RETURN STROKE 

Figure 1.16 Further development of the flash. 

17 
appear to be during this period that the intermediate current (Component D of 

Figure 1.1) is developed. As the discharge continues to spread through the cloud, 

for times on the order of fractions of a second, currents on the order of a few 

hundred amperes continue to flow in the lightning flash. These are referred to as 

continuing currents (Component E of Figure 1.1). As one may expect, there is 

no clear-cut demarcation between the tail of the return stroke and the 

intermediate current, or between the end of the intermediate current and the 

start of the continuing current. 

Eventually, and usually, the developing discharge within the cloud reaches 

into a different cell of the cloud or, at any rate, into a region where there is 

another localized body of electrical charge. At this stage there occurs what is 

called a restrike (Component F of Figure 1 .I). The restrike starts with additional 

charge being lowered from the cloud to form a new leader, or, more properly, to 

recharge the central portion of the old leader. Presumably, because of the 

residual ionization in the channel, this charging process occurs smoothly, not in 

the step-by-step process by which the initial leader penetrates into the virgin air. 

Accordingly, this is called a dart leader instead of a step-leader. Unlike the initial 

step-leader, the dart leader is seldom branched. When the dart leader reaches 

ground level, a return stroke again occurs. The amplitude of this return stroke is 

again high, since the current comes from an intensely ionized region close to the 

ground. While the amplitude is usually not as high as that of the first return 

stroke, the current rises to crest more rapidly than does that of the initial return 

stroke, presumably because the upward leader from the ground does not have to 

propagate into virgin air. 

1.4 Lightning Polarity and Direction 

Most Iightning flashes originate in the cloud and lower negative charge to 

earth. The question of direction of the lightning flash is sometimes confusing. 

With the intent of clarifying matters, the statement is sometimes made that 

lightning strikes upward and not downward. This is at least partially true; the 

return stroke that produces the high peak currents, thunder, and the highest 

intensity light, in fact does start near the ground and grow upward into the 

ionized channel previously established by the step-leader, thus tapping the 

charge in the step-leader. The step-leader, nevertheless, originated at the cloud. 

The source of energy is in the cloud, and the lower amplitude and longer 

duration currents have their origin in the charge stored in the cloud. Thus, in 

terms of the engineering definition of current, these flashes result in the 

direction of current flow from the earth to the cloud. This type is commonly 

called a negative polarity flash. 

When tall buildings or mountain tops are involved, the lightning flash often 

does originate at the ground; the step-leader starts at ground level and propagates 

upwards into the cloud. Such flashes seem to be triggered by the high electric 

field concentrated around the top of the building or mountain. They may be 

recognized by the upward direction of branching, as mentioned earlier and 

shown in Figure 1.9(b). This type of flash therefore results in current flow from 
the cloud to the ground and is called a positive polarity flash. Positive polarity 

flashes usually have lower peak currents than do flashes that originate at the 

cloud. 

About 10% of all flashes are positive polarity flashes, and a fraction of 

these involve the highest peak currents and charge found associated with 

lightning. Examples of some of these strong positive flashes are shown in Figure 

1.17 (Reference 1.19). The positive flashes typically have only one high current 

stroke; they lack the restrike phase generally noted on flashes of negative 

polarity. 

Figure 1.1 7 Examples of strong positive strokes. Currents are recorded on timescales t 1 and t2; Qs electric charge (coulombs) within 2 ps from 

the origin; Qk electric charge (coulombs) in the continuing 

current after 2 ps. 

1.5 Intracloud Flashes 

Intracloud flashes occur between charge centers in the cloud. A distinguishing characteristic of intracloud flashes is that they seem to lack the intense 

return stroke phase typical of flashes to the ground, or at least that the electrical 

radiation associated with true intracloud discharges lacks the characteristics 

associated with the return stroke of cloud-to-ground flashes. Discharges between 

charge centers take place during cloud-to-ground flashes as well, and, to an 

observer within the cloud, it may be difficult to tell whether or not a flash to 

ground occurred. 
With regard to aircraft the matter may be academic. Aircraft are struck 

underneath clouds by clear-cut cloud-to-ground flashes and by flashes within 

clouds. Based on the damage observed, the peak current sometimes is very high. 

Whether the high current was associated with the upper end of a cloud-to-ground 

flash or with a true intracloud flash makes little difference. 

In temperate regions about two-thirds of all flashes are intracloud flashes. 

In tropical regions, where there is more lightning activity, the ratio is higher. 

1.6 Measured Characteristics of the Lightning Flash 

Lightning flashes are quite variable from one to another. Peak currents, 

total duration, waveshapes, number of strokes in the flash, charge transferred, 

etc., may all vary over wide limits, and only in general terms can one find a 

correlation between different parameters. Data on the characteristics of lightning 

are best presented in statistical terms, the mode that will be used in the 

following sections. 

One item that needs to be emphasized is that virtually all the data on 

lightning comes from measurements made at ground level, and these measurements may be influenced by the growth of an upward leader. Very few 

measurements have been made of the amplitude and waveshape of lightning 

currents passing through aircraft. Most of the measurements that have been 

made were of strokes with lower peak currents and longer times to crest than 

those often observed at ground level. In part this may be explained by chance 

and in part by the fact that many of the flashes intercepted may have been 

intracloud flashes and not cloud-to-ground flashes. As noted earlier, intracloud 

flashes often lack the well-defined high-amplitude return stroke of cloud-toground flashes. 

The best summary of the statistical characteristics of lightning is that 

compiled by Cianos and Pierce (Reference 1.20). They observed that many of 

these characteristics were nearly linear when plotted as a log-normal distribution. 

They then made a judgment as to the linear distribution that was the best fit to 

the experimentally observed data. The figures that follow are reproduced from 

their report. 

Figure 1 .I 8 (Reference 1.2 I) shows data on the peak current amplitude in 

lightning strokes. Regarding the damage that may be caused by lightning, this is 

one of the most important parameters. There are two curves shown, one for the 

first return stroke in a flash and one for subsequent return strokes. The first 

return stroke is generally of the highest amplitude. For engineering analysis, 

Cianos and Pierce have determined that subsequent return strokes may be 

represented as half the amplitude of the first return stroke. Marked on the curves 

are the amplitudes corresponding to the 2, 10, 50,90, and 98% probabilities. 

The peak value of the current is related to the explosive, or blasting, effect 

of lightning. It is also relative to the maximum voltage developed across ground 

resistance and hence to the risk of side flashes occurring in the vicinity of objects 

struck (or, related to the maximum voltage developed across loading resistors 

struck (or, related to the maximum voltage developed across bonding resistance 

and hence to the possibility of sparking at structural interfaces). 

20 
SUBSEQUENT RETURN 

STROKES 

0.1 1 2 5 10 20 30 40 50 60 70 80 90 95 98 99 99.9 

% > ORDINATE 

Figure 1.18 Distribution of peak currents for first return stroke and subsequent strokes. 

loo & I II I I I11III I I I I 3 - - 

- - 

- 

- 

- 

- 

- - I- - - u - E 

- - E - - 3 - u - 

Y 

- 

$ - - 

4 

0 

7 

"J 1- - 

I - - - - - - - - - - 

- - 

0.1 i I IL 1 I I IILII I 

001 0.1 1 2 5 10 20 yl 40 50 60 70 80 90 95 98 99 99.9 99.99 

% > ORDINATE 

Figure 1.19 Distribution of time to peak current. 

21 
Figure 1.19 (Reference 1.22) gives a distribution of the time for the 

current to reach its peak amplitude. This time is subject to considerable 

interpretation for any particular lightning stroke, since there is seldom a clearly 

definable time at which the stroke starts. Lightning strokes typically have a 

concave front, starting out slowly and then rising faster as the current gets 

higher. Thus, the effective rate of rise of the lightning current is not directly 

obtained by dividing the peak current by the front time. The best summary of 

the effective rates of rise is given on Figure 1.20 (Reference 1.13). The rate of 

rise of a lightning current is an in~portant factor in determining how much 

voltage is induced into electrical equipment, and in determining how many 

lightning conductors are needed and how they sliould be placed. 

0.01 0.1 1 2 5 10 20 30 40 50 60 70 80 90 95 98 99 99.9 99.99 

% > ORDINATE 

Figure 1.20 Distribution of rates of rise. 

The duration of the stroke current affects the distance across which side 

flashes may develop, and affects how severely metal structures may be deformed 

by magnetic forces or the explosive libelation of energy. This distribution 1s 

shown in Figure 1.2 1 (Reference 1.24). The duration of the stroke, which is 

measured in tens of microseconds, should not be confused with the total 

duration of the lightning flash. The total duration, shown in Figure 1.27 

(Reference 1.25), is frequently on the order of a second. The duration of the 

total flash is influenced by the number of return strokes in the flash (Figure 1.23 

[Reference 1.261) and the time interval between strokes (Figure 1.24 

[Reference 1.271 ). 

Figure 1.25 (Reference 1.28) gives information on the total charge 
1OOO. I II I I I IIIII I I 111 I - - - - 

- 

- 

- 

- 

- 

- 

I - 

W 

- -

:: - - - - 

_I 

- 

a - I - 

I- - z 

- 

LY - a 

u 

U 3 

o 10 r- - 

- YI 

z - 

- 

- 

- - 

- - 

1 - I I1 I l.iIII 11 11 I 

0.01 0.1 1 2 5 10 20 30 40 50 60 70 80 90 95 98 99 99.9 99.99 

% > ORDINATE 

Figure 1.21 Distribution of time to current half value. 

10,oW~ I II I I I IIIII I I 111 I =I - 

- 

- - 

- - 

- - - 

- 

- 

- - I 

- 

- 

- 

- 

Y - - 0 

z - - 

0 

loo, 0 - 

- 

- 

- 

- 

- 

- 

10 - I I1 I I I IiIII 1 11 11 

0.01 0 1 1 2 5 10 20 30 40 50 60 70 80 90 95 98 99 99.9 99.99 

% > ORDINATE 

Figure 1.22 Distribution of duration of flashes to earth. 

23 
Figure 1.23 Distribution of the number of return strokes/flash. 

18 

16 

1 14 $ 

2 12- x 

B 10- 

Z E 

3 + w 8- 

a 

u 

0 

u 6- Y 

m 

I 

z 4- 

2 

0 

- I I1 I I I IIII) I 1, (1 I 

- - 

- - 

- 10-1 1 STROKES 

- 

- 5-6 STROKES 

- 

- 

2-3 STROKES 

- - 

I II I I IIIIIi~.I 111 I 

Figure 1.24 Distribution of time interval between strokes. 

0.01 0.1 1 2 5 10 20 30 40 50 60 70 80 90 95 98 99 99.9 99 99 

% > ORDINATE 

loo0 

E 

I 100 

cn u 

Y 

B 

t, 

Z 

YI 

i 

m 

u 10 

I 

1 

- II I I I III(I ( I - 

- I II I - 

- - 

- - 

- - 

- 

- 

- 

- - 

- 

- - - - - 

- - 

- 

- - 

- - 

- - 

=- 

- 

- 

- 

- 

- 

- 

- - 

I I 1-UU-l 1 I 

0.01 0.1 1 2 5 10 20 30 40 50 60 70 80 90 95 98 99 99.9 W99 

% > ORDINATE 
% > ORDINATE 

Figure 1.25 Distribution of chargelflash. 

loo0 I I I I I IIIII I I 111 I - 

- - - 

- - 

- - - 

- 

- 

100 - - 

- 

U - 

- 

- 

I - 

- 

- 

Y 

- - 

(3 - - a 

a 

I 

U 

- - 

10 =- 

- 

- 

- 

- 

- 

- 

- 

1 I 11 I1 1 IlIII I '1 11 I 

0.01 0.1 12 5 10 20304050607080 90 95 9899 999 99.99 

% > ORDINATE 

Figure 1.26 Distribution of charge in continuing current. 
Figure 1.27 Distribution of duration of continuing currents. 

1o.m !- 

% > ORDINATE 

Figure 1.28 Distribution of amplitude of continuing current. 

26 

E 

I 

Iz 

I I I I I I-- - 

- 

- - 

- - 

- 

- - 

- - 

- 

- - 

- 

- - 

- 

- 

- 

- 

I- 

- 

- 

Q - 5 

3 - a - 

- - 

10 I L~-I LI 1 11111 I I I 11 I 

0.01 0.1 12 5 10 20304050607080 90 95 9899 99.9 99.99 

% > ORDINATE 
transferred in the flash. Little of the charge is transferred by any one stroke. 

Instead, most of it is transferred by the continuing currents. The total charge 

transfer and the amplitude and duration of the continuing currents largely 

govern the thermal effects of lightning. Data on the characteristics of these 

continuing currents are shown in Figures 1.26, 1.27, and 1.28 (References 1.29, 

1.30, and 1.3 1). 

1.7 Thunderstorm Frequency and Lightning-Flash Density 

One of the major factors to consider in determining the probability of 

lightning damage is the number of lightning flashes to earth in a given area and 

for a given time. Since precise quantitative data do not exist (except at a few 

specifically instrumented structures), a secondary measure, the frequency of 

thunderstorms, is used. 

For many years, weather bureau stations have recorded thunderstorm days 

(the number of days per year on which thunder is heard). This index, called the 

isokeraunic level, is shown for the continental United States in Figure 1.29' 

(Reference 1.32). It should be noted that the information so collected is of 

limited value for several reasons. First of all, no distinction is made between 

cloud-to-cloud discharges and cloud-to-ground flashes. Also, there is no 

allowance for the duration of a storm. A storm lasting an hour would be counted 

as heavily as one lasting several hours. A better indicator of lightning frequency 

would be thunderstorm hours per year. Some weather bureau records are now 

being made of thunderstorm hours, but not much data has yet been 

Figure 1.29 Thunderstorm days (isokeraunic level) within the continental 

United States as reported by U.S. Weather Bureau. 

2 7 
accumulated. Despite its limitation, the isokeraunic level is broadly useful and 

can be correlated at least partially with lightning strokes to earth-based objects. 

Pierce has summarized some of the available data, shown in Figure 1.30 

(Reference 1.33). 

0 PIERCE 

- 

WESTINGHOUSE 

JAPANESE 

--- EUROPEAN 

- 

lo-= 10-I 1 10 1 02 

Figure 1.30 Relationship between annual thunderstorm day (Ty) and flash 

density (ay) values. 

This flash density includes both flashes between clouds and flashes to 

ground. There is some evidence that the proportion of flashes that go to ground 

is related to the geographical latitude of the point under study. This relation 

would exist because the proportion of flashes to ground depends partly on the 

average height of clouds, and this, in turn, depends on the type of storm 

formation. Plerce (Reference 1.34) has proposed the relat~on shown in Figure 

I .3 I). 

The number of lightning flashes to any particulal object depends strongly 

on the terrain upon which the object is situated. Objects on the crests of hills are 

more prone to be struck than are objects in valleys. All othel things considered, 

however, the probability that a given object will be struck depends upon the area 

covered by that object. For objects flat upon the ground- -for example, a long 

group of cables laid on the ground surface--the strike interception area can be 

taken as directly equal to the area covered. Figure 1.37 shows such a group of 

cables. Equating the stroke interception area to the actual area covered implies 

that the stroke interception probability is unity for stroke 1 and zero for stroke 

7 falling somewhat to the side. This may not be strictly true, since the ultimate 

contact point of a lightning flash seems to depend upon the junction of the 

downcoming lightning leader and induced leaders which are drawn from the 
ground when the electric field strength at the ground surface becomes high 

enough. Such induced leaders could well appear somewhat earlier from a group 

of cables lying on the ground than from the undisturbed ground surface. 

0.15 I 

25 30 35 40 45 50 

DEGREES LATITUDE 

Figure 1.31 Relationship between geographicaI Iatitude and proportion of 

flashes to ground. 

If there is a protruding object on the ground, the stroke interception area 

depends upon the height of the object. While there is no uniformity of judgment 

on the effective stroke interception area of a protruding object, one can at least 

glean some information from the observed pattern of damage to objects near a 

tall building or on a building protected by lightning rods. It has long been 

observed that lightning very seldom strikes within a one-to-one cone of 

protection of a protruding object. The significance of this fact is shown on 

Figure 1.33. A lightning flash that would ordinarily continue directly to the 

ground at point r will instead be diverted to the protruding object. This implies 

that if the distance to point r is less than h, the probability of the protruding 

object being struck at a distance less than h is very little less than unity. It stands 

to reason also that the probability of striking the ground falls off from unity 
, 

///I//////// 

.......... 

(a) GEOMETRY 

(b) PROBABILITY 

Figure 1.32 Probability of striking a flat area 

only gradually as one moves farther and farther away from the protruding 

object. For purposes of calculations of lightning-stroke incidence rate, it can 

then be assumed that a structure of height h will intercept all flashes that would 

ordinarily strike the ground over a circle of radius 2h. 

It is not as easy to determine analytically how often aircraft in flight will 

be struck. Field experience seems to be the only reliable guide. Commercial 

aircraft in regularly scheduled service in the United States are struck about once 

per year, frequently while in takeoff, landing, or holding patterns, and usually 

while flying at less than 15 000 ft altitude. Transport aircraft are seldom struck 

while at cruising altitudes and speeds, partly of course as a result of the fact that 

storms usually can be, and are, avoided. Aircraft constrained to operate at lower 

altitudes and along fixed corridors tend to be struck more often. 
LIGHTNING FLASH 

(a) GEOMETRY DISTANCE 

DISTANCE 

(b) PROBABILITY 

Figure 1.33 Probability of striking a protruding object. 

Military aircraft tend to be struck less often than commercial aircraft, 

since they are flown in training flights, which are usually scheduled during good 

weather, much more often than in combat. 

1.8 Engineering Models of Lightning Flashes 

For purposes of analysis it is helpful to have models of the current flowing 

in both typical and severe lightning flashes. Pierce (Reference 1.35) gives several 

3 1 
j 20 TlME TO PEAK CUARENT - 1.5 pr 

1 10 + 

0 z 

a 5 a 

3 

u 1 

0.1 

(a) FLASH WITHOUT ANY CONTINUING CURRENT STAGES 

TlME TO PEAK CURRENT = 1.5 us 

0.3 rnr 

(b) FLASH WITH FINAL STAGE CONTINUING CURRENT 

Figure 1.34 Time history of typical (basic) lightning models. 

TlME TO PEAK CURRENT - 1.5 ps 

TIME TO HALF VALUE - 40 #S ib-i---A r, A /, A A I- 

? CONTINUING 

I CURRENT FINAL STAGE 

4W A 200 A UI 

a TIME a 

a0 lorn 60rn 11Omr 160ms 460 mr 520 ms 580 rnl MO rnl 7W rnr 860 mr 

Figure 1.35 Time history of severe (basic) lightning model. 

models, shown in Figures 1.34 and 1.35 (References 1.36 and 1.37). Other 

models may be devised. One that has received wide prominence in the aerospace 

field is the Space Shuttle Lightning Protection Criteria waveform (Figure 1.36 

[Reference 1.381). These, of course, are models for design purposes. As such 

they duplicate the effects (usually worst case effects) of lightning, but the 
chances of any real lightning flash producing currents of this exact shape is 

virtually zero. 

- First stroke 

- - - Second return strnke 

j i 39 ms i i 2 ps 100 55 ms ,5 5 ms i 

+PI= 17ms TO 230 ms 1 

Tbme ~nrrearlnp- 'j 1. = 2 lO!.S 

t t 5 mr 

t5 - 

Figure 1.36 Diagrammatic representation of lightning model. 

(Note that the diagram is not to scale.) 
REFERENCES 

1.1 Sir Basil Schonland, "Lightning and the Long Electric Spark," Advancement of Science 19 (November 1962): 306-3 13. 

1.2 C. T. R. Wilson, "Some Thunderstorm Problems," Journal of The Franklin 

Institute 208, 1 (July 1929): 1-1 2. 

1.3 C. T. R. Wilson, "Investigations on Lightning Discharges on the Electrical 

Field of Thunderstorms," Proceedings of the Royal Society of London, 

Series A, 221, the Royal Society of London (1920): 73-1 15. 

1.4 Sir George Simpson, "The Mechanism of a Thunderstorm," Proceedings of 

the Royal Society of London, Series A, 114, the Royal Society of 

London (1927): 376-99. 

1.5 E. A. Evans and K. B. McEachron, "The Thunderstorm," General Electric 

Review, September 1936, pp. 4 13-25. 

1.6 G. C. Simpson, "Lightning," Journal of the institution of Electrical 

Engineers 67, 395 (November 1929): 1269-82. 

1.7 G. C. Simpson and F. J. Scrase, "The Distribution of Electricity in 

Thunderclouds," Proceedings of the Royal Society of Idondon, Series A, 

161, the Royal Society of London (1937): 309-52. 

1.8 R. H. Golde, Lightning Protection (London: Edward Arnold, I973), p. 6. 

1.9 J. L. Marshall, Lightning Protection (New York: John Wiley and Sons, 

1973), p. 12. 

1.10 C. F. Wagner and A. R. Hileman, "The Lightning Stroke---11," AIEE 

Transactions 80, Part 111, American Institute of Electrical Engineers, New 

York, New York (October 196 1): 622-42: 623. 

1.1 1 K. Berger and E. Vogelsanger, "Photographische Blitzantersuchungen der 

Jahre 1955.. . 1965 auf dem Monte San Salvatore," Bulletin des 

Schweizerischen Elektrotechnischen Vereins, 14 (July 9, 1966): 599-620: 

608. 

1.12 K. Berger, "Novel Observations on Lightning Discharges: Results of 

Research on Mount San Salvatore," Journal of The Franklin Institute 283, 

6 (June 1967): 478-525: 5 14. 

1.1 3 Wagner and Hileman, "The Lightning Stroke 11," p. 629. 

1.14 C. F. Wagner, "Lightning and Transmission Lines," Jourrzal of The 

Franklin Institute 283, 6 (June 1967): 558-594: 560, and C. F. Wagner, 

"The Relation between Stroke Current and the Velocity of the Return 

Stroke," IEEE Transactions on Power Apparatus and Systems 82, 

American Institute of Electrical Engineers, New York, New York (October 

1963): 609-1 7: 609. 

1.15 C. F. Wagner and A. R. Hileman, "Surge Impedance and Its Application to 

the Lightning Stroke," AIEE Transactions 80, Part 3, American Institute 

of Electrical Engineers, New York, New York (February 1962): 101 1-22. 

1 .I 6 See Berger and Vogelsanger, "Photographische Blitzuntersuchungen der 

Jahre 1955.. . 1965 auf dem Monte San Salvatore," Bulletin des 

Schweizerischen Elektrotechnischen Vereins, 14 (July 9, 1966): 599-620, 

and K. Berger, "Novel Observations on Lightning Discharges: Results of 
Research on Mount San Salvatore," Journal of The Franklin Institute, 

283,6 (June 1967): 478-525. 

1.1 7 Redrawn from C. F. Wagner, "Lightning and Transmission Lines," p. 562. 

1.18 Berger, "Novel Observations," p. 492. 

1 .I 9 Berger, "Novel Observations," p. 494. 

1.20 N. Cianos and E. T. Pierce, A Ground-Lightning Environment for 

Engineering Use, Technical Report 1, prepared by the Stanford Research 

Institute for the McDonnell Douglas Astronautics Company, Huntington 

Beach, California (August 1972). 

1.2 1 Cianos and Pierce, p. 65. 

1.22 Cianos and Pierce, p. 67. 

1.23 Cianos and Pierce, p. 68. 

1.24 Cianos and Pierce, p. 69. 

1.25 Cianos and Pierce, p. 63. 

1.26 Cianos and Pierce, p. 6 1. 

1.27 Cianos and Pierce, p. 64. 

1.28 Cianos and Pierce, p. 66. 

1.29 Cianos and Pierce, p. 72. 

1.30 Cianos and Pierce, p. 70. 

1.3 1 Cianos and Pierce. p. 7 1. 

1.32 U.S. Department of Commerce, National Oceanic and Atmospheric 

Administration, Environmental Data Service, National Climatic Center, 

AsheviIle, North Carolina. 

1.33 Cianos and Pierce, p. 124. 

1.34 E. T. Pierce, "Latitudinal Variation of Lightning Parameters," Journal of 

Applied Meteorology, 9 (1 970): 194-95. 

1.35 Cianos and Pierce, A Ground-Lightning Environment, pp. 75-83. 

I .36 Cianos and Pierce, p. 76. 

I .37 Cianos and Pierce, p. 82. 

1.38 Space Shuttle Program Lightning Protection Criteria Document, 

JSC-07636, Revision A, National Aeronautics and Space Administration, 

Lyndon B. Johnson Space Center, Houston, Texas (November 4, 1975), p. 

A-5. 

CHAPTER 2 

AIRCRAFT LIGHTNING ATTACHMENT PHENOMENA 

2.1 Introduction 

Statistics on lightning strikes reported by aircraft pilots seem to indicate 

that no aircraft is likely to receive more than one or two lightning strikes in a 

year but that some types of aircraft receive more strikes than others. Compared 

with exposure to other hazards, such as hail, birds, and turbulence, which are 

also encountered during flight, the exposure of aircraft to lightning strikes seems 

relatively infrequent. Because of this relatively low incidence, inclusion of 

lightning protective measures in the designs of some aircraft has been considered 

unnecessary. The question also arises: If lightning strikes do in fact occur 

infrequently, can they be avoided altogether? Because some types of aircraft 

seem to experience more than their "fair share" of lightning strikes, a related 

question also arises: Why are some aircraft more vulnerable than others? To 

answer these questions a considerable amount of research on the effects of such 

factors as aircraft size, engine exhaust, and microwave radar emissions on 

lightning-strike formation has been undertaken during the past few years. Much 

of this research has been aimed at answering the question of whether or not an 

aircraft can in fact produce its own lightning strike, or static discharge, or if it 

can trigger an impending flash from a nearby cloud. While some of the findings 

are as yet inconclusive, others have provided definite answers to some of these 

questions. In this chapter we summarize what has been learned about the 

aircraft's influence on lightning-strike occurrence and dispel some other 

misconceptions about this phenomenon. In succeeding chapters we examine how 

other factors, such as flight and weather conditions, affect lightning strikes and 

conclude why it is important that lightning protection be incorporated into all 

new aircraft designs. 

2.2 Aircraft Influence on Flash Formation 

At the beginning of lightning-flash formation, when a stepped-leader 

propagates outward from a cloud charge center, the ultimate destination of the 

flash at an opposite charge center in another cloud or on the ground has not yet 

been determined. The difference of potential which exists between the 

stepped-leader and the opposite charge center(s) establishes an electrostatic force 

field between them, represented by imaginary equipotential surfaces. They are 

shown as lines in the two-dimensional drawing of Figure 2.1. The field intensity, 

commonly expressed in kilovolts per meter, is greatest where equipotential 

surfaces are closest together. It is this field that is available to ionize air and form 

the conductive spark which is the leader. Because the direction of electrostatic 

force is normal to the equipotentials and strongest where they are closest 

together, the leader is most likely to progress toward the most intense field 

regions. 

If an aircraft happens to be in the neighborhood, it will assume the 
\ 

Figure 2.1 Aircraft influence on stepped-leader direction. 

electrical potential of its location. Since the aircraft is a thick conductor and all 

of it is at this same potential, it will divert and compress adjacent equipotentials, 

thus increasing the electric field intensity in the vicinity of the aircraft, and 

especially between it and other charged objects, such as the leader. If the aircraft 

is far away from the leader, its effect on the field near the leader is negligible; 

however, if the aircraft is within several tens or hundreds of meters from the 

leader, the increased field intensity in between may be sufficient to attract 

subsequent leader propagation toward the aircraft. As this happens, the 

intervening field will become even more intense, and the leader will advance 

more directly toward the aircraft. 

The highest electric fields about the aircraft will occur around extremities, 

where the equipotential lines are compressed closest together, as shown in Figure 

2.2. Typically, these are the nose and wing and empennage tips, and also smaller 

protrusions, such as antennas or pitot probes. When the leader advances to the 

point where the field adjacent to an aircraft extremity is increased to about 30 

kV/cm, the air will ionize and electrical sparks will form at the aircraft 

extremities, extending in the direction of the oncoming leader. Several of these 

sparks, called streamers, usually occur simultaneously from several extremities of 

the aircraft. These streamers will cont~nue to propagate outward as long as the 

field remains above about 7 kV/crn (Reference 2.1). One of these streamers will 

meet the nearest branch of the advancing leader and form a continuous spark 

from the cloud charge center to the aircraft. Thus, when the aircraft is close 

enougii to influence the direction of the leader propagation, it will very likely 

become attached to a branch of the leader system. 
DIRECTION OF FIELD 

t EQUIPOTENTIAL LINES 

+ 

Figure 2.2 Compression of electric field around an aircraft. 

When the aircraft is attached to the leader, some charge (free electrons) 

will flow onto the aircraft, but the amount of charge which can be taken on is 

limited by the aircraft size. The measure of the aircraft's ability to store charge is 

its capacitance. The capacitance of a con~plex object, such as an aircraft, is 

impractical to calculate but may be estimated by comparing it with the easily 

calculated capacitance of a sphere of equivalent surface area in an infinite field 

in which no other charge centers are present. The capacitance of such a sphere is 

given by 

where 

C = capacitance (farads [F] ) 

a = radius (meters) 

S = surface area (meters2) 

eV = absolute dielectric constant (=10-~/36n for vacuum) 

er = relative dielectric constant with respect to vacuum 

(= 1 for air) 
For a typical transport aircraft with a total surface area of 3000 rn2, C is, 

from Equation 2.1 

= I .7 17 X 1 o-~ = 17 17 picofarads (pF) 

A similar calculation has been made by Schaeffer and Weinstock 

(Reference 2.2) for smaller fighter aircraft, of 420 pF, and by Moore (Reference 

1.3) for helicopters, of 600 pF. 

The amount of charge which can be stored by an object having a known 

capacitance is 

where 

Q = charge (coulombs) 

C = capacitance (farads) 

V = object potential (volts) 

Thus, before the amount of charge which can be stored on an aircraft can 

be found, it is necessary to determine the potential, V, which the aircraft can 

assume. For this purpose, it will be assumed that the potential of the entire 

aircraft is the same as that of a small sphere whose radius is similar to that of the 

sharp aircraft extremities from which streamers develop and from which charge 

begins to leave the aircraft. Assuming, for example, that this radius is 2 cm, the 

capacitance of a sphere of this radius is first determined directly from Equation 

2.1. 

The electric field at the surface of this charged sphere (Reference 2.4) is given by 

E = Q 7 (2.5) 

477~~~~a~ 

If E IS assumed to be equal to 30 kVIcm, which is the ionization stress of air in a 

uniform field, and if Equation 3.3 is substituted in Equation 1.5 for Q, then 
The values of C and a of Equation 2.4 are then substituted into Equation 

2.6 and the equation solved for V, as follows: 

= (30x 1 05j [(O. 1 1 x 1 o-~) (0.02)~] (2.7) 2.22~ 10-12 

- 0.00132~ = 0~000594x108 

2.22x 10-11- 

= 59.4 kV 

Since the cloud and ground charge centers already establish an ambient field, 

59.4 kV is the potential at which the aircraft must be relative to the ambient 

field for streamering to occur. If the capacitance of even a large aircraft is no 

more than about 2000 pF, as given by Equation 2.2, then the maximum amount 

of charge which the aircraft can retain is 

Q = 1 18.8~10-~ coulombs (C) 

Q = 100 microcoulombs (PC) 

This can be considered the amount of charge which a leader can deliver to 

the aircraft before streamers occur from opposite extremities. If additional 

charge flows onto the aircraft, more profuse streaming will occur, and from 

extremities of larger radii of curvature. In fact, the maximum charge which can 

be on the aircraft probably exceeds by up to 100 times the streamer initiation 

value. However, statistics on natural lightning characteristics show that a typical 

leader contains about 1 to 10 C (Reference 2.S), so there is still no room for any 

significant portion of this to accumulate on an aircraft. Thus, the aircraft merely 

becomes an extension of the path being taken by the leader on its way to an 

ultimate destination at a reservoir of opposite polarity charge. Streamers may 

propagate onward from two or more extremities of the aircraft at the same time. 

If so, the incoming leader will have split, and the two (or more) branches will 

continue from the aircraft independently of each other until one or both of 

them reach their destination. This process of attachment and propagation 

onward from an aircraft is shown in Figure 2.3. 

When the leader has reached its destination and a continuous ionized 

channel between charge centers has been formed, recombination of electrons 

and positive ions occurs back up the channel, and this forms the high-amplitude 

return stroke current. This stroke current and any subsequent stroke or 

continuing current components must flow through the aircraft, which is now 

part of the conducting path between charge centers, as shown in Figure 2.4(a). 

If another branch of the original leader reaches the ground before the 

branch which has involved the aircraft, the return stroke will follow the former, 

and all other branches will die out, as shown in Figure 2.4(b). No substantial 
Figure 2.3 Stepped-leader attachment to an aircraft. 

(a) Stepped-leader approaching aircraft 

(b) Stepped-leader attachment and continued 

propagation from an aircraft 

currents will flow through the aircraft in such a case, and any damage to the 

aircraft will be slight. A still photograph of a downward-branching flash after 

completion of the main channel is shown in Figure 2.5. Several dying branches 

are evident in the photograph. 

2.2.1 Precipitation Static 

The foregoing analysis began with the assumption that the aircraft is at the 

potential of its position in an electric field established by the cloud and ground 

charge centers. In dry air this is correct, but if the aircraft is flying through dry 
I2igure 2.4 Return stroke paths. 

(a) Return stroke through the aircraft 

(b) No return stroke through the aircraft 

precipitation in the form of sleet, hail, or snow, the impact of these particles on 

the aircraft will cause a charge to separate from the particle and join the aircraft, 

leaving the aircraft with a preponderance of positive or negative charge 

(depending on the form of precipitation), thereby changing its potential with 

respect to its surroundings. This phenomenon is known as triboelectric charging 

and has been extensively studied by Tanner and Nanevicz, and others (Reference 

2.6). It is commonly referred to asprecipitatiorz static, or P-static. 

The P-static charging process is easily capable of raising the aircraft to a 

potential of 50 kV, or more, with respect to its surroundings, a charge sufficient 

to cause ionization at sharp extremities. This ionization radiates broadband 
Figure 2.5 Downward-branching flash. 

electromagnetic radiation (EMR) throughout the low- and high-frequency 

spectrum. This EMR is often received as interference, or static, by the aircraft 

communications or low-frequency automatic direction finding (LF-ADF) 

receivers. The radiated EMR spectrum from P-static discharging at a typical 

trailing edge is shown in Figure 2.6 (Reference 2.7). 

The EMR results from a continuous series of minute streamer-like 

discharges of ionized air in the immediate (i.e., 10 cm) vicinity of sharp 

extremities. These discharges also produce a continuous ultraviolet glow visible 

at night and called St. Elmo 's Fire or corona. 

P-static discharging (corona) will occur initially from the sharpest 

extremities, where the surrounding field first reaches the ionization potential for 

air. If more P-static charge enters the aircraft than is bled off by the discharges at 

these extremities, the aircraft potential will increase until the field surrounding 

extremities of large radii also becomes intense enough to ionize air. Thus, as the 

aircraft potential increases, the radiated EMR from static discharging becomes 

more intense, and so does the associated static in communications receivers. 
FREQUENCY - Mc 

Figure 2.6 Normalized noise spectrum from trailing edge. 

To reduce the level of P-static interference, it is therefore appropriate to 

reduce the amount of charge which can be stored on the aircraft, and thus its 

potential relative to its surroundings. This has been done (Reference 2.8) by 

attaching devices called sratic dischargers to the aircraft. These are brushes or 

sharp needles with very small radii of curvature and thus of low corona inception 

potential. They are most effective in draining charge from the aircraft when 

located at regions where the surrounding field stress is already high and where 

the airflow may readily carry away discharged electrons or ions. Thus, they are 

usualIy found on trailing edges of wing tips, empennage tips, and tail cones, as 

shown in Figure 2.7. While these dischargers are sometimes struck by lightning, 

they are not capable of either attracting or preventing lightning flashes, or of 

diverting them from other attachment points on the aircraft. 

P-static persists for as long as the aircraft is being charged by impact with 

dry precipitation, such as sleet or snow. It is rarely reported in rain. It is thus a 

continuous phenomenon lasting from several seconds to many minutes. Once the 

aircraft has left such a region, the static in communications receivers quickly 

clears up, and the aircraft potential reverts to that of its surroundings-- - 

established again by its location in the ambient cloud-ground electric field. 

Because of its low capacitance, an aircraft cannot retain enough P-static charge 

to produce a startling flash of several meters (or more) in length or a loud report, 

such as is often heard when lightning strikes the aircraft. Nevertheless, pilots 

often report a "static discharge" from the aircraft, and proceed to describe the 
Figure 2.7 Precipitation static discharger on an aircraft. 

symptoms of a lightning strike: i.e., a long, bright flash or spark extending 

outward from the aircraft, usually accompanied by a loud report, such as that 

which would be produced by a shotgun going off outside the cockpit. 

Thus, the P-static process cannot contribute much to the formation of a 

lightning leader or to the process of leader attachment to an aircraft. It is true 

that an approaching streamer 1s most likely to attach to a point from wh~ch an 

opposing streamer has developed, but the intense field produced by an advancing 

leader would overcome that produced by the P-static process and draw streamers 

of its own from most locations where P-static discharges were occurring. 

The P-static discharging process may be intensified when the aircraft is in a 

region where the amblent electric field is relatively intense, as is the case when a 

lightning flash is imminent. When the flash occurs (whether or not it Intercepts 

the aircraft), the main charge centers are neutralized and the field collapses, 

thereby reducing the intensity of P-static discharging. This is the reason that 

pilots frequently report that P-static interference gradually intensifies until a 

lightning flash occurs and then diminishes instantaneously. This experience 

reinforces the pilot's impression that the flash is a sudden static discharge from 

the aircraft alone. For this reason P-statc interference should be looked upon by 

pilots as an indicator that a lightning strike may be imminent (i.e., withln a few 

seconds to a few minutes). P-static interference is, in fact, reported prlor to the 

lightning strike in about half of the Iightning-strike incidents described in recent 

airline 11ghtning-strike reports (Reference 2.9). 
2.2.2 Can an Aircraft Trigger a Lightning Strike? 

A question often asked is If an aircraft cannot produce its own lightning 

flash, can it trigger a natural one'? Stated another way the question might be 

Would the lightning flash have occurred if the aircraft were not present? 

While there is insufficient scientific data upon which to base a conclusive 

answer to these questions, the following factors suggest that the aircraft does not 

often trigger a flash. 

I. Aircraft often fly through electrified regions without being struck, 

while lightning flashes are occurring nearby. 

2. The stepped-leader must begin from a charge source capable of 

furnishing it with several coulonibs of charge. Thus, the potential 

(voltage) of this center, and the surrounding field intensity, would 

seem to be much greater than that about an aircraft, leaving the 

implication that, unless the aircraft is very close to the charge center, 

it can have little influence on the surrounding field or on the process 

of leader initiation. 

3. Laboratory breakdown tests of long high-voltage air gaps, thought to 

be similar to lightning leader formation, show that initial ionization 

always begins at one of the electrodes and not from an object 

suspended in the gap (Reference 2.10). Such an object significantly 

influences the voltage level at which breakdown begins only if it is 

close enough to one electrode to influence the field about this 

electrode. 

It is more probable that the aircraft does not become involved until after 

leader propagation has begun. If the leader happens to approach the aircraft, the 

field intensification produced by the presence of the aircraft becomes much 

more significant, and the leader may now be attracted to the aircraft. 

There is some evidence (Reference 2.1 1) that jumbo jet (wide-body) 

aircraft do trigger their own flashes, but this is not yet conclusive, since 

accumuIated flight hours are not yet nearly as great as those for conventional 

aircraft. If large-body aircraft are in fact triggering flashes, it is probably because 

their larger sizes make a more noticeable perturbation on the electric field near 

the cloud charge centers from which leaders begin. 

The aircraft motion has little influence on the propagating leader because 

the aircraft is moving much slower, about lo2 mls, than the leader, which is 

advancing at 1 o5 to 1 oh mls. Thus, the aircraft appears stationary to the leader 

during the leader formation process. 

Several other stimuli have been mentioned as possible causes of aircraft 

lightning strikes. These include engine exhaust and radiated electromagnetic 

energy (i.e., radar transmission). 

2.2.2.1 Effect of Jet Engine Exhaust 

There has been speculation that the hot jet-engine exhaust gases may 

contain a sufficient number of ionized particles to attract or trigger a lightning 
flash to the aircraft. This speculation has been heightened by the widely 

publicized launch of Apollo 12, which apparently triggered a lightning flash that 

struck the top of the vehicle when it had reached 1950 m and again at 4270 m. 

The flash exited from the vehicle exhaust plume. Subsequent studies by 

Nanevicz, Pierce, and Whitson (Reference 2.12) of this and other incidents in 

which a rocket was rapidly introduced into an intense electric field indicate that 

the exhaust plume does appear electrically conductive, making the rocket appear 

longer than its own physical length. An empirical study by Pierce (Reference 

2.1 3) of documented strikes to tall grounded and airborne conductors concludes 

that there must also be a potential discontinuity between the conductor and the 

adjacent atmosphere of up to loh V if the lightning leader is initiated from the 

conductor, and that the rapid discharge of hot ionized gas from the rocket 

engine may cause sufficient charge separation from the vehicle to increase its 

potential to lo6 V or more with respect to its surroundings. 

Shaeffer and Weinstock (Reference 2.14) have studied the conductivity of 

an aircraft jet-engine exhaust. In this case, ionized particles and free electrons in 

a jet exhaust originate in the combustion chamber as a result of chemical 

reactions taking place between the intake air and jet fuel. The ion concentration 

in a jet-engine exhaust has been measured by Fowler (Reference 2.15) to be 

between 5x10~ and 3x10~ particles per cubic centimeter ( /cm3) and the free P electron density deduced from this to be between 5 x 10 to 3 x lo5 p/cm3. 

The electron density in luminous rocket exhaust has been calculated by Pierce 

(Reference 2.1 6) to be 1 o1 p/cm3, as has that in the tip of an advancing leader. 

Conversely, the free electron density in ambient air ranges from lo0 to lo3 

p/cm3. Evidently, then, the jet-engine exhaust is only slightly more ionized than 

the ambient air and much less so than the rocket exhaust, with the result that 

the jet exhaust would not be expected to have sufficient conductivity to initiate 

or attract a lightning leader. This conclusion is supported by aircraft lightningstrike incident reports, which indicate that engine tail pipes are not often 

lightning attachment points unless they are already located at an aircraft 

extremity, where the electric field would be intense from geometrical conditions 

alone. 

There is also no evidence to suggest that jet aircraft are struck more often 

than piston-engined aircraft. Overall, the ability of the jet aircraft to operate at 

higher altitudes and spend less time climbing and descending to airports has 

probably rendered the jet less susceptible than its piston-engined predecessor to 

lightning strikes, which occur predominantly at low or intermediate altitudes. 

Recorded strikes which have hit jet engines almost always have terminated on 

the tail pipes of the aft- (tail-) mounted engines of conventional jet transport 

aircraft. However, a few recent reports of strikes to wide-body aircraft show 

evidence of strikes to the tail pipes of wing-mounted engines (Reference 2.1 7). 

The evidence to date is indirect, since no tell-tale burn marks have been found 

on the tail pipes themselves. Instead, electronic engine instruments with sensors 
mounted on the engine have experienced damage, and no attachment points 

have been found other than the flash entry point at a nose or wing tip. Lightning 

attachment points on jet tail pipes may be masked by exhaust deposits or simply 

be indiscernible because of the high melting temperature of tail pipe metals. If 

this is true, the increased volume and temperature of the jumbo jet exhaust may 

have become a region of sufficient ionization or weakened dielectric strength to 

divert a propagating leader into the exhaust channel or to enable the formation 

of outward-propagating streamers from the tail pipe, once the aircraft has been 

struck elsewhere by a leader. While this engine exhaust may divert an existing 

leader, it seems improbable that it could trigger a flash by itself. 

Viewed another way, if the jet engine were leaving behind an ionized 

exhaust, it seems probable that an imbalance of charge on the aircraft would 

eventually result, even when flying in clear weather. As with P-static charging, 

this would raise (or lower) the aircraft potential with respect to its surroundings 

and cause static discharging and corona from sharp extremities, causing 

interference in radio equipment. It is well established that this does not happen. 

2.2.2.2 Effects of Electromagnetic Radiation 

Another suggestion has been made that an aircraft's radar may trigger or 

divert lightning strikes. This possibility was also investigated by Schaeffer and 

Weinstock (Reference 2.18), who show that the transmitted power level of 

microwave radiation necessary to produce an electric field capable of ionizing air 

is about 6.7 x lo6 watts (W), which is far greater than that available from 

aircraft radars. They also point out that a radar would not be designed to ionize 

air because the energy required would reduce transmitted beam energy and 

because ionized air would cause an undesired radar return signal. 

Aircraft lightning-strike incident reports also show no evidence of radar or 

other EMR having been involved in the lightning-strike formation. There are 

cases in which radomes are punctured, but clearly these are simple cases of 

dielectric breakdown of the plastic radome material, with flash termination on 

some airframe-grounded object inside the radome or on the radar antenna itself. 

The punctures have occurred whether or not the radar set was turned on. The 

addition of conductive diverter strips to the outside of the radome usually 

prevents these punctures by enabling lightning flashes to attach directly to a 

diverter. 

2.3 Swept Strokes 

After the aircraft has become part of a completed flash channel, the 

ensuing stroke and continuing currents which flow through the channel may 

persist for up to a second or more. Essentially, the channel remains in its original 
location, but the aircraft will move forward a significant distance during the life 

of the flash. Thus, whereas the initial entry and exit points are determined by 

the mechanisms previously described, there may be other, subsequent, attachment points that are determined by the motion of the aircraft through the 

relatively stationary flash channel. In the case of a fighter aircraft, for example, 

when a forward extremity such as the pitot boom becomes an initial attachment 

point, its surface moves through the lightning channel, and thus the channel 

appears to sweep back over the surface, as illustrated in Figure 2.8. This 

occurrence is known as the swept-stroke phenomenon. As the sweeping action 

occurs, the type of surface can cause the lightning channel to attach and dwell at 

various surface locations for different periods of time. If part of the surface, 

such as the radome, is nonmetallic, the flash may continue to dwell at the last 

metallic attachment point (aft end of the pitot boom) until another exposed 

metallic surface (the fuselage) has reached it; or the channel may puncture the 

nonmetallic surface and reattach to a metallic object beneath it (the radar dish). 

Whether puncture or surface flashover occurs depends on the amplitude and rate 

of rise of the voltage stress created along the channel, as well as the 

voltage-withstand strength of the nonmetallic surface and any air gap separating 

it from enclosed metallic objects. When the lightning arc has been swept back to 

one of the trailing edges, it may remain attached at that point for the remaining 

duration of the lighting flash. An initial attachment point at a trailing edge, of 

course, would not be subjected to any swept-stroke action. 

Tn FLASH 

TI T2 

Figure 2.8 Typical path of swept-stroke attachment points. 

The aircraft cannot fly out of, or away from, the channel. This is because 

the potential difference between charge centers (cloud and earth or another 

cloud) is sufficient to maintain a very long channel until the charges have 

neutralized each other and the flash dies. The aircraft is a very small (and highly 

conductive) part of the channel and cannot move sufficiently far away from the 

vicinity of the channel to become detached from it during the brief lifetime of a 

lightning flash. 
2.4 Aircraft Lightning Attachment Zones 

Since there are some regions on the aircraft where lightning will not 

attach, and others which will be exposed to attachment for only a small portion 

of the total flash duration, it is appropriate to define the zones on the aircraft 

surface which will be exposed to different components of the flash and therefore 

receive different types and degrees of effects. For purposes of aircraft fuel 

system protection, the Federal Aviation Administration (FAA) has defined in its 

advisory circular AC 20-53 (Reference 2.1 9) the following zones: 

Zone I. 

(a) All surfaces of the wing tips located within I8 inches of the tip 

measured parallel to the lateral axis of the aircraft, and surfaces 

within 18 inches of the leading edge on wings having leading edge 

sweep angles of more than 45 degrees 

(b) Projections such as engine nacelles, external fuel tanks, propeller 

disc, and fuselage nose 

(c) Tail group: within 18 inches of the tips of horizontal and vertical 

stabilizer, trailing edge of horizontal stabilizer, tail cone, and any 

other protuberances 

(d) Any other projecting part that might constitute a point of direct 

stroke attachment 

Zone 2. Surfaces for which there is a probability of strokes being swept 

rearward from a Zone I point of direct stroke attachment. This zone 

includes surfaces which extend I8 inches laterally to each side of 

fore-and-aft lines passing through the Zone 1 forward projection points of 

stroke attachment. All fuselage and nacelle surfaces, including 18 inches of 

adjacent surfaces not defined as Zone I, are included in Zone 2. 

Zone 3. Surfaces other than those covered by Zones I and 2. Ignition 

sources in these areas would exist only in the event of strearnering. 

These definitions confine zone boundaries to the 18-inch distances 

described above. Even though they have proved adequate for protection of 

present-day transport category aircraft, it has been recognized that the lightning 

attachment zones of new aircraft designs in fact may not fit these rigid 

definitions. Therefore, an industry committee, designated special Task F of the 

Society of Auton~otive Engineers (SAE) Committee AE-4 on Electromagnetic 

Compatibility, has recently clarified the definition of these zones and broadened 

their application by removal of the arbitrary 1Sinch description. The Task F 

recommended definitions are as follows (Reference 2.20): 

Zone 1. Surfaces of the vehicle for which there is a high probability of 

direct iightning-flash attachment or exit 
Zone 2. Surfaces of the vehicle across which there is a high probability of 

a lightning flash being swept by the airflow from a Zone 1 point of direct 

flash attachment 

Zone 3. Zone 3 includes all of the vehicle areas other than those covered 

by Zone 1 and Zone 2 regions. In Zone 3 there is a low probability of any 

direct attachment of the lightning-flash arc, but Zone 3 areas may carry 

substantial amounts of electrical current by direct conduction between 

some pairs of direct or swept-stroke attachment points in other zones. 

The Task F definitions further divide Zones I and 2 into A and B regions 

depending on the probability of the flash hanging on for any protracted period 

of time. An A-type region is one in which there is low probability that the arc 

will remain attached, and a B-type region is one in which there is a high 

probability that the arc will remain attached. Some examples of zones are as 

follows: 

Zone 1.4. Initial attachment point with low probability of flash hang-on, 

such as a nose 

Zone IB. Initial attachment point with high probability of flash hang-on, 

such as a tail cone 

Zone ZA. A swept-stroke zone with low probability of flash hang-on, such 

as a wing mid-span 

Zone 2B. A swept-stroke zone with high probability of flash hang-on, 

such as a wing trailing edge. 

The Task F definitions are not in conflict with the FAA criteria but, 

rather, are broader in scope and consider the complete aircraft in addition to 

those areas of importance to the fuel system, which is the system primarily 

addressed by FAA AC 20-53. 

In accord with the Task F criteria, typical lightning effects and test criteria 

(described in subsequent chapters of this book) have been categorized according 

to each of these zones. The actual boundaries of each of these zones on a 

particular aircraft can never be exactly determined, but they can be established 

with sufficient accuracy to enable adequate lightning protection to be designed 

for the aircraft. 

Initial lightning attachment points (Zone IA) can usually be established by 

inspection and comparison with other aircraft in-flight experience. Locations of 

the other zones can then be deduced from the locations of Zone lA, in 

accordance with the definitions. Alternately, simulated lightning flashes can be 

fired in a laboratory to a model of the aircraft, and the Zone 1 (A and B) 

attachment points determined photographically; however, the validity of such 
scale model testing is somewhat questionable because of the nonlinear behavior 

of air breakdown phenomena. Model test results have shown good comparison 

with aircraft of conventional wing and empennage configuration but may be less 

capable of accurately predicting the attachment points of more modern designs 

of blended wing-fuselage construction or high wing sweep-back angles. An 

example of lightning-strike attachment zones determined from a model test of 

the Lockheed S-3A aircraft is shown on Figure 2.9. 

LEGEND 

. . . , . . . . . . F:..fc.'.;:j ZONE 1 A ...,.... ::\: 

I I I I I I 2A 

- ZONES lB AND 28 

Figure 2.9 Lightning-strike attachment zones predicted from model test results. 
REFERENCES 

2.1 C. T. Phelps, "Field Enhanced Propagation of Corona Streamers," Journal 

of Geophysical Research 76, 24 (August 20, 1971): 5799-5806. 

2.2 J. F. Shaeffer and G. L. Weinstock, Aircraft Related Lightning Mechanisms, Technical Report AFAL-TR-386, prepared by the McDonnell 

Aircraft Company, McDonnell Douglas Corporation, for the Air Force 

Avionics Laboratory, Air Force Systems Command, Wright-Patterson Air 

Force Base, Ohio (January 1973), p. 2 1. 

2.3 K. A. Moore, "Precipitation Static Noise Problems on Operational 

Aircraft," Lightning and Static Electricity Conference, 3-5 December 

1968, Part I[: Conference Papers, AFAL-TR-68-290, Air Force Avionics 

Laboratory, Air Force Systems Command, Wright-Patterson Air Force 

Base, Ohio (May 1969), pp. 168-78. 

2.4 M. A. Uman, Lightning (New York: McGraw-Hill, 1969), pp. 75-81. 

2.5 R. Plonsey and R. E. Collin, Principles and Applications of Electromagnetic Fields (New York: McGraw-Hill, 1961), p. 50. 

2.6 R. L. Tanner and J. E. Nanevicz, "An Analysis of CoronaCenerated 

Interference in Aircraft," Proceedings of the IEEE 52, 1, Institute of 

Electronic and Electrical Engineers, New York, New York (January 1964): 

44-52. 

2.7 Tanner and Nanevicz, "CoronaCenerated Interference," p. 47. 

2.8 Cumulonimbus with cyclonic circulation at low levels to its southwest, 

photographed from U.S. Weather Bureau DC-6B at 20,000 ft. near 

Topeka, Kansas, 1750 CST, 2 1 April 1961. 

2.9 B. I. Hourihan, Data from the Airlines Lightning Strike Reporting Project, 

June 1971 to November 1974, Summary Report GPR-75-004, High 

Voltage Laboratory, Environmental Electromagnetics Unit, Corporate 

Research and Development, General Electric Company, Pittsfield, 

Massachusetts (March 1975), p. 5. 

2.10 R. H. Golde, I,lghtning Protection (London: Edward Arnold, 1973), p. 25. 

2.1 1 E. T. Pierce, "Triggered Lightning and Some Unsuspected Lightning 

Hazards," American Association for the Advancement of Science, 138th 

Annual Meeting, 1971, Stanford Research Institute, Menlo Park, California 

(January 1972), pp. 14-28. 

2.12 J. E. Nanevicz, E. T. Pierce, and A. L. Whitson, Atmospheric Electricity 

and the Apollo Series, Note 18, Stanford Research Institute, Menlo Park, 

California (June 1972). 

2.1 3 Pierce, "Triggered Lightning." 

2.14 Shaeffer and Weinstock, Aircraft Related Lightning Mechanisms, pp. 

33-45. 

2.15 R. T. Fowler, Ion Collection by Electrostatic Probe in a Jet Exhaust, 

graduate thesis, Air Force Institute of Technology, Wright-Patterson Air 

Force Base, Ohio (June 1970). 

2.16 E. T. Pierce, Atmospheric Electric and MeteorologicaI Environment of 
Aircraft Incidents Involving 1,ighttting Strikes, Special Interim Report I, 

Stanford Research Institute, Menlo Park, California (October 1970). 

2.1 7 Hourihan, Data /rum the Airlines, pp. 40-4 I. 

2.18 Shaeffer and Weinstock, Aircraft Related 1,ighttzing Mechanisms, pp. 

65-66. 

2.19 Protection of Aircraft Fuel Systems Against Lightning, Federal Aviation 

Agency Advisory Circular AC 10-53, Federal Aviation Administration, 

Department of Transportation, Washington, D.C. (Octoher 6, 1967), pp. 

2-3. 

2.20 Lighfnij~g Test Wave forms and TecAniqltes for Aerospace Vehicles atld 

IJard\vare, Society of Automotive Engineers, Warrendale, Pennsylvania (5 

May 1976). 

CHAPTER 3 

LIGHTNING STRIKE EXPERIENCE 

The atmospheric and flight conditions under which aircraft have been 

struck by lightning have been of interest since the beginning of powered flight 

because lightning and other thunderstorm effects, such as turbulence and icing, 

are to be avoided if possible. To learn about these conditions, various 

lightning-strike incident reporting projects have been implemented. Beginning in 

1938, the Subcommittee on Meteorological Problems of the National Advisory 

Committee for Aeronautics (NACA) prepared and distributed a sixteen-page 

questionnaire to airlines and the Armed Forces (Reference 3.1). Pilots filled out 

one of these questionnaires after each lightning-strike incident and forwarded it 

to the NACA subcommittee for analysis. This questionnaire was evidently too 

long for widespread use, however, and, although it could have been shortened, 

was discontinued by 1950. Nevertheless, the program provided important data 

for the first time on the meteorological conditions prevailing when strikes 

occurred and the resulting effects on the aircraft. For one thing, the data 

obtained from the NACA program showed that some lightning-strike conditions 

arc common to many incidents. 

Subsequent programs were conducted by the Lightning and Transients 

Research Institute (LTRI) (Reference 3.2) and the FAA (Reference 3.3). 

Projects currently in operation are being conducted by Plumer and Hourihan of 

General Electric (Reference 3.4), Anderson and Kriininger of South Africa 

(Reference 3.5), Perry of the British Civil Aviation Authority (Reference 3.6), 

and Trunov of the USSR National Research Institute for Civil Aviation 

(Reference 3.7). Strike incidence data, based largely on turbojet or turboprop 

aircraft, is usually summarized according to the following categories: 

Altitude 

Flight path (i.e., climb, level flight, descend, etc.) 

Meteorological conditions 

Outside air temperature 

Lightning-strike effects on the aircraft 

Altitude, flight path, meteorological conditions, and air temperature are 

topics discussed in this chapter. Lightning-strike effects on aircraft will be 

discussed in the succeeding chapter. 

3.2 Altitude and Flight Path 

Figure 3.1 shows the altitudes at which these reporting projects show 

aircraft are being struck, as compared with a typical cumulonimbus (thunder) 

cloud. The turbojet and turboprop data from the four summaries are in close 

agreement. For comparison, the data from the earlier piston aircraft survey of 

Newman (Reference 3.8) are also presented. Cruise altitude for jet aircraft is 

considerably higher (10 km) than that of earlier piston aircraft, which flew at 
about 4 to 5 km; yet Figure 3.1 shows the altitude distribution of 

lightning-strike incidents to be nearly the same. This fact indicates (I) that there 

are more lightning flashes to be intercepted below about 6 km than above this 

altitude, and (2) that very likely jet aircraft are being struck at lower than cruise 

altitudes: that is, during climb, descent, or hold operations. Flight regime data 

obtained from the jet projects shown in Figure 3.2 (Reference 3.9) confirm this. 

THUNDERCLOUDS MAY RISE TO60 000 FEET 

t PERCENT OF STRIKES 

Cf TURBOJET AND TURBOPROPTON ] 

Figure 3.1 Aircraft lightning-strike incidents vs altitude. 

If the strike altitudes shown in Figure 3.1 are compared with the electrical 

charge distribution in the typical thundercloud shown in the figure, it is evident 

that strikes which occur above about 3 km result from intracloud flashes 

between positive and negative charge centers in the cloud (or between adjacent 

clouds), whereas strikes below about 3 km probably result from cloud-to-ground 

flashes. Strike incidents occurring above 6 km are rare because of the absence of 
concentrated charge centers at the higher altitudes and because aircraft at these 

altitudes can more easily divert around thunderclouds than can aircraft at lower 

altitudes. 

FLIGHT CONDITION 

1 STRIKE = 0.51 % 

lwr 

Figure 3.2 Flight conditions when struck. 

m 

LU 80 

Y - 

z 

I- * 60 

8 

8 40- 

20 

3.3 Synoptic Meteorological Conditions 

- 

LEVEL 

- FLT. 

CLIMB 

- 

1- 

DESC. 

OR 

GND APPR. 

0- 

Data discussed thus far might imply that an aircraft must be within or 

beneath a cloud to receive a strike and, since electrical charge separation is 

brought about by precipitation, that most strikes would occur when the aircraft 

is within a cloud in regions of precipitation. Strike incident reports show that 

these conditions often do exist, but other lightning strikes occur to aircraft in a 

cloud when there is no evidence of precipitation nearby, or even to aircraft 

flying in clear air a supposedly safe distance from a thundercloud. FAA and 

airline advisory procedures instruct pilots to circumvent thunderclouds or 

regions of precipitation evident either visibly or on radar, but strikes to aircraft 

flying 25 miles from the nearest radar returns of precipitation have been 

reported. Occasionally a report of a "bolt from the blue," with no clouds 

anywhere around, is received. It is highly improbable, however, that these 

reports are correct because it does not seem possible for electrical charge 

separation of the magnitude necessary to form a lightning flash to occur in clear 

air. In most well-documented incidents, a cloud is present' somewhere (i.e., 

within 25 miles) when the incident occurs. 

Perhaps of most interest to aircraft operators are the area weather 

conditions which prevailed at the time of reported strikes. There is no universal 

59 
data bank for this type of data, but a summary has been made by H. T. Harrison 

(Reference 3.10) of the synoptic meteorological conditions prevailing for 99 

United Air Lines lightning-strike incidents occurring between July 1963 and 

June 1964. Table 3.1 lists the synoptic type and the percentage of incidents (z 

number of cases) occurring in each type. Examples of the most predominant 

synoptic conditions are presented in Figure 3.3 (a), (b), (c), (d), and (e) 

(Reference 3.1 1). 

Table 3.1 SYNOPTIC TYPES INVOLVED WITH 

99 ELECTRICAL DISCHARGES 

JULY 1963 TO JUNE 1964 

Harrison has summarized this data by saying that any conditions which 

will cause precipitation may also be expected to cause electrical discharges 

(lightning), although he adds that no strikes were reported in the middle of 

warm front winter snowstorms. Data from the projects of Plumer and Perry 

(Reference 3.1 2) presented in Figure 3.4 (Reference 3.13) show that lightning 

strikes to aircraft in the United States and Europe occur most often during the 

spring and summer months, when thunderstorms are most prevalent. 

It is also important to note that many strike incidents have been reported 

where no bona fide thunderstorms have been visually observed or reported. L.P. 

Harrison (Reference 3.14), for example, rcports that in 1946 active thunderstorms were manifested in less than half of the 150 strike incidents which he 

studicd. Later, from July 1963 to June 1964, United Air Lines flight crews 

reported 99 cases of static discharges or lightning strikes in flight (Reference 

3.1 5). Correlation of thesc incidents with weathcr conditions prevailing in the 

- 

Synoptic type 

Airmass instability 

Stationary front 

Cold front 

Warm front 

Squall line or instability line 

Orographic 

Cold LOW or filling LOW 

Warm sector apex 

Complex or intense LOW 

Occluded front 

Pacific surge 

Percentage 

2 7 

18 

17 

9 

9 

6 

5 

3 

3 

1 

1 
vicinity and in the general area at the time of strike gave the results shown in 

Table 3.2. 

LBF 

FL 350 (9%) 

0-720 WARMFRONT 

Figure 3.3 Examples of most frequent synoptic meteorological conditions when 

aircraft have been struck. Tip of arrow indicates position of 

aircraft when struck (continued). 

61 
Figure 3.3 (Continued). 

Table 3.2 PERCENTAGES OF STRIKE INCIDENTS VS 

REPORTED THUNDERSTORMS 

Thunderstorms Reported in Vicinity 33% 

Thunderstorms Reported in General Area 14% 

No Thunderstorms Reported 4 274 

3.4 Immediate Environment at Time of Strike 

Figures 3.5, 3.6, and 3.7 show the immediate environment of the aircraft 

at the times of the 214 strikes reported in the project of Plumer and Hourihan 

62 
(Reference 3.16). In over of the strikes reported, each aircraft was within a 

cloud and was experiencing precipitation and some turbulence. With precipitation present, turbulence is to be expected, since vertical air currents acting 

against precipitation are thought to be the cause of electrical charge separation. 

Incident reports also show that most aircraft strikes occur when an aircraft 

is in the freezing level of 0 "C. Figure 3.8 (Reference 3.17) from Newman's 

project (Reference 3.18), for example, shows the distribution of lightning strikes 

to aircraft as a function of outside air temperature. Freezing temperatures (and 

below) are thought to be required for the electrical charge separation process to 

function. Of course, strikes to aircraft at temperatures higher than +I0 "C have 

occurred when the aircraft was close to (or on) the ground, where the ambient 

air temperature may be as high as about +25 "C. 

Figure 3.4 Occurrence of lightning strikes relative to months, in Europe and 

the United States. 

63 
- 

TURBULENCE 

100 

CLOUDS PRECIPITATION 

0 

Figures 3.5, 3.6, 3.7 Environmental conditions at time of strike. 

100 

80 

60 % 

40 

20 

0 

IN A CLOUD 

? 

- 

- 

- 

IN 

- THE 

CLEAR 

0% 

-20 -15 -10 -5 0 5 10 15 20 

TEMPERATURE (OC) 

Figure 3.8 Lightning strikes to aircraft as a function of temperature. 

64 

7 

100 - 

80 

60 % 

- 

- - RAIN 

40 - SLEET 

OR 

HAIL 20 - NONE 

0 
3.5 Thunderstorm Avoidance 

Clearly, whenever it is possible to avoid the severe environments which 

thunderstorms present, it is desirable to do so. For, even if the aircraft is 

adequately protected against lightning effects, the turmoil caused by wind and 

precipitation in or near thunderstorms presents a serious hazard to safe flight. 

Consequently, the operating procedures of commercial airlines and those of 

other air carriers strongly advise against penetration of thunderstorms. 

In attempts to avoid thunderstorm regions, pilots use three indicators: 

Visual sighting of thunderclouds (cumulonimbus) in daytime and of 

lightning at night 

Airborne radar patterns of precipitation areas 

Ground-base radar patterns, if available, relayed by Air Traffic 

Control (ATC) to aircrews as instructions for thunderstorm 

avoidance. 

Methods of thunderstorn avoidance in common use are, in order of 

preference 

Circumvention of thunderclouds, ideally by 25 miles or more 

Flying over the tops of thunderclouds 

Flying beneath the bases (bottoms) of thunderclouds. 

Obviously, when other conditions are equal, the degree to which any of 

these measures is successful depends on the accuracy of the information received 

by the pilots. 

Aside from visual observation (which has obvious limitations), the most 

common method of detecting thunderstorms is using airborne weather radar. 

Radar, however, cannot detect clouds themselves; it can detect only the 

associated precipitation (if any), which is capable of producing an echo; only the 

rain that may be present in the cloud will produce a radar echo. 

A typical C-band airborne weather radar presentation of a thunderstorm 

(cumulonimbus cloud) with active precipitation ar?d frequent lightning is shown 

in Figure 3,9 (Reference 3.19). The pictures were taken during a research project 

carried out by Beckwith of United Air Lines to determine the weather detection 

capability of airborne radar. The photographs shown were taken during a United 

Air Lines flight from Chicago to Denver on August 3, 1960. 

Figure 3.9 (Reference 3.19) shows the northern end of a line of severe 

thunderstorms, developed from a cold front, in Illinois. A detour to the north 

was planned and successfully executed with the aid of this radar presentation. 

The flight remained in clear and generally smooth air while making the detour. 

The strong echos were easily detected with a slight upward tilt of the radar 

antenna to eliminate ground clutter. 

It is possible to obtain more information regarding the intensity of a storm 

by use of the contour circuit provided on most radars. This circuit provides a 

means of eliminating any reflected signal the intensity of which is above a 

certain level. Figure 3.10 (Reference 3.20) shows the same radar return as that in 

Figure 3.9 (Reference 3.19) one minute later with the contour circuit employed. 

65 
The thin, distinct outlines which now appear in place of the original echo 

indicate a narrow boundary across which the intensity of rain varies from no rain 

(outside the white outline) to intense rain (inside the white outline). This change 

in return intensity is called the rain gradient; and the narrower the white outline 

of the return, the more abrupt, or steeper, is the gradient. 

TIME: 00262 TIME: 00282 

ALTITUDE: 28 000 FT. ALTITUDE: 28 000 FT. 

RANGE MARKS: 10-MILE SOUTH OF DUBUQUE, IOWA, LOOKING SE 

ANT.: 1 'hO UP TO S AND W EDGE OF THUNDERSTORM LINE. 

HEADING: 255O TRUE THIS CUMULONIMBUS WAS BUILT TO AN ESTISTRONG ECHOES HERE ARE THE N END OF MATED ALTITUDE OF 45 000 FEET. ECHO ON 

, A LINE OF THUNDERSTORMS WHICH WERE PRO- PHOTO 17-R2 CORRESPONDS TO THIS VISUAL. 

DUCING SEVERE WIND AND LIGHTNING DAMAGE BELOW. LESS THAN ONE HOUR EARLIER. 

HAILSTONES OF GOLF BALL SIZE WERE REPORTED IN THE AREA OF THE NEAREST LARGE 

ECHO. STORM PICTURED ON PHOTO 17-C2. 

Figure 3.9 Radar presentation and subsequent photograph of a thunderstorm. 

The amount of electric charge separation and lightning activity is known 

to be directIy related to the degree of interaction between precipitation and 

vertical air currents (turbulence). Further, the severity of turbulence is also 

related to the amount of temperature difference that exists between different 

masses of air. Thus, turbulence and electrical activity are likely to exist at 

well-defined boundaries, such as those indicated by steep rain gradients on 

contoured radar. These boundaries derived from contour radar are often used in 

planning a detour. 

Considering the variable nature of thunderstorms and the limited 

information as to their whereabouts and severity available to pilots, it is not 

surprising that there are varying opinions as to what detour distance is adequate 

to avoid turbulence and lightning. Primarily, a pilot is advised to use distances 

commensurate with his radar's specific capability. 

The specifications and policies of one of the major airlines (Reference 

3.2 1) follow. 
TIME: 00272 

ALTITUDE: 28 000 FT. 

RANGE MARKS: 10-MILE 

ANT.: 1 " UP 

HEADING: 255" TRUE 

CONTOUR: ON 

Figure 3.10 Same as Figure 3.9 but with contour to show very steep rain 

gradient of each cell. 

Specifications: 

Wavelength: 3.2 cm (X-Band) through 5.6 cm (C-Band) 

Antenna Size: 12-inch or larger (X-Band); 25-inch or larger (C-Band) 

Power (Peak): 10 kW (X-Band) or higher; 75 kW (C-Band) or higher 

Policies: 

1. When the temperature at flight level is 0 "C or higher, avoid all 

echoes exhibiting sharp gradients by 5 nautical miles. 

2. When the temperature at flight level is less than O°C, avoid all 

echoes exhibiting sharp gradients by 10 nautical miles. 

3, When flying above 23 000 feet avoid all echoes, even though no 

sharp gradients are indicated, by 20 nautical miles. 

Weather radar, however, is not a foolproof means of detecting and 

avoiding thunderstorms because situations exist in which radar is not capable of 

distinguishing a thunderstorm return from ground or other precipitation returns 

in the same vicinity. Such a case is illustrated in Figure 3.1 1 (Reference 3.22). 
TIME: 01532 TIME: 01552 

ALTITUDE: 29 MX) FT. ALTITUDE: 29 000 FT. 

RANGE MARKS: 25-MILE DECAYING THUNDERSTORM BUILT TO 

ANT.: O0 ABOVE FLIGHT ALTITUDE FOR WHICH NO 

HEADING: 060° TRUE. OVER ST. GEORGE, MATCHING ECHO WAS VISIBLE IN SCOPE PICUTAH. TURE 11 -R-3. 

A THUNDERSTORM ECHO IS NOT DISTINGUISHABLE FROM TERRAIN AT THIS OR OTHER 

ANTENNA SETTINGS TO MATCH THE VISUAL 

SIGHTING AT 9:30 BEARING IN PHOTO 11 -C2. 

Figure 3. I I Thunderstorms not distinguishable on radar scope. 

In this case, returns froin the ground (ground clutter) obscured the return 

from the storm. However, if ground clutter does not obscure a storm return and 

an aircraft is successful in avoiding all thunderstorms by the recommended 

distances of up to 25 miles, the severe turbulence associated with thunderstorms 

is usually also avoided. Nevertheless, lightning flashes may extend farther 

outward from the storm center than does turbulence and for this reason are not 

as easily avoided. Indeed, there are several reports each year of aircraft receiving 

strikes "in the clear" 25 or more miles from the nearest storm. That lightning 

flashes can propagate this distance is evident from ground photographs of very 

long, horizontal flashes. 

A decaying thunderstorm, moreover, may not present a distinctive radar 

echo. Sometimes this type of storm becomes embedded in expanding anvils or 

cirrus clouds in such a way that it is not visible. In-flight measurements 

conducted by the Air Force and the Federal Aviation Administration, and 

reported by Fitzgerald (Reference 3.23), indicate that thunderstorms in their 

early stages of dissipation have sufficient charge to cause a few lightning 

discharges if a means of streamer initiation becomes available; an aircraft 

entering such a region may initiate, or trigger, such a flash. Thus, in normal IFR 

operations in regions where an active thundercloud is merged with decaying 

thunderclouds and other cloudy areas, those diversions from the normal course 

that are taken to avoid the active cloud may redirect flight through a decaying 

area, where a lightning strike is possible. 

Attempts have been made from time to time to develop an airborne 
instrument capable of warning pilots of an impending lightning strike and 

providing information to the pilot for use in avoidance. Most such instruments 

are based on the principle of detecting the ambient electric field which would 

exist when a lightning flash is imminent (Reference 3.24). None of these 

instruments have been successfully tested in an aircraft, and because of the 

apparent wide variation of electric field magnitudes and directions which may 

exist at the aircraft surface just before a strike occurs, the prospects for success 

seem remote. The situation is further complicated by the problems of field 

interpretation and translation into advisory information to aid the pilot in 

deciding on an avoidance maneuver. 

Perhaps the most effective warning of imminent strikes available to flight 

crews is that which is readily available-the buildup of static discharging and (at 

night) St. Elmo's Fire (corona). Static discharging causes interference (instability) in Low Frequency Automatic Direction Finding (LF-ADF) indicators, or 

audible "hash" in most communications receivers. St. Elmo's Fire is visible at 

night as a bluish glow at aircraft extremities where the discharging is occurring. 

Pilot responses to Lightning strikes (which may also be called static discharges or 

electrical discharges) vary. Typical answers by pilots of one airline (Reference 

3.25) to the question "Do you have any recommendations for avoiding electrical 

discharges?" were as follows: 

"From cruise speed, a reduction of 25% to 30% in airspeed will 

often allow the static buildup to stabilize at a lower maximum and 

dissipate rather than discharge. These buildups are generally 

accompanied by a buzz-type static on VHF [very high frequency] 

and ADF {automatic direction finding] and a random swinging of 

the ADF needles though I have observed the ADF needles to hold a 

steady error of up to 90O as the static level stabilized at or near its 

peak, generally just prior to the discharge or beginning of dissipation. " 

"Climb and descend through the freezing level as quickly as 

possible. " 

"Avoid all precipitation. I know of no way to predict accurately 

where a discharge wiN occur." 

"Slow down to minimum safe speed, change altitude to avoid 

temperature of 20 OF ro 35 OF." 

'Not without excessive detour, both route and altitude." 
"The static discharges I have encountered have built up at a rate 

which would preclude any avoidance tactics (3 to 15 seconds)." 

"No, I have never known when to expect this until just prior to the 

discharge. " 

"No, not in the modern jets. Once the static begins the discharge 

follows very quickly. " 

'Y11 information received at the . . . training center applicable to 

static discharges and their avoidance has been completely accurate 

and helpful. " 

"No, hang on!" 

"Lead a clean life. '" 

Thus, there is a wide divergence of pilot opinion regarding the best way to 

avoid lightning strikes. However, from this and many other sources, it is possible 

to list the symptoms most often present just prior to experiencing a lightning 

strike, and the actions (if any) which most pilots take to reduce the possibility 

of receiving a strike. 

A lightning strike is imminent when a combination of some of the 

symptoms which follow is present. 

Symptoms: 

1. Flight through or in the vicinity of the following: 

Unstable air 

Stationary front 

Cold front 

Warm front 

Squall line 

2. Within a cloud 

3. Icy types of precipitation 

4. Air temperature near 0 "C 

5. Progressive buildup of radio static 

6. St. Elmo's fire (when dark) 

7. Experiencing turbulence 

8. Flying at altitudes between 1.5 and 4.5 km (5000 and 15 000 ft); 

most prevalent: 3.35 km (1 1 000 ft) 

9. Climbing or descending in the vicinity of airports 

70 
Actions: 

1. Circumvent areas of heavy precipitation. 

2. Reduce speed (or rpm with piston-engine aircraft). 

3. Change altitude to avoid temperature near 0 "C. 

4. Turn up cockpit lights. 

5. Have one pilot keep eyes downward. 

Since air traffic congestion often precludes circumvention of precipitation 

and since diversion often poses hazards, avoidance, while desirable, is neither a 

dependable nor an adequate means of protecting the aircraft against lightning 

strikes. The aircraft, therefore, must be designed to safely withstand lightning 

strike effects. 

3.6 Frequency of Occurrence 

The number of lightning strikes which actually occur as compared with 

flight hours for piston, turboprop, and pure jet aircraft is tabulated in Table 3.3 

based on the data of Newman (Reference 3.26) and Perry (Reference 3.27). 

From this data it follows that an average of one strike can be expected for each 

3000 hours of flight for any type of commercial transport aircraft, with slightly 

greater exposure expected for piston aircraft only. The probable reason for this 

difference is that piston aircraft are limited to lower cruise altitudes, where 

lightning flashes are more prevalent. 

Unlike commercial airlines, military and general aviation aircraft need not 

adhere to strict flight schedules or c~ngested traffic patterns around metropolitan airports. The result is that these aircraft do not experience as many 

strikes as do corrimercial aircraft, as is evident from Table 3.4, which shows U.S. 

Air Force experience for the years 1965 to 1969. 

Comparison of the experience reported in Tables 3.3 and 3.4 indicates that 

the probability of commercial aircraft being struck is anywhere from 10 to 200 

times greater than for U.S. Air Force aircraft for the same number of flight 

hours. The discrepancy is most certainly a result of variations in flight 

operations, rather than a result of physical differences in the types of aircraft 

flown. Whereas the Air Force may curtail flight operations in adverse weather, 

commercial air lines usually continue operations, and their aircrafts' exposure to 

lightning is further increased in bad weather by traffic congestion and holding 

requirements near airports. 

Statistics such as these, which apply to a broad category of aircraft and 

include data from a variety of different operators in varying geographic 

locations, may be misleading, however. For example, whereas Table 3.4 shows 

that there is an average of 99 000 flying hours between lightning strikes to U.S. 

Air Force fighter-type aircraft, the strike experience in Europe is known to be 

about 10 times more frequent than strike experience in the U.S. and in most 

other parts of the world. Weinstock and Shaeffer (Reference 3.28) report 10.5 

strikes per 10 000 hours for certain F-4 models flying in Europe, which rate is 

about 5 times greater than the world-wide exposure rate for these aircraft. A 

similar situation pertains to commercial aircraft operating in Europe, as 
indicated by Perry's summary of United Kingdom and European strike data 

(Reference 3.29), for example. This unusually high lightning-strike exposure 

seems Lo result both from the high level of lightning activity in Europe as 

compared with that in many other regions and fronl the political constraints 

placed on flight paths in this multinational region. 

There are several trends in commercial and general aviation which are 

likely to cause even greater exposure of aircraft everywhere to lightning strikes 

in the future: 

More time in holding patterns as a result of increased traffic: at major 

airports 

Increases in the number of intermediate stops along former nonstop 

routes, resulting in more time in descent-hold-climb patterns at 

lower altitudes 

Increasing use of radar and other navigation aids in general aviation 

aircraft, permitting IFR flight under adverse weather conditions. 

These factors warrant continued diligence in the design and operation of aircraft 

with respect to the possible hazards which lightning may present. 

Table 3.3 INCIDENCE OF REPORTED LIGHTNING STRIKES 

TO COMMERCIAL AIRCRAFT 

b 

Newrnan Perry 

(1950 - 1961) (1959 - 1974) TOTALS 

No. hours 

Strikes Hours Strikes Hours Strikes Hours per strike 

Piston 808 2 000000 - - 808 2000000 2475 

Turboprop 109 415 000 280 876 000 389 1 291 000 3320 

Pure Jct - 4 1 427000 - 480 -- 1314000 521 1741000 3340 

ALL 958 2 842000 760 2 190000 1718 5 032000 2930 

Table 3.4 INCIDENCE OF LIGHTNING STRIKES TO 

U.S. AIR FORCE AIRCRAFT 

68 000 140 000 1 I2 000 124 006 76 000 104 000 

ighter 141 000 105 000 11 2 000 65 000 73 000 99 100 
REFERENCES 

3.1 L. P. Harrison, Lightning Discharge to Aircraft and Related Meteorological 

Conditions, Technical Note 1001, National Advisory Committee for 

Aeronautics, Washington, D.C. (May 1946). 

3.2 M. M. Newman and J. D. Robb, Aircraft Protection from Atmospheric 

Electrical Hazards, ASD Technical Report 61-493, L and T Report 374, 

Lightning and Transients Research Institute, Minneapolis, Minnesota 

(December 196 1). 

3.3 "Lightning Strike Survey Report for the Period of January 1965 through 

December of 1966," Federal Aviation Agency Report of the Conference 

on Fire Safety Measures for Aircraft Fuel Systems, Appendix 11, 

Department of Transportation, Washington, D.C. (December 1967). 

3.4 B. I. Hourihan, Data from the Airlines Lightning Strike Reporting Project, 

June 19 71 to November 1974, Summary Report GPR-75-004, High 

Voltage Laboratory, Electromagnetics Unit, Corporate Research and 

Development, General Electric Company, Pittsfield, Massachusetts (March 

1975). 

3.5 R. B. Anderson and H. Kroninger, "Lightning Phenomena in the 

Aerospace Environment: Part 11, Lightning Strikes to Aircraft," Proceedings of the 1975 Conference on Lightning and Static Electricity at Culham 

Laboratory, England, 14-1 7 April 1975, Session I: Fundamental Aspects 

and Test Criteria, the Royal Aeronautical Society of London (December 

1975). 

3.6 J. A. Plumer and B. L. Perry, "An Analysis of Lightning Strikes in Airline 

Operation in the USA and Europe," Proceedings of the 1975 Confererzce 

on Lightr~ing arzd Static Electricity at Cltlham Laboratory, England, 14-1 7 

April 1975, Session IV: Aircraft Applications, the Royal Aeronautical 

Society of London (December 1975). 

3.7 0. K. Trunov, "Conditions of Lightning Strikes in Air Transports and 

Certain General Lightning Protection Requirements," Proceedings of the 

1975 Conference on Lightning and Static Electricity at Culham Laboratory, England, 14-1 7 April 1975, Session IV: Aircraft Applications, the 

Royal Aeronautical Society of London (December 1975). 

3.8 Newman and Robb, Aircraft Protection, pp. 97, 99. 

3.9 Hourihan, Data from the Airlines, p. 1 1. 

3.10 H. T. Harrison, UAL Turbojet Experience with Electrical Discharges, UAL 

Meteorological Circular No. 57, United Air Lines, Chicago, Illinois 

(January 1, 1965), pp. 2748. 

3.1 1 H. T. Harrison, UAL Turbojet Experience, pp. 37, 39, 47,43, 30. 

3.17, PLumer and Perry, "An Analysis of Lightning Strikes," pp. 2, 10. 

3.13 Plumer and Perry, "An Analysis of Lightning Strikes," p. 10. 

3.14 L. P. Harrison (1946) summarized by H. T. Harrison in UAL Turbojet 

Experience, p. 4. 

3.1 5 H. T. Harrison, UAL Turbojet Experience, p. 10. 

3.1 6 Hourihan, Data from the Airlines, p. 1 1. 
3.1 7 Newman and Robb, Aircraft Protection, p. 99. 

3.18 See Newman and Robb, Aircraft Protection, Appendix I: Lightning Strike 

Statistics Extended to Jet Aircraft. 

3.19 W. B. Beckwith, me Use of Weather Radar in Turbojet Operations, UAL 

Meteorological Circular No. 53, United Air Lines, Denver, Colorado (April 

1, 1961), pp. 34, 35. 

3.20 Beckwith, The Use of Weather Radar, p. 35. 

3.21 H. R. Hoffman and G. W. Peckham, The Use of Airborne Weather Radar, 

United Air Lines, Denver, Colorado (1968), p. 12. 

3.22 Beckwith, The Use of Weather Radar, p. 25. 

3.23 D. R. Fitzgerald, "Probable Aircraft Triggering of Lightning in Certain 

Thunderstorms," Monthly Weather Review, December 1967, pp. 835-42. 

3.24 G. A. M. Odam, An Experimental Automatic Wide Range Instrument to 

Monitor the Electrostatic Field at the Surface of an Aircraft in Flight, 

Technical Report 692 18, Royal Aircraft Establishment, Farnborough 

Hants, England (October 1961). 

3.25 Harrison, UAL Turbojet Experience, pp. 74-77. 

3.26 Newman and Robb, Aircraft Protection, p. 100. 

3.27 Plumer and Perry, "An Analysis of Lightning Strikes," pp. 1-2. 

3.28 J. F. Shaeffer and G. L. Weinstock, Aircraft Related Lightning Mechanisms, Technical Report AFAL-TR-72-386, prepared by the McDonnell 

Aircraft Company, McDonnell Douglas Corporation, for the Air Force 

Avionics Laboratory, Air Force Systems Command, Wright-Patterson Air 

Force Base, Ohio (January 1973). 

3.29 Plumer and Perry, "An Analysis of Lightning Strikes, " p. 2. 
CHAPTER 4 

LIGHTNING EFFECTS ON AIRCRAFT 

4.1 Introduction 

"We had just taken off from Presque Isle, Maine, and had been 

in cruise power for 50 minutes, when a large thunderhead cumulus 

was observed directly on course. Lightning could be seen around the 

edges and inside the thunderhead. All cockpit lights were on and the 

instrument spotlight was full on, with the door open. I had just 

finished setting the power and fuel flows for each engine. As the ship 

approached the thunderhead, there was a noticeable drop in 

horsepower arzd the airplane lost from 180 MPH airspeed to I68 

MPH, and continued to lose airspeed due to power loss as we 

approached the thunderhead. . . . A few seconds before the 

lightning bolt hit the airplane all four engines were silent and the 

propellers were windmilling. Simultaneous with the fish of lighr- 

?zing, the engines surged with the original power. The lightning flash 

blinded the Captain and me so severely that we were unable to see 

for approximately eight minutes. I tried several times during this 

interval to read cockpit instruments and it was impossible. lie First 

Officer was called fiom the rear to watch the cockpit. Of course, 

turbulent air currents inside the cumulus tossed the ship around to 

such an extent that, had the airplane not been on auto-pilot when 

the flash occurred and during the interval of blindness by the 

cockpit occupants, the ship could have easily gone completely out of 

control The Captain and I discussed the reason for all four engines 

cutting simultaneously prior to the lightning &sh and could not 

explain it, except for the possibility of a magnetic potentialaround 

the cumulus affecting the primary or secondary circuit of all eight 

magnetos at the same time. " First Officer N.A. Pierson's experience 

on a flight from Presque Isle, Maine, to the Santa Maria Islands on 

July 9, 1945 (Reference 4. I). 

It wasn't long after the beginning of powered flight that aircraft began 

being struck by lightning--sometimes with catastrophic results. The early 

wooden aircraft with metal control cables and guy wires were not capable of 

conducting lightning-stroke currents of several thousand amperes or more. 

Wooden members and even the control cables exploded or caught fire. Even if 

severe structural damage did not occur, pilots were frequently shocked or 

burned by lightning currents entering their hands or feet via control pedals or 

the stick. Sometimes fuel tanks caught fire or exploded. These effects, coupled 

with the air turbulence and precipi-tation also associated with thunderstorms, 

quickly taught pilots to stay clear of stormy weather. 

With the advent of all-metal aircraft, most of the catastrophic effects were 

eliminated, but thunderstorms continued to be treated with respect. Nonethe- 
less, because a few bad accidents attributed to lightning strikes continued to 

happen, in 1938 the Subcommittee of Aircraft Safety, Weather and Lightning 

Experts was formed by the National Advisory Committee for Aeronautics 

(NACA) to study lightning effects on aircraft and determine what additional 

protective measures were needed. Dr. Karl B. McEachron, Director of Research 

at the General Electric High Voltage Laboratory, was a key member of this 

committee, and during its twleve-year existence he performed the first 

man-made-lightning tests on aircraft parts. During and subsequent to this period, 

other organizations, such as the U.S. National Bureau of Standards, the 

University of Minnesota, and the Lightning and Transients Research Institute, 

also began to conduct research into lightning effects on aircraft. 

For a long time the physical damage effects at the point of flash 

attachment to the aircraft were of primary concern. These included holes burned 

in metallic skins, puncturing or splintering of nonmetallic structures, and 

welding or roughening of movable hinges and bearings. If the attachment point 

was a wing tip light or an antenna, the possibility of conducting some of the 

lightning current directly into the aircraft's electrical circuits was also of 

concern. Today, these and other physical damage effects are called the direct 

effects Since present-day military and commercial aircraft fly IFR (Instrument 

Flight Rules) in many kinds of weather, protective measures against direct 

effects have been designed and incorporated into these aircraft so that hazardous 

consequences of lightning strikes are rare. 

In recent years it has become apparent that lightning strikes to aircraft 

may cause other effects, or indirect effects, to equipment located elsewhere in 

the aircraft. For example, the operation of instruments and navigation 

equipment has been interfered with, and circuit breakers have popped in electric 

power distribution systems when the aircraft has been struck by lightning. The 

cause of these indirect effects are the electromagnetic fields associated with lightning currents flowing through the aircraft. Even though metallic skins provide a 

high degree of electromagnetic shielding, some of these fields may penetrate 

through windows or seams and induce transient voltage surges in the aircraft'b 

electrical wiring; these surges in turn may damage electrical or electronic 

equipment. 

To date, few aircraft accidents can be attributed positively to the indirect 

effects of lightning, but there are two trends in aircraft design which threaten to 

aggravate the problem unless new protective measures are developed and 

utilized. The first of these trends is the increasing use of miniaturized, solid state 

components in aircraft electronics and electric power systems. These devices are 

more efficient, lighter in weight, and far more functionally powerful than their 

vacuum tube or electromechanical predecessors, and they operate at much lower 

voltage and power levels. Thus, they are inherently more ~nsitive to overvoltage 

transients, such as those produced by the indirect effects of ;ightn~ng. 

The second trend is the increasing use of reinforced plastics and other 

nonconducting materials in place of aluminum skins, a practice that reduces the 

electromagnetic shielding previously furnished by the conductive skin. This 

reduced shielding may greatly increase the level of surges induced in wiring not 

76 
protected by other means. Because electronic systems were being increasingly 

depended upon for safety of flight, the National Aeronautics and Space 

Administration, the Federal Aviation Agency, and the Department of Defense 

initiated research programs, beginning in 1967, to learn how to measure or 

predict the levels of lightning-induced voltages and how to protect against them. 

A considerable amount of research has followed these initial programs 

Since the indirect effects originate in the aircraft's electrical wiring, their 

consequences may show up anywhere within the aircraft, such as at equipment 

locations remote from the lightning-flash attachments. The direct effects, on the 

other hand, occur primarily at the points of arc attachment. This comparison is 

illustrated in Figure 4.1. 

INDIRECT 

Figure 4.1 Areas of direct and indirect effects. 

Before discussing the techniques for protecting aircraft against either type 

of lightning effect, it is worthwhile to review some of the common examples of 

each which occur on today's aircraft. The purpose of this review is to remind the 

aircraft designer of areas needing particular attention and to alert pilots to what 

to expect when lightning strikes occur in flight. Detailed discussion of the causes 

of each effect, including the lightning-flash characteristic most responsible and 

77 
its quantitative relationship to damage severity, is deferred until protective 

measures are discussed, beginning with Chapter 5. 

4.2 Direct Effects on Metal Structures 

Metal structures include the outer skins of the aircraft together with 

internal metallic framework, such as spars, ribs, and bulkheads. Because lightning 

currents must flow between lightning entry and exit points on an aircraft and 

because these currents tend to spread out as they flow between attachment 

points, using the entire airframe as a conductor, the aluminum with which most 

of these structures are fabricated provides excellent electrical conductivity. As a 

result, the current density at any single point in the airframe is rarely sufficient 

to cause physical damage between entry and exit points. Only if there is a poor 

electrical bond (contact) between structural elements in the current flow path is 

there Iikely to be physical damage. On the other hand, where the currents 

converge to the immediate vicinity of an entry or exit point, there may be a 

sufficient concentration of magnetic force and resistive heating to cause damage. 

Damage at these points is further compounded by the lightning arc, from which 

intense heat and blast forces emanate. Discussion of individual effects follows. 

4.2.1 Melting and Burnthrough 

If a lightning arc touches a metal surface for a sufficient time, melting of 

the metal will occur at the point of attachment. Common evidences of this are 

the successive pit marks often seen along a fuselage or empennage, as shown in 

Figure 4.2 (Reference 4.2) or the holes burned in the trailing edges of wings or 

empennage tips, as shown in Figure 4.3 (Reference 4.3). Most holes are melted 

in skins of 1 rnrn (0.040") thick, or less, except at trailing edges, where the 

lighting arc may hang on for a longer time and enable holes to be burned 

through much thicker pieces. Since a relatively large amount of time is needed 

for melting to occur, the continuing currents are the Iightning-flash components 

most conducive to melting and burnthrough. Melting or burnthrough of skins is 

usually not a safety-of-flight problem unless this occurs in an integral fueI tank 

skin. 

4.2.2 Magnetic Force 

Metal skins or structures may also be deformed as a result of the intense 

magnetic fields which accompany concentrated lightning currents near attachment points. It is well known that parallel wires with current traveling in the 

same direction are mutually attracted to each other. If the structure near an 

attachment point is viewed electrically as being made up of a large number of 

parallel conductors converging to this attachment point, then as lightning 

current flows from the point, forces occur which tend to draw these conductors 

closer together. If a structure is not sufficiently rigid, pinching or crimping may 

occur, as shown in Figure 4.4 (Reference 4.4). The amount of damage created is 
Figure 4.2 Successive pit marks extending backward from leading edge of 

vertical stabilizer. 

proportional to the square of the lightning-stroke current amplitudes and is 

directly proportional to the length of time during which this stroke current 

flows. Thus the high amplitudes of return stroke and intermediate stroke are the 

lightning-flash components most responsible for magnetic force damage. 

Besides the main airframe, other parts which may be damaged by magnetic 

forces include bonding or diverter straps, pitot probes, or any other object which 

may conduct lightning-stroke currents. Magnetic force damage is usually not, by 

itself, significant enough to require abortion of a flight, and may not even be 

detected until the aircraft is on the ground. However, since overstress or severe 

bending of metals is involved, aircraft parts damaged by this phenomenon are 

not often repairable. 

4.2.3 Pitting at Structural Interfaces 

Wherever poor electrical contact exists between two mating surfaces, such 

as a control surface hinge or bearing across which lightning currents may flow, 

melting and pitting of these surfaces may occur. In a recent incident, for 

example, the jackscrew of an inboard trailing edge flap of a jet transport was so 
Figure 4.3 Hole burned in trailing edge corner of ventral fin. 

damaged by a lightning flash that the flap could not be extended past 1 So. Since 

this jackscrew is located on the inboard side of the flap, the flash must have 

reached it after sweeping along the fuselage from an earlier attachment point 

near the nose, as shown on Figure 4.5. Instead of continuing to sweep aft along 

the fuselage, the flash apparently hung on to the jackscrew long enough to melt 

a spot on it. The event did not cause difficulties in landing the aircraft, and the 

damage. in fact, was not discovered until after the aircraft was on the ground. 

The damage, however, was extensive enough that the jackscrew had to be 

replaced. 

It should be noted that the jackscrew in this instance was not an initial, or 

Zone 1, attachment point (See Figure 2.9). It subsequently became an 

attachment point only by being in the path of a swept stroke and is therefore in 

Zone 3B, a swept-stroke zone with high probability of hang-on. 

A second illustration of pitting is the damage caused to the seals of the 

hydraulic jack operating the tail control surfaces of another jet transport 

aircraft. In this case the jack was shunted by a jumper of adequate cross section 

to carry lightning-stroke currents but of excessive length, which caused most of 

the current to flow through the lower inductance path directly through the jack 

body and across the seals, resulting in leakage of hydraulic fluid. 

Earlier aircraft, especially those with nonmetallic skins, experienced more 

troublesome consequences from lightning currents, nearly all of which had to 

flow through the control cables. In these cases, lightning currents entering a 
Figure 4.4 Example of magnetic pinch effect at lightning attachment points. 

JACKSCREW 

(FINAL AflACHMENT 

NOSE 

(INITIAL ATTACHMENT 

POINT) 

Figure 4.5 Swept-stroke attachment and inboard flap jackscrew attachment. 
control surface have been conducted all the way through the aircraft via the 

control cables, sometimes with very damaging results. There is at least one case 

on record (Reference 4.5) of a wooden glider lost as a result of lightning current 

which entered the aircraft at the left aileron pulley support and then flowed via 

the aileron cables to the right pulley, from which point it exited the aircraft. The 

cables were disintegrated, and the wreckage indicated that extensive damage to 

the wooden airframe had occurred before impact, probably as a result of the 

cables exploding inside. There are no records of loss of a powered aircraft from 

this effect, since these aircraft have nearly always been of metallic construction. 

The high-amplitude stroke currents are primarily responsible for pitting, 

but continuing currents may also contribute to this damage, as illustrated by the 

jackscrew incident. 

4.2.4 Resistive Heating 

The glider accident mentioned above is also an example of another direct 

effect: resistive heating of conductors. When the resistivity of a conductor is too 

high or its cross-sectional area too low for adequate current conductance, 

lightning currents flowing in it may deposit appreciable energy in the conductor 

and cause an appreciable temperature rise. Since the resistivity of most metals 

increases with temperature rise, a given current in a heated conductor will 

deposit more energy than it would in an unheated, less resistant conductor; this 

process in turn increases the conductor temperature still further. 

Resistive energy deposition is proportional to the lightning current action 

integral (li2dt), and for any conductor there is an action integral value at 

which the metal will melt and vaporize, as shown in Figure 4.6. The result is the 

1 o2 1 o3 1 0' loa 

(AMPERE' - SECONDS) 

Figure 4.6 1 i2dt vs temperature rise in a conductor. 

82 

TEMPERATURE 

RISE 

OC 7 VAPORIZATION 

4- MELTING 

I I I 
exploding conductor apparent in the glider incident described in Section 4.2.3. 

Other consequences of resistive heating and explosive vaporization of 

conductors are shown in Figures 4.7 and 4.8 (References 4.6 and 4.7). The 

damage is usually most severe when the exploding conductor is within an 

enclosure, which contains the explosion until the pressure has built up to a level 

sufficient to rupture the container. 

Figure 4.7 Lightning damage to radome-probably as a result of exploding 

pitot tube ground wire. 

4.2.5 Shock Wave and Overpressure 

When a lightning-stroke current flows in an ionized leader channel (as 

when the first return stroke occurs), a large amount of energy is delivered to the 

channel in 5 to 10 /IS, causing the channel to expand with supersonic speed. Its 

temperature has been measured by spectroscope techniques to be 30 000 OK and 

the channel pressure (before expansion) about 10 atmospheres (Reference 4.8). 

When the supersonic expansion is complete, the channel diameter is several 

centimeters and the channel pressure is in equilibrium with the surrounding air. 

Later, the channel continues to expand more slowly to the equilibrium situation 

of a stable arc. The cylindrical shock wave propagates radially outward from the 

center of the arc, and, if a hard surface is intercepted, the kinetic energy in the 

shock wave is transformed into a pressure rise over and above that in the shock 

wave itself. This results in a total overpressure of several times that in the free 
Figure 4.8 Resistive heating and explosive vaporization of conductors. 

shock wave at the surface. Depending on the distance of the arc from the 

surface, overpressures can range up to several hundred atmospheres at the 

surface, resulting in implosion-type damage, such as that shown in Figure 4.9 

(Reference 4.9). The arc does not have to contact the damaged surface but may 

simply be swept alongside it, as was evident in the case shown in Figure 4.9. 

Apparently a return stroke or restrike occurred as the tip of the propeller passed 

just below the leading edge of the wing, positioning a cylindrical shock wave 

horizontally beneath the wing, as in Figure 4.10. This hypothesis by Hacker 

(Reference 4.1 0) is supported by scorching of the paint on the imploded panels, 

an indication of a nearby heat source. 

If an arc is contained inside a structure, such as would occur when a 

nonmetallic assembly is punctured, its overpressure may cause additional damage 

to the structure. This may have been responsible for some of the damage to the 

radome shown in Figure 4.7. 

Other examples of shock wave implosion damage include cracked or 

shattered windshields and navigation light globes. Modern windshields, especially 

those aboard transport aircraft, are of laminated construction and evidently of 

sufficient strength to have avoided being completely broken by arc blast and 

overpressures. Broken windshields resulting from a lightning strike, however, are 

considered a possible cause of the crash of at least one propeller-driven aircraft. 

(Reference 4.1 I). 

84 
Figure 4.9 Implosion damage from lightning-flash overpressure. Flash swept aft 

beneath wing from propeller. 

4.3 Nonmetallic Structures 

Early aircraft of wood and fabric construction would probably have 

suffered more catastrophic damage from lightning strikes had it not been for the 

fact that these aircraft were rarely flown in weather conducive to lightning. The 

all-aluminum aircraft which followed were able to fly in or near adverse weather 

and receive strikes, but because aluminum is an excellent electrical conductor, 

severe or catastrophic damage from lightning was rare. There is a trend again, 

however, toward use of nonmetallic materials in aircraft construction. These 

include fiber-reinforced plastics and polycarbonate resins, which offer improvements in cost and performance. However, some of these materials have begun to 

appear at aircraft extremities, such as nose, wing and empennage tips, and access 

door covers, where structural loads are moderate but where lightning strikes 

frequently attach. Often the nonmetallic material is used to cover a metallic 

structure. If this material is nonconducting, such as is the case with fiberglass, 

electric fields may penetrate it and initiate streamers from metallic objects 

inside. These streamers may puncture the nonmetallic material as they propagate 

outward to meet an oncoming lightning leader. This puncture begins as a 

pinhole, but, as soon as stroke currents and accompanying blast and shock waves 

follow, much more damage occurs. An example of a puncture of a fiberglasshoneycomb radome is shown on Figure 4.1 1 (Reference 4.12). In this case a 

85 
SHOCK WAVES 

mr INCIDENT 

nnnnnnn REFLECTED 

TO GROUND 

ARC PATH CYLINDRICAL 

(a) SHOCK WAVE 

WING 

Figure 4.1 0 Orientation of lightning path and shock wave with respect to lower 

side of aircraft wing shown in Figure 4.9. 

(a) Chordwise plane 

(b) Fore-aft plane: perpendicular to lightning path at point P of part (a). 

streamer evidently propagated from the radar dish or some other conductive 

object inside the radome, puncturing the fiberglasshoneycomb wall and rubber 

erosion protection boot on its way to meet an oncoming lightning leader. Most 

of the visible damage was done by the ensuing stroke current. 

Other materials, such as boron- and graphite-reinforced composites, do 

have some electrical conductivity, and, because of this, their behavior with 

respect to lightning is considerably different from that of nonconductive 
Figure 4.1 I Puncture of a fiberglass-honeycomb radome. 

materials. At the present time, boron and graphite composites are only beginning 

to see in-flight service in Zone 1 and Zone 2 regions (Figure 4.9, where lightning 

attachments may occur; because of widespread concern regarding possible 

lightning vulnerability, these zones have been conservatively protected with 

conductive strips or coatings. No reports exist as yet of composite parts damaged 

by natural lightning. Simulated lightning tests which have been performed on 

composites in the laboratory, however, have shown (Reference 4.13) that 

unprotected composites are likely to be vulnerable. The reason is that there is 

sufficient conductivity in the reinforcing fibers or filaments to prevent electric 

field penetration and puncture of the composite but not enough conductivity to 

safely carry away the lightning flash currents, which in this case tend to flow 

into the fibers or filaments themselves. When carrying even minute portions of 

the total flash current, these poor conductors overheat, and damage to 

themselves or to the surrounding resin matrix results. An example of what 

happens to the filaments in a boron-reinforced composite is shown in Figure 

4.1 2. The lightning current entered the composite at the damaged areas shown 

and flowed to a conducting plate at one end of the panel. There is little or no 

damage to the resin matrix, but the damage to the reinforcing filaments results 

in loss of overall material strength (Reference 4.14). 

Because of their greater electrical conductivity and other differences, 

87 
TYPICAL INDIVIDUAL 

FILAMENT TRANSVERSE 

AXIAL FAILURE PLANES 

\-FRONT OF PANEL 

TYPICAL PANEL CROSS SECTION 

SHOWING FILAMENT DAMAGE 

Figure 4.12 Fifty-six kiloampere simulated lightning current damage to boron 

composite laminate panel. Seventeen-ply, 0~-90~-0~ orientation. 

graphite fibers can withstand more lightning current than can boron fibers. At 

higher current levels the graphite fibers remain intact but become hot enough to 

boil and ignite the plastic resin. Figure 4.13 (Reference 4.1 5) shows a small 

"popsicle stick" graphite composite laminate through which 10 kA of lightning 

current have been passed. Nearly all of the resin has burned away, and the 

graphite fibers are left in disarray. As a whole, however, graphite-reinforced 

composites are able to withstand higher amounts of lightning current without 

damage than can boron-reinforced composites. Often some damage to either 

material can be accepted or repaired, but sometimes a lightning-damaged 

composite part will need replacement. 

Transparent acrylics or polycarbonate resins are often utili~ed for canopies 

and windshields. These materials are usually found in Zone 1 or Zcne 2 

locations, where either direct or swept-lightning flashes may occur. Most of the 

polycarbonates are very good insulators, however, and so will successfully resist 

punctures by lightning or streamers. The electric field will penetrate them and 

induce streamers from conducting objects inside, but these streamers are not 

usually able to puncture a polycarbonate. Thus, fighter pilots beneath polycarbonate canopies have often reported electric shocks indicative of strealnering 

off their helmets, but the current levels involved have not been harmful because 

the streamers have not come in direct contact with the lightning flash. Leaders 

approaching the outside of a canopy travel along its surface to reach a metallic 
P 

AFTER 

BEFORE 

Figure 4.13 Unidirectional graphite composite sample before and after conduct- . ing 10 kA of lightning current end-to-end. 

skin, or those initially attached to a forward metal frame may be swept aft over 

a canopy until they reattach to an aft metallic point. Sometimes this occurrence 

will leave a scorched path across the canopy, as shown in Figure 4.14 (Reference 

4.16). Scorches like this can usually be polished away. 

While harmless to a canopy itself, flashes passing just outside frequently 

cause electrical shock or flash blindness to the pilot. In at least one case shock or 

blindness to the pilot caused him to lose control of the aircraft at low altitude 

and resulted in a fatal accident. 

In addition to the direct effects described in the preceding paragraphs, 

replacement of metallic skins with nonmetallic materials removes the inherent 

protection against electromagnetic field penetration that is an important 

by-product of aluminum skins. Electrical wiring and electronics components 

enclosed inside nonmetallic skins are therefore likely to be much more 

susceptible to the indirect effects of lightning than those inside metallic skins 

unless specific measures are taken to reduce this susceptibility. 

4.4 Fuel Systems 

Potentially, aircraft fuel systems represent the most critical lightning 

89 
Figure 4.14 Evidence of lightning attachment to canopy fastener and scorching 

of canopy. 

hazard to flight safety. An electric spark produced by only 0.2 millijoule (mJ) of 

energy is sufficient to ignite a propagating flame in a near stoichiometric mixture 

of hydrocarbon fuel and air (Reference 4.17); yet lightning-flash currents may 

deposit several thousand joules of energy in an aircraft. 

There are several jet and turbojet transport accidents on record which have 

been attributed to lightning ignition of fuel. Although the exact location of 

ignition in each case remains obscure, the most prevalent opinion is that 

lightning ignited fuel vapor at the wing tip vent outlets of these aircraft 

(References 4.1 8 and 4.19). It is also possible that sparking occurred somewhere 

inside a fuel tank as lightning currents flowed through the aircraft. The inflight 

loss of at least two military aircraft also has been attributed to lightning ignition 

of fuel, and there is a report of a lightning strike igniting fuel in another military 

aircraft parked on the ground (Reference 4.20). 

In addition to the direct effects described above, there are several instances 

in which indirect effects have evidently accounted for ignition of fuel. Lightninginduced voltages in aircraft electrical wiring are believed to have resulted in 

sparks, for example, across a capacitance-type fuel probe or some other electrical 

object inside fuel tanks of several military aircraft, resulting in loss of external 

tanks in some cases and the entire aircraft in others. Capacitance-type fuel 
probes are designed to preclude such occurrences, and laboratory tests 

(Reference 4.21) have shown that the voltage required to spark a typical 

capacitance-type probe is many times greater than that induced in fuel gauge 

circuits by lightning. However, other situations involving unenclosed circuits, 

such as externally mounted fuel tanks, exist wherein induced voltages may be 

much higher than those found in circuits completely enclosed by an airframe. 

The accidents mentioned above prompted extensive research into the 

lightning effects on and protection of aircraft fuel systems. Tmproved bonding, 

lightning-protected filler caps and access doors, active and passive vent flame 

suppression devices, flame-retardant foams, and safer (i.e., less flammable) fuels 

are examples of developments which have resulted from this research. In 

addition, government airworthiness requirements now include lightning protection for aircraft fuel systems and specify requirements and tests that must be 

passed to demonstrate compliance prior to aircraft certification. As a result of 

these safety measures, lightning strikes present fewer hazards to the fuel systems 

aboard modern transport aircraft than to those of older aircraft, and properly 

certificated aircraft may expect to experience lightning strikes with no adverse 

effects on fuel systems. Continued changes in airframe designs and materials, 

however, make it mandatory that care and diligence in fuel system lightning 

protection not be relaxed in the future. 

4.5 Electrical Sys terns 

If an externally mounted electrical apparatus, such as a navigation lamp or 

antenna, happens to be at a lightning attachment point, protective globes or 

fairings may shatter and permit some of the lightning current to enter associated 

electrical wiring directly. 

In the case of a wing tip navigation light, for example, lightning may 

shatter the protective globe and light bulb. This may in turn allow the lightning 

arc to contact the bulb filament so that lightning currents may flow into the 

electrical wires running from the bulb to the power supply bus. Even if only a 

fraction of the total lightning current enters the wires, they may be too small to 

conduct the thousands of amperes involved and thus be melted or vaporized. 

The accompanying voltage surge may cause breakdown of insulation or damage 

to other electrical equipment powered from the same bus. At best, the initial 

component affected is disabled, and, at worst, enough other electrical apparatus 

is disabled along with it to require evacuation of the crew and loss of the 

aircraft. There are many examples of this effect, involving both military and civil 

aircraft. Externally mounted hardware most frequently involved includes 

navigation lights, antennas, windshield heaters, pitot probe heaters, and, in 

earlier days, the trailing long-wire antennas that were deployed in flight for 

high-frequency radio communications. The latter were quite susceptible to 

lightning strikes, and, since these wires were too thin to conduct the ensuing 

currents, they were frequently burned away. The high-frequency radio sets 

feeding these antennas were also frequently damaged, and cockpit fires were not 

uncommon. 
Damage may be increased when an electrical assembly is mounted on 

nonmetallic portions of the airframe because some lightning current may have to 

use the assembly ground wire as a path to the main airframe. That the resulting 

damage can be extensive was exemplified by a recent strike to a small 

single-engine aircraft with fiberglass wing tips which included fuel tanks, the 

type pictured in Figure 4.1 5. The details of this incident will illustrate several of 

the strike effects described in this section. 

NAVIGATION LIGHT 

Figure 4.15 General avaiation aircraft with plastic wing tips. 

This aircraft, flying at about 900 m (3 000 ft), was experiencing light rain 

and moderate turbulence when it was struck by lightning. The pilots had seen 

other lightning flashes in the vicinity before their aircraft was struck, and 

embedded thunderstorms had been forecast enroute, but there had been no cells 

visible on the air traffic control (ATC) radar being used to vector the aircraft, 

which, of course, had no weather radar of its own. 

The strike entered one wing tip and exited from the other. It sounded to 

the pilot reporting like a rifle going off in the cabin, and the cabin immediately 

filled with smoke. Other effects follow. 

The No. 1 VHF communication set burned out. 

Seventy-five percent of the circuit breakers were popped, of wh~ch 

only SO? could be rebet later. 

The left wing tip fuel tank quantity indicator was disabled. 

The right main fuel tank quantity indicator was badly damaged. 

Several instrument lights were burned out. 

The navigation light switch and all lights were burned out. 
The aircraft, nevertheless, was able to land at a nearby airport. Subsequent 

inspection showed extensive damage to the right and left wing tips and to their 

electrical wiring. The attachment points and direct effects are pictured in Figure 

4.16 (A-F) (Reference 4.22) and are represented by a diagram in Figure 4.17. 

The evidence suggests that the flash included two or more strokes separated by a 

few milliseconds of continuing current. Assuming, for purposes of explanation, 

that the original lightning flash approached the right wing tip, the probable 

sequence of events was as follows: the initial point of attachment was the right 

wing tip navigation light housing (Figure 4.16 [A]). Current from this stroke 

entered the housing ground wire and exploded both sections of it on the way to 

the right outboard metallic rib, as evidenced by the absence of these wires and 

the blackened interior shown in Figure 4.16(B). Current continued through the 

airframe to the left outboard rib and out the sender unit ground wire to the 

sender unit, the base of which is shown in Figure 4.16(C). From there, the 

current followed the filler cap ground braid and exited the aircraft at the filler 

cap (Figure 4.16[D]). The current exploded the sender unit ground wire but not 

the heavier filler cap ground braid, which was only frayed. Sparksundoubtedly 

occurred inside the fuel tank along the ground braid and between the filer cap 

and its receptacle, but the fuel-air mixture in the ullage of these half-full tanks 

was probably too rich to support ignition. 

Blast forces from stroke No. 1 at the right navigation light housing also 

shattered the lamp globe and bulb, as shown in Figure 4.16(A). This shattering 

allowed a portion of the first stroke current to enter the right navigation light 

power wire, exploding it between the lamp and the outer rib, where the current 

jumped to the rib and continued through the rest of the airframe to the left 

sender unit ground wire. 

Lightning current flowing in the navigation lamp power wire elevated its 

voltage to several thousand volts with respect to the airframe, a voltage high 

enough to break down the insulation at the outer rib feed-through point, as 

shown in Figure 4.17. Until breakdown occurred here (a few microseconds after 

the first stroke began), the wire was at sufficiently high voltage to break down 

the insulation to the neighboring sender wire. This breakdown occurred all along 

the wire inside the right wing. The portion of current arcing into the sender wire 

caused a large voltage to build up across the right wing tip tank fuel gauge 

magnet inductance, to which this wire connects. This voltage in turn sparked 

over the gap between the gauge terminal and the nearest grounded housing wall, 

the arcing badly damaging the gauge unit. While the left navigation light power 

wire was also exploded, it is probable that this did not occur until the second 

stroke. 

Since the aircraft was moving forward, the entry and exit points of the 

second stroke were farther aft on both wing tips than the points of the first 

stroke. Since no other metallic components were present aft of the first stroke 

entry point on the right wing tip, the second stroke punctured a hole in the 

fiberglass trailing edge and contacted the metallic outboard rib, as shown in 

Figure 4.16(E). As shown in Figure 4.17, current from this stroke proceeded 
Figure 4.1 6 Attachment points and direct effects on plastic wing tips. 

through the airframe to the left wing tip, where by this time the stroke had 

swept aft adjacent to the navigation lamp (shown in Figure 4.16 [F]), from 

which point the stroke current exited. Current from stroke No. 2 thus probably 

arced between the left outer rib to the navigation lamp power wire (the ground 

wire having been vaporized by the first stroke), which it followed to the lamp 

housing. The power wire was vaporized by the second stroke current 

flowing in it. 

Both left and right navigation lamp power wires were connected together 

in the cabin and to both the 12 V dc bus and the tail light. The voltage and 

current surges which entered the lamp power wires inboard of the outer rib 

feedthroughs were also conducted to the tail light, burning it out, and to the 12 

V do bus. The surge on the bus, of course, was immediately imposed on all of 

the other electrical equipment powered from this bus, or all of the electrical 

equipment in this aircraft. Arcing undoubtedly occurred in a number of 
PLASTIC PLASTIC 

TIP TIP 

- - - - - - - I---------- 

Figure 4.1 7 Plastic wing tips and associated electric circuits and locations of 

lightning effects. 

components, causing circuit breakers to pop. Because circuit breakers, however, 

react much too slowly to prevent passage of a lightning surge, at least one piece 

of equipment (the No. 1 VHF communication set) and several instrument lamps 

were burned out. 

There have been several similar incidents (References 4.23 and 4.24), and 

together these have stimulated the design and verification of protective measures 

(Reference 4.25) for general aviation aircraft with fiberglass components such as 

these wing tip fuel tanks, Many of these aircraft, however, are still flying without 

adequate protection. 

The foregoing incident is an example of how a change in materials can 

increase the vulnerability not only of the airframe but also of other systems 

which previously had the inherent protection of conventional aluminum skins. 

The lightweight lamp and sender unit electrical wires were quite adequate for an 

installation in which a metal skin was available to carry away lightning currents, 

but they are woefully inadequate when used unmodified inside a plastic wing 

tip, where they become the only conducting path available to lightning currents 

trying to enter the main airframe. 
4.6 Engines 

With the exception of a few incidents of temporary malfunction similar to 

the incident reported in the introduction to this chapter, there have been no 

reports of adverse lightning effects on reciprocating engines. Metal propellers and 

spinners have been struck frequently, of course, but effects have been limited to 

pitting of blades or burning of small holes in spinners, as shown in Figures 4.18 

and 4.19 (References 4.26 and 4.27). Lightning currents must flow through 

Figure 4. I8 Lightning-strike damage to a propeller. 

96 
propeller blade and engine shaft bearings, but these are massive enough to carry 

these currents with no harmful effects. Wooden propellers, especially ones 

without metal leading edges, could probably undergo more damage, but these 

are seldom used on aircraft which fly in weather conditions where lightning 

strikes occur. 

Figure 4.19 Lightning-strike damage to a spinner. 

Reported lightning effects on turbojet engines show that these effects also 

are limited to temporary interference with engine operation. Flameouts, 

compressor stalls, and roll-backs (reduction in turbine rpm) have been reported 

after lightning strikes to aircraft with fuselage-mounted engines. This type 

includes military aircraft with internally mounted engines and fuselage air 

intakes, or other military and civil aircraft with engines externally mounted on 

the fuselage. There have been no attempts to duplicate these events with 

simulated lightning in a laboratory, and there has been no other qualitative 

analysis of the interference mechanism; however, it is generally believed that 
these events result from disruption of the inlet air by the shock wave associated 

with the lightning arc channel sweeping aft along a fuselage. This channel may 

indeed pass close in front of an engine intake, and if a restrike occurs, the 

accompanying shock wave is considered sufficient to disrupt engine operation. 

The steep temperature gradient may also be important. These effects have been 

reported as occurring more often on smaller military or business jet aircraft than 

on larger transport aircraft. Thus, smaller engines are probably more susceptible 

to disrupted inlet air than are their larger counterparts. 

In some cases a complete flameout of the engine results, while in others 

there is only a stall or roll-back. There is no case on record, however, in which a 

successful restart or recovery of the engine to full power was not made while still 

in flight. Perhaps because of this, together with the impracticality of a 

laboratory simulation, there has been little research into the problem. 

Nevertheless, operators of aircraft with engines or inlets close to the fuselage 

should anticipate possible loss of power in the event of a lightning strike and be 

prepared to take quick corrective action. 

There are no reports of lightning effects on wing-mounted turbojet 

engines, since lightning strikes do not often occur near the inlets of these 

engines, and there are no reports of power loss of turboprop engines as a result 

of lightning strikes. 

4.7 Indirect Effects 

Even if the lightning flash does not directly contact the aircraft's electrical 

wiring, strikes to the airframe are capable of causing voltage and current surges 

in this wiring which may be damaging to aircraft electronics. 

The mechanism whereby lightning currents induce voltages in aircraft 

electrical circuits is illustrated in Figure 4.20. As lightning current flows through 

an aircraft, strong magnetic fields which surround the conducting aircraft and 

change rapidly in accordance with the fast-changing lightning-stroke currents are 

produced. Some of this magnetic flux may leak inside the aircraft through 

apertures such as windows, radomes, canopies, seams, and joints. Other fields 

may arise inside the aircraft when lightning current diffuses to the inside surfaces 

of skins. In either case these internal fields pass through aircraft electrical 

circuits and induce voltages in them proportional to the rate of change of the 

magnetic field. These magnetically induced voltages may appear between both 

wires of a two-wire circuit, or between either wire and the airframe. The former 

are often referred to as line-to-line col€ages and the latter as common-mode 

voltages 

In addition to these induced voltages, there may be resistive voltage drops 

along the airframe as lightning current flows through it. If any part of an aircraft 

circuit is connected anywhere to the airframe, these voltage drops may appear 

between circuit wires and the airframe, as shown in Figure 4.20. For metallic 

aircraft made of highly conductive aluminum, these voltages are seldom 

significant except when the lightning current must flow through resistive joints 
EXTERNAL - 

MAGNETIC 

FLUX 

INTERNAL .----- MAGNETIC 

FLUX 

'AIRCRAFT WIRING 

NAVIGATION LIGHTS 

e =-/ 

dt 

WING 

Figure 4.20 Magnetic flux penetration and induced voltages in electrical wiring. 

or hinges. However, the resistance of titanium is 10 times that of aluminum, and 

that of composite materials many hundred times that of aluminum, so the 

resistive voltages in future aircraft employing these materials may be much 

higher. 

Upset or damage of electrical equipment by these induced voltages is 

defined as an indirect effect. It is apparent that indirect effects must be 

considered along with direct effects in assessing the vulnerability of aircraft 

electrical and electronics systems. In situations like that of the light aircraft 

described in Section 4.5, the direct effects are clearly the most severe. Other 

aircraft exist, however, whose electrical systems are well protected against direct 

effects but not so well against indirect effects. 

Until the advent of solid state electronics in aircraft, indirect effects from 

external environments, such as lightning and precipitation static, were not much 

of a problem and received relatively little attention. No airworthiness criteria are 

available for this environment. There is increasing evidence, however, of 

troublesome indirect effects. Incidents of upset or damage to avionic or 

electrical systems, for example, without evidence of any direct attachment of 

the lightning flash to an electrical component are showing up in airline 

lightning-strike reports. Table 4.1 summarizes the reports of interference or 

outage of avionic or electrical equipment reported by a group of U.S. airlines for 

the period June 1971 to November 1974 (Reference 4.28). 
Table 4.1 EVIDENCE OF INDIRECT EFFECTS IN 

COMMERCIAL AIRCRAFT 

(2 14 strikes) 

Interference Outage 

HF communication set - 5 

VHF communication set 2 7 3 

VOR receiver 5 2 

Compass (all types) 22 9 

Marker beacon - 2 

Weather radar 3 2 

Instrument landing system 6 - 

Automatic direction finder 6 7 

Radar altimeter 6 - 

Fuel flow gauge 2 - 

Fuel quantity gauge - 1 

Engine rpm gauges - 4 

Engine exhaust gas temperature - 2 

Static air temperature gauge 1 - 

Windshield heater - 2 

Flight director computer 1 - 

Navigation light - 1 

ac generator tripoff (6 instances 

of tripoff) 

Autopilot 1 - 

The incidents reported in Table 4.1 occurred in 20% of the total of 214 

lightning-strike incidents reported during the period. U.S. military aircraft have 

had similar experience. This experience is probably a result of the increasing 

sensitivity of miniaturized solid state electronics to transient voltages, a trend 

which necessarily would not have posed a problem in older, less sophisticated 

equipment. In any one incident, only a few electronic components are affected; 

others are not. Yet laboratory tests (Reference 4.29) have shown that 

lightning-induced voltages appear in all aircraft electric wiring at once. Thus it is 

evident that surges reach higher values in some circuits than in others or that 

some electronics are Iess tolerant of such surges than others. 

Indirect effects were evident in about 2W/r of all the incidents reported, 

while outages were reported in only about 10% of all incidents. Since severe 

strokes also occur in only about 10 to 20% of all flashes, it is probable that a 

severe stroke may be required to cause noticeable effects. 

While the indirect effects are not presently a major safety hazard, there are 

four trends in aircraft design and operations which could increase the potential 

problem. These include the following: 

Increasing use of plastic or composite skin 

Further miniaturization of solid state electronics 
Greater dependence on electronics to perform flight-critical 

functions 

Greater congestion in terminal airways, requiring more frequent 

flight through adverse weather conditions at altitudes where lightning strikes frequently occur. 

Design of protective measures against direct effects is the subject of the 

next three chapters. Design of protective measures against indirect effects is 

treated in the remainder of this book. 
REFERENCES 

4.1 Answer to questionnaire Effects of Electrical Pherzomena upon Airplanes 

in Right, National Advisory Committee for Aeronautics, Washington, D.C. 

(July 7, 1945). 

4.2 U.S. Air Force photograph. 

4.3 National Aeronautics and Space Administration photograph. 

4.4 National Aeronautics and Space Administration photograph. 

4.5 B. L. Peny, "British Researches and Protective Recommendations of the 

British Air Registration Board," Lightrzirzg and Static Electricity Conference, 3-5 December 1968, Part II, Conference Papers, Technical Report 

AFAL-TR-68-290, Air Force Avionics Laboraiory, Air Force Systems 

Command, Wright-Patterson Air Force Base, Ohio (May 1969), pp. 

81-103: 96. 

4.6 U.S. Air Force photograph. 

4.7 U.S. Air Force photograph. 

4.8 M. A. Uman, Lightning (New York: McCraw-Hill, 1969), p. 230. 

4.9 Paul T. Hacker, Lightning Damage to a General Aviation Aircraft- - 

Description and Analysis, NASA TN-7775, National Aeronautics and 

Space Administration, Lewis Research Center, Cleveland, Ohio (September 

1974). 

4.10 Hacker, Lightning Damage. 

4.1 1 Report of the Investigation of an Accident Int.*ohing Aircraft of U.S. 

Registry, NC 21 789, Which Occurred near Love ttsville, Virginia, on 

August 31, 1940, Civil Aeronautics Board, Washington, D.C. (1941). 

4.12 United Air Lines photograph. 

4.13 L. C. Walko, Current Status of Composites Vulnerability to Lightning, 

SRD74-090, Corporate Research and Development, General Electric 

Company, Schenectady, New York (August I, 1974). 

4.14 F. A. Fisher and W. M. Fassell, Lightning Effects Relating to Aircraft, Part 

I: Lightning Effects on and Electromagnetic Shielding Properties of Boron 

and Graphite Reinforced Composite Materials, Technical Report AFALTR-72-5, Air Force Avionics Laboratory, Air Force Systems Command, 

Wright-Patterson Air Force Base, Ohio (January 1972), pp. 48-80. 

4.15 From studies of A. P. Penton, J. L. Perry, and K. J. Lloyd. See 

"Fundamental Investigations of High Intensity Electric Current Flow, 

Processes and Resultant Damage in Advanced Composites," 1970 Lightning and Static Electricity Conference, 9-1 1 December, sponsored jointly 

by the Air Force Avionics Laboratory and the Society of Automotive 

Engineers, Air Force Systems Command, Wright-Patterson Air Force Base, 

Ohio (December 1970), pp. 253-297. 

4.16 General Electric Company photograph. 

4.1 7 Fire Protection Research Program for Supersonic Transport, APL 

TDR-64-105, Air Force Aero Propulsion Laboratory, Wright-Patterson Air 

Force Base, Ohio (October 1964), pp. 1 1 1-1 13. 
4.1 8 Aircraft Accident Report: Boeing 707-121, N709PA, Pan Ameriran World 

Airways, Inc., near Elkton, Maryland, December 8, 1963, File No. 1-001 5, 

Civil Aeronautics Board, Washington, D.C. (February 25, 1965). 

4.1 9 Aircraft Accident Report: TWA Lockheed 1649A near Milan, Italy, File 

No. 1-0045, Civil Aeronautics Board, Washington, D.C. (November 1960). 

English translation of report by Italian Board of Inquiry. 

4.20 C-14 1 on-ground incident. 

4.2 1 J. A. Plumer, "Lightning-Induced Voltages in Electrical Circuits Associated 

with Aircraft Fuel Systems," Report of Second Conference on Fuel 

System Fire Safety, 6 and 7 May 1970, Federal Aviation Administration, 

Washington, D.C. (1970), pp. 171-92. 

4.22 General Electric Company photographs. 

4.2 3 Don Flagg, "Night Flight," Aero Magazine, January /February 1972, pp. 

18-21. 

4.24 Page Shamberger, "Learning About Flying the Hard Way," Air Progress, 

February 1971, p. 64. 

4.25 J. A. Plumer, Guidelines for Lightning Protection of General Aviation 

Aircraft, FAA-RD-73-98, Federal Aviation Administration, Washington, 

D.C. (October 1973). 

4.26 National Aeronautics and Space Administration photograph. 

4.27 National Aeronautics and Space Administration photograph. 

4.28 J. A. Plumer and B. L. Perry, "An Analysis of Lightning Strikes in Airline 

Operation in the USA and Europe," Proceedings of the I975 Confererlce 

on Lightning and Static Electricity at Culham Laboratory, England, 14-1 7 

Apn'l 1975, Session 111: Aircraft Applications, the Royal Aeronautical 

Society of London (December 1975). 

4.29 J. A. Plumer, F. A. Fisher, and L. C. WaIko, Lightning Effects on the 

NASA F-8 Digital Fly-By-Wire Airplane, NASA CR 2524, prepared by the 

High Voltage Laboratory, Environmental Electromagnetics Unit, Corporate Research and Development, General Electric Company, Pittsfield, 

Massachusetts, for the National Aeronautics and Space Administration, 

Lewis Research Center, Cleveland, Ohio (March 1975), p. 138. 

CHAPTER 5 

DIRECT EFFECTS PROTECTION 

Successful protection of an aircraft against the direct effects of lightning 

depends upon protection of each of its various systems or coniponents that may 

be susceptible either to direct lightning attachment or to current flow between 

lightning attachment (entry and exit) points. Components located in different 

sections of the aircraft are likely to experience different degrees of susceptibility 

to lightning, and they may be vulnerable to different components of the 

lightning flash. In this chapter we present the basic steps which should be 

followed in eslablishing lightning protection by discussing the varieties of 

lightning susceptibility and the designing of suitable protective measures against 

arc entry and current flow-through damage. We explain how the lightning-strike 

zones and lightning current environments are established, since environmental 

conditions in the zones are those under which specific protective measures must 

perform. We then call attention to those airworthiness regulations which apply 

to lightning protection. In subsequent chapters we present specific protection 

techniques for aircraft fuel and structural systems exposed to the direct effects 

of lightning strikes. 

5.2 Basic Steps in Protectio~i Design 

5.2.1 Establishment of the Ligh tnilig Susceptibility 

Design of successfill protection usually involves taking the following basic 

steps. As described in Chapter I, lightning strikes co~nmonly attach to an aircraft 

nose, wing tips, vertical fin tip, horizontal stabilizer tips, and to other 

appendages, such as propellers, pitot booms, antennas, or pylon-mounted 

external stores. These are the places where the hot lightning arc attaches directly 

to the aircraft and thus the places of initial concern as far as direct effects are 

concerned. However, lightning currents must flow through the aircraft skin and 

structural members between these arc entry and exit points, and it is possible for 

these currents to damage components along the way if safe conduction paths are 

not provided. Thus, the lightning protection designer's job is not completed 

when he has adequately treated only the attachment points. Instead, he must 

also look throughout the aircraft at other components that might be vulnerable 

in one way or another to direct effects from current flow through the aircraft. 

The susceptibility (degree of exposure to lightning effects) of components 

depends greatly on their location on or within the aircraft. Thus it is helpful to 

utilize the lightning-strike zone definitions of Chapter 2 to establish the lightning 

environment which each particular component experiences. Once this has been 

established, the vulnerability (if any) of the component to its own environment 
can be determined, and protective measures designed if necessary. 

The lightning attachment zone definitions proposed by SAE Special Task 

F (Reference 5.1) are repeated below: 

Zone IA: An initial attachment point with a low probability of flash 

hang-on, such as a leading edge 

Zone IB: An initial attachment point with a high probability of flash 

hang-on, such as a trailing edge 

Zone 2A: A swept-stroke zone with a low probability of flash hang-on, 

such as a wing mid-span 

Zone2B: A swept-stroke zone with high probability of flash hang-on, 

such as a wing inboard trailing edge 

Locations of these zones can be determined either from comparison with 

actual lightning attachment points on existing aircraft or by laboratory tests in 

which a scale model of the aircraft is subjected to simulated lightning strikes. 

Perry (Reference 5.2) has shown the locations of actual lightning attachment 

points on several types of propeller and turbojet aircraft in use in the United 

Kingdom and Europe. His strike patterns for three aircraft are shown on Figures 

5.1, 5.2, and 5.3 (Reference 5.3). From these figures it is evident that initial 

attachment points (Zone IA and 1B) are at or very near the extremities, such as 

nose, wing tips, or empennage tips. if an initial attachment point is at a trailing 

edge, the arc must hang on there until the flash dies naturally; thus more damage 

may occur. On the other hand, if the initial attachment point is a forward 

extremity, such as the nose, the aircraft can fly through the flash channel, 

allowing the arc to reattach at subsequent points along the line of flight, as 

evident from the lines of successive burn marks dong the fuselage in the three 

figures. Thus, forward initial attachment points are commonly in Zone 1A and 

the subsequent burn marks in Zone 3A. Once in a while, the flash may remain 

attached to a forward-most extremity or to a small protrusion, such as an 

antenna or even a rivet head along the fuselage. In such cases the attachment 

point becomes a B-Zone point with high probabiIity of hang-on. Since most 

lightning flashes persist for one-fourth of a second or more (up to about 1 

second), the aircraft may move forward its entire length, or more, in this period 

of time, and its entire length may be exposed to swept-lightning flashes. Its 

entire length must therefore be considered to be in Zone 2. 

If a new aircraft design is of the same general configuration (that is, 

relative position and size of fuselage, wings, empennage, etc.) the general 

location of its lightning attachment zones can be determined by comparison 

with known attachment points on similar aircraft. Knowledge of the exact 

boundaries between one zone and another is important, however, and these 

boundaries must often be determined by a laboratory test in wh~ch a full-size 

replica of the aircraft assembly in question (such as a wing tip or nose radome) is 

subjected to simulated lightning flashes from various directions. Knowledge of 

these zone boundaries is particularly important in designing protection for wing 

fuel tanks, for example. Recogni7ing that laboratory tests may not be practical, 

the Federal Aviation Administration in its advisory circular on protection of 

aircraft fuel systems against lightning defined Zone 1 as including "[a] ll surfaces 

106 
SIDE ELEVATION fi c7-JJ 

lop& 

a 

SURFACES 

INDICATES PENETRATION 

OF METAL SKIN UNDER : SURFACES 

. INDICATES BURN MARKS 

, ' WITHOUT PENETRATION 

Figure 5.1 Position of lightning strikes on Viscount aircraft - March 1959 to 

June 1964. 

of the wing tips located within 18 inches of the tip measured parallel to the 

Lateral axis of the aircraft, and surfaces within 18 inches of the leading edge on 

wings having leading edge sweep angles of more than 45 degrees." Also included 

are " [p] rojections such as engine nacelles, external fuel tanks, propeller disc, 

and fuselage nose"; in the tail group all surfaces "within 18 inches of the tips of 

horizontal and vertical stabilizer, trailing edge of horizontal stabilizer, tail cone, 

and any other protuberances," as well as "[a] ny other projecting part which 

might constitute a point of direct stroke attachment." 

FAA defines Zone 2 as including "[s]urfaces for which there is a 

probability of strokes being swept rearward from a Zone 1 point of direct stroke 

attachment. This zone includes surfaces which extend 18 inches laterally to each 

side of fore-and-aft lines passing through the Zone I forward projection points 

of stroke attachment. Ail fuselage and nacelle surfaces, including 18 inches of 

adjacent surfaces, not defined as Zone I are included in Zone 2" (Reference 

5.4). (FAA does not subdivide Zone 1 or Zone 2 into A and B zones as 

recommended by SAE Task F.) 

Thus, the establishment of the lightning attachment zones for aircraft of 

conventional configuration is relatively straightforward and can be done with 

reasonable accuracy. Prediction of the lightning attachment zones for new 

aircraft of unconventional design is not as easily accomplished by reference to 

inflight data or to the FAA 18-inch criteria. Such an aircraft might look like the 

one in Figure 5.4 (Reference 5.5). The gradual blending of fuselage and wing 

lines makes identification of the regions of maximum electric field stress (and 

lightning attachment) less easy to distinguish. With the exception of the pointed 
SIDE NVATION 

INDICATES PENETRATION 

OF METAL SKIN 

DICATES BURN MARKS 

ITHOUT PENETRATION 

Figure 5.2 Position of lightning strikes on Boeing 707 aircraft January 1962 

to December 1967. 

nose and vertical wing tips, the regions of highest electric field stress are less 

evident on this aircraft than on the conventional aircraft of Figures 5.1 through 

5.3. Thus, in this case strike attachment tests on a scale model in a laboratory 

are desirable. The SAE Task F report (Reference 5.1) describes how this test 

should be performed. 

5.2.2 Protection of Exposed Systems or Components in Zones 1 

and 2 Against Arc Entry Damage 

The most obvious protection task is to protect systems or components 

located in Zone 1 or Zone 2 from damage that results from direct arc 

attachment. In some cases, complete protection against burning and erosion is 

impractical, since the consequences of permitting some of this damage to occur 

are of a maintenance nature only. In other cases, such as those involving integral 

fuel tank skins, it is very important that complete protection be designed and 

applied. 

108 
INDICATES PENETRATION UNDER SURFACE 

OF METAL SKIN . :. - . ,:;.$:;;:$. INDICATES BURN MARKS 

;...... '2;: WITHOUT PENETRATION 

June 1968. 

Figure 5.3 Position of lightning strikes on Trident aircraft -- May 1964 

Figure 5.4 Possible aircraft configuration of the future. 

5.2.3 Provision of a Safe, Controlled Path for Lightning Currents 

Since the aircraft is part of the lightning current path between two 

external charge centers, lightning currents must flow through the airframe 

between attachment points. In most cases aluminum skins and structural 

elements provide an excellent conductive path without the need of additional 

provisions for lightning currents. Nonmetallic materials, however, such as 

fiber-reinforced plastics are finding increased use in newer aircraft. At present, 
use of these materials has been limited to outer wing tips or radomes, for 

example; therefore, in these cases additional conductive paths need only be 

added across or through the nonmetallic section. However, an aircraft made 

entirely of fiberglass has been fabricated. In this case positive steps must be 

taken to provide the necessary conducting path(s) for lightning currents to flow 

through the entire aircraft. 

5.3 The Lightning Environment as Related to Attachment Zones 

For guidance in the lightning protection design and qualification testing of 

aerospace vehicles and hardware, the Society of Automotive Engineers Committee AE4, Special Task F, has formulated an idealized representation of the 

current components of a severe lightning flash. This representation incorporates 

the important aspects of both positive and negative flashes. Four current 

components, A, B, C, and D, represent the four current characteristics found in 

most lightning flashes. These are shown together on Figure 5.5 (Reference 5.6). 

This model represents a very severe flash wherein intensities are exceeded less 

than 1% of the time. It may be used to define the environment that aircraft 

systems and components must withstand or against which they must be 

protected. At the present time, this model is not included in any government 

specification or standard dealing with lightning protection; however, the model 

is expected to be incorporated into forthcoming revisions of most of these 

requirements. The lightning environment described in Figure 5.5 is equal in 

severity to the lightning currents which are described in present specifications, or . 

greater than such currents. The model is also in general agreement with new 

b 

Component A (Initial Stroke) 

Peak amplitude - XIO kA f 10% 

Action Integral - 2 x 10%' - seconds f 20% 

Time duration - < 500 us. 

- Component B (Intermediate current) PI Component D (restrike) -6 Maximum Charge Transfer - 10 Coulom 

g :: Average amplitude - 2 kA f 10% 

k 3 Component C (Continuing Action integral = 0.25 x 3-, 

- c Current) Charge T ransfer = 10' A' - seconds 

200 Coulombs 

f 20% Amplitude = 

t + Time I ~X~O-~SM 0.25 seclT<l sec 

Figure 5.5 Current test waveform components for evaluation of direct effects. 

lightning protection requirements being formulated in the United Kingdom and 

Europe (Reference 5.7). 

In order to design protection for a particular system or piece of hardware, 

its susceptibility to the various lightning current characteristics must be known. 
For this purpose SAE Task F has also defined the lightning current component(~) likely to strike the aircraft in each of the four lightning attachment 

zones. The results are shown on Table 5.1. Further references will be made to 

these Task F criteria during discussion of protection techniques in succeeding 

chapters of this book. 

Table 5.1 LIGHTNING CURRENT COMPONENTS 

EXPERIENCED IN EACH ZONE 

A B C D 

Current Component First Return Intermediate Continuing 

Zone Stroke Stroke Current Restrike 

1A: initial attachment X X 

point with low 

probability of 

flash hang-on 

1B: initial attachment 

point with high 

probability of 

flash hang-on 

2A: swept-stroke 

zone with low 

probability of 

flash hang-on 

2B: swept-stroke 

zone with high 

probability of 

flash hang-on 

3: low probability X 

of direct attachment 

*Only if flash hang-on time is greater than 5 ms 

5.4 Applicable Airworthiness Regulations 

There are a number of government airworthiness regulations that apply 

specifically to aircraft lightning protection. These are applicable in the United 

States and other countries and generally must be complied with as part of the 

aircraft certification requirements. Table 5.2 lists the regulations presently in 

effect. These are concerned primarily with fuel system protection, since this 

system has presented the most serious hazard in the past. Other areas such as 

nonmetallic structures and aircraft electronics systems are of increasing concern, 

however, and are beginning to appear in updated requirements. Applicable 

portions of the regulations in Table 5.2 will be referred to and discussed in 

succeeding chapters. 
Table 5.2 GOVERNMENT AIRWORTHINESS REQUIREMENTS AND RELATED DOC

LIGHTNING PROTECTION OF AIRCRAFT 

Document 

Identification Title 

Federal Aviation Airworthiness Standards: 

Regulations, Part 25 Transport Category Airplanes 

Applicable 

Paragraphs 

25.581, US Fede

25.954 Dept. of 

Advisory Circular Protecrion of Aircraft Fuel Systems AU US Fede

AC 20-53 Against Lightning Dept. of 

6 October 1967 

- MSC-07636 C 

Space Shuttle I, ightning Protection All US Nati

N June 1973 CFiteriu Document Admini

Lyndon 

MIL-B-5087B (ASG) Military Specification: Bonding, Electrical, 3.3.4, 6.3.1, US Air 

15 October 1964 and Lightning Protection, For Aerospace 6.3.9 

Systems 

MIL-C-38373A(ASG) MiIitarySpecification:Cap, FluidTunk 3.5.1,4.6.14 USAirF

26 February 1969 Filler 

MIL-A-9094D ( ASG) Military Specification : Arrester, I,ightning, All US Air 

17 March 1969 General Specification For 
Table 5.2 GOVERNMENT AIRWORTHINESS REQUIREMENTS AND RELATED DOCU

LIGHTNING PROTECTION OF AIRCRAFT (CONT.) 

Document 

Identification Title 

Applicable 

Paragraphs 

MIL-1-83456 (USAF) Miritary Specification: Installation of All US Air Fo

2 June 1976 Segmented Lightning Diverter Strips on 

Aircraft Radomes, General Specification For 

BCAR D4-6 British Civil Airworthiness Requirements All Civil Aero

1 November 1963 Sub-Section 04 - Design and Construction 

Chapter 04-6 Electrical Bonding and c. 

+ Lightning Discharge Pro tectiotz (A ircraft) W 

BCAR G4-6 British Civil Airworthiness Rey uirements All 

7 November 1975 Sub-Section G4 - Design and Construction 

Chapter G4-6 Electrical Bonding and 

Lightning Discharge Protection (Rotocraft) 

Civil Aero

TSS Standard No. 8-6 Electrical Bonding and Lightning Discharge All 

29 May 1973 Protection (Supersonic Transport) 

Anglo-Fr
REFERENCES 

5.1 Lightning Test Wave forms and Techniques for Aerospace Vehicles and 

Hardware, Society of Automotive Engineers, Warrendale, Pennsylvania (5 

May 1976). 

5.2 B. L. Perry, "British Researches and Protective Recommendations of the 

British Air Registration Board," Lightning and Static Electricity Conference, 2-3 December 1968, Part 11, Conference Papers, AFAL-TR-68-290, 

Air Force Avionics Laboratory, Air Force Systems Command, WrightPatterson Air Force Base, Ohio (May 1969), pp. 8 1-103. 

5.3 Perry, "British Researches and Protective Recommendations," pp. 90, 92, 

91. 

5.4 Protection of Aircraft Fuel Systems Against Lightning, Federal Aviation 

Agency Advisory Circular AC 20-53, Federal Aviation Administration, 

Department of Transportation, Washington, D.C. (6 October 1967), p. 2. 

5.5 Rockwell International designed research vehicle, Aviation Week and 

Space Technology, November 17, 1975, p. 38. 

5.6 Lightning Test Waveforms and Techniques for Aerospace Vehicles and 

Hardware, p. 9, 

5.7 J. Phillpott, "Simulation of Lightning Currents in Relation to Measured 

Parameters of Natural Lightning." Proceedings of the 1975 Conference on 

Lightning and Static Electricity at Culham Laboratory, England, 14-17 

April 1975, Session I: Fundamental Aspects and Test Criteria, the Royal 

Aeronautical Society of London (December 1975), pp. 6-7. 
CHAPTER 6 

FUEL SYSTEM PROTECTION 

6.1 Introduction 

Because of the hazardous nature of most fuels and the potential for 

catastrophe if ignition should occur, the design of adequate lightning protection 

for the aircraft fuel system is probably the most important lightning protection 

job to be accomplished. This is the system that government airworthiness 

certification requirements stress most heavily, for the cause of the most tragic of 

lightning-related aircraft accidents has been attributed to the vulnerability of 

fuel systems. 

Elements of the fuel system are typically spread throughout much of an 

aircraft and occupy much of its volume. They include the fuel tanks themselves, 

as well as associated vent and transfer plumbing, and electrical controls and 

instrumentation. Careful attention must be paid to all of these elements if 

adequate protection is to be obtained. 

The main objective of fuel system protection is of course to prevent any 

direct or indirect source of ignition of the fuel by lightning. Accomplishment of 

this goal is quite challenging when it is remembered that thousands of joules of 

energy must be conducted through the airframe when the aircraft is struck by 

lightning, that every metallic structural element in and on the aircraft is involved 

to some degree in this conduction process, and that a tiny spark of - 2 x lo4 

joule may be all of the energy that need be released inside a fuel tank to ignite a 

fire or initiate an explosion. 

Prevention of fuel ignition from lightning must be accomplished by at least 

one of the following approaches: 

(a) removing one or more of the requirements for combustion 

(b) eliminating all sources of ignition. 

In this chapter we review typical aircraft fuel flammability characteristics 

and possible ways to control fuel flammability and ignition. Research in the field 

of fuel system protection distinguishes the areas of adequate knowledge and 

practice from areas of inadequate knowledge and practice. Because current 

guidelines for controlling lightning-related ignition derive in part from this 

research, we refer in some detail to several important studies, by way of 

background, before presenting guidelines for effective protection of fuel 

systems. 

6.2 Fuel Flammability 

Ignition of fuel cannot occur until some of the fuel has become vaporized 

and mixed in a compatible proportion with oxygen or air. Thus, the ability of 

fuels to vaporize and mix with air in typical aircraft fuel tanks is of primary 

concern. 

The flammability of the vapor space in a fuel tank varies according to the 

concentration of evaporated fuel in the available air. Reducing the fuel-to-air 
ratio below a definite minimum value produces a vapor and air mixture that is 

too lean in fuel to burn. Likewise, there is a maximum fuellair ratio which, when 

exceeded, results in a vapor space mixture too rich in fuel to be flammable. In 

between these extremes there is a range of mixtures that will burn. When only 

equilibrium conditions are considered, the particular fuellair ratio that can exist 

is determined by the temperature and altitude of the fuel system. The 

temperature determines the quantity of the fuel by controlling its vapor 

pressure, and the altitude determines the quantity of air. Therefore, by a suitable 

combination of temperature and altitude, under equilibrium conditions, the 

ullage of a fuel tank can be made either flammable or nonflammable by varying 

combinations of temperature.and altitude. 

Much research into the fuel vapor and flammability characteristics that 

exist within aircraft fuel systems has been accomplished by many organizations 

concerned with aircraft propulsion and the resulting need to safely transport 

large quantities of fuel aloft in a vehicle with severe space and weight limitations. 

This research has included laboratory investigations of the flammability 

characteristics of the fuel vapors that exist within laboratory-type containers as 

well as those that exist during flight. Perhaps the most comprehensive laboratory 

work is that of Nestor (Reference 6.1), who has paid particular attention to the 

behavior of fuels in aircraft tanks and flight environments. Nestor shows that, as 

a result of the wide variety of temperatures, pressures, and motions that can 

exist in flight, there are correspondingly wide variations in the amount of fuel 

that can exist in the vaporized state. One reason for this variation is that, in 

addition to fuel already present as a vapor, the aircraft motion and vibration also 

cause the liquid to be dispersed in the vapor space in the form of mists and 

spray. An ignition source may preheat these droplets of fuel to a point where 

sufficient vapors are evolved to support a flame, even if insufficient vapor has 

been released from the bulk fuel surface itself. The heat from the developing 

flame then causes more fuel to vaporize and continue to feed and enlarge the 

flame. Temperature and altitude, however, are the primary characteristics that 

relate to flammability. 

The relationship of temperature and altitude to tank vapor flammability is 

illustrated by the flammability envelope. A typical flammability envelope is 

shown in Figure 6.1 (Reference 6.2). 

Turbine engine fuels are placed in two broad categories based on their 

distillation temperature ranges: wide-cut turbine engine fuels and aviation 

kerosenes. Aircraft with reciprocating engines are commonly fueled with 100 

octane gasoline. Turbine- Cjet-) powered aircraft fueled in the United States and 

in some foreign countries utilize aviation kerosenes. These include Jet A and Jet 

A-1 fuels. U.S. military aircraft and commercial transports in some other foreign 

countries are fueled with wide-cut fuels, such as Jet B or JP-4, which are 

kerosene-type fuels with a higher percentage of volatile components. The 

flammability limits for the turbine fuels as determined in the laboratory by 

Nestor are presented in Figure 6.2 (Reference 6.3). 

The flammability envelopes are defined as lines beyond which no 

flammable points were observed in the laboratory tests. In Figure 6.2 these lines 
P 

TEMPERATURE 

Figure 6.1 Typical flammability envelope of an aircraft fuel. 

have been drawn to enclose those data points at which ignition and flame 

propagation occurred but exclude some of the data points where only ignition 

occurred with no flame propagation. This results in a random scatter of ignition 

points about the lines, as shown by Nestor in Figure 6.2. The flammability 

envelopes of the aviation kerosene-type fuels (Jet A and Jet A-I) differ from the 

wide-cut fuels (JP-4 and Jet B) primarily in their temperature ranges. The 

aviation kerosenes have an average lean limit at sea level of about 38 "C and a 

rich limit of approximately 80 OC. For the wide-cut fuels, the lean and rich 

limits at sea level are about -20 "C and I0 "C, respectively. 

Figure 6.3, taken from 3 Lockheed study on flammability of advanced 

flight vehicles (Reference 6.4) shows the flammability envelopes of aviation 

gasolines (AVGAS) along with a number of turbine fuels. The Reid vapor 

pressure of each of these fuels is also presented, and here the relationship 

between this parameter and volatility becomes apparent. Fuels with higher vapor 

pressures will release sufficient vapor at lower temperatures to reach a flammable 

mixture, thus lowering the flammability envelope as compared with the envelope 

of a fuel with a lower vapor pressure, which must be at a higher temperature to 

release sufficient vapor to form an ignitable mixture. 

The temperature vs altitude profiles for standard and tropical atmospheres 

are also presented in Figure 6.3 to show the mean air temperature at various 

flight altitudes. 

There are important factors which may alter the flammability limits of the 

vapor inside an aircraft fuel tank from those shown in Figures 6.1 and 6.3, 

however. One of these is the mixing of one type of fuel with a small amount of 
TEMPERATURE, OC 

FLAME PROPAGATION OBSERVED 

r NO FLAME PROPAGATION OBSERVED 

FOUR INCH TUBE. HIGH ENERGY SPARK 

TEMPERATURE. OC 

Figure 6.2 Flammability limits for typical aircraft turbine fuels as a function of 

altitude and temperature. 

another. Figure 6.4 (Reference 6.5) shows, for example, the flammability limits 

determined by Nestor of a mixture proportionately comprised of 85% Jet A and 

15% Jet B fuels. Such a mixture might have occurred if the aircraft had 

originally been fueled with Jet B and later refueled with Jet A. The flammability 

envelope of the resulting mixture has been lowered such that it encompasses the 

altitudes and temperatures where most lightning strikes to aircraft have 

occurred. 

Agitation is another way that the flammability envelopes might be altered 

from those shown in Figure 6.2, for the latter are valid only when the mixture is 

stabilized. Agitation of a fuel or spray from a pump or pressurized fuel line can 

result in extending the lower temperature flammability limit below the curves 

shown in Figures 6.2 and 6.4. Thus, even though Figures 6.2 and 6.4 indicate 

that the vapor-air mixture in the tanks of an aircraft fueled with Jet A would be 

too lean to support combustion, there can be no assurance that this is always so, 

since agitation of the fuel may occur in flight. 

A number of methods have been developed to provide more positive 
TEMPERATURE, OC 

Figure 6.3 Flammability envelopes of fuels of different volatility. 

TEMPERATURE, 'C 

Figure 6.4 Flammability envelope of the fuel blend 85% Jet A/15% Jet B, and 

altitudeltemperature envelope enclosing most lightning strikes. 

inerting of the ullage in aircraft fuel tanks--inerting sufficient to inhibit the 

enlargement and propagation of flames away from a point of ignition, so that an 

ignition source could occur yet not result in a catastrophic fire or explosion. 

119 
These methods include inerting of the fuel tank ullage with gases, such as 

nitrogen or carbon dioxide, and mechanicaI subdivision of the tank volume with 

reticulated foams. Nitrogen-inerting systems have been developed and flight 

tested on the USAF C-141 and C-135 (Reference 6.6) among others, but they 

have not been used extensively on any fleet of commercial aircraft because of 

excessive cost and weight penalties. Inerting systems have been used on several 

military aircraft for protection against fuel ignition resulting from enemy 

gunfire. 

Even if inerting systems were to come into common use, it would be 

advisable to take positive steps to assure that no ignition source could occur in 

the event of a lightning strike because there may never be complete assurance 

that a flammable mixture did not exist somewhere in the fuel system. Therefore, 

while the use of low-volatility fuels such as Jet A or while the inclusion of 

inerting systems to reduce the flammability of the atmosphere in a fuel tank will 

undoubtedly increase flight safety in the event of a lightning strike, these 

measures should not be relied upon to prevent all in-flight fires or explosions 

resulting from lightning. Positive steps must still be taken to prohibit ignition 

sources within the fuel system. 

6.3 Sources of Ignition 

While there have been a number of in-flight fuel tank explosions 

undoubtedly caused by lightning strikes, the exact location or source of ignition 

has not been positively identified in any of these accidents. This failure is 

perhaps a result of the fact that so small an amount of energy in an electrical 

spark is required to cause ignition that such a spark would leave little or no other 

evidence of itself (like pitting of a metal surface). It may be possible that 

ignition has begun outside of the fuel system itself--for example, at a fuel vent 

outlet--and then propagated inside, although there has been no proof of this. 

Laboratory studies involving simulated lightning strikes to fuel tanks or 

portions of an airframe containing fuel tanks have demonstrated several possible 

ignition mechanisms, and investigations of the accidents involving fuel tank fires 

and explosion have raised other possibilities. Possible ignition sources can be 

divided into the following two broad categories. 

Thermal Ignition Source 

1. Direct contact of the lightning arc with the fuel-air mixture, as at a 

vent outlet 

2. Hot spot formation or complete meltthrough of a metallic tank skin 

by lightning arc attachment 

Electrical Ignition Sources 

1. Electrical sparking between two pieces of metal conducting lightning 

current, such as poorly bonded sections of a fuel line or vent tube 

2. Sparking from an access door or filler cap (which has been struck) to 

its adapter assembly 

3. Sparking among elements of a capacitive-type fuel quantity probe, 
caused by lightning-induced voltages in the electrical wires leading to 

such a probe 

The balance of this chapter considers each of the above ignition sources 

and methods of their elimination. 

6.4 Vent Outlets 

The temperature of a lightning arc far exceeds that required to ignite a 

flammable fuel-air mixture; therefore, any direct contact of the lightning flash 

with such a mixture must be considered an ignition source. Since fuel tank vents 

are the primary means by which a flammable fuel vapor can be exposed to the 

outside of an aircraft, a considerable amount of research has been undertaken to 

evaluate the possibilities of lightning ignition of fuel vent vapors. One of the first 

studies of this possibility is that of the Lockheed-California Company and the 

kghtning and Transients Research Institute in 1963 for the National Aeronautics and Space Administration (Reference 6.7). In this program, fuel-air 

concentrations in the vicinity of aircraft vent outlets of several configurations 

were measured and mapped under various conditions of tank vapor fuellair ratio 

and effluent velocity. The tests were run in a wind tunnel producing an airflow 

of up to 100 knots. Mast vents discharging into wakes and free air streams were 

tested, as well as flush vents discharging into boundary layers, as shown in 

Figure 6.5 (Reference 6.8). The study showed that a vent discharging into a free 

stream exhibits the greatest dilution and thus the smallest flammable region, but 

that flammable mixtures can exist in the immediate vicinity of each type of vent 

outlet. 

A typical mixture concentration profile aft of a flush vent outlet is shown 

in Figure 6.6 (Reference 6.9). Depending upon the richness of the original 

effluent, dilution of it by the addition of air to 3096, or more, might well lean it 

out of the flammability envelope. Thus, dilution to a nonflammable mixture 

probably occurs farther than about one outlet diameter away from the outlet. 

This finding suggests that a lightning strike or associated streamer must occur 

very near to the edge of a vent outlet for an ignition to occur. 

The above prediction was confirmed in the second part of the 1963 

Lockheed-LTRI study, which included application of simulated lightning strikes 

to a flush vent outlet with flammable fuel-air mixtures exiting from it into an 

airstream moving at 100 knots, both with and without flame arresters installed. 

These flame arresters, which were being evaluated as possible protective devices, 

consist of a parallel bundle of small-diameter metal tubes inserted into the vent 

line near the vent outlet, as pictured in Figure 6.7(a) (Reference 6.10). An 

alternate construction utilized a series of baffles extending into the fuel vapor 

flow, as shown in Figure 6.7(b). The object in either case is to quench 

(extinguish) the flame by cooling it, and the small tubes, or baffles, are intended 

to accomplish this by providing additional surface area to conduct heat away 

from the flame. The test results showed that it was possible to get ignition of the 

vent effluent in the 100-knot airflow only when the lightning arc was delivered 
- AIRFLOW WAKE RESULTING FROM 

NOTE: THIS CONFIGURATION MAY HAVE FLAP, 

AILERON, OR BOTH, ADJACENT TO THE 

VENT MAST TO DISTURB FLOW PATTERN. 

Figure 6.5 Three general classes of fuel vent exits. 

directly to the lip of the vent outlet. Arcs delivered to spots as little as 2.5 cm 

away did not ignite the effluent. The results also showed that flames ignited by 

the strikes to the outlet could propagate inward through the flame arresters 

when these were instaIled near the vent outlet. While the flame arresters did 

extinguish some flames ignited at the outlet, they did not stop all such flames. 

Evidently the intense heat and blast pressures of the lightning arc act to heat the 

arrester and force the flame through. 

As a result of the crash of a Boeing 707 aircraft near Elkton, Maryland, on 

December 8, 1963, after being struck by lightning (Reference 6.1 I), another 
BOUNDARY LAYER THICKNESS 0.50 VENT DIAMETERS 

VENT EXIT VELOCITY 0.10 X FREE STREAM VELOCITY 

LONGITUDINAL DISTANCE - VENT DIAMETERS 

6 8 10 

I I I I I 

-10- 

VERTICAL DIAMETERS 2- 

1 - PERCENTAGE INITIAL CONCENTRATION 

Figure 6.6 Typical profile of vent effluent in air stream aft of a flush vent 

outlet. (Profiles vary according to boundary layer thickness, effluent 

and air stream velocities, etc.) 

investigation of the possibility of lightning-initiated fuel tank vent fires was 

undertaken by Bolta and others of the Atlantic Research Corporation with the 

support of Newman and others of the Lightning and Transients Research 

Institute (Reference 6.12). This work was sponsored by the FAA and focused on 

the Boeing 707 wing tank and vent system, with the objective of determining the 

conditions under which ignition of fuel vent effluent allows flames to propagate 

back through the vent duct and surge tank and from there into the reserve tank. 

Another objective was the evaluation of various protective measures, including 

flame arresters and flame-extinguishing systems. 

Unlike the earlier Lockheed program, the vent was tested in still air, the 

rationale being that, if an ignitable effluent is assumed to be at the outlet, 

attachment of the lightning arc to the lip of the vent outlet is the governing 

factor in obtaining ignition. The ignitable mixture was a 1.15 stoichiometric 

mixture of propane and air. 

Thermocouples were installed along the vent line to determine the time at 

which a flame passed by, thus enabling calculation of flame front velocities. The 

latter information was important because an automatic extinguishing system 

under consideration depended for its operation on sufficient time elapsing 

between the initial sensing of a flame at a vent outlet and the activation of an 

extinguisher in the surge tank located about 1 m down the vent line. The 

simulated lightning tests confirmed the earlier result--that a strike must occur 

very close to the vent outlet for ignition to occur. Average flame velocities of up 

to 45 m/s were recorded between the vent outlet and the surge tank when 

simulated lightning strokes of 175 kA and 1.5 x lo6 ampere2-seconds (A~-s) 

were applied. These strokes simulated a severe return stroke such as that which 

might be experienced if the vent outlet were located in Zone 1A or 1B. Other 

tests were run with the high current delivered through a 10 cm aluminum foil 

tape which would explode as the current passed through it. The resulting flames 

reached higher velocities (up to 126 m/s), but the exploding tape may have 

caused the higher flame speeds, creating conditions more severe than those in a 

natural case, where an arc alone is the ignition source. 

When lower amplitude currents of 44 kA and only 0.001 x lo6 A~-S were 

123 
OUTLET 

OUTLET 

(b) 

Figure 6.7 Flame arrester configurations. 

(a) Tubular construction 

(b) Baffle construction 

applied, ignitions still occurred, but the highest average flame velocity was 17.4 

m/s, which indicates that the intensity of the lightning discharge is one of the 

factors that affect the flame velocity. The 44 kA current is described by Bolta as 

a "high-voltage" discharge (Reference 6.13) because it was applied with a 

Marx-type, high-voltage generator, but the ignitions must have occurred from the 

ensuing 44 kA current discharge, since the 13 cm or 30 cm air gap between 
discharge electrode and vent outlet was too short to simulate the in-flight 

environment of the aircraft in a high electric field and allow formation of 

significant prebreakdown streamers at the vent outlet. 

Various flame arresters fabricated from corrugated aluminum and stainless 

steel, a ceramic material, and various copper screens were tested. None of these 

arresters was capable of stopping a flame ignited by a simulated stroke current 

when the arrester was installed near the outlet of the vent line, but arresters 

wound from corrugated stainless steel 1.27 cm or 2.54 cm deep did stop the 

flames when these arresters were installed about 1 m upstream from the vent 

outlet near the surge tank (as shown in Figure 6.8 [Reference 6.14]), even when 

these flames traveled in the vent tube at an average speed as high as 122 m/s (the 

highest measured). Other arresters made of screens and ceramics did not stop 

flames ignited by the simulated strokes. 

LIGHTNING 

STRIKE 1, OUTLET TO OUTLET 

1.27 OR 2.54 cm DEEP 

ONE ARRESTER HERE 

DID STOP FLAMES 

CONSTRUCTION 

TYPICAL CELL 

DEPTH: 2.54 cm 

Figure 6.8 Successful flame arrester installation in simulated B-707 vent tube. 
A flame-suppression system developed by Fenwal, Inc. (Reference 6.15) 

for industrial applications was also tested in this program. This system consisted 

of a fast-acting sensor for detecting the presence of a flame and a set of canisters 

containing a quantity of liquid suppressing agent for release into the surge tank 

by an electric detonator. When the detector sees the light of the flame, it sends a 

pulse to the detonator, which disperses the extinguishant into the tank within a 

few milliseconds of flame sensing, before the flame itself has reached the tank. 

This system effectively suppressed those flames which traveled slowly enough 

(30 mls or less) to give the system time (18 ms) to react, but did not stop those, 

of course, which had gone past the surge tank by the time the extinguishant was 

released. In the latter case the extinguishant might still extinguish the fire in the 

surge tank, but the flame front on its way through interconnecting vent lines to 

the fuel tanks would have passed out of the reach of the extinguishant. 

The Atlantic Research-LTRI program also included an investigation of arc 

plasma propagation into the vent line, which was inconclusive because of 

instrumentation difficulties. The question of icing of the successful flame 

arrester was also considered (Reference 6.16), resulting in the conclusion that 

unacceptable icing would occur only when the worst combination of atmospheric and flight conditions existed. This conclusion, however, is based on 

analysis only and should be verified by flight tests if this arrester or any other is 

to be incorporated into an aircraft fuel system. 

6.4.1 Airflow Velocity Effects 

Lightning attachment to the lip of a vent outlet was assumed to be 

possible in the Atlantic Research-LTRI and Lockheed programs reviewed in the 

preceding paragraphs. No attempt, however, was made in these programs to 

establish whether or not this phenomenon could in fact occur to an aircraft in 

flight or, if it could, how often. Yet neither the Elkton airplane nor any other 

airplane known to have been struck by lightning has shown evidence of lightning 

attachment to a vent outlet, and the outlet in the B-707, while located near the 

wing tip, is not located at the very tip of the wing, where lightning is known to 

attach. It is therefore improbable that flush-mounted vents, such as those on the 

B707, will receive direct strikes. These questions then arise: Could an arc sweep 

across the outlet from another point and ignite the effluent?; if so, How close to 

the outlet must such an arc pass to cause ignition?; and Is it possible for 

streamering to be induced at the vent outlet and ignite an effluent? Answers to 

these questions were sought by Newman and others, who undertook an 

experimental program (Reference 6.17) during 1966 and 1967. Simulated 

lightning strikes were delivered to a B-707 wing tip and vent assembly to learn 

more about these possibilities, as well as the degree to which air flowing past at 

realistic flight speeds would make ignition unlikely even if an arc did in fact 

become attached to the outlet. 

Whereas in earlier programs ignitions were obtained nearly 100% of the 

time when the vent outlets were in still air, Newman found that ignition of a 1.5 
stoichiometric propane-air mixture by a 48 kA, 0.009 x lo6 A~-S simulated 

direct stroke occurred only once in 34 shots with a 90-knot (46 m/s) airflow 

over the outlet and not at all after 200 shots in a 200-knot (100 m/s) airflow. 

When strokes of longer duration were swept across the vent outlet by the 

90-knot windstream, the effluent was ignited 11 times out of 15 runs, but when 

the airflow was increased to 200 knots, only 2 ignitions occurred in 46 runs. 

Nearly all of Newman's tests were performed under the most vulnerable 

effluent condition, which was found to be a 1.5 mls flow out of the vent outlet, 

such as might exist when an aircraft is climbing. Since more than half of all 

reported lightning strikes occur when the aircraft is either in level flight or 

descending and since most aircraft climb at well over 90 knots, the probability 

of an in-flight ignition from a direct stroke to a vent outlet must be remote. 

Newman's investigation, however, showed that a flash sweeping across the vent 

outlet might have a greater (2 in 46) chance of igniting an effluent, even under 

climb conditions at the more realistic speed of 250 knots. This result 

demonstrates the importance of locating vent outlets away from both direct and 

swept-stroke zones on the aircraft. 

6.4.2 Explosive Ignitions 

In the program just described, Newman and his colleagues conducted a test 

(Reference 6.18) in which a stroke to the vent outlet produced indications of 

unusually high flame velocities and severe deformation of the vent outlet, 

indicating much higher pressures than normal. They cite a similar case in another 

program in which flames traveling in excess of 300 m/s (1000 ftls) were actually 

measured. The implication of these findings is serious because an arrester or 

surge tank protection system capable of extinguishing the Iower velocity flames 

may not be able to stop flames traveling as fast as 300 mls. 

Kester and others (Reference 6.19) attempted to reproduce such speeds in 

a 14 cm simulated vent line, shown in Figure 6.9, but did not measure flame 

velocities higher than 20 mls (65 ftls) in this system, even when severe, 180 kA, 

1.0 x lo6 A~-S strokes were applied. These velocities were, in fact, comparable 

to those measured in the Atlantic Research-LTRI program of 1964. Kester and 

his colleagues also reported one explosive ignition, when a stroke of 195 kA was 

delivered to the vent outlet. It was found that the 195 kA stroke current had 

induced a voltage in instrument wiring sufficiently high to spark over the 

insulation around several pressure probes inside the vent line. The vent outlet 

and parts of the surge tank were badly deformed, even though these parts were 

made of 0.64 cm (0.25 in) steel. Again, much higher than usual pressures were 

indicated. The explosion in the Kester program serves as a warning that a similar 

consequence might conceivably result from inductive voltages in an actual fuel 

system unless, by means of design, care is taken to eliminate such situations. 

The question of whether or not these explosive ignitions could occur in 

actual fuel tank vent systems was of such importance that the Federal Aviation 

Administration undertook yet another study of flame propagation in vent 
PHOTO DIODES 

TOTAL LENGTH FROM 

OUTLET TO SURGE TANK 

OUTLET A 

FUEL 

1 

AND 

w 

AIR ENTRANCE 

Figure 6.9. Simulated vent system--flames up to 20 m/s. 

systems. The work, conducted by Gillis (Reference 6.20), expanded upon earlier 

research by including the study of flame behavior in the long vent Iines leading 

inboard from the surge tanks to the fuel tanks, which comprise a typical, 

complete, vent system (Figure 6.10) (Reference 6.21). Gillis did not use 

simulated lightning arcs for an ignition source but, instead, discharged 100 J of 

electrical energy into a spark plug at the vent outlet. This is much less energy 

than would be released by a lightning arc of the same length. Nevertheless, Gillis 

recorded flames (Reference 6.22) that had accelerated to 300 mls (1000 ftls) far 

inboard (shown in Figure 6.10) when the aircraft was in a climb condition. 

When fresh air was made to flow into the vent outlet, as in a descent 

condition, supersonic velocities like this also occurred as near to the surge tank 

as probes PI and P2 in Figure 6.10. These supersonic flames expanded some of 

the vent lines from a rectangular to an elliptical cross section, a result similar to 

those occurring in the earlier programs of Newman and Kester. 

The total number of authentic tests run by Gillis was 13, of which I1 

resulted in flame velocities of 150 m/s or higher. The occurrence of such high 

speeds is perhaps best explained in Gillis' own words (Reference 6.23): - - 

When an explosive gas is confined in a channel and ignited, the flow 

induced by the thermal expansion of the gas in the combustion wave 

is restricted by the channel wall. Consequently, the flow attains 

much higher velocities than under conditions of free expansion in an 
open flame and flame and flow commonly augment each other by a 

feedback mechanism as follows: stream turbulence, however slight it 

may be initially, produces a wrinkling of the combustion wave 

surface; the resulting increase of surface increases the amount of gas 

burning per unit time, namely, the flow of gas in the channel; this in 

turn produces more turbulence and hence, increased wrinkling of the 

wave, and so on, so that the progress of the combustion wave 

becomes nonsteady and self-accelerating. In addition, the burning 

velocity increases as the unburned gas ahead of the flame is 

preheated and precompressed by the compression waves that are 

generated by the mass acceleration in the combustion wave. The 

compression wave is initially a comparatively weak pressure wave, 

which is overtaken and reinforced during its travel by numerous 

other pressure waves originating in the combustion zone. The 

coalescence of these pressure waves into a strong shock front in a 

configuration which is dead-ended can result in a reflection of the 

shock wave back toward the conlbustion zone. The effect of the 

passage of this reflected shock wave through the combustion wave is 

similar to the effect of a sudden release of pressure by a rupture of a 

diaphragm. A rarefaction wave propagates backward into the 

unburned gas and a jet of unburned gas develops which penetrates 

deeply into the burned gas. The shear between burned and unburned 

gas in this flow configuration produces extreme turbulence so that a 

sudden large increase in the burning rate occurs. 

In three of the 13 tests mentioned above, localized pressures of sufficient 

intensity to distort 3- to 5-foot sections of the duct were developed in the 

rectangular vent duct. A subsequent hydrostatic pressure test of a 1 m section of 

similar duct showed that a pressure of approximately 475 psig was required to 

produce similar distortion. This pressure exceeds the structural limitations of 

typical aircraft fuel tank and vent structures. 

Gillis concluded that the surge tank located just inboard of the vent outlet 

is a factor contributing to the high flame speeds because, when a flame reaches it 

from the vent outlet, the pressure permitted to build up in it serves as a force to 

drive flames rapidly on down the vent lines towards the fuel tanks. This creates 

turbulence in these ducts, which further serves to accelerate flames down the 

ducts. Gillis' work is perhaps the most convincing demonstration to date that 

flames traveling at sonic velocities and with damaging overpressures can occur in 

typical transport aircraft vent systems. Since flame arresters or STP systenls 

reliably capable of stopping such flames are not yet available, the importance of 

preventing any source of ignition within or near the vent system is very clear. 

6.4.3 Summary of Vent Flame Research and Protection 

Considerations 

Table 6.1 summarizes the ignition and flame velocity results for each of 
EFFLUENT 

6.7 m/s 

AS IN 

MAXIMUM 

CLIMB 

SCOPE OF 

PREVIOUS lNVESTlGATlONS 

-C -/-

------ //-- 

AVG. 1.4 rn/s ----------- 1----- 

MAIN TANK 

Figure 6.10 Flame speeds in mock-up vent system. 
Table 6.1 SUMMARY OF RESULTS OF SIMULATED LIGHTNING STRIKES TO FU

(STROKES TO LIP OF VENT OUTLETS ONLY. NO IGNITIONS W

OBTAINED FROM STROKES AWAY FROM VENT OUTLET.) 

Simulated Lightning 

Amplitude Action lntegral Airstream Attachment 

Program (kA (10"'s) Velocity (knots) Point 

1963 Luckheed-LTRI 100 0.069 100 Lip of vent 100% ignitio

(50 m/s) outlet gating with a

arrester insta

1964 Atlantic Research-LTRI 1 75 1.5 0 Lip of vent 100% ignitio

(blta, et al) outlet gatlng up to 

m/s when ign

ing foil. Flam

these flames 

upstream fro

1964 Atlantic Research-LTW 44 0.00 1 0 Lip of vent 100% ignitio

(Bolta, et al) outlet gating up to 

arrester). 

1966 LTRl 

(Newman, et al) 

1%6 LTRl 

(Newman. el al) 

1966 LTRI 

(Newman, et al) 

1966 LTRl 

(Newman. et al) 

1966 LTRl 

(Newman, et al) 

1966 Dynamic Scien 

(Kester, et at) 

1969 Fenwal 

(G~llis) 

48 0.009 200 

( 100 mls) 

58 0.172 90 

(45 m/s) 

58 0.1 72 200 

( 1 00 mis) 

58 o.ln 250 

(1 30 mls) 

100 1 spark 0 

Lip of vent 

outlet 

Lip of vent outlet 

Swept 

access outlet 

Swept 

access outlet 

Swept 

access outlet 

Lip of vent outlet 

Lip of vent 

outlet 

1 ignition an

out of 34 dir

outlet (no ar

0 ignitions o

strokes to ve

1 l ignitions 

strokes acros

2 ignitions o

strokes acros

No ignitions 

strokes acros

10046 ignitio

gating up to 

11 ignitions 

over 150 rn/

tests. Sever

(300 mls). 
the research programs just discussed. While not all of the answers to 

lightning-related vent flame questions are in hand, a number of important 

conclusions and protection considerations can be drawn from the research. 

1. As of this writing there has been no positive evidence that a natural 

lightning strike ever ignited a vent effluent on a transport-type aircraft, although 

there have been several in-flight explosions after lightning has struck wing tips at 

places within a meter or so of the vent outlet. 

2. For ignition to occur, a lightning-type arc must attach directly to the 

edge of a vent outlet or within a few centimeters of it, or must be swept over a 

vent outlet. Neither of these situations has been reported to date. Most aircraft 

vent outlets are located in conformance with FAA Advisory Circular AC 20-53 

(Reference 6.24). This guideline says that vent outlets should not be located in 

lightning-strike Zone 1 or 2 and defines these zones by the "18 inch" rule, 

which, if interpreted literally, are as illustrated in Figure 6.1 ](a). If the points 

defined by the "fore and aft line passing through the Zone 1 forward projection 

points of stroke attachment" are taken to be at the outboard tips only, then 

Zone 1 is restricted to a small area and Zone 2 may extend inboard only a little 

more than 18 inches (0.46 m) from the tip. If the tip is sharp, strikes are indeed 

most likely to attach only to the edge, and Zone 1 would be limited to a small 

distance (several centimeters or so) inboard. However, if the tip has a large radius 

of curvature, such as that illustrated in Figure 6.1 I, it is prudent to utilize a 

more conservative interpretation and consider Zone 1 to extend inboard the full 

18 inches (0.46 m). 

If the radius of curvature of the outboard tip is greater than 18 inches, the 

lightning attachment points will be even more scattered over the surface, and 

Zone 1 should be considered to extend inward a distance equal to the actual 

radius of curvature. 

When nonmetallic skin materials are used on the wing tips, it is probable 

that direct strikes will occur to the edge of the outermost metallic element, such 

as a rib or spar, instead of to the nonmetallic tip itself. Therefore, Zone 1 must 

include not only the nonmetallic tip but also an area extending 18 inches 

inboard from the outermost metallic structure, as shown in Figure 6.1 l(b). If 

there is a metallic diverter assembly mounted on the outboard edge of a 

nonmetallic tip, this diverter instead of the outboard rib (for example) may 

receive strikes and thus permit Zone I to begin at the outboard edge, as in 

Figure 6.1 ](a). Metal diverter bars intended for this purpose have been installed 

on nonmetallic wing tips, and subsequently failed, however, in their intended 

purpose, with strikes circumventing the diverter and either puncturing the 

nonmetallic skin or attaching to an exposed metallic item inboard of the 

diverter. The most common reason for this failure is that the diverter has been 

coated with a tough, erosion-resistant paint or coating. Moreover, most of these 

paints also have good electrical insulating capability, which inhibits formation of 

the electrical streamers necessary for the diverter to intercept oncoming 

lightning flashes. Greater reliability for a diverter intended to protect a 

nonmetallic tip assembly or to enable location of avent outlet farther outboard 

132 
ZONE2 / ZONE3 

FROM ZONE 1 

/ h-181~ BIN BIN --f/ 

(48 cm) (48 cm) 

Figure 6.11 Vent outlets must be located in Zone 3. With nonmetallic tips, the 

18-inch Zone 1 and Zone 2 areas must be measured from the outboard metallic structure and not from the nonmetallic tip. 

(a) Lightning-strike zones on all-metal wing tip 

(b) Lightning-strike zones on metal wing with nonmetal tip 

can be assured if diverters are subjected to such high-voltage attachment tests as 

those described in L ighfn ing Test Waveforms and Techniques for A erospace 

Vehicles and Hardware (Reference 6.25). Since the authenticity of this test has 

not been absolutely established, it is nevertheless prudent to locate a vent outlet 

in Zone 3 as defined more conservatively in Figure 6.1 I(b). 

3. Flame arresters of the corrugated steel type shown in Figure 6.8 have 

been the most effective of those experimented with in stopping flames. The 

research to date has not turned up a completely successful arrester, however. 

Flame arresters are most satisfactory when located some distance away 

from the vent outlet, so that blast forces in the immediate vicinity of the 

lightning arc will not propel flames through the arrester. The most successful 

location seems to be at the surge tank end of the vent outlet tube, as shown in 
Figure 6.8. A flame arrester at this location will certainly reduce the possibility 

of flames entering the surge tank, although there is no assurance that the arrester 

will stop all flames here. And if an arrester is located anywhere in the vent 

system, the possibility of its free air passage becoming blocked with ice must be 

evaluated, preferably by in-flight tests under the expected environmental 

conditions. 

4. Surge tank protection (STP) systems, the purpose of which is to sense 

flames initiated at vent outlets and extinguish them before they reach fuel tanks, 

are available and should be considered if the vent outlet must be located in or 

close to Zone 1 or 2. Since the time required from sensing of the flame to 

dispersion of the extinguishant is several milliseconds, there is a possibility that 

flames traveling at sonic velocities will outrace the STP system. Thus, while an 

STP system will unquestionably improve overall safety, it must not be relied 

upon to provide absolute protection of the vent system. 

5. Ninetydegree bends in the vent lines should be avoided because they 

expand the turbulence and surface area associated with propagating flames and 

thereby increase the velocity of propagation. Instead, straight or smoothly 

curved ducts should be utilized to minimize the possibility of explosive flame 

propagation. 

6. Location of the fuel vent outlet in an ascertained Zone 3 area will 

provide the highest degree of protection of any of the methods described above. 

A recessed or flush outlet is greatly to be preferred over a protruding tube 

outlet, which could become a source of corona or streamering when the aircraft 

is flying in a strong electric field. 

6.5 Fuel Jettison and Drain Pipes 

On some aircraft provision exists for dumping (jettisoning) fuel overboard. 

Often this provision consists of a pipe extending into the airstream from a fuel 

tank, as shown in Figure 6.12. A normally closed, electrically operated valve is 

inserted into the pipe, so that it is unlikely a flame could travel past this valve 

into the fuel tank. Even if the pipe were struck by lightning while jettisoning 

fuel, it is improbable that flames could propagate through the full pipe into the 

tank. 

A more likely hazard--if not guarded against in design--is that electrical 

sparks will occur from lightning currents flowing across a poor electrical bond 

between the pipe and the fuel tank wall, since the wall is frequently coated with 

an insulating, corrosion-resistant paint. Bonding jumpers installed across such 

joints may be adequate to equalize static charge differentials which sometimes 

occur in fuel systems, but these usually have too much inductance to prevent 

some lightning current from breaking down insulation and sparking at the 

interface between the pipe and tank wall. Clearly the best way is to make the 

shortest path also the conductive one, as shown in Figure 6.13. This may be 

done by making the faying surfaces of clean, uncoated metal and/or by 

providing bare metal-to-metal contact via the rivets or bolts. If there is doubt 
WING 

FUEL JETTISON PlPE 

'HERE - 

Figure 6.12 Fuel jettison pipe--a possible lightning attachment point. 

Good electrical bond between pipe and airframe desirable at 

point a, rather than forcing all lightning current to enter fuel 

tank wall at b. 

In every case a good electrical bond must be provided at b. 

INSULATING PAINT BETWEEN MATING 

SURFACES = SPARKING HERE 

\ 

- - 

PlPE 

BARE-METAL CONTACT BETWEEN PIPE 

FLANGE AND TANK WALL = NO SPARKING 

\ 

WORST BEST 

Figure 6.13 Bonding of fuel dump pipes to fuel tanks. 

Bond straps have too much inductance to prevent sparking 

across the shorter path. 

about the adequacy of a particular bond, the bond should be tested with 

simulated lightning currents to assure that sparking does not occur at the 

interface. The test currents to be applied should be those recommended in 

Lightning Test Waveforms and Techniques for Aerospace Vehicles and Hardware 
(see Reference 6.25) for the particular lightning-strike zone in which the jettison 

or drain pipe is located. 

6.6 Hot Spot and Hole Formation in Integral Tank Skins 

Integral tank skins are those in which fuel is in direct contact with the 

outside skin of the aircraft. Tanks of this type are commonly found in the wings 

of transport and some general aviation aircraft, and in the wings and fuselages of 

modern fighter aircraft. External fuel tanks of the type carried on pylons or 

wing tips by military aircraft are also of the integral type. 

If integral tank skins are located in direct or swept-lightning attachment 

zones, the possibility of hot spot formation or meltthrough of the skin must be 

protected against. For the aluminum skins in common use, making the skin thick 

enough to dissipate the heat from the lightning arc before the skin can be melted 

completely through will assure protection because ignition will occur only with 

complete penetration and the resulting arc contact with the fuel vapor. This 

phenomenon depends on the fact that the ignition temperature of most 

hydrocarbon fuel vapors (-1300 "C) is higher than the melting temperature of 

aIuminum (-500 "c); thus, ignition will not occur unless the skin is melted 

completely through to the fuel vapor by the hot Iightning arc, which can reach 

upwards of 30 000 "C. On the other hand, the melting temperature of titanium 

(-1700 OC) and stainless steel skins is higher than the fuel ignition temperature, 

so that, where this type is concerned, the skin need not be melted completely 

thkough for ignition to occur. 

The amount of lightning current required to erode or melt holes in aircraft 

skins has long been of interest, first, for the purpose of estimating how much 

lightning current actually was involved in the damage sustained by aircraft in 

flight and, second, for determining the minimum skin thickness required to 

prevent melt through of an integral fuel tank skin. Quantitative relationships 

between electric arcs and typical aircraft skin metals were reported in 1949 by 

Hagenguth (Reference 6.26), who made laboratory measurements of the amount 

of continuing current necessary to melt holes of various sizes in various metals. 

He found a nearly linear relationship to exist between the amount of charge (Q) 

delivered to an arc attachment spot and the amount of skin metal melted away. 

And, although he reported that "the type of metal appears to have very little 

influence" (Reference 6.27) (a fact later found to be untrue), he determined 

that the size of a hole melted in a metal sheet of a given thickness can be 

expressed by the following two equations: 

A = 2.53 Q (-4-0.9 for D(t<0.089 cm 

A=245 Q(&) -1.54 for 0.089<t<0.38 1 cm 

where A = area of hole melted (square millimeters) 

136 
Q = charge (coulombs) delivered to the point by the arc 

t = thickness of metal sheets (centimeters) 

Results of Hagenguth's laboratory tests are presented in Figure 6.14 (Reference 

6.28). By comparing the size of holes produced by actual lightning strikes with 

those produced in the laboratory, Hagenguth was able to estimate the amount of 

charge that had entered the aircraft through damaged parts, such as those shown 

in Figure 6.1 5. 

Taken at face value, the Hagenguth equations would predict a hole of 

some size (A) through skin metals of any thickness (t) up to 0.38 1 cm (0.15 in) 

40 80 120 160 200 

COULOMBS IN ARC 

Figure 6.14 Relation between coulombs in the arc and the size of hole burned 

in metal sheets. Arc current amplitude ranges between 50 and 

I000 A. 

1. Stainless steel - 10 mil (0.254 mm) 

2. Galvanized iron - 15 mi1 (0.381 mm) 

3. Copper - 20 mi1 (0.508 mm) 

4. Stainless steel - 40 mi1 (0.1 6 mm) 

5. Aluminum - 51 mi1 (1 2954 mm) 

6. Aluminum - 100 mi1 (2.54 mm) 
Figure 6.15 Natural lightning damage sustained by Airplane NC25629 about 

August 14, 194 1. Top views. 

Upper: trailing edge of left wing; material - 24st Alclad, 0.89 cm 

(0.35 in) thick. Note numerous small pitted spots. 

Lower: rudder tip bow; material - 17st aluminum alloy, 2.54 cm 

(1 in) O.D. tubing, 0.07 cm (0.028 in) wall thickness. 

Coulombs estimated from laboratory tests on metal sheets and 

tubing. 

Wing edge: large hole 137 C Rudder bow: 300 C 

small hole 6 C 

150pits 150C 

total 293 C 

no matter how small an amount of charge (Q) is delivered, but it can be 

concluded logically that, if the charge and resultant heating were too low, 

138 
damage would be limited to partial meltthrough and no hole would be formed. 

Robb, Hill, Newman, and Stahmann reported this result in 1958 (Reference 

6.29), based on tests in which a three-component simulated lightning current 

consisted of the following: 

1 IOOkA 0.069 x 1 o6 A~-S Return stroke 

2 5 kA 0.172 x lo6 A2-s 30 C intermediate stroke 

3 200kA 0.040 x lo6 A2-s 200 C continuing current 

This current was applied to commonly used samples of 0.05 1 cm (0.020 in), 

0.1 02 cm (0.040 in), 0.1 63 cm (0.064 in), 0.206 cm (0.08 1 in), and 0.259 cm 

(0.102 in) aluminum skin materials. Robb reported that this discharge, which is 

about as severe as a similar combination of SAE (Reference 6.30) current 

components D, B, and C (in that order, which would be specified for tests in 

Zone 2B), did not puncture the skins which were 0.206 cm or greater. Kester, 

Gerstein, and Plumer confirmed this result by showing that over 200 C 

(delivered at 288 A) are required to melt through 0.163 cm aluminum 

(Reference 6.3 1). 

The t hree-component discharge including a 200 C continuing current 

would almost certainly appear only at a trailing edge (Zone 2B) after being 

swept there from an initial attachment point somewhere on a leading edge or 

forward extremity. Fuel is rarely in direct contact with trailing edges of aircraft 

or within direct stroke zones (Zone IB), where extended flash hang-on is likely; 

yet airworthiness certification criteria, such as those listed in Table 5.1, generally 

require that integral fuel tank wall thicknesses be at least 0.203 cm (0.080 in) 

wherever lightning strikes are expected. 

Robb also observed (Reference 6.32) that swept-flash pit marks from 

several centimeters to about 60 cm (2 ft) apart are commonly found, and that a 

slow, worst case aircraft velocity of 62 mls (140 mph, 205 ftls) would result in 

the lightning arc dwelling for 9.7 ms at each point. 

- 0.6m =9.7ms 'dwell - 62 - mls 

If this is so, the amount of charge delivered to each dwell point across a typical 

3 m wing surface by a lightning flash similar to Robb's three-component discharge would be as shown in Figure 6.16 (Reference 6.33). 

Since the amount of charge entering any of the Zone 2A points in Figure 

6.16 is considerably less than the 230 C applied by Robb to melt through 0.1 63 

cm aluminum skins, it might be argued that a skin thinner than 0.203 cm would 

be adequate to prevent meltthrough. On the other hand, it has been found that 

the rate of charge delivery (current amplitude) is also important in determining 

whether sufficient energy is available to melt completely through, or not, and 

that smaller amounts of charge will also cause meltthrough of 0.163 cm 

aluminum if the amounts are delivered at a sufficiently high rate. Since any 

meltthrough of aluminum skins must be considered a possible source of fuel 
ZONE 2A 

ZONE 1A 

V) LY 

g looooo 

i lol= 

L loo 

= 10 E 

3 

U 1 20 4b 100 loo0 

Time - MILLISECONDS 

Figure 6.16 Charge delivered at successive dwell points across a wing surface. 

Only a small amount of charge enters Zone 2A dwell points. 

ignition, it is imperative that the minimum amount of charge and current within 

the range of possible lightning characteristics be established for proposed integral 

fuel tank skins. This has been termed by Plumer the coulomb ignition threshold 

(Reference 6.34). 

6.6.1 Coulomb Ignition Thresholds for Clean Skins 

The amounts of charge and current required to melt through aluminum 

and titanium skins of various thicknesses and to cause fuel ignition were 

reported by Brick in 1968 (Reference 6.35) and by Oh and Schneider in 1972 

(Reference 6.36) to depend heavily on current amplitude as well as charge. 

Whereas earlier work had shown that over 200 C, when delivered by a current of 

200 A, were necessary to burn through 0.206 cm aluminum skins, the work of 

Brick, Oh, and Schneider shows that only about 10 C, when delivered by a 

current of 500 A, are required to melt completely through a 0.206 cm aluminum 

skin; and in laboratory tests as little as 2 C, when delivered by about 130 A, will 

melt a hole completely through 0.102 crn of aluminum. Oh and Schneider's 

meltthrough thresholds for these and other skin thicknesses are shown in Figure 

6.17 (Reference 6.37). The close (0.24 to 0.48 cm) proximity of their test 

electrode to the skins may have restricted natural movement of the arc on the 

surface of the skin and been responsible for these unusually low coulomb 

ignition thresholds. 

Work by Kester, Gerstein, and Plurner (Reference 6.38) with an L-shaped 

electrode spaced 0.64 cm above the skin, permitting greater arc movement, 

showed that 20 C or more, when delivered at 130 A, are required to burn 

through a 0.102 cm aluminum skin. Since a natural lightning arc is not restricted 

by an electrode, it is probable that the aluminum skin puncture threshold curves 

of Figure 6.17 are conservative, at least for unpainted surfaces. On the other 

hand, the electrical insulating properties of most paints tend to restrict arc 

movement and so would concentrate the heating effects at a smaller point, 

decreasing the amount of thermal energy required to melt completely through. 
Oh and Schneider have similarly determined the coulomb ignition 

thresholds for titanium skin materials of various thicknesses, as shown in Figure 

6.18 (Reference 6.39). They and other researchers have confirmed that, since 

the melting point of titanium is higher than the fuel ignition temperature, it is 

not necessary for a hole to be melted completely through for ignition to occur. 

For ignition to occur only a hot spot need be formed on the inside surface of a 

titanium skin. The lower thermal conductivity of titanium prevents rapid heat 

transfer away from the arc attachment point and accounts for the generally 

lower coulomb ignition thresholds than those for aluminum. 

Figure 6.1 7 Coulomb meltthrough and ignition threshold for aluminum skins. 
Coulomb ignition thresholds for stainless steel have not been studied as 

thoroughly as have those for aluminum and titanium, but they have been 

determined for currents of 80 to 130 A by Kester, Gerstein, and Plumer 

(Reference 6.40) to be 4.4 C for 0.051 cm (0.02 in) skins, 12 C for 0.102 cm 

(0.04 in), and 41 C for 0.152 cm (0.06 in) stainless steel skins. A comparison of 

these thresholds with those found by Kester and his colleagues for aluminum 

and titanium skins of the same thicknesses is shown in Table 6.2. The times 

required for ignition to occur in several of the tests reported in Table 6.2 are 

Listed in Table 6.3 (Reference 6.41). 

Table 6.2 COULOMB IGNITION THRESHOLDS FOR VARIOUS SKIN 

METALS. 1.5 STOICHIOMETRIC PROPANE-AIR MIXTURE 

Material Titanium Aluminum Stainless steel 

Thickness (Ti4AI-1 MO-IV) (2024T-3) (Series 304) 

Coulombs Coulombs Coulombs 

0.051 cm 

(0.020 in) 5.5 

0.102 cm 

(0.040 in) 5.75 

0.152 cm 

(0.060 in) 1001125 

not determined 4.4 

Table 6.3 DISCHARGE AND IGNITION TTMES AT 

COULOMB IGNITION THRESHOLD 

Thick- Couness lomb 

0.05 1 cm 

(0.020 in) 6.0 

0.102 cm 

(0.040 in) 8.0 

0.152 cm 

(0.060 in) 107.0 

Dis. Ign. Cou- Dis. Ign. Cou- Dis. Ign. 

time time lomb time time lomb time time 

(ms) (ms) (ms) (ms) (ms) (ms) 

*Above threshold level 

The charges and times given in Tables 6.2 and 6.3 for ignition to 

commence are somewhat greater than those associated with the lowest coulomb 

ignition thresholds (bottoms of the curves) of Figures 6.17 and 6.18 because the 
currents (rates of charge flow into the skins) of 80 to 130 A were generally 

lower in the tests of Tables 6.2 and 6.3 than the 100 to 1000 A currents which 

produced the lowest coulomb ignition thresholds of Figures 6.17 and 6.18. But 

since intermediate currents (as in component B) and continuing currents (as in 

component C) of from a few amperes to 1000 A are common in natural 

lightning, both data sources must be considered authentic. Clearly, the lower 

currents must dwell longer at a single attach point to produce the same damage 

and possibility of ignition that higher currents produce in shorter times. An 

illustration by Plumer (Reference 6.42) of the resulting wide range of charge and 

dwell time combinations necessary to melt through aluminum skins and ignite 

fuels beneath is shown in Figure 6.1 9. 

DATA SOURCES: 

PLUMER 

O PHILLPOTT t=ls 

500 + BRICK / 

400 - 

AVERAGE = 15ms 

(.020) (.060) (.loo) (.140) (.180) (.220) (.260) (.300) 

0.51 0.153 0.25 0.36 0.46 0.56 0.66 0.76 

SKIN THICKNESS - cm (INCHES IN PARENTHESES) 

Figure 6.19 Range of minimum combinations of lightning dwell time and 

charge necessary to ignite fuels beneath bare aluminum skins. 

6.6.2 Dwell Times 

The data of Figures 6.17 and 6.18 and of Tables 6.2 and 6.3 can be used 

to determine the possibility of meltthrough or hot spot formation sufficient to 

cause ignition if the amount of time a lightning arc may dwell at a particular 

143 
point and the amplitude of the current that flows into this point are known. 

Figure 6.20 (Reference 6.43) shows a time exposure photograph of an actual 

lightning arc attaching to successive dwell points along the radome-mounted 

pitot boom of a fighter aircraft, and Figure 4.2 shows what similar points look 

like along a typical aircraft surface. 

Knowledge of dwell times at swept-stroke (Zone 2A) attachment points 

like these is of most importance because integral fuel tank skins are often found 

in these zones. Existing regulatory documents (Table 5.5) have specified that 

aluminum skins in lightning-strike zones should be at least 0.203 cm (0.080 in) 

thick to withstand meltthrough because the simulated lightning current of 

Figure 6.16 would not melt through this thickness of aluminum, even if all of 

this current is delivered to one point. The regulatory documents also imply that 

other materials should be of thicknesses capable of withstanding an equivalent 

lightning current. 

Figure 6.20 Time exposure of a lightning arc at successive dwell points. 

Until recently, structural design demands also made it necessary to use 

skins of at least 0.203 cm, so the need for this thickness for lightning protection 

was not challenged. But in some recent designs, including the designs of 

supersonic transports, thinner skins would have met structural requirements and 

have permitted a savings in weight and cost were it not for the lightning 

144 
protection requirement for a 0.203 cm thickness. Therefore, Brick, Oh, and 

Schneider (Reference 6.44) studied dwell times of 400 A decaying arcs blown by 

the exhaust from a wind tunnel over aluminum and titanium skin panels with 

several surface treatments to see how long such an arc might actually dwell at 

one point before reattaching to the next. Wind speeds of 67 m/s (150 mph) and 

112 m/s (250 mph), representing approach speeds for typical aircraft, were 

utilized. The test current they used was representative of lightning continuing 

current, and would have delivered an average charge of about 0.2 C/ms of dwell. 

The researchers reported that the arcs dwelled for 2 ms or less on uncoated 

surfaces of both metals, and for 4.8 ms on an anodized aluminum surface. At 0.2 

Clms of dwell these dwell points would have received 0.4 and 0.96 C, 

respectively. By Figures 6.17 and 6.18 this would not have been enough charge 

to melt through even the 0.05 cm thicknesses of either metal. If the lightning 

current amplitude experienced in Zone 2A never exceeded 400 A, this result I would indicate that skins as thin as 0.05 cm would be safe. The problem is, of 

course, that the occurrence of higher currents has now been recognized as 

possible in Zone 2A. 

At the end of the dwell at one point, the arc will reattach at a point 

farther aft on the aircraft, as shown in Figure 6.21 and explained in Section 2.3. 

During the dwell times the aircraft would be moving, and the distance covered 

would be D=v td (6.4) 

Where D = distance arc is drawn along the aircraft surface (m) 

v = aircraft velocity (m/s) 

td = dwell time (s) 

Thus, at velocities of 67 and 112 m/s and with the dwell times of 2.0 and 

4.8 ms reported by Brick, spacings between successive dwell points would be as 

given in Table 6.4. 

EARLIER LATER 

DIRECTION 

OF FLIGHT 

Figure 6.21 Basic mechanism of swept-stroke reattachment. 
Table 6.4 DWELL POINT SEPARATIONS FOR TYPICAL LIGHTNING 

DWELL TIMES AND AIRCRAFT VELOCITIES 

Dwell 

Times (td) 

4.8 ms I D=32.2cm D=53.8cm 

Painted (12.7 in) 1 (21.1 in) I 

Aircraft Velocities (v) 

67 mls I 112 mls 

2.0 ms 

Unpainted 

surface 

surface I 

Because spacings of this order are often observed along aircraft surfaces, 

the dwell times predicted by the Brick tests are realistic. The later work of Oh 

and Schneider (Reference 6.45) basically confirms these results for uniform 

airflow conditions but shows that conditions which cause the airstream to leave 

the surface may force the arc to dwell longer at the last attachment point before 

the airflow is diverted. Oh and Schneider also demonstrate that higher aircraft 

velocities result in shorter arc dwell times, since the arc is stretched greater 

distances, allowing sufficient voltage to build up along it to break down the 

insulation at its heel (see Figure 6.21) at an earlier time. 

Validity of the wind tunnel technique for simulating swept strokes along 

metallic surfaces has sometimes been questioned because the airstream is used to 

blow the arc along the surface of the test object, whereas in flight both the air 

and the arc are stationary. It might be expected that the moving air would carry 

the arc along at an equal velocity, but in practice this doesn't quite happen. In 

addition, there must be cooling of the arc by the faster moving air, which allows 

it to build voltage that in turn might create a new attachment point sooner than 

had cooling not occurred. Also, the behavior of the upper terminus of the arc as 

it moves along the electrode probably has some effect on the behavior of its 

lower terminus at the test object. The scope of these effects and the extent to 

which they cause departure from an authentic simulation of natural lightning has 

not been investigated. 

A more realistic simulation would undoubtedly result if the test object, 

like an aircraft in flight, could be moved through a stationary arc. Plumer 

(Reference 6.46) attempted this by driving a wing tip fuel tank beneath a 

high-voltage electrode at a velocity of 15.5 mls (35 mph) A flash was triggered 

when the approaching tank, carried atop a truck, sufficiently closed the airgap 

between the electrode and ground. Limitations of this technique were the low 

velocity and the low (4A) test current amplitude. Nonetheless, subsequent 

in-flight lightning strikes to two fuel tanks of this type have confirmed the 

146 

D=13.4cm 

(5.3 in) 

D=22.4cm 

(8.8 in) 
occurrence of the predicted attachment points and breakdown paths. Despite 

the lower velocity and current amplitude, the arcs in Plumer's tests also dwelled 

for times of between 1 and 4 ms on unpainted aluminum surfaces, results closely 

parallel to those of the wind tunnel work. These parallel conclusions seem to 

indicate that the arcs in the wind tunnel tests may not have moved as fast as the 

wind itself, and, also, that current amplitude has relatively little to do with dwell 

time. 

Robb, Stahmann, and Newman (Reference 6.47) have utilized the wind 

tunnel technique to determine arc dwell times on various painted or coated 

surfaces. Most of these coatings are electrically insulating and thus require that 

the arc be further lengthened to allow the greater voltage buildup necessary to 

puncture the insulation and form the next attachment point. The reasonable 

agreement of test methods with the in-flight strike evidence makes it possible to 

show, with some confidence, in Table 6.5 the dwell times to be expected on 

several surfaces commonly found in swept-stroke zones (Zone 2A). 

Table 6.5 LIGHTNING DWELL TIMES ON TYPICAL 

AIRCRAFT SURFACES lN ZONE 2A 

Note 1 - Plumer (See Reference 6.46) 

Note 2 - Oh and Schneider (See Reference 6.45) 

Note 3 - Robb, Stahmann, and Newman (See Reference 6.47) 

Note 4 - Brick, Oh, and Schneider (See Reference 6.44) 

Aircraft Velocities 

There are many different coatings and paints of various thicknesses found 

on aircraft surfaces, and the effect of many of these on arc dwell time has not 

yet been determined. It may therefore be advisable to perform swept-stroke tests 

on metallic skiits by considering the particular coatings for a particular design. 

Guidance for performance of such tests is given in Section 4.3 of the SAE report 

Lightrlirig Test Wale forms and Techniques for Aerospace Vehicles arld Hardware 

(May 1976). 

Once the expected dwell times are established, next it will be necessary to 

determine the amplitude of current that might flow during these periods. 

Surface Type 

Aluminum and 

titanium unpainted 

Aluminum anodized 

Aluminum painted 

15.5 m/s 

(35 mph) 

' 1 to 4 ms 

58 m/s 

(1 30 mph) 

'2.0 ms 

44.8 ms 

3~p to 20 ms 

103 m/s 

(230 mph) 

1.0 ms 

'2.6 ms 
6.6.3 Current Amplitudes 

A wide range of current amplitudes appears in typical lightning flashes. 

These are discussed in Chapter 1; Figure 1.1, for example, shows that most of 

the charge in a flash is delivered by continuing currents at a rate of about 200 A 

and by intermediate currents at a rate of about 5 kA. The higher amplitude 

strokes are usually of too short a duration to deliver appreciable charge. 

Therefore, emphasis should be pIaced on the intermediate currents, which can 

occur to surfaces in Zone 2A and which can deliver large amounts of charge at 

appreciable current amplitudes. Section 4.4.2.3 of the SAE report specifies that 

current components B, C, and D should be applied in Zone 2A, with D applied 

first, since it is a restrike, which would be expected to create a new dwell point. 

Components B and C follow for whatever dwell time is expected--or for 50 ms 

if the dwell time is unknown and has not been determined from swept-stroke 

tests. This current is shown in Figure 6.22. If a dwell time of less than 5 ms is 

expected, an average current of 2 kA should be applied for the actual dwell time 

only. 

B AVG 2 kA DELIVERING 10 C IN 5 ms 

,/ 

look* 

/;= C 400 A, DELIVERING 0.4 oms 

I 

I 

f 

I 

I 8 I I I 

0 5 10 15 20 25 30 35 40 45 50 

time - MILLISECONDS 

Figure 6.22 Current and charge expected at a Zone 2A dwell point. 

Delivers 10 C in first 5 ms at any dwell point 

Delivers .04 C/ms thereafter (From Figure 3.3 SAE report) 

Delivers 400 A for all 50 rns if dwell time not known (Drawn 

with straight lines for explanation purposes only) 

The current amplitudes of Figure 6.22 are based on the SAE study of 

natural lightning data and in-flight damage reported by aircraft operators over 

the years. As such, these current amplitudes are a good representation of what to 

expect in a severe natural lightning flash, the type exceeded only 1 or 2% of the 

time. It is recognized, however, that more severe currents, like those which 

sometimes occur in a positive polarity, cloud-to-ground flash, could appear at 

Zone 2A dwell points. An example of such flashes, recorded by Berger 

(Reference 6.48), is shown in Figure 6.23. 

The current measurement was made at an instrument tower on the ground, 

and the long, upward-moving leader which extended from this tower to the 

cloud base accounted for the (approximate) 11.5 ms which elapsed before the 
Figure 6.23 Positive flash of type 4. 

Above: photograph on fast-moving film 

Below: oscillogram of stroke current 

return stroke appeared at the ground. Not all of this delay, therefore, would 

have occurred had the measurement been made on an aircraft in flight 

intercepted by this flash. But the several milliseconds which still would have 

elapsed between initial leader attachment to a forward extremity of the aircraft 

and the occurrence of the return stroke may have allowed the arc to move over 

Zone 2A. The positive polarity stroke of Figure 6.23 would have delivered about 

20 C to this dwell point during the minimum of 2 ms of time it would have 

dwelled there. According to the data of Figure 6.17 meltthrough of a 0.1 6 cm 

(0.063 in) aluminum skin is possible under these conditions. There are reports of 

holes melted in 0.102 cm (0.040 in) skins from in-flight strikes in Zone 2A areas, 

but holes in 0.16 cm aluminum in Zone 2A have not been reported, a fact that 

demonstrates the apparent rarity of the high-energy positive polarity flash in 

Zone 2A areas. Because of the lack of positive evidence that these flashes have 

struck aircraft in critical zones, they have not been fully accounted for in 

present test current waveforms. 

6.6.4 Determination of Skin-Thickness Requirements 

With the establishment of methods to predict the lightning current 

amplitude and charge, as well as dwell time, sufficient information is available to 

utilize the charts of Figures 6.1 7, 6.18 or Tables 6.2 and 6.3 to determine the 

skin thicknesses which will or will not be melted through. 

For example, let us assume that a bare aluminum skin is planned for an 

integral tank in Zone 2A, and that it is desired to determine how thick this skin 

should be to prevent meltthrough. Further, let us say this aircraft will fly at 
velocities as low as 58 m/s (130 mph). From Table 6.5 the expected dwell time 

for this unpainted skin would be 2 ms. From Figure 6.22, an average of 2 kA 

would flow into the dwell point during this period, delivering 4 C of charge. 

According to Table 6.17 these parameters intersect at a point about half-way 

between the coulomb ignition threshold curves for 0.051 cm (0.020 in) and 

0.102 cm (0.040 in) aluminum skins, indicating that 0.102 cm is the thinnest 

skin that should be considered. Since there would be little margin of safety if a 

skin of this thickness were actually used, it would be prudent to select instead 

the next larger thickness-0.16 cm (0.064 in). 

If one notes that the airworthiness certification criteria of Table 5.6 

generally call for an even thicker skin--0.203 cm (0.080 in)--in lightningstrike zones, one must also remember that these criteria make no distinction 

between painted and unpainted surfaces, and that dwell times on painted 

surfaces may be considerably longer than the 2 ms experienced on bare 

aluminum skins. 

Different dwell points, different coatings, and different thicknesses will 

cause correspondingly different dwell times. Consider, for example, the longest 

dwell time, 20 ms, recorded by Robb, Stahmann, and Newrnan (Reference 6.47) 

for a painted surface. During this period the current of Figure 6.21 would deliver 

16 C. These parameters intersect at a point just above the 0.229 cm (0.090 in) 

curve in Figure 6.17, indicating that even the 0.203 cm (0.080 in) thickness 

required by the certificating agencies is insufficient to prevent ignition where 

certain paints are used. 

Therefore, if paints must be used, it is advisable to perform swept-stroke 

tests to establish the actual dwell time. The procurement or certificating agencies 

have generally been willing to accept test data in support of an alternate skin 

thickness to the aforementioned 0.203 cm (0.080 in) thickness, and such tests 

may lead to the use of either thinner or thicker skins for ensuring adequate 

protection. 

6.6.4.1 Titanium and Stainless Steel Skins 

By using the chart of Figure 6.1 8, it is possible to determine titanium skin 

thicknesses in a manner similar to that for determining aluminum skin 

thicknesses. The coulomb ignition threshold for titanium occurs when the back 

side of the skin reaches 1320 "C (2400 OF), a temperature sufficient to ignite a 

fuel-air vapor. Because titanium will not melt at this temperature, no hole will be 

formed when the coulomb ignition threshold is reached. 

No charts of the type in Figure 6.17 and 6.1 8 have been generated for 

stainless steel or other metals, since, with the exception of certain space vehicles, 

these materials seldom appear as integral tank skins. Tables 6.2 and 6.3, while 

not accommodating as wide a variation in parameters, provide some guidance. 

Certain other materials, such as fiber-reinforced plastics (composites) 

behave differently from metallic materials under lightning stress. These are 

treated separately in 3ection 6.2 and Chapter 7. 
A comparison of the physical and electrical properties of titanium with 

those of aluminum has been made by Kosvic, Helgeson, and Gerstein (Reference 

6.49) and is presented in Table 6.6. 

Table 6.6 PHYSICAL AND ELECTRICAL PROPERTIES 

OF TITANIUM AND ALUMINUM 

Property Titanium Aluminum 

Density p (g/cm3) 4.5 2.7 

Thermal conductivity k (calls cm OK) 0.040 0.37 

Specific heat Cp (cal/g OK) 0.13 0.23 

Thermal diffusivity a (cm2/s) = k/p Cp 0.07 0.60 

Electrical Resistivity 01 CL-cm) 199 6.3 

Fusion point (OK) 1950 930 

Heat of fusion (callg) 77 93 

Vaporization point (OK) 3550 2750 

Heat of vaporization AHva, (caljg) 2140 2580 

As noted in this table, titanium has a markedly lower thermal conductivity 

and higher resistivity than has aluminum. These properties have given rise to the 

belief that the backside of titanium skins would reach the coulomb ignition 

threshold with a lower lightning current amplitude and charge transfer than 

those necessary for aluminum, raising the possibility that titanium skins might 

have to be thicker than aluminum skins to provide the same degree of lightning 

protection. But, at the same time, it was evident that structural requirements 

with titanium could be met with skins thinner than the 0.203 cm (0.080 in) 

which was necessary to meet strength requirements' with aluminum. The 

question of lightning effects on titanium skins has also been of concern because 

this material is used in swept- or delta-shaped wings in some supersonic aircraft, 

and the swept-stroke zones may extend over internal fuel tanks in such shapes. 

Therefore, several in-depth studies of titanium behavior under simulated 

lightning arcs have been conducted. One of the most important of these is the 

work of Kofoid (Reference 6.50), who conducted both experimental and 

analytical studies of the lightning current amplitude and dwell times necessary to 

reach the ignition temperature (coulomb ignition threshold) on the inside 

surfaces of titanium skins. Kofoid's experimental results closely follow an 

empirical heat diffusion equation: 

where J = lightning current amplitude (A) (assumed constant) 

d = skin thickness (cm) 

tc = time to reach ignition temperature on inside surface (s) 
The constant (480) is empirical. Curves according to Equation 6.5 are plotted in 

Figure 6.24, together with test data for simulated lightning continuing and 

intermediate currents of up to about 1900 A. 

The experimental data adhere closely to the curves, and both are also in 

good agreement with the data of Figure 6.18 and Tables 6.2 and 6.3. Figure 6.24 
(Reference 6.51) shows that the higher the lightning current amplitude, the 

shorter the time, tc, required to reach the coulomb ignition threshold, and that 

for thicknesses as great as 0.102 cm (0.040 in), only 5 to 6 C of charge are 

necessary to reach this threshold if the charge is delivered in 10 ms or more by a 

current of 500 A or less. This means that if a titanium skin is proposed for use in 

Zone 2A, a lightning arc should not be permitted to dwell for more than about 5 

ms at any one point on the surface. According to the dwell time data of Table 

6.5, this means, in turn, that a 0.102 cm (0.040 in) skin may not have its outside 

surface painted. Anticorrosion paints or sealants on the inside surface of a 

titanium skin may ease the problem by providing thermal insulation between the 

fuel and the hot spot, but the degree of actual protection thus obtained should 

be determined by laboratory tests before such coatings are judged reliable in a 

design. 

6.6.4.2 Ways to Reduce Dwell Time 

The foregoing discussion makes clear the importance of lightning arc dwell 

time in establishing whether or not ignition from a strike to a certain skin is 

likely to occur. And dwell time is perhaps the only lightning characteristic over 

which the integral fuel tank designer may have any control. The objectives, of 

course, should be to have as short as possible a dwell time at any one spot and to 

spread thz arc attachment among many different spots. A bare metallic external 

finish will best achieve these aims, since most paints and other coatings act to 

concentrate the attachment at more widely separated points for correspondingly 

longer times than does a bare surface. 

If a paint must be used, lightning dwell times may be reduced by making 

the paint partially conductive. Robb and others (Reference 6.52) have 

demonstrated that aluminum powder is effective in increasing the conductivity 

of polyurethane paints, thereby increasing the ability of the arc to reattach to 

new points as the aircraft surface moves beneath the arc. Since no parametric 

data relating dwell time to amount of additive are yet available, it is advisable to 

determine in the laboratory the degree of improvement afforded by particular 

combinations. 

6.6.5 Integral Fuel Tanks in Trailing Edges 

If an integral tank skin is the aft-most trailing edge of an aircraft wing tip 

or other structure (Zones IB, 2B), metal skins of almost any thickness will be 

melted completely through because the flash is likely to hang on there for most 

of its life. This conclusion derives from the data presented; the application 

follows. Figure 1.25 shows that 99% of all flashes deliver 200 C or less, and SAE 

current component C reflects this in its requirement for transfer of 200 C in 1 s 

or less. Figure 6.1 7 shows that this much charge would melt through aluminum 

skins up to 0.8 cm (0.3 13 in) thick, and Figure 6.18 shows that it would 

produce the fuel ignition temperature of 1320 OC (2400 OF) on the inside 

surface of titanium skins over 0.318 cm (0.125 in) thick. Therefore, because of 

153 
the excessive flash hang-on times, it is not safe to have fuel in direct contact with 

trailing edge skins of practical thicknesses. If it is desired to have fuel in trailing 

edges, an extended, solid section capable of sustaining up to about 8 cm3 of 

melting at an attachment point about 2.5 cm wide should be provided, as shown, 

for example, in Figure 6.25. 

Externally mounted or tip tanks of the cylindrical variety may be 

protected by closing out the aft-most volume of the tank (as shown in Figure 

6.7,6), so that if meltthrough occurs no fuel vapor will be contacted. 

LIGHTNING TRAILING 

METAL SKINS 

FASTENERS 

Figure 6.25 Trailing edge construction to avoid fuel ignition as a result of 

extended flash hangon. 

Volume of at least 8 cm2 available for melting away 

Fuel or vapor does not contact hot arc or 1320 OC hot spot 

LIGHTNING ZONE 2A 

FROM LEAKING TO VOID AREA 

ZONE 

1 A 

Figure 6.26 Trailing edge construction to avoid fuel ignition from extended 

flash hang-on. 

Hole melts into aft void area, but no vapor is present to ignite. 

Void sections removable for replacement 

Forward void area not as necessary because leading edge dwell 

times are short 

6.7 Integral Fuel Tanks with Nonmetallic Skins 

In some cases nonmetallic materials such as fiberglass-reinforced plastics 

(FRP) are used instead of metals for integral fuel tank skins. For example, the 

light aircraft of Figure 4.15 has FRP wing tips which include fuel tanks 

containing aviation gasoline; these tanks have suffered several in-flight explosions 

154 
when lightning strikes occurred. The fact that the fuel in these tanks is in direct 

contact with the inside surface of the tank skin and the fact that portions of this 

skin are located in Zones 1 A and 2A undoubtedly accounts for these explosions. 

The basic differences between the lightning effects which may occur to 

metallic and nonmetallic skins result from the fact that nonmetallic skins are not 

electrically conductive. This means that the meltthrough or hot spot ignition 

mechanisms associated with attachment of the lightning arc to the skin will not 

occur. Instead, the lightning flash will either 

Puncture the skin and attach to an electrically conductive object 

inside the skin (tank), path (a) on Figure 6.27, or 

Divert around the nonmetallic skin and attach to an adjacent 

metallic skin or other object, path (b). 

i 

METAL 

WING TIP 

Figure 6.2 7 Possible lightning-flash attachment to nonconducting skins. 

If the flash reaches the skin but remains on its outside surface, path (b), it 

is not likely that the skin will be burned through or that its inside surface will 

become hot enough to ignite fuel. The reason for this is that none of the current 

is being electrically conducted into the skin, and the hot arc does not lie close 

enough to or long enough against the skin to burn through it. In fact, its effects 

are usually limited to burning of external paints, if present, or to slight singeing 

of the nonmetallic skin. 

The occurrence of punctures may be explained by the basic lightning 

attachment mechanism described in Chapter 2. An approaching leader induces 

streamers from conducting extremities on the aircraft--including objects 

located beneath nonconducting skins. These streamers are often restricted by the 

insulating skin, but if the electric field producing them is strong enough, the skin 

155 
may suffer dielectric breakdown (puncture) and allow the streamer to reach the 

approaching leader. This is most likely to occur when there is no external 

conducting object nearby from which another streamer could propagate and 

reach the leader first. Obviously, if a puncture occurs through an integral fuel 

tank wali, fuel is placed in direct contact with the lightning arc, and ignition is 

probable. 

It is believed possible that, even if a puncture does not occur, an internal 

streamer can ignite fuel by itself. Protection against ignition therefore requires 

that two basic criteria be met: 

1. Punctures be prevented 

2. Internal streamers be prevented. 

These criteria are, of course, in addition to other requirements such as 

elimination of ignition from strikes to vent outlets, or sparking across joints 

discussed elsewhere in this chapter. 

Puncture of nonmetallic skins can be prevented, or at least greatly 

minimized, by placement of conductive diverters on the outside surface of the 

tank. These diverters must be placed close enough to each other or to other 

conductors to prevent an internal streamer from puncturing the skin. In 

theory this means that the diverters must be close enough together to greatly 

reduce the internal electric field and prevent internal streamer formation in the 

first place, because such streamers, once formed, will intensify the field about 

themselves as they propagate outward. Prevention of internal streamers by 

diverters is called electrostatic shielding, and successful accomplishment of this 

function is necessary to prevent skin punctures and, equally important, to 

prevent internal streamers from forming and becoming a source of ignition 

themselves. Types of diverters which have proved successful and criteria for 

placement and spacing of them are given in Chapter 7 (Structures Protection). 

Tests which may be used to verify diverter adequacy are also described in 

Chapter 7. 

Because streamers form most readily from sharp conducting edges and 

corners, the number of metallic parts inside a nonmetallic fuel tank should be 

minimized wherever possible, and those that remain should be located as far 

from nonmetallic skins as possible and designed with smooth, rounded edges 

instead of sharp points. These concepts are illustrated in Figures 6.28 and 6.29. 

Also metallic parts should be located as far from nonmetallic skins as possible. 

If a diverter protection system such as that shown in Figure 6.28 is used, 

the diverter should not be painted or covered with any other nonconducting 

material. The electrical insulation provided by most aircraft paints is nearly as 

good as that provided by the fiberglass skin itself, so the paint will mask the 

diverter and greatly reduce its effectiveness. Painting of the fiberglass skin itself, 

of course, is to be encouraged, as the added insulation provided by the ~aint will 

further reduce the probability of skin puncture. 

Plumer (Reference 6.53) has successfully tested a diverter arrangement of 

the configuration shown in Figure 6.27, and similar arrangements have been 

flown on some light aircraft with nonmetallic wing tip tanks. 
FUEL QUANTlN UNIT 

MOUNT AS FAR INBOARD 

AS POSSIBLE 

NAVIGATION LIGHT 

BOND TO DlVERTER STRIP 

RUN POWER WIRES INSIDE METAL 

CONDUIT WHICH IS SOLIDLY 

BONDED TO INBOARD WALL 

AND TO LAMP HOUSING 

FUEL OR VENT LINE 

NONMETALLIC AND OUTLET ASSEMBLY 

WING TIP PERHAPS MAKE THIS OF 

NONMETALLIC MATERIAL 

IF IT MUST BE METALLIC, 

KEEP AS FAR INBOARD AS 

POSSIBLE 

KEEP AS FAR FROM SKINS 

AS POSSIBLE 

Figure 6.28 Location of metallic parts within nonmetallic fuel tanks. 

6.7.1 0 ther Methods of P r o t e c t i n g N o n m e t a l l i c Tanks 

If a diverter arrangement is not practical for some reason, a conductive 

coating may be applied to the tank instead. Such a coating may be a metal foil 

or a paint heavily doped with metal or carbon particles. The major disadvantage 

of these coatings is that a relatively large portion of the coating may be melted 

or burned away when the flash attaches to it. A coating would have to be 

prohibitively thick (one approaching the thickness of conventional metallic skins 

themselves) to avoid these results. 

On the other hand, a conductive coating has the important advantage of 

providing an overall electrostatic shield that will virtually eliminate internal 

electric fields and streamering. Thus, if metallic objects must be located inside 

the tank in such a manner that streamers are of concern, it may be advisable to 

consider a conductive coating over the entire nonmetallic skin. This precaution is 

particularly appropriate if occasional burnoff of part of the coating is acceptable 

from a maintenance standpoint. 

Coating materials which have been tested for such applications include 

aluminum foils 5- to 15 mm thick, braided and woven aluminum wire meshes of 

between 40 and 80 wires per centimeter, and various conducting paints. These 

coatings are discussed more thoroughly in Chapter 7. 

6.8 Current Flow Through Structures and Components 

In each of the aircraft accidents that were attributed to lightning and that 

occurred as a result of in-flight explosion of fuel, the exact cause of fuel ignition 

has remained obscure. Ignition at fuel vent outlets or by meltthrough of integral 

tank skins has been suspected in several cases, but no conclusive evidence to this 

effect has ever been found. While it is not correct to eliminate vents and skins 

from further consideration as the prime factors, it may be possible that minute 
PROTRUDING 

WlRE 

AVOID J 

FLOAT 

TERMINATE 

WlRE INSIDE 

FLOAT OR 

USE 

NONMETALLIC 

FLOAT-TYPE FUEL TANK SENDER UNIT 

SHARP 

PROTRUDING 

EDGES 

g j 

EXPOSED ELECTRICAL TERMINAL BOARD 

SHARP 

EDGES 

ROUNDED 

EDGES 

(FUEL OR VENT LINE) 

Figure 6.29 Avoid sharp edges on metallic parts inside nonmetallic fuel tanks. 
sparks resulting from lightning current having been conducted through the fuel 

tanks caused some of these disasters. Whereas the relatively large amount of 

thermal energy released at the points of concentrated lightning current entry 

into the aircraft would be expected to leave tell-tale evidence, the marks left by 

a several-millijoule spark are very difficult, if not impossible, to detect. 

In the past, much attention has been given to keeping lightning currents 

out of the interior of the aircraft by providing conductive and tightly bonded 

skins. While most lightning currents, by virtue of their short duration, do not 

have sufficient time to diffuse completely to interior structural elements such as 

spars and ribs, it is not correct to assume that no current flows in these interior 

structures. Nor is it correct to assume that none of the lightning current flows in 

other internal conductors such as fuel and vent lines. Since so little energy is 

needed for an igniting spark, the behavior of these internal currents is of great 

importance to fuel system safety. Of course, when the lightning currents which 

do flow in the skin encounter discontinuities at access doors, filler caps, and the 

like, a possibility of sparking exists if electrical bonding is inadequate; therefore, 

the behavior of lightning currents remaining in the skins-is important. 

Electrical wires entering fuel tanks from other locations in the aircraft 

pose another potential problem if they are routed in areas which may be reached 

by high electromagnetic fields when lightning strikes occur. 

Figure 6.30 shows the possible lightning current paths in a typical fuel 

tank and calls attention to these areas of greatest concern. In the following 

paragraphs we discuss each of these and present guidelines for protection design, 

referring to related research or testing experience that exists. We also call 

attention to other areas of concern for which little or no quantitative data is yet 

available. 

6.8.1 Filler Caps and Access Doors 

The need to provide a liquid-tight seal around fuel tank filler caps and 

access doors has led to use of various gaskets and seals between the cap or door 

and its mating surface in the tank. Most of these have little or no electrical 

conductivity, leaving the metallic screws or other fasteners as the only 

conducting paths into the cover. If these fasteners are inadequate or if they 

present too much inductance, lightning currents may build sufficient voltage 

along such paths to spark across the nonconducting seals, creating shorter paths. 

Some of these sparks may occur at the inside surfaces of the joints and be a 

source of fuel ignition. 

Newman, Robb, and Stahmann were among the first-to recognize and 

evaluate this possibility in the laboratory. They demonstrated (Reference 6.54) 

that direct strikes to filler caps or access doors of (then) conventional design 

would cause profuse spark showers inside the tank--spark showers of ample 

energy to ignite fuel. They applied simulated strokes to typical fuel filler caps 

and access doors. Their test strokes ranged in energy from a very mild 35 kA, 

0.006 x 1 o6 A2-s stroke to a very severe 180 kA, 3 x I o6 A2-s stroke; profuse 

sparking occurred under all conditions. When an ignitable fuel mixture was 
DRY ~LECTRICAL APPARATUS 

AND WIRING 

Figure 6.30 Lightning current paths in a fuel tank and potential probIem areas. 

placed in the tanks, these spark showers readily produced ignition. 

Newman and his colleagues did not report sparking or fuel ignition when 

the same access doors or filler caps were not directly struck but were 

nevertheless located in a tank skin, which instead was conducting the same 

simulated lightning currents distributed throughout its surface. Nor has sparking 

under such a condition been reported elsewhere in the literature. The possibility 

of such an occurrence, however, has perhaps not been evaluated as thoroughly as 

it should be. 

6.8.1 .1 Lightning-Pro tected Filler Caps 

After observing the sparking from direct strikes to the original access doors 

and filler caps, Newman evaluated several design modifications to prevent this 

sparking. Figure 6.3 1 shows the locations of sparking at the inside surfaces of an 

unprotected filler cap and a modified design intended 'to prevent internal 

sparking. It incorporates a plastic insert that precludes any sparking at interior 
faying surfaces and replaces the previous ball chain with a nonconducting plastic 

strap. Newman successfully tested a mockup of this modified cap with direct 

strikes (Reference 6.55). Since then lightning-protected fuel tank filler caps 

embodying this principle have become commercially available and are being used 

in many aircraft where filler caps must be located in direct or swept-lightningstrike zones (Zones IA, lB, 2A, 2B), and at least one specification, 

MIL-C-38373A (ASG), has been written prescribing a lightning-protected cap. 

LBALL CHAIN 

FUEL 7 

---- - 

FUEL PLASTIC STRAP 

Figure 6.31 Fuel filler cap designs. 

(a) Unprotected 

(b) Protected 

161 
Lightning-protected caps should always be used if there is any possibility 

that the cap may receive a lightning strike. The consequences of not doing so 

have been demonstrated by Plumer (Reference 6.56) in tests of a light aircraft 

nonmetallic wing tip fuel tank equipped with an automotive-type fuel filler cap 

located in a region where it would be subject to direct and swept strokes. The 

cap sparked to its grounded adapter even when struck by a very low-amplitude 

(-5 kA) stroke, igniting a flammable mixture of AVGAS and air. The resulting 

flame exhausted through a blowout hole that had been cut in the underside of 

the tank, as shown in Figure 6.32. The same type of cap on the same type of tip 

tank was also struck during the in-flight incident described in Section 4.5 (Figure 

4.16); however, no ignition occurred, presumably because the tank contained 

sufficient fuel to keep the ullage vapor overly rich, as would be predicted by the 

AVGAS flammability envelope limits of Figure 6.3. 

Figure 6.32 Direct strike to unprotected fuel filler cap resulting in internal 

sparking and ignition of fuel. 

The flammability limits for a particular fuel should not be relied upon, of 

course, to establish adequate lightning protection because there are factors-- 

such as fuel quantity, misting, and venting--which may cause the actual 

conditions prevailing in a particular tank to deviate from the limits shown in 

Figure 6.2 and Figure 6.3. Major emphasis should always be placed upon 
eliminating possible sources of ignition, regardless of the expected flammability 

conditions in the tank. 

If any doubt about the protection capability of the particular design 

exists, the cap should be tested as described in the SAE report Lightning Test 

Waveforms and Techniques for Aerospace Vehicles and Hardware. (Reference 

6.57). 

6.8.1.2 Ligh tning-Pro tec ted Access Doors 

Newman also reported sparking at access door interfaces when the mating 

surfaces had O-ring seals or phenolic or anodized clamp ring surfaces which were 

electrically insulating. It was found (Reference 6.29) that the application of a 

Boeing Service Bulletin modification (SB-1955) successfully prevented this 

sparking when the access doors were directly struck. The modification consisted 

of the following: 

(1) Removing phenolic and anodized coatings from the clamp ring 

(2) Removing anodized coating from faying surfaces of the access door 

(3) Sealing of the door to the surrounding skin adapter with conductive 

grease consisting of 35% by weight of aluminum powder and 65% by 

weight of Aeroshell 14 grease. 

Other modifications in access door seal design have proved equally 

effective in preventing sparking. In some cases anodized clamp rings or insulating 

gaskets have been acceptable when sufficient metal-to-metal conductivity exists 

via the bolts or fasteners alone. The door requiring the SB-1955 modification 

described above to resist sparking had only 23 bolts, whereas other doors having 

over 40 bolts have been found to resist sparking even if insulating finishes or 

paints such as anodize or zinc chromate remain on the mating surfaces. 

A recent survey of different manufacturers' access door designs showed a 

wide range of different materials and sealing practices in use and relatively little 

documented lightning test data upon which to judge their adequacy. If access 

doors are located in Zones 1 or 2, where direct or swept strokes may attach 

directly to them, they should be tested with direct strokes in accordance with 

the SAE report Lightning Test Waveforms and Techniques for Aerospace 

Vehicles and Hardware (See Reference 6.57). Even if these doors are not located 

in direct or swept-stroke zones, they are likely to be involved in conducting 

some lightning current, since this current passes along the aircraft surface 

between other attachment points. This report also describes appropriate tests to 

be performed (for Zone 3) to evaluate this possibility. 

The importance of adequate electrical conductivity between access doors 

and their surroundings can not be overemphasized, especially in cases where the 

access door is itself a large part of the aircraft skin and where this skin encloses 

fuel. Such a case is illustrated in the small fighter aircraft of Figure 6.33, in 

which several large access doors cover a fuselage fuel tank. The doors cover a 

large area and are located in Zone 2A, a zone that requires them to be designed 

to safely accept strokes up to 100 kA, 0.25 x lo6 A~-s. 

163 
Figure 6.33 Large access doors in Zone 2A over fuselage fuel. 

Swept-lightning strokes will attach to these doors. 

Adequate electrical conductivity and protection can not be verified by 

measuring the dc resistance between the door and the surrounding airframe 

because this resistance may change under high-current stress, permitting a 

voltage rise, and also because the inductance of current paths through fasteners 

is equally capable of causing the voltage rises that cause sparking. Structural 

bonds that do not spark have often had dc resistances below 2.5 ma, but there 

is no universal correlation between this resistance and the degree of sparking. It 

is, in fact, quite possible that a joint with less than 2.5 mR of resistance may 

nevertheless have enough inductance to cause a spark. 

Some guidelines to follow in designing a protected door include the 

following: 

1. Provide as much bare-metal-to-bare-metal contact as possible. 

2. If bare-metal-to-bare-metal contact at the mating or faying surfaces 

is not practical because of corrosion problems, provide as much bare 

metal contact via screws or fasteners as possible, and make the 

current paths through these fasteners as short as possible. 

3. Seal door with conductive greases or conductive gaskets if possible. 

The adequacy of an3 design should be verified by performing the 

simulated lightning tests described in the SAE report Lightning Test Waveforms 

and Techniques for A erospace Vehicles and Hardware. 
6.8.2 Joints Between Skins, Spars, and Ribs 

The large numbers of rivets or fasteners typically employed to attach a 

metal skin to a metal rib, spar, stiffener, or other internal structural element 

have usually also provided adequate electrical conductivity for lightning currents. 

The most common situations found are those where both mating surfaces are 

painted with an insulating zinc chromate paint and sometimes further insulated 

with environmental sealers, as shown in Figure 6.34. There are, however, usually 

enough rivets or fasteners which make bare metal contact with the mating 

surfaces to provide the necessary electrical conductivity for lightning currents. 

High-density patterns of rivets or fasteners, such as are commonly utilized to 

join fuel tank skins to stringers, ribs, and spars, have been found to be capable of 

conducting 200 kA stroke currents, even when nonconductive primers and 

sealants are present between the surfaces, as in Figure 6.34 (Reference 6.58). 

Short, for example, reports 200 kA stroke tests of skin-to-stringer joint samples 

fastened with double rows of taper-lock fasteners. The sample tgsted had 16 

fasteners on each end of the joint, and no sparking was detected anywhere on 

the joint (Reference 6.59). 

CONTACT BETWEEN 

EDGES OF FASTENER 

CURRENT 

PATH 

/ 

AND STRUCTUR 

NONCONDUCTIVE PRIMER 

AND/OR SEALANT 

Figure 6.34 Bonding through mechanical fasteners. 

165 
There is no hard-and-fast rule for the number of fasteners per meter of 

joint (density) which are necessary to avoid sparking. After all, a single No. 8 

AWG copper or aluminum wire is capable of conducting 200 kA, 2 x lo6 A~-S 

strokes, so a single fastener of this or larger diameter would also, by itself, be 

capable of conducting this much current. Any sparking usually occurs at the 

interfaces between the fastener and surrounding metal, and the occurrence of 

such sparking depends on other physical characteristics, such as skin metal 

thickness, surface coatings, and fastener tightness. Tests in which simulated 

lightning currents are conducted through the joint should always be made on 

samples of joints involving new materials or designs. The type and amount of 

lightning current to be applied for these tests and diagnostics to be used, should 

be as described in Lightning Test Waveforms and Technques for Aerospace 

Vehicles and Hardware for Zone 3. 

The foregoing relates to the case where the joint is in Zone 3 and must 

conduct some portion of the lightning current, but other zones can be, and are 

likely to be, struck directly by the lightning arc itself. In direct and swept-stroke 

zones (1 A, 2A) and hang-on zones (IB, 2B), it is possible for the lightning arc to 

attach directly to a single fastener or rivet. If it persists, the hot arc may melt or 

otherwise damage surrounding skin. Figure 6.35 (Reference 6.60) shows an 

example of the damage that can be done at such a place. Here, a 1 cm-diameter 

rivet head and much of its surrounding aluminum skin have melted away. If an 

arc persists long enough at such a spot, it may melt completely through the skin, 

beyond the inside member. Such penetration occurred at the point shown in 

Figure 6.35. The danger involved may require that, if ignitable fuel vapors can 

exist beneath any joint in a swept-stroke zone (]A, 2A), the skin may have to be 

thicker than that required on the basis of swept-stroke dwell times only; or some 

other provision may have to be made to prevent complete meltthrough near a 

rivet. 

Some other guidelines which should be noted in designing integral tank 

joints are as follows: 

1. Avoid use of anodizing in joints which may have to carry lightning 

currents. Anodizing is usually applied to prevent corrosion and results in a hard 

oxide surface which is nonconductive and difficult to break down at less than 

about 300 V. Thus, when breakdown eventually does occur, the resulting spark 

will be more energetic than if a poorer insulator were present. The best coating 

of all for lightning protection is, of course, no coating. 

2. Design short, direct lightning current flow paths because lightning 

currents want to flow in straight, direct paths between entry and exit points on 

the aircraft. Voltages which may cause sparking will build up wherever diversions 

in these paths exist. The diversions through the fasteners and stringer of Figure 

6.34 are acceptable. More extended paths may not be. 

3. Do not depend on resistance measurements to confirm the adequacy of 

lightning current conductivity in a joint. The inductance of the lightning current 

flow path through the joint plays an equally important part, but this inductance 

is rather small (a few hundredths of a microhenry) and difficult to measure. 

Resistance measurements (ac or dc) may be useful as a production quality 

166 
control tool, but they are not useful for establishing the adequacy of the 

lightning current path through the joint. 

4. Account for aging and mechanical stress which may cause reduced 

electrical conductivity. Continued flexing of structures under flight load 

conditions may eventually loosen a joint to the point where sparking could 

occur. To evaluate this possibility, perform simulated lightning tests on joint 

samples previously given fatigue or environmental tests. 

Figure 6.35 Typical lightning damage to a met. 

Meltthrough to inside beyond rib surface. 

167 
6.8.3 Joints and Interfaces in Pipes and Couplings 

Because they are metallic and come in contact with structure, the myriad 

of fuel, hydraulic and vent pipes, and other plumbing inside a fuel tank will have 

to conduct some, however small, amounts of current when lightning currents are 

flowing through the structure containing them. This current may cause sparking 

at pipe joints and couplings where there is intermittent or poor electrical 

conductivity. Some pipe couplings, for example, are designed to permit relative 

motion between the mating ends of a pipe so as to relieve mechanical stresses 

caused by wing flexure and vibration. Also, anodized or other electrically 

insulating coatings are often applied to the pipe ends and couplings to control 

corrosion. Relative motion and vibration may wear this insulation away, 

providing unintentional and intermittent conductive paths. To date, almost no 

quantitative information has been developed to describe the amounts of current 

or voltage necessary to cause a spark, nor has this possibility been related to the 

design or condition of a particular system. Yet this sparking may have been a 

factor in the cause of the several in-flight explosions associated with lightning 

strikes cited earlier in this book. This aspect should, therefore, be given 

particular attention in the design of aircraft fuel systems. 

6.8.3.1 Lightning Currents in Plumbing 

Because of their relatively short duration, the high-amplitude, shortduration stroke currents will not have time to spread very deeply into interior 

structural elements or other interior conductors. Most of this current will remain 

in metallic skins, but this current does begin immediately to diffuse into 

underlying metallic structure. Appreciable current levels may be reached in 

spars, stringers, pipes, and other internal conductors if the lightning stroke 

persists long enough. 

The basic penetration, or current redistribution, process is described in 

Chapter 8, and it is shown that penetration times for typical aircraft skins are of 

the order of several tens of microseconds. If the stroke current were to remain at 

its peak amplitude for a time longer than this, dc conditions would be 

approached and the current would be distributed resistively, with a significant 

portion of it flowing in the internal structure. Fortunately, the stroke current 

does not remain at its peak level for more than a few microseconds. The amount 

of this stroke current which, under typical lightning-stroke conditions, can 

actually flow on internal pipes is not yet known, and determination of the 

magnitude of this current cannot yet be done confidently by analysis. 

Laboratory tests in which simulated lightning currents are forced through 

complete fuel tank structures can be performed, however, and the current in 

particular pipelines can be measured. 

Intermediate and continuing currents do persist for times long enough for 

the penetration process to be completed, and it is possible to estimate, as 

follows, the amount of intermediate current that may flow in a pipe after dc 

conditions are reached. 
Assume that the leading and trailing edge sections of the wing in Figure 

6.30 are nonconductive or sufficiently isolated as to be unavailable for 

conduction and that the remaining wing box comprised of skins and spars has 

chordwise dimensions as shown in Figure 6.36. The cross-sectional area of the 

spars and skins forming this box is 135 cm2. The tank also contains a 0.05 cm 

I.D., 10 cm O.D. aluminum vent pipe which is electrically bonded to the 

structure at each end of the tank. The cross-sectional area of this tube is 1.57 

cm2. 

Figure 6.36 Hypothetical wing box with integral fuel tank. 

Assuming an intermediate stroke whose average amplitude is 2 000 A for a 

duration of 5 ms (in accordance with component B of the SAE report Lightning 

Test Wavefoms and Techniques for Aerospace Vehicles and Hardware) the 

current in the pipe can be calculated as follows: 

1.57 cm - (2 000 A) = 22.9 A 'tube -137- 

This current would flow for as long as the lightning intermediate current 

flows, which is 5 ms if component B is assumed to be representative. The ability 

of currents of the order of 20 A to cause a spark at pipe interfaces and couplings 

presently used in aircraft is not known, but the sparking that commonly occurs 

at similar current levels between other conductors in relative motion (such as 

battery clamps and motor commutators) raises a distinct possibility of sparking 

at couplings where relative motion and/or poor electrical conductivity exists. 

Electrical bond straps are sometimes installed across poorly conducting 

pipe couplings, as shown in Figure 6.37. 

These bond straps are usually intended to equalize static electrical charges 

such as those carried into the tank during refueling operations. 

169 
ELECTRICAL 

BONDING 

SAFETY WIRE 

COUPLING 

Figure 6.37 Electrical bonding jumper across insulated coupling. 

BONDING JUMPERS SUCH AS THOSE SHOWN IN FIGURE 6.37 

SHOULD NOT BE RELIED UPON TO PREVENT SPARKING FROM 

LIGHTNING CURRENTS! 

The reason that jumpers like this are unsatisfactory is that, when 

fast-changing lightning currents begin to flow in these jumpers, voltage rises 

occur along the inductance of the strap equal to 

where E = inductance voltage (volts) 

L = bond strap inductance (henries) 

I = current in tube (amperes) 

This voltage appears across all other possible current paths through the coupling. 

If there exists another possible but unintentional path, such as a safety wire, a 

retaining chain, or the coupling interface itself, this inductive voltage may be 

sufficient to force current through these paths and cause sparking if the 

conductivity is poor or intermittent. For example, if the 20 A portion of the 

intermediate current were to begin to flow in the bonding jumper of Figure 6.37 

and if this current were to reach its peak in 100 ps, its rate of rise would be 

If the inductance of the bonding jumper is 0.1 pH, then the inductive 

voltage developed along the jumper would be 

By itself this voltage is too low to break down coupling insulation such as 

170 
anodized or zinc chromate primer, but if it is applied across preexisting surfaces 

where these treatments have been worn through, this voltage may be sufficient 

to cause current flow and sparking. 

As stated earlier, the fraction of stroke current which enters a bonding 

jumper and the inductive voltage it produces are not known. It is quite probable 

that at least some stroke currents would produce even higher rates of current rise 

in the bonding jumpers than do the intermediate currents. 

6.8.3.2 Conductivity through Couplings and Interfaces 

Here again is an area where little electrical data is available. The extensive 

use of anodized coatings to provide noncorrosive mating surfaces in pipe 

couplings would seem to preclude arcing across the pipe interface, but relative 

motion between these surfaces can wear through the anodized coating, causing a 

conductive path. If there happens to be bare metal-to-metal contact within the 

coupling, this would probably represent the shortest path, and it is altogether 

possible that no sparking would occur anywhere else in the coupling. A slight 

change in the relative position of the mating surfaces or in some other aspect, 

such as the tightness of its threaded assembly or introduction of dirt or residue, 

may drastically change the electrical conductivity of a coupling. In fact, it is 

probable that the electrical conductivity of a typical pipe coupling changes many 

times during a flight as a result of relative motion caused by structural vibrations 

and flexing. 

In the absence of definitive data on the electrical conductivity of pipe 

couplings under in-service conditions, it is advisable to take the following 

approach: 

1. Determine the fraction of lightning current expected to flow in a 

particular pipe, by analyses similar to the analysis of Section 6.8.3.1, 

or, if available, a more authentic means. 

2. Inject this expected current through a sample of the desired 

coupling, under simulated in-flight vibration and contamination 

conditions. 

3. Perform this test in a darkened enclosure to enable inspection for 

sparks. Repeat the test until a reliable result is established. 

If sparking does occur in a coupling, a redesign is likely to be required for 

assured prevention; however, it may be possible to provide adequate conductivity with the addition of two or more short, wide braided jumpers or a braided 

sleeve, as shown in Figure 6.38(a) and (b). 

Alternatively, completely insulated couplings might be designed. If such an 

insulated coupling were used, a jumper as shown in Figure 6.38(c) could be used 

to conduct lightning currents and static charges, provided the length of the path 

through the jumper is not much longer than the path through the air along the 

surface or through the insulated coupler. 

Before using any one of these or any other protective method that has to 

be added to an existing coupling design, its adequacy should be verified by 

impulse current tests to check for sparking. 
/ COPPER BRAID 

\COPPER BRAID 

SLEEVE COVERS 

ENTIRE COUPLING 

SHORT 

COUPLING CLAMPS CLOSE 

MADE OF TO END OF 

INSULATING CONDUCTING PIPES 

MATERIAL 

Figure 6.38 Possible methods to reduce sparking at pipe couplings. 

6.8.4 Electrical Wiring in Fuel Tanks 

Another result of lightning current flow through a structure is that the 

fields it produces may induce voltages in electrical wiring. If this wiring enters a 

fuel tank, the possibility of these induced voltages being high enough to cause a 

spark in the fuel tank is a matter for concern. 

Electrical wires found inside fuel tanks are typically those that run to 

172 
capacitance-type fuel quantity probes or small electric motors which operate 

pumps or valves. If these wires are totally enclosed by metal skins and ribs or 

spars, the internal magnetic fields and induced voltages will be relatively low. 

Electrical devices, such as fuel probes, have been intentionally designed to 

withstand comparatively higher voltages without sparking. The fuel system 

designer, however, must be continually alert for structural design or materials 

changes that may alter this situation and permit excessive induced voltages to 

appear in fuel tank electrical circuits. 

A number of measurements have been made of the voltages which 

simulated lightning currents may induce in fuel probe wiring found in 

conventional aircraft. Measurements have also been made of the voltages 

required to cause a spark between the elements of typical capacitance-type fuel 

quantity probes. The sparkover voltages have always been found to be much 

higher than those found to be induced in the wiring. Newman, Robb, and 

Stahmann (Reference 6.61) found that at least 3 000 V of dc voltage were 

required to cause a spark between the active and grounded cylinders of a 

capacitance-type fuel probe found in a KC-135 aircraft wing, and that even 

higher voltages were needed to spark over the other gaps in this probe. Plumer, 

(Reference 6.62) ran a similar test using the 1.2 x 50 ,us impulse voltage 

waveform specified as Voltage Waveform C of SAE Report Lightning Test 

Waveforms and Techniques for Aerospace Vehicles and Hardware and found that 

12 kV were required to cause a spark between the inner and outer cylinders of 

the probe. Small gaps like this will withstand more impulse than dc voltage, and 

the impulse test more realistically represents an induced voltage. 

The tests reported above were made at atmospheric pressure on the 

ground; the same gap would break down at a lower voltage when at the lower air 

pressures associated with flight altitudes. The amount of reduction to be 

expected in impulse sparkover voltage at various altitudes can be determined 

from the curves of Figure 6.39(a) and (b) (Reference 6.63). The product, pd, of 

pressure (mm Hg) and gap distance (mm) is first determined by obtaining 

pressure at altitude from the chart of Figure 6.39(c). If it is desired to find the 

sparkover voltage of a 5 mm gap at 10 000 m altitude, for example, the pressure 

at this altitude is first determined from the chart to be 198.16 mm Hg. The 

product, pd, would then be (5 mm) (198.16 mm Hg) = 990.8 mm Hg mm. From 

Figure 6.39(b), the gap would spark over at about 5500 V. 

The sparkover voltages of Figures 6.39(a) and (b) should be considered 

approximate, since these curves were determined for plane parallel electrodes 

only. Other electrode configurations, such as a sharp point-to-plane, would spark 

at voltages perhaps 25% more or less than these. Impulse sparkover tests should 

be made on particular gap configurations to determine the actual sparkover 

voltages. 

Contaminants on the electrode surfaces may also act to reduce the 

breakdown voltage; therefore, it is wise in any case to reduce voltage levels 

measured by laboratory tests by a factor of at least 2 when assigning a withstand 

voltage to a fuel probe or other electrical gap that might appear in a fuel tank. 
Pressure 

mm of Hg 

When determining the breakdown voltages of an object such as a fuel 

quantity probe, care must be taken to apply test voltage to all of the electrode 

combinations which may exist (this need not all be done simultaneously). In 

most cases both the "high" and "low" electrodes in such a device are insulated 

from "ground" (the airframe), as shown in the simplified drawing of Figure 

6.40, but there is always a mounting bracket which brings the probe in 

proximity to the airframe. Thus, the test voltage should be applied across each 

of the basic electrode combinations shown in Figure 6.40. Usually, the gaps to 

the mounting bracket or airframe will be larger than the others and require a 

higher voltage to cause sparkover. Typical impulse sparkover voltages to be 

expected across these gaps, at a flight altitude of 3000 m, might be as shown in 

Table 6.7, where the "high" and "low" electrodes are defined as shown in Figure 

6.40. 
CONDUCTIVE TUBES J 

---------- 

TERMINALS -+ 

Figure 6.40 Possible breakdown of gaps in capacitance-type probe. 

1. High to Low 

2. High to Airframe 

3. Low to Airframe 

Table 6.7 TYPICAL EMPULSE SPARKOVER VOLTAGES 

FOR A CAPACITANCE-TYPE PROBE 

Figure 6.40 illustrates that the proximity of the probe to the airframe 

structure on which it is mounted is very important in determining breakdown 

voltages to the airframe. Because of this, design of adequate insulation between 

the active elements of the probe and the airframe may not be entirely within the 

probe designer's control. 

CARE MUST BE TAKEN BY THE FUEL SYSTEM DESIGNER TO SEE 

THAT ADEQUATE GAPS ARE MAINTAINED BETWEEN THE ACTIVE 

ELECTRODES OF THE PROBE AND THE AIRFRAME. 

It is particularly important that sufficient insulation be provided between 

Figure 6.40 

Electrode Pair 

High to Low 

High to 

Airframe 

Low to 

Airframe 

Sparkover Voltage at 3 000 m Altitude 

3 kV 

6 kV 

8 kV 
active electrodes and the airframe because the highest induced voltages usually 

appear between all of the incoming wires and the airframe. The reason is 

illustrated in Figure 6.41. The wires are most closely referenced to the airframe 

at the electronics end, which is usually in the fuselage. If one or more of the 

wires are shielded, usually only the electronics ends of the shields are grounded 

to the airframe because grounding of both ends permits stray ac fields to induce 

circulating currents in these shields. These currents are a source of steady state 

electromagnetic interference (EMI) sufficient to interfere with the operation of 

the fuel quantity electronics. This EM1 has been avoided by leaving the shields 

ungrounded at the fuel tank end. 

INDUCED 

VOLTAGES 

ARE 

FJENVEEN 

SHIELDS 

AND 

AIRFRAME 

I 

/am 

SHIELDED . 

- 

SHIELDS 

GROUNDED 

ONLY AT 

ELECTRONICS 

1-27 :I / / END 

mc*sAcITANcE-npE PROBE 

WING 

FUEL 

TANK 

Figure 6.41 Typical fuel probe wiring. 

Most magnetic flux and induced voitaie appears between either 

wire (or shield) and the airframe at probe end. 

Less flux and induced voltage exists between any two wires. 

Even less flux and voltage exists between a wire and its shield. 

The performance of shielded circuits within magnetic fields is discussed in 

Chapter 13, Section 13.2, where it is shown that a shield grounded at only one 

end actually provides no magnetic shielding of an enclosed wire whose return 

path (the airframe) is outside of the shield. If a shieId is grounded at only one 

end, the magnetic flux encircled in the loop this shield forms with the airframe 

will induce a voltage in this loop of 

where 4 = magnetic flux 

Most of this induced voltage will appear between the open end of the shield and 

the airframe. The same field will induce a similar voltage between the enclosed 

wire and the airframe. Only if the shield is grounded to the airframe at both ends 

will it prevent these voltages from being induced. If the ungrounded end of the 

shield is in a fuel tank, the highest induced voltages will appear there. 
Induced voltages will also appear between two or more wires or shields at 

the probe end, but any one of these wire-to-wire voltages will usually be less 

than the voltage between either conductor and the airframe because the area 

between wires in a cable bundle is usually less than the area between any one of 

them and the airframe. 

The routing of the fuel probe wires, therefore, has a lot to do with how 

much induced voltage appears at apparatus inside fuel tanks. Since fuel tanks are 

often completely enclosed with metallic walls which are electrically bonded and 

well sealed, the magnetic fields from lightning (or any other source) are likely to 

be lowest inside the tanks. Thus, wires that must eventually enter the tanks 

should be routed through them from the start. Newman, Robb, and Stahmann, 

for example, have measured up to 11 V, or less, in the fuel quantity probe 

circuits routed inside a KC-135 wing fuel tank (Reference 6.64) through which 

50 kA strokes with a 50 kA/ps rate of rise were being conducted. This voltage 

would extrapolate to a maximum of 44 V at a severe stroke amplitude of 200 

kA. Plumer (Reference 6.65) measured up to 4 V in a fuel probe circuit routed 

inside the tanks of an F-89J wing. The simulated lightning current in this case 

was 40 kA, with a rate of rise of 8 kA/ps. From these results Plumer predicted 

that a 200 kA, 100 kA/ps lightning stroke would induce 185 V in the same 

circuit. At the time these measurements were made, most researchers looked for 

induced voltages of the same duration as that of the lightning current and paid 

little attention to other signals, suspected to be "noise," appearing within the 

first microsecond of an induced voltage trace on an oscilloscope. These "noise" 

signals, however, often had higher amplitudes than the more familiar, longer 

duration voltages considered to be authentic. With the advent of better 

instrumentation, these high-frequency voltages became known to be legitimate, 

induced perhaps by the higher frequency traveling wave currents that may 

appear in the airframe as a result of the mechanism described in Chapter 8, 

Section 8.2. Thus, higher voltages than those reported by Newman and Plumer, 

may in fact be induced in fuel quantity probe circuits, although the literature 

shows no recent measurements of induced voltages in fuel tank electrical circuits 

to establish what the amplitudes of these voltages might actually be. 

If it is not possible to route wiring inside the fuel tank, it may be routed 

outside of a tank as pictured in Figures 6.30 and 6.41; but, if this is done, great 

care must be taken to minimize the amount of magnetic flux that may link these 

wires or their shields. This precaution is particularly pertinent for wings where 

wires are routed within leading and trailing edges, which are often exposed to 

higher magnetic field levels than Asewhere. Fiberglass has replaced aluminum in 

some leading edge skins, and, in some aircraft, control surface openings allow 

magnetic flux to penetrate directly to the aft wing spar, along which electrical 

cables are often routed. 

If fuel system wiring must be routed outside of the tanks, there are several 

methods that can be used to control induced voltages, often to levels far below 

those required to cause a spark at a fuel probe. These involve routing 

improvements, grounding of the shields to the airframe at both ends, and use of 
an overall shield grounded at both ends to enclose a bundle of individually 

shielded wires whose shields are left ungrounded at one end for EM1 purposes. A 

relative comparison of six possible wiring methods is shown in Figure 6.42. It is 

recommended that methods a and b of this figure not be used, and that c and d 

be used only if the (leading or trailing edge) enclosure has an electrically 

conductive skin or coating that can provide some electromagnetic shielding 

against the outside world. Methods e and f are preferable because they will 

provide the lowest induced voltages at the probe. Method f is to be used if the 

individual shields must be left ungrounded at one end for EM1 control purposes. 

The comparisons shown in Figure 6.42 apply also for wiring to fuel tanks in 

other parts of the aircraft. 

For whatever method is used, assurance should be obtained that the 

expected induced voltages are lower by a factor of 10 to 1 than the probe 

sparkover voltage at altitude. An estimate of the induced voltages to be expected 

in the fuel tank wiring can be obtained by analysis as described in Chapter 12 or 

by lightning transient analysis tests, as described in Chapter 17. 

6.9 Fuel System Checklist 

A considerable amount of information on various aspects of fuel systems 

has been presented in the preceding paragraphs. With reference to these 

paragraphs, the following checklist is provided to summarize key points and to 

focus the designer's attention on the important aspects of fuel system lightning 

protection: 

1. Are the fuel tanks located in Zones 1 and 2, where lightning strokes 

may directly contact their skins? (Chapter 5, Sections 5.2 and 5.3) 

If so, 

a. What is the expected arc dwell time on these skins? (Section 

6.6) 

b. Are the skins thick enough to avoid meltthrough? (Section 

6.6) 

c. Can a lightning arc attach directly to an access door or filler 

cap?; if so, can the lightning current be conducted into the 

surrounding skin without sparking inside the tank? (Section 

6.8) 

2. Are there "dry bay" areas into which fuel may leak from adjoining 

tanks or plumbing? If so, these are also subject to questions la, lb, 

and 1 c. 

3. Are fuel vent outlets or jettison pipes located in direct strike zones 

(Zones 1 A or 1 B) or swept-stroke zones (Zones 2A or 2B) where the 

arc may attach or sweep close to the vent outlet? (Chapter 5, 

Sections 5.2 and 5.3; Sections 6.4 and 6.5) 

a. If so, is an effective flame arrester or surge tank protection 

(STP) system used?; has its effectiveness been verified by test? 

(Section 6.4) 
amT ( BETTER 

Figure 6.42 Relative comparison of six system wiring methods. 
b. Is the response time of the STP system shorter than the 

possible flame propagation time from the STP sensor to the 

extinguisher? (Section 6.4) 

c. Is the STP system protected from false trips resulting from 

lightning-induced voltages in its electrical wiring or light from 

nearby flashes? 

4. If nonmetallic skins are used, is an adequate diverter or conductive 

coating system that will prevent skin puncture or internal streamering provided? (Section 6.7) 

5. Is the fuel tank structure capable of mnducting Zone 3 lightning 

currents even if the tank itself is not located in a direct or 

swept-stroke zone? Has this been demonstrated by a test in which 

simulated lightning currents are conducted through a complete tank 

structure? (Section 6.8) 

6. If a simulated lightning test of the complete tank assembly is not 

feasible, have all of its individual joints, seams, access doors, filler 

caps, drains, vents, plumbing, and electrical systems been tested for 

their ability to conduct simulated lightning currents and found to be 

free of sparking? (Sections 6.5 and 6.8) 

7. Are electric circuits entering the fuel tanks protected against high 

induced voltages? (Section 6.8) Have they been routed away from 

other wiring, such as navigation lamp circuits, which may be 

susceptible? 

8. Are clearances between exposed electrical parts and each other or 

between any of them and the airframe sufficient to withstand the 

induced voltages that may occur without sparking? (Section 6.8) 

9. Have the electrical circuits inside nonmetallic tanks been adequately 

shielded so that excessive induced voltages will not occur? (Section 

6.8) 

10. Have applicable government airworthiness certification regulations, 

military standards, or other specifications pertaining to lightning 

protection been adhered to? (Section 5.4) 
REFERENCES 

6.1 L. J. Nestor, Investigation of Turbine Fuel Flammability within Aircraft 

Fuel Tanks, Final Report DS-67-7, prepared by the Aeronautical Engine 

Department, Naval Air Propulsion Test Center, for the Engineering and 

Safety Division, Aircraft Development Service, Federal Aviation Administration, Department of Transportation, Washington, D.C. (July 1967). 

6.2 Adapted from Nestor, Investigation of Turbine Fuel Flammability, p. 8. 

6.3 Adapted from Nestor, Investigation of Turbine Fuel Flammability, pp. 12, 

13. 

6.4 W. Q. Brookley, "USAF C-141 and C-135 Fuel Tank Nitrogen Inerting 

Tests," Report of the Second Conference on Fuel System Fire Safety, 6 

and 7 May 1970, Federal Aviation Administration, Department of 

Transportation, Washington, D.C. (1970), p. 75. 

6.5 Adapted from Nestor, Investigation of Turbine Fuel Flammability, p. 23. 

6.6 Brookley, "USAF C-141 and C-135 Fuel," pp. 75-82. 

6.7 Investigations of Mechanisms of Potential Aircraft Fuel Tank Vent Fires 

and Explosions Caused by Atmospheric Electricity, Final Report under 

Contract No. NASr-59, prepared by the Lockheed-California Company, 

for the National Aeronautics and Space Administration, Federal Aviation 

Administration, Department of Transportation, Washington, D.C. (May 

31, 1963). 

6.8 Redrawn from Investigations of Mechanisms of Potential Aircraft Fuel 

Tank Vent Fires, p. 84. 

6.9 Redrawn from Investigations of Mechanisms of Potential Aircraft Fuel 

Tank Vent Fires, p. 90. 

6.10 Redrawn from Investigations of Mechanisms of Potential Aircraft Fuel 

Tank Vent Fires, p. 123. 

6.1 1 Civil Aeronautics Board Aircraft Accident Report, Boeing 70 7-121, 

N709PA, Pan American World Airways, Inc., near Elkton, Maryland, 

December 8, 1963, File No. 1-0015, adopted February 25, 1965. 

6.12 C. C. Bolta, R. Friedman, G. M. Griner, M. Markels, Jr., M. W. Tobriner, 

and G. von Elbe, Lightning Protection Measures for Aircraft Fuel Systems, 

Phase 11, Technical Report ADS-18, prepared by the Atlantic Research 

Corporation with the Lightning and Transients Research Institute for the 

Federal Aviation Agency, U.S. Department of Commerce, Washington, 

D.C. (May 1964). 

6.13 Bolta et al, Lightning Protection Measures, pp. 39-43. 

6.14 Based on BoIta et al, Lightning Protection Measures, pp. 22 and B-3. 

6.15 Bolta et al, Lightning Protection Measures, pp. 70-79 and C-1 to C-3 

(excerpted from Proposal No. PS-139, Fenwal, Inc., Ashland, 

Massachusetts [January 23, 19641). 

6.16 Bolta et al, Lightning Protection Measures, pp. 1 16-149. 

6.17 M. M. Newman, J. R. Stahmann, and J. D. Robb, Airflow Velacity Effects 

on Lightning Ignition of Aircraft Fuel Vent Efflux, FAA Final Report 

DS-67-9, prepared by the Lightning and Transients Research Institute, 
with Consultant Staff Cooperation of the Atlantic Research Corporation, 

for the Federal Aviation Administration, Department of Transportation, 

Washington, D.C. (July 1967). 

6.18 Newman, Stahmann, and Robb, Airflow Velocity Effects, pp. 1@20. 

6.19 Redrawn from F. L. Kester, M. Gerstein, and J. A. Plumer, A Study of 

Aircraft Fire Hazards Related to Natural Electrical Phenomena, NASA 

CR-1076, prepared by Dynamic Science for the National Aeronautics and 

Space Administration, Department of Transportation, Washington, D.C. 

(June 1968), p. 75. 

6.20 J. P. Gillis, Study of Flame Propagation through Aircraft Vent Systems, 

Final Report NA-69-32, prepared by Fenwal, Inc., Walter Kidde and Co., 

for the National Aviation Facilities Experimental Center, Federal Aviation 

Administration, Department of Transportation, Atlantic City, New Jersey 

(May 1969). 

6.2 1 Redrawn from Gillis, Study of Flame Propagation, pp. 2 I, 22. 

6.22 Gillis, Study of Flame Propagation, p. 23. 

6.23 Gillis, Study of Flame Propagation, p. 35. 

6.24 Federal Aviation Agency, Washington, D.C. (October 6, 1967). 

6.25 Society of Automotive Engineers, Warrendale, Pennsylvania, (5 May 

1976). 

6.26 J. H. Hagenguth, "Lightning Stroke Damage to Aircraft," AIEE Transactions 68, American Institute of Electrical Engineers, New York, New 

York (1949): 1-1 1. 

6.27 Hagenguth, "Lightning Stroke Damage," p. 2. 

6.28 Redrawn from Hagenguth, "Lightning Stroke Damage," p. 2. 

6.29 J. D. Robb, E. L. Hill, M. M. Newman, and J. R. Stahmann, Lightning 

Hazards to Aircraft Fuel Tanks, Technical Note 4326, National Advisory 

Committee for Aeronautics, Washington, D.C. (September 1958), pp. 

12-1 4. 

6.30 Lightning Test Waveforms and Techniques for Aerospace Vehicles and 

Hardwarp, Society of Automotive Engineers, Warrendale, Pennsylvania (5 

May 1976). 

6.3 1 Kester, Gerstein, and Plumer, A Study of Aircraft Fire Hazards, p. 40. 

6.32 Robb, et al, Lightning Hazards, p. 9. 

6.33 Society of Automotive Engineers-designated zones. 

6.34 Kester, Gerstein, and Plumer, A Study of Aircraft Fire Hazards, pp. 5 1-69. 

6.35 R. 0. Brick, "A Method for Establishing Lightning-ResistanceiskinThickness Requirements for Aircraft," Lightning and Static Electricity 

Conference, 3-5 December 1968, Purt 11: Conference Papers, AFA L-TR68-290, Air Force Avionics Laboratory, Air Force Systems Command, 

Wright-Patterson Air Force Base, Ohio (May 1969), pp. 295-3 17. 

6.36 L. L. Oh and S. D. Schneider, "Lightning Strike Performance of Thin 

MetaI Skin," Session 111: Structures and Materials, Proceedings of the 19 75 

Conference on Lightning and Static Electricity, 14-1 7 April 1975, at 
Culham Laboratory, England, the Royal Aeronautical Society of London, 

(1975). 

6.3 7 Redrawn from Oh and Schneider, Lightning Strike Performance, p. 12. 

6.38 Kester, Gerstein, and Plumer, A Study of Aircraft Fire Hazards, p. 39. 

6.39 Redrawn from Oh and Schneider, Lightning Strike Performance, p. 12. 

6.40 Kester, Gerstein, and Plumer, A Study of Aircraft Fire Hazards, pp. 55-68. 

6.41 Kester, Gerstein, and Plumer, A Study of Aircraft Fire Hazards, pp. 57,6l. 

6.42 J. A. Plumer, Guidelines for Lightning Protection of General Aviation 

Aircraft, FAA-RD-73-98, Federal Aviation Administration, Department of 

Transportation, Washington, D.C. (October 1973), p. 11. Data sources 

include J. Phillpott, "Simulation of Lightning Currents in Relation to 

Measured Parameters of Natural Lightning," Proceedings of the 1975 

Conference on Lightning and Static Electricity, 14-1 7 April 19 75, at 

Culham Laboratory, England, Session I: Fundamental Aspects and Test 

Criteria, the Royal Aeronautical Society of London, &gland (1975); and 

Brick, "A Method for Establishing Lightning-Resistance/Skin-Thickness 

Requirements." 

6.43 Air Force Cambridge Laboratory Photograph from Rough Rider Program. 

6.44 R. 0. Brick, L. L. Oh, and S. D. Schneider, "The Effects of Lightning 

Attachment Phenomena on Aircraft Design," 1970 Lightning and Static 

Electricity Conference, 9-11 December, sponsored jointly by the Air Force 

Avionics Laboratory and the Society of Automotive Engineers, Air Force 

Systems Command, Wright-Patterson Air Force Base, Ohio (December 

1970), pp. 139-156. 

6.45 Oh and Schneider, "Lightning Strike Performance," pp. 6-7. 

6.46 J. A. Plumer, Lightning Effects on General Aviation Aircraft, FAA-RD73-99, Federal Aviation Administration, Department of Transportation, 

Washington, D.C. (October 1973), pp. 21-44. 

6.47 J. D. Robb, J. R. Stahmann, and M. M. Newman, "Recent Developments 

in Lightning Protection for Aircraft and Helicopters," 19 70 Lightning and 

Static Electricity Conference, 9-11 December, sponsored jointly by the Air 

Force Avionics Laboratory and the Society of Automotive Engineers 

(December 1970), pp. 25-35. 

6.48 K. Berger, "Development and Properties of Positive Lightning Flashes at 

Mount S. Salvatore with a Short View to the Problem of Aviation 

Protection," Proceedings of the 1975 Conference on Lightning and Static 

Electricity, 14-1 7 April 19 75, at Culham Laboratory, England, Session I: 

Fundamental Aspects and Test Criteria, the Royal Aeronautical Society of 

London (l975), p. 7. 

6.49 T. C. Kosvic, N. L. Helgerson, and M. Gerstein, Ignition of Fuel Vapors 

Beneath Titanium Aircraft Skins Exposed to Lightning, NASA CR 

120827, National Aeronautics and Space Administration, Lyndon B. 

Johnson Space Center, Houston, Texas (September 1971), p. 19. 

6.50 M. J. Kofoid, Lightning Discharge Heating of Titanium Aircraft Skins, 

Dl-82-0752, Plasma Physics Laboratory, Boeing Scientific Research 

Laboratories, Seattle, Washington (September 1968). 

6.5 1 Redrawn from Kofoid, Lightning Discharge Heating, p. 33. 
6.52 J. D. Robb, J. R. Stahmann, T. Chen, and C. P. Mudd, "Swept Lightning 

Stroke Effects on Painted Surfaces and Composites of Helicopters and 

Fixed Wing Aircraft," Proceedings of the 1975 Conference on Lightning 

and Static Electricity, 14-1 7 April I9 75, at Culham Laboratory, England, 

Session 111: Structures and Materials, the Royal Aeronautical Society of 

London (1 975). 

6.53 Plumer, Guidelines for Lightning Protection of General Aviation Aircraft. 

6.54 M. M. Newman, J. D. Robb, and J. R. Stahmann, Lightning Protection 

Measures for Aircraft Fuel Systems--Phase I, FAA ADS-17, prepared by 

the Lightning and Transients Institute for the Federal Aviation Agency, 

Office of Technical Services, U.S. Department of Commerce, Washington, 

D.C. (May 1964), pp. 22-44. 

6.55 Newmann, Robb, and Stahmann, Lightning Protection Measures, pp. 

47-68. 

6.56 Plumer, Lightning Effects on Gerzeral Aviation Aircraft, pp. 11-22. 

6.57 pp. 19-21. 

6.58 L. E. Short, "Electrical Bonding of Advanced Airplane Structures," 

Lightning and Static Electricity Conference, 3-5 December 1968: Part TI, 

Conference Papers, AFAL-TR-68-290, Air Force Avionics Laboratory, Air 

Force Systems Command, Wright-Patterson Air Force Base, Ohio (May 

1969), pp. 425-441: 433. 

6.59 Short, "Electrical Bonding," pp. 435. 

6.60 General Electric Company photograph. 

6.6 1 Newman, Robb, and Stahmann, Lightning Protection Measures, pp. 17-35. 

6.62 J. A. Plumer, "Lightning-Induced Voltages in Electri:al Circuits Associated 

with Aircraft Fuel Systems," Report of Second Conference on Fuel 

System Fire Safety, 6 and 7 May 1970, Flight Standards Service, 

Engineering and Manufacturing Division, Federal Aviation Administration, 

Washington, D.C. (1970), pp. 171-191. 

6.63 J. D. Cobine, Gaseous Conductors: Theory and Engineering A,7plication 

(New York: Dover Publications, 1958), p. 164. 

6.64 Newman, Robb, and Stahmann, Lightning Protection Measz,res, p. 15, 

2 1-44. 

6.65 Plumer, Lightning Induced Voltages, pp. 176-178, 187-191. 
CHAPTER 7 

STRUCTURES PROTECTION 

7.1 Introduction 

It was not easy, however, to escape from Crete, since Minos 

kept all his ships under military guard, and now offered a large 

reward for his apprehension. But Daedalus made a pair of wings for 

himself, and another for Icarus, the quill feathers of which were 

threaded together, but the smaller ones held in place by wax. Having 

tied on Icarus' pair for him, he said with tears in his eyes: "My son, 

be warned! Neither soar too high, lest the sun melt the wax; nor 

sweep too low, lest the feathers be wetted by the sea!" Then he 

slipped his arms into his own pair of wings and they flew off. 

'rFollow me closely," he cried, "do not set your own course." 

As they sped away from the island in a north-easterly 

direction, flapping their wit~gs, the fishermen, shepherds, and 

ploughmen who gazed upward mistook them for gods. 

They had left Naxos, Delos, and Paros behind them on the left 

hand, and were leaving Lebynthos and Colymne behind on the right, 

when Icarus disobeyed his father's instructions and began soaring 

towards the sun, rejoicing in the lift of his great sweeping wings. 

Presently, when Daedalus looked over his shoulder, he could no 

longer see Icarus; but scattered feathers floated on the waves below. 

The heat of the sun had melted the wax, and Icanrs had fallen into 

the sea and drowned. Daedalus circled around, until the corpse rose 

to the surface, and then carried it to the nearby island now called 

Icaria, where he buried it. Robert Graves, The Greek Myths I 

(Baltimore: Penguin Books, 1955): 3 12-1 3. 

The materials of which an aircraft is made and the methods used to hold 

these materials together, forming the aircraft structure, are factors as important 

in protecting a modern aircraft from hazardous natural environments as they 

were for the mythical Icams. Conventional aluminum airframes of riveted 

construction have, by virtue of their excellent electrical conductivity, rarely 

suffered critical damage from lightning strikes; and these structures have 

provided excellent protection for more vulnerable systems, as well as for 

personnel, carried within. But the day of all-metal aircraft is ending. Taking their 

place will be aircraft constructed partly of new, fiber-reinforced plastics with 

desirable light-weight and highstrength properties but with poor electrical 

conductivity. Some of these materials will be fastened together with equally 

nonconductive adhesives, the strength of which may be lost if the adhesives are 

exposed to lightning currents. Whereas the designer of the all-metal structure v 

needed to add little, if anything, to his design to achieve adequate lightning 

protection, the designers of nonmetallic structures must pay particular attention 

to the lightning environment, taking positive measures to protect against its 
adverse effects lest safety of flight be endangered. 

In this chapter we first review lightning effects on metal structures and 

discuss how to design protection for the few situations where special protection 

is necessary. We then review the expected lightning effects on nonmetallic 

materials such as fiberglass and the advanced composites. 

Concern for the potential hazards has prompted a large number of 

laboratory studies of lightning effects and protection techniques for these 

materials. We summarize the important results of these studies and discuss those 

protective techniques and devices which appear most promising. Our discussion 

emphasizes the electrical performance, leaving to the structures designer (who 

himself is best equipped to answer them) questions of compatibility with other 

requirements or manufacturing methods. 

7.2 Metallic Structures 

Because of their inherently good electrical conductivity, conventional 

aluminum aircraft structures have rarely suffered serious damage from lightning 

strikes, and few protective measures have had to be added. The metal skins, ribs, 

spars, stringers, and other substructures may themselves be utilized to provide 

the safe conducting path for lightning currents between possible attachment 

points called for as a basic protection principle in Chapter 5. No special 

conductor dedicated only for lightning currents need be added to conduct 

lightning currents through the aircraft, as long as care is taken in the design of 

basic skins and structures. It would, in fact, be highly impractical, and even 

dangerous, to design a separate "lightning conductor" to be run through a metal 

aircraft. No further mention of such a conductor will be made. 

Thus, in accord with the principle of providing a safe conductive path for 

lightning currents, metallic skins and substructures should be designed to provide 

this conductive path and to minimize, or safely tolerate, the following lightning 

effects: 

Melting at lightning attachment points 

Resistive temperature rise 

Magnetic force effects 

Arcing across bonds, hinges, and joints 

Ignition of vapors within fuel tanks. 

Techniques to assess these effects and design protection, if needed, are discussed 

in the following paragraphs. 

7.2.1 Melting at Lightning Attach~iient Points 

The quantity of metal melted away at a lightning attachment point has 

been shown to be most closely related to the charge carried into the point by the 

lightning flash together with the type of metal and its thickness. Most of the 

charge is delivered by intermediate and continuing current components of the 

flash, and the amount of this charge that is necessary to initiate a hole (i.e., a 

pinhole) in a metal skin has been discussed in Section 6.6 of Chapter 6. The 
charge required to initiate a hole, however, is of little concern except where the 

skin is over a fuel tank. Of concern from a structural integrity point of view is 

the eventual quantity of metal that can be melted away. This aspect is of 

primary concern for skins and structures to which the arc can hang on for 

prolonged periods, as it may in Zones 1 B and 2B. 

Plumer (Reference 7.1) has performed laboratory tests to determine the 

area of holes which may be melted through aluminum or titanium skins of 

various thicknesses, by varying amounts of continuing current delivered at rates 

of between 200 A and 800 A. Figure 7.1 shows the hole areas to be expected in 

typical aluminum and titanium skin metals. The volume of metal melted away 

may be found by multiplying the area by the thickness of the particular skin 

material. 

900- A10.051 cm (0.020 in) 

800- 

0 40 80 120 160 200 240 280 

LIGHTNING FLASH CHARGE (COULOMBS) 

Figure 7.1 Area of holes melted through aluminum and titanium aircraft skins 

by lightning flash charge. 

Normally, holes may be permitted to occur anywhere on the aircraft as 

long as safety of flight is not impaired. The conditions under which safety of 

flight might be endangered from hole formation follow. 

187 
When the hole is melted through a fuel tank skin or other enclosure 

of flammable materials 

When the hole is melted through the wall of a pressurized enclosure 

When the hole sufficiently degrades the mechanical strength of a 

flight-critical component to cause failure 

The problem of hole formation in fuel tank skins has been dealt with in 

Chapter 6. There have been no reports of holes having been melted through the 

walls of a pressurized container, except for the very small (less than 0.1 cm dia) 

holes sometimes found at swept-stroke dwell points along the surface of a 

pressurized fuselage--holes that have been far too small to cause depressurization problems. Since trailing edge areas are usually not pressurized, the larger 

holes melted by longer duration currents hanging on to trailing edges have also 

not posed a depressurization problem, although the possibility of this event 

should be considered in any design that involves pressurization of a trailing edge 

section. 

The structural integrity of a traiIing edge closeout member or other 

loadcarrying part may be degraded if a significant portion of metal is melted 

away by a lightning arc hanging on for a prolonged period of time. For design 

purposes the worst case should be taken as 200 C, in accordance with the SAE 

report Lightning Test Waveforms and Techniques for A erospace Vehicles and 

Hardware (Reference 7.2). Estimates of the amount of metal to be melted away 

by 200 C of charge entering a hang-on point may be made from Figure 7.1 or by 

the use of Equations 6.1 and 6.2 of Section 6.6. 

7.2.2 Mechanical Property Degradation of Structural Metals 

Aside from the burning of holes at attachment points, the flow of 

lightning current will not degrade the mechanical strength properties of metal 

skins and other basic structures unless temperature rises approaching the melting 

temperature occur. Except within a centimeter or two of the points of entry, 

lightning currents will normally spread out sufficiently so as not to produce a 

noticeable temperature rise. 

A case in which an excessive temperature rise may occur is one in which 

lightning current remains in a single conductor, such as a bond strap or pipe. In 

this case, the action integral (~1~dt) of the lightning current and the resistance of 

the conductor may both be high enough to allow sufficient energy to be 

deposited in the conductor to raise its temperature appreciably. Since the 

resistance of most metal conductors increases with temperature, an even higher 

amount of energy will be deposited in the conductor as its temperature rises 

during current flow, and this, in turn, will increase the temperature even further. 

If the action integral is high enough, the conductor will be unable to absorb all 

of the energy and will explode. This will happen for currents whose action 

integrals are sufficient to raise the conductor temperature above its melting 

point. 

It is therefore important to be able to determine the expected temperature 

rise which lightning may produce in conductors of small cross-sectional area and 
to obtain assurance that the melting point will not be reached. The temperature 

rise in a current-carrying conductor is expressed by the basic equation (units in 

parenthesis): 

JIZ dt (AZ -s) Resistivity (Rcm) 

AT("C) = (7.1) 

Specific heat (g*) ( ) Density Areaz (cm4 ) g-"C cm 

It is evident that temperature rise is directly proportional to the square of 

the lightning current and inversely proportional to the square of the crosssectional area of the conductor. As it stands, however, Equation 7.1 has two 

shortcomings: 

1. It assumes no energy (heat) loss resulting from radiation during the 

time current is flowing 

2. It assumes that resistance does not change with temperature. 

Since little energy is lost during the short time duration of lightning 

strokes, the first shortcoming has a negligible effect on the accuracy of the 

predicted temperature rise. Resistance does depend on temperature, however, 

and, since resistance also affects the amount of electrical energy that can be 

deposited in the conductor, it is important to account for this dependency. This 

can be done by expressing resistivity as a function of temperature, using the 

following expression: 

where X is the temperature coefficient of resistivity. Incorporating Equation 7.2 

into the expression of Equation 7.1 gives the following equation: 

~1'dt (A' -s) Resistivity (R-cm) [ 1 + h ($)AT('C)] 

"TCC) = 

Specific heat (g@) Density (A) ~rea' (cm4) g-°C cm3 (7.3) 

Since everything but is a material property constant, this equation is 

most easily solved by combining the conductor dimensions and material 

properties in a constant, k, in the following expression: 

where k is equal to the right side of Equation 7.1. 
Physical and electrical properties needed in Equations 7.5 and 7.6 for 

common structural metals and electrical conductors are provided in Table 7.1. 

The melting points of these materials are also provided. 

Table 7.1 PHYSICAL AND ELECTRICAL PROPERTIES 

OF COMMON METALS 

Stainles I 

Aluminum Copper ntanium Steel (304) Magnesium Silver 

Resistivity, 2.8 x lod 1.72 x lo6 42 x lo4 72 x lod 4.45 x lo6 1.59 x lo6 

(0-cm) 

Temperature 0.00479 0.00393 0.0035 0.001 0.0165 0.0041 

coefficient of 

resistance, h 

(11°C) 

Thermal 0.254 x lo4 0.164 x lo4 0.085 x lo4 0.120 x lo4 0.025 x lo4 0.019 x lo4 

coefficient 

of linear 

expansion 

(]PC) 

Specific heat 0.2 15 0.092 0.124 0.120 0.245 0.056 

(g-~al/~?C) 

Density 2.70 8.89 4.51 7.90 1.74 10.49 

(dcm3) 

Melting point 660 1084 1670 1150 6 50 962 

ec) 

Because of their high amplitudes, the return stroke and restrike currents 

have the highest action integrals and may be expected to produce higher 

temperatures than the other components of the lightning flash do when they are 

conducted through structural metals or other electrical conductors. For design 

purposes, the action integrals associated with current waveform Components A 

andlor D of the SAE report Lightning Test Wuveforms and Techniques for 

Aerospace Vehicles and Hanlwarc should be used for determining temperature 

rise in conductors, depending on the lightning-strike zone in which the particular 

conductor 1s located. 

It should be remembered that if several parallel conductors are available to 

share the stroke current, the current in each will be divided by the number of 

conductors but that the actiorl integral in each will be reduced by the square of 

the number of conductors. Consider, for example, a case in which a NAV light is 

mounted on a plastic vertical fin cap and grounded to the airframe via two bond 

straps, as shown in Figure 7.2. If this case is considered to be located in Zone 

IB, its bond straps must be able to conduct the entire flash current. If it is 

assumed that the arc will not touch the bond straps, the limiting lightning design 

criterion for these straps is that together they must be able to carry safely 

current Components A and D, which have a total of 2.25 x lo6 A~-s. However, 
the one-half of the total current flowing in each strap will produce only 

one-fourth of the total action integral, or about 0.51 x lo6 A~-S ineach strap. 

NAVIGATION 

STRUCTURE 

Figure 7.2 Bond straps for NAV light. 

If each strap is made of copper with a cross-sectional area of 0.1 cm2 

(0.015 in2), the temperature rise produced by 0.5 1 x lo6 A~-S in a strap would 

be determined from Equation 7.4, using the physical and electrical properties for 

copper given in Table 7.1. Solving for k first 

This, of course, would be the actual temperature rise if the resistivity remains 

constant. But the resistivity actually increases as temperature increases; 

therefore, Equation 7.5 must be used as follows to calculate the actual rise: 

ATCC) = (25.6 "C) 

A conductor with this cross-sectional area is about the same size as an 

AWG 15 wire. Since a 28 "C temperature rise would not damage the wire, two 

19 1 
bonding straps or wires of this cross-sectional area could, theoretically, be 

utilized to conduct current Components A and D. If one of them had worked 

loose, however, all of the current would have to flow through the other one. If 

the same equations are used in this circumstance, the total action integral of 

2.25 x lo6 A~-S would cause a calculated temperature rise of over 200 "C, a 

temperature that should be considered excessive. 

In fact, it is prudent to keep conductor temperature rises below 100 "C in 

order to avoid explosion where flammable materials or vapors may exist and also 

to avoid mechanical stresses resulting from thermal expansion (See Section 

7.2.3). The temperature coefficients of resistivity which are given in Table 7.1 

remain reasonably constant only through the range of -50 "C to +I00 "C; thus, 

calculations that are based on these coefficients and which produce temperatures 

above 100 "C should not be considered accurate. 

Consideration of temperature rise effects in conductors carrying lightning 

currents is most important where all or a large fraction of the stroke currents 

need to be carried in individual conductors. Examples include radome diverter 

straps, bonding straps (or jumpers), conduits or cable shields, ground wires 

passing through plastic radomes or wing tips, and certain hydraulic or control 

lines which may become exposed to lightning-strike currents. 

Figure 7.3 presents calculated temperature rises for conductors of the 

metals of Table 7.1 as a function of cross-sectional area for action integrals of 

0.25, 2.0, 2.25, and 3 x lo6 A~-s. These curves may be used to determine the 

expected temperature rise in structural elements, diverter bars, bond straps, or 

other conductors expected to carry lightning currents. The temperature rises 

determined from Figure 7.3 should be added to the ambient temperature of the 

- conductor to determine the actual temperature. 

If the reader desires to calculate the temperature rise for action integrals 

other than the four of Figure 7.3, he may do so manually by using Equations 7.1 

and 7.5, or he may use the computer program written for the Texas Instruments 

SR-52 programmable calculator included in Appendix 1. The program requires 

the action integral, resistivity, specific heat, density, and cross-sectional area of 

the conductor as inputs from which it calculates temperature rise. The program 

also calculates the elongation that takes place in the metal as a result of the 

temperature rise. 

A separate SR-52 program is provided in Appendix 2 for calculation of 

temperature rise and elongation in tubes and pipes. This program accepts 

dimensions of the outside diameter and wall thickness, thereby saving the user 

the need to calculate manually the cross-sectional area of the pipe wall. 

If several parallel conductors of the same length and cross-sectional area 

are present, the current may be expected to divide equally among them; the 

action integral associated with each portion of the current will be 

JI' dt (each conductor) = $r2 dt (total) 

d 

where N is the number of parallel conductors. This distribution would suggest 

192 
1w 

AT 

OC 

10 

1 5 

AREA - cm' 

that the use of several conductors of small cross section would be advantageous, 

since the total amount of metal required for all of them would still be less than 

that required for one large conductor. This conclusion is sound, but there are 

two factors that should be given careful consideration before any decision to 

utilize several thin conductors instead of one thick one to conduct lightning 

currents is reached: 

1. As a result of vibration or corrosion stresses, thin conductors may 

break sooner, leaving a situation in which the remaining conductor(~) may be greatly overstressed by lightning currents 

2. The magnetic forces which act upon parallel current-carrying 

conductors may cause excessive damage to these conductors. 

On the other hand, magnetic forces may sometimes be reduced sufficiently 

by separating the conductors a sufficient distance from each other (See Section 

7.2.4). 

193 
7.2.3 Thermal Elongation 

Most metals expand when their temperatures rise. This phenomenon will 

not be a problem for flexible conductors which can deflect when expanded, but 

large stresses may arise in some rigidly held conductors, such as the diverter strap 

or the pitot probe air pressure tube shown in Figure 7.4. By virtue of its location 

at the nose of the aircraft, the pitot probe will be in Zone I A and must therefore 

be designed to conduct return stroke currents of 2 x lo6 A~-s. The purpose of 

the air pressure tube is to transmit pitot-static pressure back to the flight 

instruments, but the tube may also provide a safe path for lightning currents to 

enter the airframe if the tube is designed to safely withstand the resistive 

heating, thermal elongation, and magnetic force effects. 

NOSE -- RADOME 

AIR GROLINnINC: PRESSURE AND/OR TI IRF 

DIVERTER 

Figure 7.4 Examples of rigidly held lightning current conducton. 

Cross-sectional area should be sufficient to prevent excessive 

temperature rise and thermal expansion. 

For example, if the air pressure tube in the radome of Figure 7.4 is made 

of copper and has an outside diameter of 0.476 cm (3116 in) and a wall 

thickness pf 0.124 cm (0.049 in), it will have a cross-sectional area of 0.138 cm2 

(0.021 in2). According to the curves of Figure 7.3(c) an action integral of 7- x lo6 

A~-S will raise the temperature of a conductor of this cross-sectional area 67 OC. 

The amount of thermal expansion to be expected along any dimension of 

a part is dependent on the temperature rise and the thermal coefficient of linear 

expansion of the metal, according to the following relation (dimensions in 

parenthesis): 

AL = Expansion coefficient (7 -9) 

where L is the length of the dimension of concern. Of course all dimensions of a 
homogeneous material will elongate proportionately, but frequently only one 

dimension is of much concern. The most significant expansion in a part such as 

the pressure tube would be along its length. 

Based on the coefficient of linear expansion for copper given in Table 7.1 

as 0.1 64 x 1 o'~ (1 1°C) and the assumption that the tube has a length, L, of 2 m 

when "cold," a 67 "C temperature rise would cause an elongation of 

If the tube is fastened rigidly, this elongation may cause sufficient stress to break 

clamps or force the tube to bend. Thus, the tube supports should be designed to 

accept this elongation without undue stress. 

Thermal elongation calculations may be made for other conductors by 

using Equation 7.9 or the SR-52 programmable calculator programs in 

Appendices 1 and 2. These programs calculate the elongation in millimeters per 

centimeter of conductor length for the basic conductors and in inches per foot 

of conductor length for tubes, since dimensions of tubes are commonly 

presented in inches. 

The temperature rise and elongation expressions described in this and in 

the previous section have been closely confirmed up to 100 "C by laboratory 

tests in which simulated lightning currents of known action integral are passed 

through conductors whose temperature and elongation are measured. Temperature rise is most easily measured by temperature-sensitive tapes or by an 

optical pyrometer. Elongation can be measured by a mechanical deflection 

indicator attached to the specimen. Calculated results above 100 "c, however, 

will begin to diverge from the actual because of physical property changes at 

higher temperatures. 

7.2.4 Magnetic Force Effects 

As mentioned earlier, there are situations in which several conductors in 

parallel may carry lightning currents and, in such cases, each conductor will be 

acted upon by a magnetic force proportional to the product of the currents in 

each conductor. One such case is shown in Figure 7.5. 

CONDUCTOR 1 , dl,, 

0 $ 

7 44 dP, 

D 

CONDUCTOR 2 - ) 

dl2 ' J. 2 

Figure 7.5 Magnetic forces among parallel current-carrying conductors. 
Based on Ampere's Law, the force per unit length on wire 2 is 

where pa = permeability of media between conductors 

= 10'~ for air 

PI ,2 = force on conductor 1 or conductor 2 (N/m) 

Q1 ,2 = unit of length of conductor 1 or conductor 2 (m) 

D = center-to-center distance between conductors (m) 

I1 ,2 = currents in conductor 1 or conductor 2 

These forces act to pull the conductors together when the currents in them 

flow in the same direction and to repel the conductors when the currents are in 

opposite directions. In either case, the magnetic forces are strongest when the 

conductors are close together. One such case would exist, for example, in a 

radome-mounted pitot system in which two air pressure tubes are needed to 

convey both static and dynamic air pressure to the flight instruments. If these 

tubes are mounted 2 cm apart and half of the lightning current is assumed to 

flow in each, the peak force in each conductor, as a result of the currents in 

both, will be 

= 100 000 N/m of tube length 

= 6 854 Ib/ft of length 

This peak force, of course, would exist only at the instant when the lightning 

current is at its peak, with lower forces active when the current is lower. During 

the short duration of the higher amplitude stroke currents, the mass of the 

conductor may prevent its being deflected appreciably because of its relatively 

slow mechanical response time, although it can absorb appreciable kinetic energy 

in that time and continue deflecting afterward. The resulting deflection also 

depends on the degree of restraint provided by supports, but forces of this 

magnitude have been shown to cause tubes in one typical installation to slam 

together, leaving permanent deformations, as shown in Figure 7.6. Deformations 

like those in Figure 7.6 (Reference 7.3) are sufficient to disrupt the transmission 

of pitot pressure to flight instruments. 

Mathematical calculation of expected deformations from impulsive forces 

such as these is difficult, and such calculation is further complicated by 

unknown mechanical factors. James and Phillpott, however, have defined an 

effective transient force, Pt which, if applied slowly but continuously, results in 

the same stress as that produced by the peak magnetic force, Po, calculated by 

Equation 7.1 1. This effective force, Pt, is a function of the angular frequency, 
Figure 7.6 Air pressure tubes deformed by magnetic forces (after carrying a 

120 kA, 2 x lo6 A~-S stroke). 

w, of the meclzanical system being acted upon and of the electric current decay 

time constant, 7. In most practical applications the angular frequency will be 

between about lo3 for heavy, flexible systems and lo4 for light, stiff systems. 

James and Phillpott give relative values of Pt/Po (Table 7.2 [Reference 7.41) for 

each of these frequencies and three lightning current waveforms. Their 300 kA 

and I kA currents represent a return stroke and continuing current of the same 

magnitude as those represented by Components A and C of Ligl?miizg Test 
Wavefornzs and Techniques for Aerospace Vehicles and Hardware. However, the 

50 kA intermediate current would deliver much more energy than would the 2 

kA intermediate current, represented by Component B of Lightning Test 

Waveforms and Techniques. A 50 kA current decaying in as long as 2 ms would 

be found only in the severest of positive polarity flashes, which occur very 

seldom. In any event, Table 7.2 shows that the 200 kA stroke current creates a 

much higher effective transient force, Pt, than does either of the other currents. 

Table 7.2 EFFECTIVE TRANSIENT MAGNETIC FORCE Pt 

An example of how the table is used may be found in the case of the two 

parallel air pressure tubes 2 cm apart, for which a peak force, Po, of 100 000 

N/m of length was calculated. Assuming that these tubes have an angular 

frequency of lo4, Table 7.2 shows that the effective transient force, Pt, would 

be 

Pt = 0.45 Po (7.13) 

or 45 000 N/m (3084 Ib/ft) of length. This is the amount of force which should 

be used in performing mechanical response calculations or mechanical strength 

tests. 

James and Phillpott also conclude from Table 7.2 that 

(a) Continuing current pulses do not give rise to a high effective force 

because l2 and therefore the peak magnetic forces are too low. 

1 

(b) If 7<- as for relatively heavy flexible systems (a = lo3) the 

W 

effective force is roughly proportional to ~~t. Thus the fast and 

intermediate components give about the same effective force 

although the peak forces are very different. 

1 

(c) If 7% - as for relatively light stiff systems (a = lo4), the effective 

W 

force equals twice the peak force and is therefore proportional to I~. 

For this doubling of the peak force to occur it is assumed that the 

1 

current rise time is < -. 

1 W 

(d) If TZ - the effective force is 45% of the peak force for the case of 

a' 
the fast component with w = I@. This alternative also gives the 

highest effective force because the high value of more than 

compensates for the effects of a small value of 7. (Reference 7.5). 

These authors conclude that "[tlhe effect of magnetic forces on 

currentcarrying components is influenced by a large number of parameters, but 

the contribution from the continuing current is negligible." Concurring with the 

authors of this book, they also conclude that future studies of or tests on 

components likely to be damaged by magnetic forces should be performed with 

simulated lightning currents having values of I and 12t similar to Component A 

of Lightning Test Wavefoms and Techniques. 

Outer portions of the airframe and control surfaces sometimes become 

badly damaged by magnetic force effects. One example is the wing tip trailing 

edge shown in Figures 7.7 and 4.4, in which the upper and lower surfaces of the 

wing were wrinkled, pinched, and pulled forward. The upper and lower surfaces 

of this wing were made respectively of 0.71 mm (0.028 in) and 0.80 mm (0.031 

in) aluminum. Hacker (Reference 7.6) has made a very interesting analysis of the 

magnetic forces acting upon these surfaces and found, as illustrated in Figure 

7.7, that the directions of the calculated forces correIate well with the damage 

pattern. Hacker calculated that a 100 kA stroke would have created 21 500 N/m 

(1476 Ib/ft) on a 4 cm wide path across the skins. This amounts to 5.4 x lo5 

~/m~, or 78psi, an extremely high compressive pressure for a wing structure 

reportedly designed for a steady state loading of 1.7 x 103 ~/m~ (0.25 psi). The 

reader is referred to Hacker's report for a more thorough discussion of the 

procedure followed in making this analysis. 

\ ,LIGHTNING PATH \ I 

~2.4~ 

\\' I E 

B \ 

\ 

\- ------- REFERENCE LINE ---- A C t 

-1.6' D 

Figure 7.7 Approximate geometry of chordwise cross section of wing tip trailing edge of Figure 4.4. 
The foregoing examples have illustrated the case in which the magnetic 

forces are attractive (when parallel currents flow in the same direction). Cases 

exist, however, where currents in adjacent conductors, or in adjacent legs of the 

same conductor, flow in opposite directions. When this happens, the magnetic 

force acts to separate the conductors, a fact that may create a problem for the 

unwary designer. One such case is the folded protective foil splice shown in 

Figure 7.8. 

Figure 7.8 Reverse current paths through joint. 

A splice of this sort was considered (Reference 7.7) for joining two plastic 

laminates protected with aluminum foils. An aluminum edge wrap was utilized 

to provide electrical conductivity between the foils on the lower and upper 

laminates. This arrangement, however, forced the lightning current to travel in 

opposite directions through the edge wrap, as shown in Figure 7.8. When a 100 

kA simulated lightning current was passed through a sample of this joint, the - 

edge wrap pulled away from the fasteners as shown in Figure 7.9. 

Jacobsen (Reference 7.8) gives the force, P, per unit length of current 

path, d, acting to repel two parallel plates whose separation, X, is small compared 

to width, b, and which carry current, I, in opposite directions as follows: 

where p=4nx IO-~H/~ (7.1 5) 

Assuming the aluminum edge wrap of Figure 7.8 was 0.1 m (4 in) wide, 

the peak force for the 100 kA test current would have been 

For a length of current path, d: of 0.013 m (0.5 in) through the edge wrap 

200 
This force was evidently adequate to blow off the upper surface of the edge 

wrap. 

Figure 7.9 Aluminum edge wrap torn away from fasteners by magnetic force. 

Another, more common, example of the forces created by currents 

flowing in opposite directions is the bent bond strap shown in Figure 7,10(a). 

Even if the strap has a cross-sectional area sufficient to conduct the current, it 

will frequently break, as shown in Figure 7.10(a), if it forms a bend of more 

than about 45". Such straps should be installed, whenever possible, in a straight 

line, as is the strap in Figure 7.10(b). 

20 1 
PARALLEL CURRENT 

COMPONENTS 

IN OPPOSITE 

DIRECTIONS, 

CREATING 

OPPOSING 

FORCES IN 

STRAP 

- 

NO MAGNETIC ( 

FORCE \ 

Figure 7.10 Magnetic force on a bent bond strap. 

Another reason for keeping the bond strap straight--and as short as 

possible--is that the inductive voltage rise occasioned by lightning currents 

flowing in a longer strap may be sufficient to spark over a direct, shorter path 

across the air gap. Some other do's and don'ts regarding the design of bond 

202 
straps intended to carry lightning currents are presented in Figure 7.1 1. The 

basic rules to follow are these: 

Use conductors with sufficient cross-sectional area to conduct 

intended lightning current action integral 

Keep bond straps as short and close to the air gap length as possible 

Avoid bends of more than 45", or other features that result in 

reversal of current directions 

Avoid all sharp turns 

If two or more parallel straps are used, separate them sufficiently 

(usually 30 cm or more) to minimize magnetic force effects. 

Of course, the above rules should also be followed not only for bond straps but 

for any other structural element that must carry lightning currents. 

Because of weight limitations, the strength and rigidity of some metallic 

components, such as wing tips, flaps, and ailerons, may not be sufficient to resist 

deformation as a result of the magnetic forces from lightning currents 

concentrated in these extremities. Such deformations will not normally impair 

safety of flight, but they may require repairs or replacement. Normally, only 

severe lightning currents cause this deformation; therefore, increased reinforcement of the extremity, if it were deemed economical, could minimize or prevent 

deformation in many cases. 

Since determination of magnetic force effects by mathematical analysis for 

all but the most elemental geometries is very difficult, laboratory tests may 

prove the most straightforward and economical way to determine whether or 

not magnetic force effects are likely to cause deformation of prospective 

structures. For such tests, simulated lightning strokes in accordance with the 

SAE criteria of Lightning Test Waveforms and Techniques (Reference 7.9) 

should be passed through the component being evaluated. 

7.2.5 Arcing across Bonds, Hinges, and Joints 

Basically, the riveted skins and substructures found in conventional 

aluminum aircraft are adequate to safely conduct lightning currents through the 

aircraft, and, aside from the special considerations relating to fuel tank areas 

discussed in Chapter 6, no provisions need be made to improve this conductivity 

for lightning protection purposes. Even with the profusion of electrically 

insulating primers and wet sealers now used to coat joined surfaces, the great 

number of rivets or fasteners needed to meet mechanical strength requirements 

provides the necessary electrical path between mating surfaces. This generalization has proved to be reliable even when the fasteners and holes themselves have 

been coated with primers or wet sealants. 

If hinges or bearings are located where lightning currents may pass through 

them (such as on control surfaces in Zone 1B or 2B), they should be able to 

safely withstand these currents with no impairment of their function. Otherwise 

suitable means should be provided to divert much of the lightning current 

around them. In a few cases, passage of lightning currents through some hinges 
or bearings may roughen or even weld the mating surfaces together, occasionally 

impairing further movement. 

BOND STRAP BOND STRAP 

SAME LENGTH LONGER THAN 

SHARP 

BENDS 

SHARP BENDS 

AT LUGS 

GOOD BAD 

PRACTICE PRACTICE 

Figure 7.11 Design of bonding straps, or jumpers. 
Damage such as this occurs only when a hinge or bearing has a single point 

of contact through which most of the lightning current must pass. In such cases 

the point(s) of hinge contact may be damaged. Hinges with multiple points of 

mechanical contact are usually able to safely withstand lightning currents with 

only minor pitting or erosion. For example, a piano-type hinge such as that 

shown in Figure 7.12 may be considered as self-bonded provided the dc 

resistance across the hinge is less than 0.01 R (Reference 7.10). In hinges where 

excessive damage occurs, additional conductivity should be provided to carry a 

major portion of the lightning current across the hinge to relieve it from 

excessive lightning current flow. Flexible bonding jumpers of the type shown in 

Figure 7.1 1 have often been installed across hinges for this purpose. In most 

cases, however, these jumpers have provided little or no increase in conductivity 

for lightning currents. The reason is that they almost always offer a longer and 

more inductive path than do the paths directly through the hinges, even though 

the resistances of the latter are high. Because of their fast rates of change, 

lightning current paths are governed much more by inductance than by 

resistance. Stahmann (Reference 7.1 1) has demonstrated that bonding jumpers, 

in fact, make little or no difference in the amount of superficial pitting that 

occurs on typical piano-type control suifaces or door hinges, such as those 

shown in Figure 7.12. Stahmann found the same to be true of ball joints and 

other hinges utilized on control surfaces or landing gear doors, even when the dc 

resistance through the hinge was as high as several ohms. No binding or other 

adverse consequences were found to occur from the small amount of pitting that 

did occur in Stahmann's tests of piano-type hinges and ball joints. 

AIRFRAME 

I 

MOVABLE 

CONTROL SURFACE 

Figure 7.12 Piano-type hinge. 

Often capable of conducting lightning stroke currents without 

need for bonding jumpers. 

To be sure of the ability of a particular hinge design to safely conduct 

lightning currents, a test similar to Stahmann's, in which simulated lightning 

currents were conducted through a prototype hinge, should be performed. The 

test waveforms and current described in Lighbzirzg Test Waveforms and 

Techniques for the zone in whlch the hinge is located should be used. If 

excessive pitting, binding, or welding of the hinge is found to occur, additional 

conductivity may be necessary. The most satisfactory way of providing this 

additional conductivity may be to provide additional areas of contact in the 
hinge itself, or else to provide additional hinges. 

Bond straps across hinges are sometimes required to prevent the 

electromagnetic interference that arises from precipitation static charges being 

conducted through hinges with loose or resistive contact. The low currents 

involved are sometimes unable to follow uninterrupted paths through hinges, 

with the result that minute sparking occurs. Bond straps are usually able to 

provide enough conductivity to reduce this sparking. If bond straps are applied 

for this purpose, the guidelines of Figure 7.1 1 should be followed. 

7.3 Nonmetallic Structures 

With the use of nonmetallic materials in aircraft structural and fuel system 

components, the question of lightning vulnerability becomes important. An 

all-metallic airframe is in many ways quite compatible with basic lightning 

protection requirements; often only minor design modifications have been 

needed to provide adequate lightning protection. Nonmetallic structural components, however, may be considerably more vulnerable to lightning from lack 

of electrical conductivity. Thus, protective measures must usually be applied. If 

this is not done, hazardous consequences, such as those illustrated in the 

example described in Section 4.5 of Chapter 4, may result. 

There are three basic types of nonmetallic materials found in aircraft. 

These are fiberglass-reinforced plastics (commonly called "fiberglass"), boron- or 

graphite-reinforced plastics (commonly called "advanced composites" or simply 

"composites"), and polycarbonate resins such as Lexan' and Plexiglas2 . The 

latter materials are used for windshields or canopies. Fiberglass is used most 

commonly to achieve light weight in wing tips, fin caps, access doors, and in 

other applications where high mechanical strength is not required. The advanced 

composites, on the other hand, are capable of sustaining very high tensile and 

shear loads; for this reason they are being considered for primary structural 

applications. 

Since the electrical conductivity and high-voltage withstand capability of 

each of these types of nonmetallic materials vary widely from those of the 

others, in the following paragraphs we shall discuss the important lightning 

effects problems and design of protective measures for structures made of each 

basic material. 

7.3.1 Fiberglass Structures 

Fiberglass-reinforced plastics have essentially no electrical conductivity. 

Moreover, they are usually incapable of withstanding the high-voltage stresses 

created by the approaching lightning leader and frequently will undergo 

dielectric breakdown, allowing the flash to puncture them and terminate on 

some conducting object inside. The ensuing return stroke current can blast a 

' Trademark of the General Electric Company 

2~rademark of the Rohm and Haas Company 
large hole in fiberglass skins such as the skin of the nose radome shown in Figure 

4.1 1. Each of the component materials, glass and resin, has relatively good 

voltage withstand capability; however, the many interfaces which exist between 

the glass fibers and the plastic resin when the two are put together create voids 

and seams which will withstand much less voltage stress before ionization and 

charge mobility (breakdown) begin. Thus, an external conductor, or diverter, 

must usually be placed on the outside surface of a fiberglass skin to provide a 

place to which lightning flashes can attach. The need for such a diverter and the 

guidelines for its proper design and location can best be understood by reviewing 

the lightning attachment mechanism itself. 

7.3 .l. 1 Lightning Attachment to Fiberglass Materials 

In Chapter 1, Section 1.2.4, we described the basic lightning breakdown 

process, and in Chapter 2, Section 2.2, we considered how this process causes 

attachment to an aircraft in flight. As noted there, when the intense electric field 

created by the oncoming leader reaches the ionization potential of air in the 

neighborhood of extremities and sharp edges on the aircraft, ionization occurs 

and streamers propagate outward from the aircraft in the direction of the 

oncoming leader. Since fiberglass has no electrical conductivity, the electric field 

passes directly through it, which phenomenon can cause streamers to originate 

from objects inside as well as outside of the fiberglass structure. What happens 

may be viewed as a race between streamers propagating from conducting objects 

inside and outside of the fiberglass structure, as shown in Figure 7.13. 

FIBERGLASS 

LEADER 

Figure 7.13 Streamers induced by approaching leader. 

Streamers occur from objects inside and outside the fiberglass 

structure. 

Examples are wing tips, external fuel tanks, and radomes. 

Both the voltage withstand capability of the fiberglass and the distances 

along alternate breakdown paths are important in establishing whether puncture 

or external flashover will occur. Usually, of course, a higher intensity electric 

field is necessary to permit the internal streamer to puncture the fiberglass wall 

and contact the leader than would be necessary merely to draw the external 

streamer through the air to the leader. As the leader approaches, the field 

intensity increases until one of the streamers reaches and joins the Ieader. The 
electric field about this extremity of the aircraft then collapses (breaks down), 

and the leader proceeds from another extremity of the aircraft (as described in 

Chapter 2). 

Figure 7.14 (Reference 7.1 2) shows the results of two simulated lightning 

attachment tests on a typical fighter aircraft nose radome performed in 

accordance with SAE Report Lightning Test Waveforms and Techniques for 

Aerospace Vehicles and Hardware. The approaching leader is represented by the 

striped electrode, to which a high voltage was applied by a Marx-type generator. 

When the field between the probe and the radome became high enough, 

streamers began from conducting points both inside and outside the radome and 

also from the high-voltage electrode. In Figure 7.14(a) the streamer drawn from 

the external conducting base reached the leader first; thus the base became the 

attachment point. In 7.1 qb), however, an internal streamer, emanating perhaps 

from the radar antenna, punctured the fiberglass wall and reached the leader 

before the external streamer from the conducting base. The antenna, therefore, 

became the attachment point. 

Figure 7.14 Simulated lightning attachment point tests of a nose radome. 

(a) Outside streamer intercepts leader. 

(b) Internal streamer punctures fiberglass wall and intercepts 

leader. 

Since these photographs were taken outdoors in daylight, only those 

streamers that reached the leader and were brightened by the generator discharge 

current are visible. Figure 7.15 (Reference 7.13) shows similar lightning 

attachment tests performed on a fiberglass section of an external fuel tank. The 
tests were made in a darkened area indoors; thus in Figure 7.15(a), in addition to 

the main discharge, streamers emanating from alternate points can be seen. 

Streamers also occurred at the same time from objects inside of the tank, and 

one of these succeeded in puncturing the tank wall, as is evident from Figure 

7.15(b). Here, both an external and air internal streamer reached the leader, 

causing some of the generator discharge current to follow the internal streamer, 

as evidenced by the reduced brightness of the external flashover. Since the tank 

had no external diverters, the result illustrates the poor insulating capability of 

fiberglass alone. 

Figure 7.15 Simulated lightning attachment tests of fiberglass fuel tank section. 

(a) External flashover 

(b) Simultaneous external and internal flashovers 

If the test voltage applied between the electrode and the tank were to rise 

at a slower rate, the electric field might not become intense enough to puncture 

the fiberglass wall before a streamer coming from an external point reached the 

leader and collapsed the field. Thus, the rate of voltage rise (dV/dt) is also 

important in determining whether puncture or external surface flashover will 

occur. 

This important factor can be further explained by comparing the 

breakdown timelag characteristics of solids and air. All insulating materials, 

whether solids or gases, respond to high voltages and break down according to a 
timelag curve of the shape shown in Figure 7.16. Timelag effect means simply 

this: the shorter the time for which a voltage is applied across a given insulation, 

the higher that voltage must be to cause breakdown; and conversely, the longer 

the time, the lower the voltage necessary to cause breakdown. There is, of 

course, a voltage level below which breakdown will not occur at all, even if the 

voltage is applied for a long time. Most solids show a flatter timelag 

characteristic than do air or surface flashover paths, as shown in Figure 7.16. 

Time TO BREAKDOWN 

Figure 7.16 Breakdown timelag curves for solids and air. 

In general, for fiberglass components, such as the radome of Figure 7.14 

and the fuel tank section of Figure 7.15, an oncoming lightning flash has 

alternate paths to a metallic conductor. One is via a puncture of the nonmetallic 

skin to an internal metallic component, such as the radar dish. The other is via 

surface flashover to the nearest exposed metal. While the path via the puncture 

may be much shorter than that along the outside surface, the added insulation 

provided by the solid skin often compensates to some degree for this, making 

both paths viable alternatives. The significance of voltage rate of rise now 

becomes evident. Because the timelag curve for the alternate paths cross each 

other, there are voltage waveforms that will intersect either timelag curve, as 

shown in Figure 7.17, where both a "fast" and a "slow" voltage waveform are 

superimposed on the breakdown timelag curves of Figure 7.16. 

From Figure 7.17 it is evident that the faster rising test voltage is the more 

severe in terms of increased probability of puncture. Thus, for attachment tests, 

such as those of Figures 7.14 and 7.15, a test voltage with the fastest rate of rise 

expected from a natural lightning leader is desirable. 

No time-domain measurements have as yet been made of the electric fields 

surrounding an aircraft during the lightning-strike formation process because 

such measurements present formidable instrumentation problems. However, in 

Chapter 1, Section 1.2.3, we state that the leader advances at about I to 2 x I o5 

m/s, which would result in an average of 5 to 10 ys for the leader to travel a 

distance of 1 m. Since about 500 kV are required to break down a 1 m air gap in 

210 
5 ps, an appropriate rate of voltage rise to use for attachment tests might be the 

following: 

APPLIED VOLTAGE WAVEFORMS 

z 

g 8 0s A 

a 0 E> 

m 

BREAKDOWNS 

Time TO BREAKDOWN 

Figure 7.1 7 Fast and slow applied voltage waveforms superimposed on breakdown timelag curves for solid and air breakdown paths. 

Higher voltage rate of rise intersects solid curve and puncture 

occurs. 

Lower voltage rate of rise intersects air curve and surface flashover occurs. 

On the other hand, if it is remembered that the actual breakdown of a single step 

of the leader is itself a series of smaller step breakdowns and pauses, it is likely 

that the rate of voltage rise across segments only a few meters long might be 

faster (or slower) than the average. Thus, to encompass the worst case, a rate of 

voltage rise 10 times as fast, or 1000 kV/p, has been prescribed in SAE Report 

Lightning Test Waveforms and Techniques for attachment tests of aircraft 

components. 

In practice, the timelag curves representing different paths through solid 

and air insulation are flatter and closer together than those drawn in Figures 

7.1 6 and 7.17, and the point where the solid and air curves cross is less clearly 

defined. Thus, a relatively wide difference in applied voltage rate of rise exists 

between waveshapes causing breakdowns 100% of the time along one path as 

compared with waveforms causing all breakdowns to occur through the other 

path. This fact is additional support for the selection of 1000 kVlps as the rate 

of rise for qualification test purposes. This waveform should create a faster 

(but not excessively fast) rising electric field than most natural lightning flashes 

create. In cases where a comparison between the results of laboratory tests using 

this rate of rise and subsequent in-flight lightning attachments to the same 

fiberglass component has been made, the laboratory tests have accurately 

predicted the in-flight attachment points and breakdown paths. 
Figure 7.18 shows the high voltage applied between the test electrode and 

the fuel tank of Figure 7.15. Tn this test the electrode was at negative polarity 

with respect to the tank; thus the waveform has a negative deflection. Both 

polarities, of course, should be applied in the course of any complete test 

program. The rate of rise of this test voltage was 1000 kV/ps; it had reached 

1000 kV when complete flashover of the 1 m air gap and fiberglass tank 

occurred. 

250 kV/div. 0.5 ps/div. 

Figrtre 7.18 Typical high-voltage test waveform for attachment tests. 

d V 

- = 1000 kV/ps dt 

Flashover on rise at 1000 kV 

The 1000 kV reached by this test voltage (Figure 7.18) was of course 

applied across the total breakdown path between the high-voltage electrode and 

the metallic parts of the tank, shown in Figure 7.1 5. That amount of this voltage 

which must develop across the fiberglass wall for puncture to occur can be 

determined by applying test voltage directly across the wall. Figure 7.19 shows 

such a test of the wall of the fuel tank of Figure 7.1 5, using only 2.54 cm (1 in) 

of air gap between the electrodes and the wall. Each test voltage was allowed to 

reach its peak and to decay through the test circuit if breakdown did not occur, 

and the tests were repeated with successively higher peak voltages until 

breakdown occurred. 

In the tests of Figure 7.19(a), peak impulse voltages of up to 125 kV were 

withstood by the wall, but partial breakdown occurred when a peak of 152 kV 

was applied. Complete breakdown occurred at 173 kV. From this test it might 

be concluded that the path involving puncture of the tank wall would require 

173 kV more voltage than would the path along the outside of the tank, and 

that if the external and internal paths were otherwise the same, the external path 

would always be fully ionized before the wall was punctured. In practice, 

212 
however, the weaker segments of a path such as this through different dielectric 

media break down first and allow the total voltage to become concentrated 

across the remaining, stronger, segments. Thus, in such situations, where both 

the internal and external surface flashover paths are long compared with the wall 

thickness, the presence of the wall makes little difference: punctures occur 

randomly together with external flashovers. 

FIBERGLASS WALL 

(FILAMENT-WOUND) 

(a) SWEEP: 0.5ps/div. VERTICAL DEFLECTION 90 kV/div. 

Figure 7. I9 Impulse voltage test of typical fiberglass wall. 

(a) Withstands at 110 kV and 125 kV 

(b) Partial breakdown at 141 kV and 157 kV; complete puncture 

at 173kV 

An important point to remember is that the voltage withstand capability 

of a path through or along several different insulating materials, including air, is 

less than the sum of the separate segment withstand voltages. Thus, the 

breakdown voltage of a complex structure cannot be determined reliably by 

testing its components individually and adding up the individual withstand 

voltages. Laboratory tests of the complete structure, such as those of Figures 

7.14 and 7.15, offer the best method of determining breakdown voltages and, as 

a result, the need for protection. 

7.3.1.2 Protection for Fiberglass Components 

The fiberglass punctures of Figures 7.14 and 7.15 indicate a need for 

external conductors that will inhibit internal streamer formation and provide 

external points from which streamers can originate. Such external conductors, 

called diverters, also shorten the surface flashover paths. If properly placed, 

diverters make puncture of a fiberglass skin unlikely. The following guidelines 

and tests should be followed to orient and space diverters properly on a 

fiberglass structure : 

213 
1. Orient the diverters as nearly as possible in the line of flight so that 

flashes which originally strike the diverter can reattach farther aft on 

the same diverter as the aircraft moves forward. 

2. While the fore-aft arrangement recommended in Guideline 1 is 

desirable for reattachment purposes, an adequate path must also be 

provided for Iightning current to be conducted into the airframe. 

This requirement usually calls for some of the diverters to be 

oriented perpendicularly to the line of flight. Examples of preferred 

arrangements incorporating both of these guidelines are shown in 

Figure 7.20. 

3. The surface flashover voltage from any point over the external 

surface to the nearest diverter strap must be less than the maximum 

voltage required to puncture the skin and attach to a conducting 

object beneath. Thus, the diverter must be within a maximum 

displacement distarzce from a point on the skin directly opposite the 

enclosed conductor. This relationship is defined as follows: 

Maximum Skin puncture voltage (kV) displacement < (7.19) 

distance (cm) Surface flashover voltage (kV/cm) 

This criterion should be met for lightning voltage stresses of either 

polarity applied at up to 1000 kV/ps rate of rise. 

4. The maximum voltage drop from the original attachment point to 

any other point on a lightning arc swept aft directly above a 

nonconducting surface must not exceed the skin puncture voltage 

from that point through the skin to any conducting objects inside. 

The maximum arc voltage drop will occur during a restrike 

formation in a multiple-stroke flash and can be assumed to be equal 

to or less than the free air breakdown voltage, or about 500 kV/m of 

arc length. Since this voltage will be applied from a point directly 

above the surface, instead of from an approaching leader, this 

swept-stroke environment may be more severe than one produced by 

an initial strike. 

5. The inductive voltage rise, VL, along any diverter segment carrying 

lightning-stroke currents to conducting structure must be less than 

the skin puncture voltage between the diverter and the nearest 

conducting object inside the structure. The inductive voltage rise, 

VL, may be expressed as follows: 

where L = diverter segment inductance (henries) 

IL = lightning current (amperes) 

t = time (seconds) 

In practice, L may be assumed to be 1 pH/m for most diverter 
straps or foils, and dIljdt may be assumed to be 100 000 Alps (I x 

lo1 A/s). 

NONMETALLIC 

NOSE CONE 

WITHOUT 

RADAR / 

DIVERTER STRXPS CAN BE SOLID, FOIL. 

OR SEGMENTED STRIPS; MUST O' onL'nCn \ 

TO OUTSIDE SURFACE 

\ TO DIVERTER SYSTEM 

DIVERTER STRAPS CAN BE SOLID, FOIL, 

OR SEGMENTED BUT MUST NOT IMPAIR 

SATISFACTORY RADAR PERFORMANCE 

(SEGMENTED STRIPS ARE 

RECOMMENDED FOR THIS 

APPLICATION) 

Figure 7.20 Arrangement of diverter straps on nonmetallic structural componen ts. 

The intent of Guideline 3 is to assure that sufficient diverters are utilized 

to prevent punctures resulting from initial strikes. Guideline 4 extends this 

criterion for swept strokes, and Guideline 5 is aimed at preventing punctures 

resulting from inductive voltages that arise when lightning currents flow through 

the diverter. This means that a diverter strap can, in fact, be too close to the 

internal conductor it is supposed to protect. If the maximum inductive voltage 

along the diverter is greater than the skin-puncture voltage, the diverter will have 

to be moved a minimum displacement distance away from a point on the skin 

directly above the enclosed conductor. This distance may be written quantitatively in the following way: 
If 

(Diverter inductive voltage) <(Skin puncture voltage), (7.21) 

then 

Minimum Diverter inductive voltage (kV) - Skin puncture voltage (kV) displacement = 

distance (cm) Surface flashover voltage (kV/cm) (7.22) 

This distance may also be expressed in terms of the diverter inductance and 

lightning current rate of rise as follows: 

LGH) dIL/dt (kA/ps) - Skin puncture voltage (kV) displacement = 

distance (cm) Surface flashover voltage (kV/cm) (7.23) 

If a minimum displacement distance applies, then the diverter, to be 

effective, must be positioned somewhere between the minimum and the 

maximum displacement distances from a point directly above the enclosed 

conductor. 

Clearly, the skin-puncture voltages of nonmetallic skins must be known 

before protective diverter systems can be designed. The skin-puncture voltage for 

a particular nonmetallic skin construction is obtainable by impulse voltage 

breakdown tests, and is a function of the material type and thickness, as well as 

of layup patterns, core fillers, surface treatments, etc. Examples of skin puncture 

voltages determined from 1000 kV/ps impulse tests are given in Table 7.3. 

Table 7.3 TMPULSE BREAKDOWN VOLTAGES OF TYPICAL 

FIBERGLASS SKIN MATERIALS 

Skin Construction 

Total Breakdown 

Thickness Voltage 

One fiberglass sheet 0.163 cm 

(0.064 in) 2 1 kV 

Two filament-wound fiberglass tape skins 1.27 cm 

enclosing polyimide foam filler (0.500 in) 150 kV 

Two fiberglass skins enclosing foam filler 0.99 cm 

(0.388 in) 70 kV 

Surfrlce flashover voltage levels depend most significantly on the type of 

surface finish applied. Most rainerosive paint surfaces will flash over at an 

average stress of around 10 kV/cm of surface. For example, if a diverter to 

prevent punctures to the grounding tube located directly beneath the fiberglass 

skin of the radome of Figure 7.4 was desired, and if this skin was found to have 

a puncture voltage of 50 kV and surface flashover voltage of 10 kVlcm, then, by 
Equation 7.19, the maximum displacement distance would necessarily be as 

folIows: 

Maximum 

displacement = 150kV = 15 cm (7.24) 

distance (cm) lo kV/cm 

If this diverter has an inductance of 1.0 pH, then the minimum 

displacement distance, by Equation 7.23, is as follows: 

Minimum 

displacement = (1fl) (luu BA/, 

distance (cm) lo kV/ 

Thus, it can be concluded that the diverter strap should be placed between 5.0 

and 15 cm away from a point directly over the protected conductor. 

The above analysis assumes that the internal conducting objects are at the 

same potential as the conducting structure to which the diverter is bonded. This 

assumption is valid because no lightning current flows through the internal 

conducting objects, and thus there are no potential differences between them 

and a common conducting structure. 

The protection of a nonmetallic structure is most important when the 

structure contains an integral fuel tank. If this combination is located in Zone 1 

or Zone 2, a protective diverter system to prevent punctures and the ignition of 

fuel should be applied. Of course the number or size of diverters that may be 

required can be minimized by making as many as possible of the internal fuel 

tank components of nonconductive materials. By so doing, some objects within 

the tank which could otherwise be the source of streamers and eventual 

attachment points will be eliminated. Any conducting parts that must remain 

inside should then be positioned as far away from the outside skins as possible, 

and they should be devoid of sharp edges, as discussed in Chapter 6, Section 6.7. 

Plumer (Reference 7.1 4) discusses a particularly challenging situation: an 

entire wing, including an integral fuel tank, is made of fiberglass; moreover, a 

metal capacitance probe, vent pipe, end panels, and filler cap are present. The 

tank is shown in Figure 7.21. The vent line and end plates of this tank could be 

made of nonconducting materials. The capacitance-type fuel quantity probe, on 

the other hand, must of necessity be made of conducting material, and the probe 

should be positioned as far as possible from the top and bottom skins. 

Figure 7.21 also shows the protective diverters that were designed for the 

bottom and top skins of this tank. Diverters will usually be required if there is 

any conducting object within the tank--and sometimes even if there are no 

conducting objects inside. If the puncture and surface flashover characteristics of 

the enclosing skin materials are known, the diverter arrangement can be designed 

using Equations 7.19 to 7.23. Frequently, however, when newer materials are 

used, puncture and surface flashover information is not available. In this case, a 

section of the proposed skin material can be given a high-voltage test to 

determine the required diverter spacings, as illustrated in Figure 7.22. The skin 

217 
material to be tested should have the same surface treatments and thickness as 

those proposed for the tank. Conducting objects simulating those in the actual 

tank shouId be positioned appropriately beneath the tank skin and together 

serve as one of the test electrodes. The other electrode should be a rod of at least 

1 cm diameter, positioned about 1 cm above the skin, as shown in Figure 7.22. 

Test voltage at up to 1000 kV/ps rate of rise should be applied between these 

electrodes. 

BOND 

WlNG 

NONMETALLIC WlNG 

THER COMPONENTS NOT SHOWN) 

FUEL FILLER CAP: MUST BE 

DIVERTER STR LIGHTNING-PROTECTED VARIETY 

AND ADAPTER ELECTRICALLY 

DED TO DIVERTER STRAP 

WlNG TIP 

DOUBLER 

15cm S55cm 

EL TANK END PLATE SHOULD BE 

8 cm OF FUEL PROBE MADE OF NONCONDUCTIVE MATERIAL 

ON OF CAPACITANCE-TYPE 

FUEL PROBE 

Figure 7.21 Protective measures for integral fuel tank within all-fiberglass wing. 

S = DIVERTER DISPLACEMENT 

STATIONARY ROD 

ELECTRODE 

DIVERTER STRAP 

L I 

SIMULATED CONDUCTING 

OBJECT 

Figure 7.22 Test setup for experimental determination of maximum diverter 

displacement distance, S. 

Proper testing is begun with the diverter strap directly beneath the test 

electrode. Since all discharges in this position will terminate on the diverter, the 

diverter can subsequently be moved from a point directly beneath the electrode 

to a spot where S equals 1 em, for example. When the test is repeated, arcs will 

flash down to the nonmetallic surface and along it to the diverter. The test 
should be repeated with the diverter moved another centimeter away, and so on, 

until a puncture occurs and the arc terminates on the conducting object beneath 

the skin instead of on the diverter. When this point has been reached, the 

diverter is no longer effective, and the largest spacing, S, at which no punctures 

occurred defines the maximum displacement distance. If more than one diverter 

is to be used, the diverters should be arranged so that the conducting object is 

between them. In this case their maximum separation must not exceed 2S, as 

shown in Figure 7.23. 

DIVERTER 

I 

DIVERTER 

STRAP r S S 4 4 STRAP 

I 

I 

NONMETALLIC SKIN I 

S = MAXIMUM DIVERTER DISPLACEMENT 6- CONDUCTING 

OBJECT 

Figure 7.23 Arrangement of multiple diverter straps for maximum separation. 

One such test was run by Plumer on a section of the wing tank shown in 

Figure 7.21, with the aluminum vent line imbedded in the upper tank wall and 

the capacitance probe along its side wall, as recorded in his report. The diverter 

was found to be effective within a spacing, S, of 6.7 cm; but a spacing of 9.2 cm 

caused the arc to puncture the skin and terminate on the'vent line beneath. 

Thus, allowing for a margin of safety, the diverter should be no more than 5 cm 

away from a line directly above the vent tube. 

From the same test performed for the capacitance probe, it was found that 

the diverters could be over 15 cm away before puncture occurred. The reason 

for this difference is that the capacitance probe was farther away from the upper 

skin; hence more voltage was required to puncture the skin and allow 

termination on the probe. Since the probe was angled, however, and did come 

closer to the tank surface farther outboard, it can be concluded that the 5 cm 

criterion determined for protection of the vent tube should also be applied for 

the capacitance probe. Since in this tank the capacitance probe and vent tubes 

are about 25 cm apart, two diverter straps (or perhaps a single strap 15 cm wide) 

would be required to provide adequate protection. Straps would be required on 

both the top and bottom of the tank for prevention of punctures to the 

capacitance probe, but only on the top surface to prevent punctures to the vent 

tube, since the vent tube is located just beneath the top surface. Fabrication of 

the vent tube from nonconducting material would therefore eliminate the need 

for a strap as wide as 15 cm. 

Each diverter strap utilized should adhere to the outer skin surface and 

should be electrically bonded to the wing tip doubler or diverter assembly at the 
outboard end and to those metallic structural components at the inboard end 

which are capable of conducting lightning current on through the airframe. 

Protection against the lightning effects on fuel systems discussed in 

Chapter 6 should also be accomplished. 

A variety of conductors for diverter applications may be considered. 

Selection of the one to use depends on such factors as weight, obstruction of the 

airstream, single- or multiple-strike capability, radar antenna pattern interference, and cost. Brief descriptions of typical diverters follow. 

External Diverter Strap 

The external diverter strap, a simple metal bar, as shown in Figure 7.24, is 

mounted on an external surface. This type of diverter is designed to receive more 

than one stroke without sustaining significant damage and is intended, therefore, 

for permanent applications. External diverter straps are commonly used to 

protect radomes and wing tips, which receive frequent strikes. The straps are 

usually made of aluminum, with a rectangular cross section of at least 0.1 crn2 

(20 000 circular mils) to permit conduction of 200 kA, 2 x lo6 A~-S stroke 

currents without excessive temperature rise or elongation. In order to remain 

intact by minimizing melting, most straps now in use have areas larger than 0.1 

crn2. Thicknesses of up to 0.64 cnl (0.25 in) and widths of up to 1.27 cm (0.5 

in) have been used. A common design is 0.32 cm (0.125 in) thick by 0.95 cm 

(0.375 in) wide. 

Figure 7.24 External diverter strap. 

The straps are usually attached to the wall with steel screws or bolts; the 

larger dimensions allow holes and countersinks to be made in the strap without 

reducing to less than 0.1 cm2 the strap cross section. In nose radomes, the 

threaded portions of the screws are often surrounded by plastic inserts to 

prevent sparking from the screws to the radar antenna during lightning 

encounters. 

220 
Since the metal straps are installed on the outside surface, there may be a 

small drag penalty. This drag can be minimized by orienting the straps parallel to 

the airstream or by shaping the cross section of the straps aerodynamically. 

Internal Strap with Protruding Studs 

Mounting of external straps flush within the nonmetallic wall not only 

creates tooling and manufacturing problems but also degrades the strength of the 

wall because of stress concentrations along the grooves required for flush 

mounting. Therefore, metal straps or tubes have sometimes been mounted on 

the inside surface structure, with metal studs protruding through the wall to 

serve as lightning attachment points, as shown in Figure 7.25. 

/ '/ 

BENEATH SURFACE 

Figure 7.25 Internal strap with protruding studs, 

Although this type of diverter installation may reduce the aerodynamic 

drag created by external mounting, it fails to take advantage of the insulation 

capability of the nonmetallic wan, and, in fact, may even intensify the electric 

field stress through the wall, promoting punctures if the protruding studs are too 

far apart. In general, the thicker the wall, the farther apart may be the studs. In 

many cases, 30 cm (1 ft) is close enough, but individual cases should be 

evaluated by high-voltage tests similar to that of Figure 7.22, or by tests of the 

complete structure similar to those of Figures 7.1 4 and 7.1 5. Appropriate test 

conditions for the latter are described in Lightning Test Wavefornzs for 

Aerospace Vehicles and Hardware. The possibility of puncture to an internal 

strap or tube can be minimized if the portion of it between the protruding studs 

is covered with an insulating sleeve. 

When lightning currents travel through any internal conductor, of course, 

there is also the possibility of the inductive voltage rise causing sparking to 

nearby conductors which are not carrying current. Covering of the lightning 

conductor with an insulating sleeve in the vicinity of nearby conductors will 

reduce the possibility; the internal strap or tube can be used without such a 

sleeve if the spacing between the strap and adjacent conductors is great enough 

to prevent sparking. For purposes of estimating what this spacing should be, 

22 1 
assume that the diverter strap has an inductance, L, of 1.0 pH/m of length. 

Assume a maximum lightning current rate of rise of 100 kA/ps, and determine 

the air-gap impulse sparkover voltage from Figure 6.39 or from impulse voltage 

tests of the particular configuration. Then make sure that the inductive voltage 

rise 

is less than the gap sparkover voltage. It is prudent to add a safety margin of 50% 

or more to the calculated inductive voltage when making this comparison. 

Thin Foil Strips 

Thin foil strips, usually made of aluminum 0.008 to 0.040 cm (0.003 to 

0.015 in) thick, are capable of conducting moderate strokes or of melting and 

vaporizing on severe strokes. In the latter case they leave an ionized channel 

through which subsequent currents in the flash may travel. Because they are 

thin, foil strips have good aerodynamic characteristics, but they usually provide 

protection against only one severe flash. The thin foil strips are generally used 

when aerodynamic drag requirements prevent thicker straps from being installed 

on external surfaces. A typical strip is shown in Figure 7.26. 

Figure 7.26 Thin foil strip. 

The skin of the structure must be thick enough to withstand safely the 

blast pressure that occurs if the strip vaporizes. Thick foil strips may also be 

used; however, they release more energy during vaporization than the thinner 

foil strips and, therefore, require a thicker radome outer skin. Wide foil strips are 

preferred because they distribute the explosive forces over a larger surface area. 

A 0.008 cm (0.003 in) thick by 0.64 cm (0.375 in) wide aluminum strip 

will protect a 0.1 cm (0.040 in) thick fiberglass wall from one 200 kA, 2 x lo6 

A~-S stroke and associated intermediate or continuing current, but the strip will 

have been vaporized and must be replaced. 

Thus, the foil strips should not be used for applications where more than 

one strike might occur in the same flight. This caution means that they should 

not be used to protect fiberglass fuel tanks or other structures in which a 

222 
puncture might result in a flight safety hazard, for, contrary to the old adage, 

lightning does strike the same place twice, even in the same flight! Thin foil 

strips are appropriate only for inexpensive protection of fiberglass fairings, wing 

tips, fin caps, or other non-flight-critical components. 

Radio frequency interference may be created if the strips are improperly 

grounded to metallic structures, or if fatigue cracks occur in the strips, which 

may result in a broken electrical connection. Fatigue cracks are often caused by 

adhesive bonding of a single strip to a surface with several panels or segments 

that may move with respect to one another. Cracks may cause severe damage 

when lightning currents flow across them, as illustrated in Figure 7.27. This 

figure shows a 2.54 cm (1 in) wide by 0.018 cm (0.008 in) thick strip of 

pressure-sensitive aluminum foil tape after a test in which a (moderate) 54 kA 

stroke, followed by a 90 A continuing current, was discharged into the strip at 

Point b. The continuing current conveyed 18 C of charge into the strip. A 

razor-thin cut had been made in the strip at Point c; the damage which resulted 

as the current arced across this confined cut was more extensive than that 

produced by the same current when it entered the strip at Point b. Unblemished 

portions of the strip itself, however, were able to conduct these currents without 

damage. In fact, a previous test had delivered a 54 kA stroke and 187 C of 

charge to Point a, causing only the erosion shown at this point. The 

cross-sectional area of foil necessary to conduct more severe stroke currents can 

be determined from the curves of Figure 7.3. 

Figure 7.27 Damage produced by moderate stroke and continuing current in 

0.018 (0.008 in) thick, 2.54 cm (1 in) wide aluminum foil with 

crack. 

Point a - erosion resulting from previous stroke 

Point b - arc entry 

Point c - location of crack 

223 
Metallic Coatings 

Coatings are thin metallic foils, paints, or plasma flame sprays that 

completely cover the nonmetallic structure to make the surface entirely 

conductive and enable it to accept strikes anywhere. Such coatings, however, 

may suffer considerable vaporization unless they are prohibitively thick. They 

are recommended only in cases where diverters are not practical. 

Woven Wire Fabrics 

Flexible metallic fabrics woven of fine aluminum, copper, or stainless steel 

wires about 80 to the centimeter, as shown in Figure 7.28, are usually more 

effective in resisting damage than are coatings because woven wire fabrics permit 

gases from arcing to be vented through the holes in the mesh instead of 

containing them, as do continuous foil and paint coating. Development of these 

fabrics came about in response to a need for the effective protection of advanced 

composites. Their use is discussed more fully in that context in Section 7.3.2.5. 

Figure 7.28 Woven wire fabric. 

224 
Segmented Strips 

Segmented strips, sometimes called "button strips," are one of several 

developments in lightning protection technology. These segmented strips, 

developed by Arnason and Cassell of Douglas Aircraft Company (Reference 

7.1 51, consist of a series of thin, conductive segments, or "buttons," fastened to 

a strap of resistive material, which is then bonded to the surface to be protected, 

as shown in Figure 7.29. Instead of attempting to conduct lightning currents, 

these segmented strips provide many small air gaps that ionize and spark when 

an intense electric field is applied along the strip, as is the case when a lightning 

leader approaches. Since the small gaps are close together, the resulting 

ionization is nearly continuous. Thus the segmented strips provide a conductive 

path, initially for streamers to follow outward from the aircraft and, after these 

reach the leader, for return stroke and other currents to flow into the airframe. 

The construction of the strip and breakdown process are pictured in Figure 7.30. 

Figure 7.29 Segmented strips. 

Laboratory tests such as the one illustrated in Figure 7.31 show that the 

segmented diverters are nearly as effective in intercepting strikes and preventing 

punctures of fiberglass walls as are the solid conductors. Since ionization of the 

strips requires that an electric field be able to exist in the direction of their 

length, the segmented strips will perform best if oriented normal to conductive 

portions of the airframe rather than parallel to such structure. Also, the strips 

should not be positioned parallel to a conducting object lying beneath the 

nonmetallic surface. For example, the grounding tube within the radome of 

Figure 7.4 will reduce the electric field gradient along its length to nearly zero, 

so that a segmented diverter positioned near the tube may not ionize and an 

approaching flash may puncture the radome wall and attach, instead, to the tube 

itself. Thus, segmented diverters should be positioned on either side of the 

grounding tube but not directly over it. This positioning is also necessary to 

prevent inductive voltage along an ionized strip from puncturing the radome 

wall, as discussed in Section 7.3.1.2. 

The maximum spacing between segmented diverters--or the minimum 

permissible spacing between them and underlying conductors--is dependent, 

among other things, upon the amount of voltage required to ionize the 
/ RESISTANCE MATERIAL CONNECTING SEGMENTS 

DIELECTRIC SUBSTRATE 

RESISTANCE MATERIAL 

\CONDUCTIVE SEGMENTS 

~ . I, SPARKS 

e CONTINUOUS 

IONIZED PATH 

FOR STROKE 

AND CONTINUING 

CURRENTS 

(b) 

Figure 7.30 Segmented strip design and breakdown process. 

(a) Segmented strip design 

(b) Segmented strip breakdown and conduction process 

segmented strips. Ideally, a very low voltage would be desirable--a voltage even 

lower than that required to ionize a path along the bare surface itself and much 

lower than that required to puncture the surface. While the developers report 

(Reference 7.16) that the strips ionize at about 10 kV/30 cm, they do not state 

what rate of voltage rise this applies to, nor have any oscillographic 

measurements showing test voltage rate of rise and strip ionization voltage yet 

226 
been reported in the literature. The authors, meanwhile, have found the strips to 

ionize at voltages of up to 70 kV/30 cm (1 ft) when tested with a voltage rising 

at 200 kV/ps. The conductivity of the surrounding surface and the proximity of 

underlying conductors undoubtedly play a role in the function of the strips, but 

no laboratory investigations of these factors have yet been reported either. 

Figure 7.31 Simulated lightning strike to a segmented diverter. 

Streamers propagate up diverter and outward to join oncoming 

leader. 

227 
The developers do report (Reference 7.17) that the segmented strips 

remain undamaged after passage of a 170 kA stroke and 200 C continuing 

currents. This capability has been confirmed bv the authors, who have found the 

strips capable of sustaining two strokes of 180 kA and 2 x lo6 A~-S (Component 

A) before destroying their ability to ionize another time. The authors have also 

found the segmented strips to be capable of sustaining a much larger number of 

the more common 50 kA strokes. Other investigators, including Conti and Cary 

(Reference 7.1 8), have reported the strips to be severely damaged after one 200 

kA discharge or one "high coulomb" continuing current discharge has been 

delivered, a finding that has aroused controversy over the effectiveness of the 

strips. The continuing current charge (500 C) delivered by Conti and Cary, 

however, is more charge than that conveyed by any natural lightning discharge 

yet measured. 

The segmented strip design might be modified to lower the ionization 

voltage or to improve durability; however, the basic principle is sound, and there 

are many applications where the presently available strips will afford lightning 

protection improved over other methods, while not interfering with radar 

performance. The present lack of strip performance data simply means that the 

proper location of the strips can not yet be determined analytically but, instead, 

must be accomplished with the aid of laboratory tests of candidate arrangements 

on the particular structure being protected. 

The selection of protective devices and the arrangement of them on a 

fiberglass structure comprise a major part of the protection task. The best design 

can be completely undone, however, if other, seemingly innocuous, features of 

the structure are left unchanged. Figure 7.32 illustrates some particularly 

troublesome situations. They usually arise when practices perfectly acceptable in 

metallic structures are continued without modification in nonmetallic structures. 

The several ways of providing protection for the fiberglass wing tip of 

Figure 7.30 should include the following: 

Remove paint from the diverters 

Electrically bond the lamp housing directly to the diverters 

Enclose the lamp power wires in a conduit or braided shield inboard 

to the metal rib 

Electrically bond the diverters to metallic structure. 

7.3.2 Advanced Composites 

Advanced composites, consisting of boron- or graphite-reinforced plastics, 

are being used increasingly often as replacements for aluminum load-bearing 

applications in aircraft. The high strength-to-weight characteristics of these 

materials make them attractive for structural applications, but because they are 

nonmetallic, they are inherently more vulnerable to lightning effects. 

In contrast to aluminum, which by virtue of its electrical and thermal 

conductivity suffers few direct effects of lightning and which provides excellent 

electromagnetic shielding, the advanced composites are neither good conductors, 

228 
nor good insulators, nor good providers of shielding against electromagnetic 

fields. If these materials are used in lightning-strike zones, it is possible that they 

will require some form of protection. The questions which must be answered are 

these: Just how vulnerable are particular advanced composites to lightning? and 

What protection, if any, may be necessary? 

LAMP HOUSING NOT 

BONDED TO DIVERTER 

ALLOWS CURRENT TO 

ENTER POWER WIRES 

Figure 7.32 Inadvertent causes of disaster. 

Unlike fiberglass-reinforced plastics, which have no electrical conductivity, 

the boron filaments or graphite fibers which are imbedded in the matrix are 

resistive conductors and will conduct some lightning currents. However, the fact 

that, like any resistor, they heat up when electric current flows in them causes a 

serious problem. For example, typical boron and gra hite composites have B resistivities of 4 to 9 x a-cm and 0.9 to 1.1 x 10 acm, respectively, as 

compared to 2.8 x a-crn for aluminum. This means that the composites 

will try to absorb between 100 and 1000 times more energy from the lightning 

current than will aluminum. Moreover, the materials may not be able to dissipate 

this much heat without some change in or destruction of physical properties. 

Simulated lightning-strike tests of typical composite laminates have demonstrated this problem, with results similar to those shown in Figures 4.12 and 

4.1 3. 

A number of careful studies of the electric current conduction and damage 

mechanisms within advanced composites have been made in order to learn the 

quantitative relationships between current intensity and the amount of 

mechanical strength degradation. It has been confirmed that damage results 

primarily from electric current flow through the boron filaments or graphite 

229 
fibers, but the mechanisms differ for these two types of composites. The current 

flow processes and degradation mechanisms for each type are reviewed in the 

following paragraphs. 

7.3.2.1 Boron Composites 

Because of initial interest in boron composites for aircraft structural 

applications and because of their higher resistivities, the lightning current 

degradation mechanisms for boron were studied first. Fassell, Penton, and 

Plumer (Reference 7.19) began with a study of the degradation process in a 

single boron filament. These filaments were formed by deposition of boron on a 

0.13 mm (0.0005 in) diameter tungsten filament called the substrate. The boron 

sheath, which provides the filament's strength, was about 1.3 mm (0.005 in) in 

diameter. The tungsten substrate is a resistive conductor, but the resistance of 

the boron is much greater by several orders of magnitude. Thus, electric currents 

entering a boron composite will want to flow in the tungsten substrates instead 

of in the boron. Tungsten has a positive temperature coefficient of resistivity, 

which means that, as its temperature increases, its resistivity increases as well, 

and it becomes even hotter as current flows. When temperature rise is calculated 

as a function of current action integral by Equation 7.5, the temperature rise 

curve of Figure 7.33 results. The quantity $12dt is, of course, equal to the unit 

energy dissipation in the substrate (per ohm of resistance) and is dependent 

solely upon the lightning current amplitude and waveform. 

Figure 7.33 shows that very little temperature rise occurs as $12dt is 

increased to 20 x A2-s. However, increases beyond this value cause 

moderate temperature increases, and increases beyond 100 x A2-s result in 

extreme temperature rise. Fassell, Penton, and Plumer injected currents beyond 

100 x A2-s into single filaments and found them to become severely 

cracked and to lose all mechanical strength. Correspondingly, filaments exposed 

to pulses with $~~dt values of 100 x lo6 A2s or less exhibited little or no 

evidence of degradation. The mechanism of filament failure was found to be 

transverse and radial cracking of the boron sheath, presumably caused by 

excessive thermal expansion of the tungsten substrate. (Examples of cracked 

filaments are shown in Figure 4.12.) Thus, the calculated temperature rise is a 

significant indication of filament failure onset. 

In comparison to natural lightning currents, it is noted that when 

distributed over several thousand filaments, a stroke of several thousand amperes 

may provide only 1 A of current per filament. If this current exists for 100 ps, it 

will indeed provide an $12dt of 100 x A2-s in a single filament; if this 

current exists for 20 ps, it will provide an jI2dt of 20 x A2-s in a single 

filament, etc. 

WhiIe the previous example illustrates the manner in which lightning 

currents may cause mechanical property degradation of composite materials, it 

also illustrates the relatively flat region in which increasing dues of $~~dt do 

not result in corresponding increases in temperature rise. For protection design, 

230 
this fact emphasizes the desirability of distributing lightning currents throughout 

as many filaments as possible. 

To explain the process of current flow into the many boron filaments in a 

composite, Penton, Perry, and Lloyd (Reference 7.20) formulated an electrical 

model of two boron filaments in a resin matrix, as shown in Figure 7.34. Each 

filament is made up of incremental portions of the tungsten substrate resistance. 

These researchers found that the filament-to-filament breakdown voltage is 

about 300 V and that the insulation breakdown at this level is represented by 

the back-to-back Zener diodes, which have a reverse conduction (Zener) voltage 

of 300 V. 

A lightning current flowing in the outer filament, they found, would 

remain in this fiIament until the resistive voltage rise dong it exceeds the boron 

and resin breakdown voltage represented by the Zener diodes, at which point the 

diodes in the electrical model conduct and permit current to enter the 

neighboring filament. In most cases this would be a filament in the next 

laminate. While some material strength degradation undoubtedly occurs from 

the breakdown of the resin and boron sheaths, the major loss of strength was 

23 1 
found to be a result of temperature rise in the fdaments and cracking of the 

boron sheaths along the entire length of current flow. The appearance of a 

filament so damaged is shown in Figure 4.12. 

END VIEW BORON FILAMENT 

Figure 7.34 Electrical model of two boron filaments within a laminate. 

When the filaments in an outer layer fail by the transverse, or axial, 

cracking, the resistances of their tungsten substrates also increase (by cracking) 

and so does the voltage rise along this layer, until the breakdown level of the 

boron sheaths and resin between it and the next layer of filaments is reached. 

When this happens, the current enters this next layer of filaments, and the 

process continues. The depth of penetration, of course, depends on both the 

amplitude and duration of the lightning current and the number of filaments 

into which this current can distribute. 

In addition to the current conduction effects just described, the high 

temperature and blast forces from the arc have been found to cause substantial 

damage to a boron composite at the point of attachment; this, in fact, may be 

the only damage which is visible. Yet, because loss of strength also results from 

cracking of the embedded filaments, there may be little or no visible change in 

the outer appearance of a boron composite away from the arc attachment point. 

An example of this is the 16-ply boron composite panel (pictured in Figure 7.35 

[Reference 7.211) after receiving an 88 kA simulated stroke. The action integral 

of this stroke was 0.2 x lo6 A~-s, one that might be expected from a restrike to 

232 
a wing or fuselage surface in Zone 2A. A damaged area approximately 7 cm wide 

and 10 cm long is visible in Figure 7.35; yet the photomicrograph of Figure 7.36 

(Reference 7.22) shows that nearly all the filaments in a cross section cut from 

the visibly undamaged area near the base of this panel are cracked. This cracking 

resulted in a 50% reduction in the shear strength of this region. 

A hole was blown completely through the panel of Figure 7.35 at the arc 

attachment point by this 88 kA stroke. Strokes of somewhat less amplitude and 

action integral would not puncture a hole; strokes of higher energy would 

produce larger holes in this 16-ply boron panel. 

The amount of damage to expect in unprotected panels of other 

thicknesses depends primarily upon the action integral of the lightning current 

and the number and orientation of the plies in the composite. An estimate of the 

visible damage to typical boron laminates as a function of lightning current 

amplitude and number of plies is presented in Figure 7.37. 

The area throughout which invisible damage will occur in a boron 

composite depends on the cross-sectional area available for conduction of 

current away from the arc attachment point. This area becomes greater with 

distance away from this point, and can be estimated from Table 7.4 in the next 

section. 

7.3.2.2 Graphite Composites 

Graphite yarns provide the reinforcement in graphite composites. These 

yarns are made up of a great many graphite fibers, each much thinner than a 

boron filament. Since the melting point of graphite (3735 "C) is much higher 

than that required to cause melting and pyrolysis of the surrounding resin 

(315 "C), damage to graphite composites occurs by pyrolysis and, usually, 

ignition of the resin, which in turn causes gross delamination. This result is in 

contrast to the initial destruction of the reinforcing fibers which occurs in boron 

composites. The burning of the resin then leaves the entire composite in 

disarray, as shown in Figure 4.1 3. 

The spectacular nature of the graphite composite damage mechanism and 

the gross disarray in which the composite is left might lead one to worry more 

about exposing graphite composites to lightning than boron composites. The 

reverse is actually the case, however, because, for the same cross-sectional area, 

the graphite composites have a much lower resistivity than has boron and 

therefore can conduct much more lightning current before dangerous temperatures are reached. 

Arcentry damage to graphite is somewhat similar, though less extensive 

than that produced by comparable arcs striking boron composites. Figure 7.38 

(Reference 7.23) shows a typical 12-ply graphite composite panel after a 98 kA, 

0.25 x lo6 A~-S stroke. Photomicrographic analysis and mechanical strength 

tests showed no damage other than that which is visible at the attachment point. 

Figure 7.39 presents an estimate of the amount of damage to expect in 

unprotected graphite laminates under other conditions. 
1 

and 

.35 Visible damage to 16-ply unprotected boron compos 

88 kA, 0.2 x lo6 A~-S 

Arc attachment to middle of panel 

Current exit from front and back plies at bottom 

iterature abounds with simulated lightning test data for 

te, but since most of these tests were conducted 

234 

ite 

bo 

be 

panel. 

th bor 

fore 1 
Figure 7.36 Boron filaments in visibly undamaged region at base of panel in 

Figure 7.35. 

Filaments are extensively cracked. 

Current has entered all 16 plies. 

Front of panel is at top. 

standardization of test conditions was attempted, the tests have produced results 

which, while valid for the particular cases of interest, are too voluminous and 

diverse to normalize and reproduce here. From tests in our own laboratory, 

however, we have learned that composite strength degradation is closely related 
8 12 16 20 24 28 32 36 

NUMBER OF PLYS 

Figure 7.37 Estimated visible damage to unprotected boron composites. 

to the lightning current action integral, even when this quantity is delivered over 

a period of up to one second, as it is in continuing currents. Table 7.4, which is 

based on these tests, compares the capability of boron and graphite composites 

to conduct lightning currents with and without degradation of their mechanical 

strength properties. 

Since this table has been prepared from a variety of tests, the numbers in it 

must be considered approximate. When results were compiled, a range of about 

one order of magnitude of action integral was found to separate samples which 

experienced no discernible loss of strength from those which lost 50% or more 

of their former tensile or shear strength. From Table 7.4 it is evident that a 

cross-sectional area of about 16 crn2 of boron composite is necessary to conduct 

a severe stroke of 2 x lo6 A~-s; yet only 2 or 3 cm2 of graphite composite 

would be needed to conduct the same current. 

In many cases it may not be practical or necessary to prevent localized 

damage at the arc attachment point, but it is important to assure that the 
surrt 

dam 

PUrF 

'.38 Visible damage to 12-ply unprotected graphite composit 

98 kA, 0.25 x lo6 A~-S 

Arc attachment to middle of panel 

Current exit from front and back plies at bottom 

ng composite can conduct away the lightning current with 

in the process. The information of Table 7.4 can be use 

s follows: 

237 

lUt t 

for 

~eing 

this 
NUMBER OF PLYS 

Figure 7.39 Estimated visible damage to unprotected graphite composites. 

For graphite composites, the damage extends only throughout 

the visibly damaged region. 

Table 7.4 COMPOSITE j12 dt CONDUCTION CAPABILITY 

Composite 11' dt (AZ s) 

Cross-Sectional Boron Graphite Area 

Non- Noncm2 in2 degraded Degraded degraded Degraded 
1. Assume that only half of the surrounding composite is available to 

conduct current, as shown in Figure 7.40. The reason for this is that 

the lightning current will want to flow in one predominant 

direction--toward its eventual exit point somewhere else on the 

aircraft. 

2. In this case the relationships among damage, radius, cross-sectional 

area, and composite thickness are given by the following: 

where d = the necessary skin thickness (cm) 

A = cross-sectional area required to conduct the required jI2dt, 

from Table 7.4 (cm2) 

r = radius of permissible damage (cm) 

GENERAL DIRECTION 

OF CURRENT FLOW 

IN AIRCRAFT 

Figure 7.40 Composite thickness vs area of damage from conducted currents. 

As an example, consider fuselage skin in Zone 2A to be made of graphite 

composite. It is desired that the area of strength loss extend no more than 5 cm 

outward from an arc attachment point. What thickness must the composite be to 

meet this requirement? 

The SAE report Lightning Test Waveforms and Techniques for Aerospace 

Vehicles and Hawlware defines the lightning current action integral for Zone 2A 

as 0.25 x lo6 A~S. Table 7.4 shows that 1.61 cm of graphite composite will 

conduct this action integral without being structurally degraded. Thus, from 

Equation 7.26, the necessary composite thickness needed to limit the degraded 

region to within a 5 cm radius from the arc attachment point may be calculated 

as follows: 

239 
If a boron composite were to be used, its thickness would have to be at 

least ten times greater to restrict the damage radius to the same 5 cm. The 

thicknesses of either composite would have to be even greater, of course, if the 

composites were located in Zone 1, where an action integral of 2 x lo6 A~-S is 

to be expected. In such a case, it is likely that either the area of damage or the 

composite thickness required to reduce this damage would be intolerable. And in 

such a case the addition of some protective device(s) capable of keeping some or 

all of the lightning current out of the composite should be considered. The 

coatings and diverter strips for fiberglass described in Section 7.3.1.2 may also 

be applicable for advanced composites; and several coatings were developed 

specifically for use on composites. The most promising of these are discussed in 

Section 7.3.2.5. 

7.3.2.3 Problems at Joints 

In addition to lightning currents causing direct effects by entering or 

flowing in composite materials, these same currents may also cause damage at 

the places where they leave the composite--for example, a place where the 

original panel is joined to another one or to substructure such as ribs and spars. 

Two basic methods of joining composites are in use. One of these is bonding, in 

which the composite is joined (glued) with an adhesive to another member. The 

other method is with bolted joints, in which bolts, rivets, or other mechanical 

fasteners are employed to hold two members together. 

Unfortunately, most adhesives used in bonding are highly resistive and by 

themselves are not able to conduct lightning currents. Instead, arcing occurs 

within the adhesive, releasing sufficient gas pressure to make the structure 

literally "come unglued." Either the adhesive must be made electrically 

conductive or a separate conducting path must be provided to carry lightning 

currents across the joint. A conductive paint or foil may be able to accomplish 

this if it is necessary only to transfer lightning current from one composite panel 

to an adjacent one. But if the composite is also fastened to substructure, 

electrical contact must be made to this substructure as well. A metal foil 

wrapped around the edge of the composite panel may suffice; however, such a 

foil must be held securely to the composite surface so that the magnetic force (a 

problem of Figure 7.8) does not cause the foil to pull away. If bolts or fasteners 

are used, they may provide the necessary electrical path, whether intentional or 

not, but the concentration of current near them may degrade the composite's 

strength in this important region. 

Relatively little work has yet been accomplished to evaluate the degree of 

strength loss caused by lightning currents in composite joints; however, further 

research is now in progress. Schneider, Hendricks, and Takashima (Reference 

7.24) are studying the effects of lightning currents on typical bonded and bolted 

joints of graphite composites, and their preliminary findings show that bonded 
joints suffer more damage from lightning currents than do bolted joints. Other 

tests (Reference 7.25) have also been made on various joints, with similar results. 

The reader is referred to these sources for further information on the specific 

joints researchers have evaluated and the results they have obtained. 

7.3.2.4 Electromagnetic Field Penetration 

The importance of keeping the electromagnetic fields produced by 

lightning currents away from the aircraft's electric wiring is by now clear. In 

conventional aircraft this is conveniently and successfully accomplished by the 

aluminum skins, and the electrical system designer benefits greatly from this 

assistance heretofore given him by the structures designer. Substitution of 

composites for aluminum, however, poses an additional threat to the electrical 

systems because composites do not possess the excellent shielding property of 

aluminum, with the result that the electrical system designer may now have to 

provide his own protection against these fields. 

The relative shielding effectiveness of composites as compared with various 

metals has been measured for magnetic (H) fields by Fisher and Fassell 

(Reference 7.26) by establishing an H-field of known intensity on one side of a 

composite surface and measuring the amount of it reaching the other side. 

Continuous sine-wave fields at several different frequencies were applied to the 

top side of a test sample in the arrangement of Figure 7.41, and the shielding 

capability of the sample was determined by measuring the voltage induced in a 

receiving coil on the other side by that portion of the field which penetrated it. 

The current, I, in the transmitting coil was kept constant during the tests to 

make sure that reflected fields did not change the level of field radiated in the 

first place. The amount of H-field attenuation was then determined by the usual 

expression: 

v 1 Attenuation = 20 Log - 

v2 

where V1 is the voltage induced in the receiving coil when no test sample is 

present and V2 is the voltage when a sample of composite or other material is 

placed between the coils. The metal enclosure was used to avoid fringing effects. 

Figure 7.42 shows the H-field attenuations for boron and graphite composites by 

themselves, and Figure 7.43 shows the improvements in shielding effectiveness 

gained by application of conductive coatings (Reference 7.27). 

The measurements of the shielding effectiveness of panels covered with 

flame-sprayed aluminum and aluminum foil indicated that most of the shielding 

came from the aluminum. Values of 16 dB and 20 dB were measured at 0.5 MHz 

and 1.0 MHz for 2.1 mm (0.083 in) boron with a coating of flame-sprayed 

aluminum. The thickness of the aluminum was about 0.4 mm. 

Many investigations into the shielding effectiveness of conductive coatings 

of enclosures have, of course, been made for other purposes. One is that of Uhlig 

(Reference 7.28), who evaluated several conductive coatings applied to 1.2 m 
cubic boxes made of wood. The coatings tested were the following: 

1. Flame-sprayed aluminum approximately 0.18 mm (0.007 in) thick 

2. Aluminum foil 0.025 mm (0.001 in) thick 

3. Two coats of conducting paint (300600 R/square) 

ELECTROSTATIC SHIELD 

- - -. . . - - - - - - T~*.lr.ll.RI.I~ -1.. 

SIGNAL 

GENERATOR 

OSCILLOSCOPE I 

3 

TEST SAMPLE L---_I 

I 1111111111,111 

1 

/ CURRENT PROBE 

\ RECEIVING COIL I I L--- -I \ [ ELECTROSTATIC SHIELD I 

METAL 

ENCLOSURE 1 I 

Figure 7.41 Arrangement for shielding effectiveness tests. 
FREQUENCY - MEGAHERTZ 

Figure 7.42 H-field shielding effectiveness of boron and graphite composite 

materials. 

-. . 

FREQUENCY - MEGAHERTZ 

Figure 7.43 H-field shielding effectiveness of boron and graphite composites 

with protective coatings. 

243 
The box was placed in an impulse-type H-field with amplitude and waveform as 

shown in Figure 7.44(a) (Reference 7.29). This field has a rise time of 

approximately 2 ys and a decay time of approximately 60 ps, such as would be 

produced by a lightning-stroke current with the same rise and fall times. 

(Component F of the SAE report). Its amplitude of 150 A/m is of the order 

produced by a typical lightning-stroke current (discussed in Chapter 9). 

ONE LAYER OF ALUMINUM FOIL 

u LLI--l _L_L_LI 

CONDUCTIVE CARBON 

1 I 1 -0.22 ~BATTENUATION I 1 

(c) (4 

Figure 7.44 H-field shielding effectiveness of various conductive coatings on a 

1.2 m cube. 

Figure 7.44(b), (c), and (d) shows the actual field strength within the 

center of the box when coated with each of the three materials. The 0.18 mm 

flame-sprayed aluminum gave an H-field attenuation of 5.2 dB produced by a 

0.025 mm thickness of aluminum foil. Conductive paint gave practically no 

attenuation. 

The numerical data presented in Figures 7.42,7.43, and 7.44 are of course 

appIicable only for the coil spacing used in these tests, since the coil spacing 

affects the impedance of the magnetic field and the shielding is determined in 

part by the impedance mismatch between the field and the shielding material. 

However, implications of the data are quite clear: 

1. The shielding effectiveness of both graphite-epoxy and boron-epoxy 

materials is much less than an equivalent thickness of aluminum. 

2. A thin piece of aluminum foil has much greater shielding effectiveness than has even a 3 mm panel of graphiteepoxy. 

244 
3. Graphite composites, by virtue of their higher conductivity, provide 

more effective shielding than do boron composites. 

4. When conductive coatings are present for control of the direct 

effects of lightning, they will have more effectiveness as electromagnetic shields than the composites they protect. 

5. When conductive coatings are not needed for direct effects protection, electrical wiring enclosed within these composites may need to 

be placed in conduits or otherwise shielded from the intense fields 

which will pas. through the composites. 

Electric fields (Efields) are also associated with lightning strikes and may 

cause as much interference in some electrical systems as do the magnetic fields, 

or more. No E-field shielding effectiveness data have as yet been reported, 

however. 

Furthermore, no measurements have yet been made of the voltages which 

lightning may actually induce in the electrical circuits inside of composite 

aircraft structures because there have been few such structures yet available. Nor 

are there analytical techniques yet available for prediction of field levels within 

composite structures. The shielding effectiveness measurements reported by 

Fisher and others have always involved small panels usually clamped to a simple 

test receptacle, like that of Figure 7.41. However, when panels are employed in a 

structure, with provisions made for conducting currents out of edges and joints, 

it is likely that some improvement in shielding effectiveness will result, and the 

situation will not be as bleak as it is indicated to be by the small panel data now 

available. 

7.3.2.5 Protection for Advanced Composites 

That it is necessary to supply lightning protection for all composites wherever they are employed on the aircraft is certainly not a foregone conclusion. 

Graphite composites greater than 4 plies thick and located in Zone 2A, for 

example, will not be completely punctured and will suffer only localized surface 

damage, as indicated by Figure 7.39. Some lightning protection may be 

necessary, however, for Zone 1 applications of graphite and for virtually arzjl 

lightning-stroke zone application of boron. 

A number of programs have been undertaken to develop suitable 

protective coatings for composites. The most important of these was conducted 

by Quinlivan, Kuo, and Brick (Reference 7.30), who developed and tested a 

wide variety of conductive coatings, and Brick, King, and Quinlivan have 

continued this work (Reference 7.31). Coatings evaluations have also been 

carried out by Fisher and Fassell (Reference 7.32) and others; the reader is urged 

to study these references to obtain the performance data necessary for the 

selection of a coating for a particular application. Coatings that have performed 

especially well include the five discussed below. 

Wire Fabrics 

Wire fabrics are flexible meshes woven or knitted from very fine aluminum 

245 
or copper wires, as shown in Figure 7.28. These perform well as lightning 

protective coatings because of their use of the skin effects for electrical 

conduction and because their transparency allows hot gasses to escape without 

damaging the composite beneath. For example, the skin area of an 80 by 80 

mesh (80 wires/cm) fabric of 0.05 mm wire is over 200 times that of the 

composite surface area actually coated by the fabric. Consequently, the fabric is 

a highly efficient conductor of current away from the arc contact point. This 

fabric has been found capable of withstanding successive 100 kA discharges at 

the same location with little visible damage to the fabric and no reduction in the 

mechanical strength of the coated laminate. At the 200 kA, 2 x lo6 A~-S level, 

the only significant damage to the coated boron laminate is that directly under 

the arc contact region, and, even in this region, the tensile strength is reduced 

only about 20%. 

One can expect paint coatings to be used on aircraft structures for various 

reasons such as camouflage, environmental protection, and marking. Because 

paints over 80 by 80 mesh wire fabric reduce the area of arc entry, damage to 

the coating increases at the point of arc entry, allowing some of the wire to be 

melted or vaporized. However, there is no additional damage to the composite. 

If mechanical conditions warrant its use, 50 by 50 wire mesh may also be used; 

however, this mesh offers somewhat less protection because the wire surface area 

is less. 

Aluminum fabrics were found by Quinlivan, Kuo, and Brick to provide the - 

best combination of lightning protection, light weight, environmental resistance, 

- and ease of application. For 80 by 80 mesh woven aluminum wire fabric, the 

area density is 0.019 lb/ft2; for a 50 by 50 mesh fabric the area density is 0.042 

1b/ft2. These weights are increased to 0.036 and 0.072 lb/ft2, respectively, if the 

resin required for encapsutation of the fabric is accounted for. Employing 

calendered wire cloth permits some weight saving. Calendering reduces the 

thickness of the cloth by flattening the intersections of the wires; weight savings 

occur because less resin is required to encapsulate the flattened mesh. As a point 

of reference, 6-mil-thick aluminum foil has an area weight of 0.084 lb/ft2. The 

weights of environmentally protective topcoats or adhesive required for bonding 

the foil (or fabric) must be added to these figures. 

Wire fabrics possess the hand and drape necessary for use as an overlay on 

complex contoured parts, and the composite matrix which fuUy encapsulates the 

fabric has been found to protect it from environments such as humidity, salt 

spray, and jet fuel. 

Aluminum Foils 

Aluminum foils from 0.05 mm to 0.2 mm thick will provide protection 

simdar to that of wire fabrics, but the mechanical properties of the foil, its 

weight, and the need for an adhesive between the foil and the composite may 

make aluminum foil less attractive as a protective device. 

246 
Flame-Sprayed Aluminum 

An equal weight of flamesprayed aluminum offers approximately the 

same protection as that of aluminum foil. It has sometimes been difficult to 

control the thickness of flame-sprayed coatings, a phenomenon that could create 

inconsistencies in the protection afforded by this type of coating, but 

flame-sprays may be easier to apply to some contoured surfaces than are foils. 

Diverter Straps and Foil Strips 

One might wonder if the metal diverter straps or foil strips recommended 

for protection of fiberglass in Section 7.3.1.2 are also applicable for advanced 

composites. As discussed in that section, these devices provide an external 

conductor from which a streamer may originate, reaching an approaching leader 

before a streamer from an internal conductor has a chance to puncture the 

fiberglass wall. The electrical conductivity of the strap, in contrast to the 

surrounding nonconductive fiberglass, is depended upon to assure that this 

streamer will come from the strip and not from some other place. 

Advanced composites, of course, have some electrical conductivity-- 

enough to supply the few milliamperes of current needed for a streamer. Thus, a 

metal diverter on a poorly conducting surface is not nearly as effective as it is on 

a nonconducting surface. This fact has been demonstrated repeatedly on nose 

radomes protected by diverters but with partially conductive antistatic paints 

also applied to their outside surfaces to reduce EM1 from P-static charge 

buildups on the radome surface. Streamers have occasionally initiated from the 

antistatic paint instead of from a diverter, causing the ensuing stroke to attach to 

the painted surface and destroy a large part of the radome while finding its own 

path to metal structure. 

Lubin and Dastin (Reference 7.33) report that, in laboratory tests, closely 

spaced aluminum foil strips did keep strikes from attaching to the boron 

composite surface of a horizontal stabilizer. Their tests--which are the most 

extensive yet reported of conducting strips on a composite--showed that 2.54 

cm (1 in) wide, 0.2 mm (0.008 in) thick aluminum foil strips had to be spaced 

within 6.35 cm (2.5 in) of each other to keep strikes from hitting the boron 

composite surface between them. A simdar spacing was necessary for 6.35 cm (2 

in) wide strips that were half as thick. Wider spacings permitted some strikes to 

terminate on the boron composite surface. 

It is likely that such strips would not work as well as this over a graphite 

composite, which is more conductive than boron. The performance of strips on 

either composite, of course, would be enhanced if the composite surface were 

covered with a dielectric (insulating) material that would inhibit streamer 

formation. A system in which a dielectric film is first applied to the entire 

composite surface, with the conducting strips then applied over the dielectric, is 

likely to work best, since this would also help to keep lightning currents in the 

foil strips from entering the composite. No thorough tests of such a system have 

been reported, however. 

247 
Dielectric Coatings 

Dielectric coatings such as polyurethanes or Kapton' film have been 

proposed by Kung, Amason (Reference 7.34), and others, as a means of 

protecting composite surfaces in swept-stroke zones (Zone 2A). These researchers acknowledge that some external conductors must still be used, but 

they suggest that the distance between such conductors could be extended by 

applying a dielectric coating to the composite surface, as shown in Figure 7.45. 

In tests of a 0.076 mm (3-mil) polyurethane coating over a graphite composite, 

Kung and Amason report that a 78 cm (31 in) separation is possible between 

conducting strips. In their tests a strike to the first conductor is initiated and 

caused to sweep aft in a wind stream across the dielectric coating until it 

reattaches to the second conductor, in the manner of Figure 6.2 1. As discussed 

in Section 6.6.2 of Chapter 6, the distance this arc will sweep depends on the 

voltage withstand capability of the surface paint versus the amount of voltage 

which can be developed along the arc. The greater the insulation capability of 

the coating, the longer the sweep to the next attachment point; the higher the 

arc voltage, the shorter the sweep. The principle of the dielectric coating concept 

is simply to make the coating from a good dielectric able to withstand high 

voltages. But there is a pitfall: if a puncture of the dielectric should occur, the 

arc will remain attached to this point much longer than it would have had no 

dielectric coating been present, and in so doing the arc will inflict more damage 

to the composite than if the arc had been uninhibited by the dielectric coating 

and allowed to move quickly from one place to another. 

METAL FOIL 

DIELECTRIC COAT 

Figure 7.45 Dielectric coating protection system. 

'ING 

'~rademark of the E. I. duPont de Nemours Company, Inc. 

248 
The voltage withstand capability of polyurethanes and similar materials 

considered for this application is very high, a characteristic that may permit large 

spans of the materials to exist between conductors in laboratory tests, such as 

those of Figure 7.45. The withstand capability of these dielectrics is drastically 

reduced, however, if their surfaces become scratched or cracked, as might occur 

in service. Thus, if a dielectric protection scheme is considered, it should be 

given the lightning tests described in Lightning Test Waveforms and Techniques 

for nonmetallic materials in Zone 2A-dfter the test sample has been exposed 

to vibration, temperature cycling, abrasion, and any other environment that is 

likely to degrade the condition of the dielectric coating. Even then, such a 

protection system may become damaged by unexpected contact with a foreign 

object and its degraded status go unnoticed until a lightning strike occurred. 

These coatings contrast with conductive coatings, which may sustain many 

abrasions or cracks with much less, if any, degradation in their protection 

capability. 

Dielectric coatings, at their present stage of development at least, should 

not be considered, therefore, for protection of flightcritical composite 

structures except where a puncture of the coating would not create an 

unacceptable hazard. Rather, dielectrics are likely to be most beneficial when 

they are used to supplement the performance of a primarily conductive 

protection system, such as the closely spaced foil strips discussed previously. 

In addition to protecting composite surfaces, it is necessary to provide a 

means to safely conduct lightning currents through joints, as discussed in Section 

7.3.2.3. If the composite is protected by a conductive coating and bolted joints 

are used, the coating should be extended into tapered bolt holes (to make good 

electrical contact with the bolts) and the bolts used to transfer current through 

the joint. If adhesive bonding joints are used instead, a conducting additive 

should be added to this adhesive; otherwise a different conducting path should 

be provided. 

If a composite is to be used in a flight-critical application, in every case the 

performance of its lightning protection system should be verified by the 

simulated lightning tests described in SAE Report Lightning Test Wavefomzs and 

Techniques for Aerospace Vehicles and Hardware. 
REFERENCES 

7.1 J. A. Plumer, Guidelines for Lightning Protection of General Aviation 

Aircraft, FAA-RD-73-98, Federal Aviation Administration, Department of 

Transportation, Washbgton, D.C. (October 1973), p. 8. 

7.2 Lightning Test Waveforms and Techniques for Aerospace Vehicles and 

Elardwvzre, Society of Automotive Engineers, Warrendale, Pennsylvania (5 

May 1976), pp. 9-1 0. 

7.3 J. A. Plumer, Lightrzir~g Protection of the F-Ill Radome Pitot System, 

76CRD018, Corporate Research and Development, General Electric 

Company, Schenectady, New York (February 1976), p. 19. 

7.4 General Electric Company photograph. 

7.5 James and Phillpott, Simulation of Lightning Strikes, p. 12. 

7.6 Paul T. Hacker, Lightning Damage to a General Aviation Aircraft-- 

Description and Analysis, NASA TN D-7775, National Aeronautics and 

Space Administration, Lewis Research Center, Cleveland, Ohio (September 

19741, p. 49. 

7.7 Advanced Development of Conceptual Hardware for the Lightweight 

Fighter, prepared by General Dynamics under Contract F 

336 15-73-C-5 130 for the Advanced Development Division, Air Force 

Materials Laboratory, Wright-Patterson Air Force Base, Ohio (15 February 

1976), p. 39. 

7.8 R. S. Jacobsen, "Magnetic Acceleration of Flyer Plates for Shock Wave 

Testing of Materials," Pulsed Electric Power Flyer Plate Notes, AFWL PEP 

2-1, Vol. I, Air Force Weapons Laboratory, Kirtland Air Force Base, New 

Mexico (July 1971): 6-5, 6. 

7.9 (5 May 1976), pp. 14-2 1. 

7.1 0 MIL-B-5087B (ASG), Military Specification: Bonding, Electrical, and 

Lightning Protection for Aerospace Systems (15 October 1964). 

7.1 1 J. R. Stahmann, "ControI Surface and Door Hinge Bonding Effectiveness 

in Modern Aircraft," Proceedings of the 1972 Lightning and Static 

Electricity Conference, 12-15 December 1972, AFAL-TR-72-325, in 

cooperation with SAE Committee AE-4 on Electromagnetic Compatibility 

and the Air Force Avionics Laboratory, Air Force Systems Command, 

Wright-Patterson Air Force Base, Ohio (1972), pp. 527-34. 

7.12 General Electric Company photographs. 

7.1 3 General Electric Company photographs. 

7.14 Plumer, Guidelines for Lightning Protection of General Aviation Aircraft. 

7.15 M. P. Amason, G. J. Cassell, J. T. Kung, "Aircraft Applications of 

Segmented-Strip Lightning Protection Systems," Proceedings of the 1975 

Conference on Lightning and Static Electricity, 14-1 7 April 1975, at 

Culham Laboratory, England, Session IV: Aircraft Applications, the Royal 

Aeronautical Society of London (1975). 

7.16 S. A. Moorefield, J. B. Styron, and L. C. Hoots, Manufacturing Methods 

for Advanced Rad~me Production, Interim Technical Report 6, prepared 

by the Brunswick Corporation for the Air Force Materials Laboratory, Air 
Force Systems Command, United States Air Force, Wright-Patterson Air 

Force Base, Ohio (November 1975), pp. 32-40. 

7.17 Amason, Cassell, and Kung, "Aircraft Applications of Segmented-Strip 

Lightning Protection," p. 9. 

7.18 D. A. Conti and R. H. J. Cary, "Radome Protection Techniques," 

Proceedings of the 1975 Conference on Lightning and Static Electricity, 

14-1 7 April 1975, at Culham Laboratory, England, Session IV: Aircraft 

Applications, the Royal Aeronautical Society of London (1975). 

7.19 W. M. Fassell, A. P. Penton, and J. A. Plumer, "The Susceptibility of 

Advanced Filament Organic Matrix Composites to Damage by Simulated 

Lightning Strikes," Lightning and Static Electricity Conference, 3-5 

December 1968, Part 11: Conference Papers, AFAL-TR-68-290, Air Force 

Avionics Laboratory, Air Force Systems Command, Wright-Patterson Air 

Force Base, Ohio (May 1969), pp. 530-69. 

7.20 A. P. Penton, J. L. Perry, and K. J. Lloyd, The Effects of High Intensity 

Electrical Currents on Advanced Composite Materials, U-4866, prepared 

by the General Electric Company for the Naval Air Systems Command, 

Department of the Navy, washington, D.C. (15 ~e~tember 1970), ppi 

2-30, 2-32. 

7.21 F. A. Fisher and W. M. Fassell, Lightning Effects Relating to Aircraft, Part 

I Lightning Effects on and Electromagnetic Shielding Properties of Boron 

arzd Graphite Reinforced Composite Materials, AFAL-TR-72-5, Air Force 

Avionics Laboratory, Air Force Systems Command, Wright-Patterson Air 

Force Base, Ohio (January 1972), p. 55. 

7.22 Fisher and Fassell, Lightning Effects, Part I, p. 79. 

7.23 Fisher and Fassell, Lightning Effects, Part I, p. 29. 

7.24 S. D. Schneider, C. L. Hendricks, and S. Takashima, Vulnerability/Survivability of Composite Structures- -Lightning Strike, D6-42673-3, prepared 

by the Boeing Commercial Airplane Company for the Air Force Flight 

Dynamics Laboratory, Aeronautical Systems Division, Wright-Patterson 

Air Force Base, Ohio (April 1976). 

7.25 Advanced Development of Conceptual Hardware for the Lightweight 

Fighter. 

7.26 Fisher and Fassell, Lightning Effects, Part I, pp. 178-204. 

7.27 Fisher and Fassell, Lightning Effects, Part I, pp. 187, 201, 198. 

7.28 E. R. Uhlig, Investigations of H-Field Shielding Effectiveness of Thin 

Conductive Coatings and Reduction of Induced Currents on Penetrations, 

submitted by author, General Electric Company, under Contract DA-49- 

129-ENG-543 and placed by Corps of Engineers, Department of the Army, 

Huntsville, Alabama. 

7.29 Uhlig, Investigations of H-Field Shielding Effectiveness, pp. 1 1, 13, 15. 

7.30 J. T. Quinlivan, C. J. Kuo, and R. 0. Brick, Coatings for Lightning 

Protection of Structural Reinforced Plastics, AFML-TR-70-303 Pt. I, Air 

Force Materials Laboratory, Nonmetallic Materials Division, Air Force 

Systems Command, Wright-Patterson Air Force Base, Ohio (March 1971). 

7.3 1 R. 0. Brick, C. H. King, and J. T. Quinlivan, Coatings for Protection of 

Structural Reinforced Plastics, AFML-TR-70-303 Pt. 11, Air Force Mate- 
rials Laboratory, Nonmetallic Materials Division, Elastomers and Coating 

Branch, AFMLILNE, Wright-Patterson Air Force Base, Ohio (February 

1972). 

7.32 Fisher and Fassell, Lightning Effects, Part I, pp. 3-1 77. 

7.33 G. Lubin and S. Dastin, "Lightning Protection for Aircraft Sandwich 

Structures with Boron/Epoxy Composite Skins," Proceedings of the 1972 

Lightning and Static Electricity Conference, 12-15 December 19 72, 

AFAL-TR-72-325, in cooperation with the SAE Committee AE4 on 

Electromagnetic Compatibility and the Air Force Avionics Laboratory, Air 

Force Systems Command, Wright-Patterson Air Force Base, Ohio (1972), 

pp. 359411. 

7.34 J. T. Kung and M. P. Amason, "Dielectric Shielding Lightning Protection 

for Composite Aircraft Structures," Proceedi~~gs of the 1972 Lightning 

and Static Electricity Conference, 12-15 December 1972, AFAL-TR-72- 

325, in cooperation with the SAE Committee AE-4 on Electromagnetic 

Compatibility and the Air Force Avionics Laboratory, Air Force Systems 

Command, Wright-Patterson Air Force Base, Ohio (1972), pp. 337-58. 
CHAPTER 8 

VOLTAGES AND CURRENTS INDUCED BY LIGHTNING 

8.1 Definition and Elementary Considerations 

The term indirect effects of lightning refers to the damage to or 

malfunction of electrical equipment that results from lightning flashes. These 

effects may range from tripped circuit breakers to computer upset, to physical 

damage to input or output circuits of electronic equipment. There may be other 

indirect effects of lightning flashes that pertain to aircraft safety, such as flash 

blindness of the crew or acoustic shock waves. These effects, however, are not 

treated in this chapter. Included in this definition and discussed here are the 

voltages and currents induced by lightning on the electrical wiring of the aircraft, 

regardless of whether or not such voltages and currents cause damage or upset of 

electrical equipment. 

The demarkation line between direct and indirect effects may be 

somewhat arbitrary in some instances. An example might involve a lightning 

flash terminating on a wing tip navigation light. In this case, the burning or 

blasting damage to the light fixture would be considered a direct effect of the 

lightning. If the lightning flash were to contact the filament of the bulb and 

inject current into the wiring, there would be produced electrical effects ranging 

from overvoltage breakdown of the insulation at the socket to tripping of 

remote circuit breakers or upset of equipment, effects of the resulting surge 

voltages on the power system. All of these effects will be considered indirect 

effects in this discussion, even though the initiating event was the direct 

injection of current into the filament. Another hypothetical possibility would 

involve a lightning flash that passed close to but did not contact the aircraft. The 

changing electromagnetic field produced by that flash might upset electronic 

equipment. The circuit upsets produced by such a flash would clearly be of an 

indirect nature, since there was no direct involvement of the aircraft with the 

lightning flash. Usually, however, the indirect effects of concern are produced by 

a flash that contacts the aircraft. 

In this section as a whole we shall attempt to explain some of the electrical 

phenomena involved and to discuss some of the analytical techniques through 

which indirect effects may be analyzed, some of the measurements that have 

been made on aircraft to determine the nature of the indirect effects, and some 

of the measures one may take to control the indirect effects of lightning. 

The arts of analysis and control of indirect effects are embryonic at 

present. Accordingly, the analytical procedures to be described are not as well 

developed as might be desired. Neither can very many examples of successful 

protective measures be given, since protective measures against indirect effects 

are as yet seldom included in aircraft design. 

The problem of analysis of indirect effects is a complex one, but, 

fortunately, analysis can be divided into several stages. These stages are 

illustrated in Figure 8.1. The overall task is to determine how the lightning 

current, IL, leads to voltages and currents on the internal wiring. The individual 
tasks are as follows: 

(a) To determine the amplitude and waveform of the current on 

the exterior of the aircraft. This external current may or may not be 

the same as the undisturbed lightning current. 

(bj To relate the internal response of the aircraft structure to the 

current flowing on its exterior. This internal response may be 

expressed either in terms of an internal electric field or an internal 

magnetic field. 

(c) To determine how the internal response of the structure affects the 

response of the aircraft's wiring. The response of the wiring may be 

expressed either in terms of voltages or currents induced on active 

electrical conductors or in terms of currents induced on shields. 

(d) To determine how these currents and voltages affect the electrical 

equipment in the aircraft. 

8.2 The Exterior Response of the Aircraft Structure 

The problem of how the aircraft as a whole responds to the passage of 

lightning current is illustrated in Figure 8.2. There are two parts of this problem, 

one involving a direct flash to the aircraft and the other involving a nearby flash. 

While it is generally true that the indirect effects produced by a direct flash are 

more severe than those produced by a nearby flash, they are not necessarily so, 

since the response of the aircraft external structure is different in the two cases. 

Consider Grst a direct hit on the aircraft. The return stroke current enters 

the aircraft at one extremity, flows through it, and exits from the other end. As 

a result of the different impedances of the lightning channel and the aircraft, the 

current flowing through the aircraft may undergo some distortion in waveform. 

The result is that the waveform of the current leaving the aircraft may be 

different from that entering the aircraft. Typically, the exit current rises to crest 

over a longer time than does input current. 

The phenomena involved are shown in Figure 8.3. A waveform is shown 

traveling along the conductor having surge impedance Z1. Surge impedance is 

defined as 

where Z = surge impedance (ohms) 

L = inductance (henries) per unit length 

C = capacitance (farads) per unit length 

at the point under consideration. At some point the current encounters a 

transition point between the conductor having surge impedance Z1 and a 

different conductor of surge impedance Z2. At this transition point, part of the 

incident current is transmitted onto the second conductor, but part of it is 

reflected back in the direction from which it came. The magnitudes of the 

transmitted and reflected components of voltage or current at this transition 
LIGHTNING CURRENT- 1, 

(b) 

I,,, -EXTERNAL CURRENT 

I (4 I 

Figure 8.1 Indirect effects. 

(a) The total system 

(b) The internal response of the structure 

(c) The response of cables 

(d) The response of piece parts 

point are given by the values of transmission and reflection operators, a and /3 

for voltage and h and 6 for current, shown in Figure 8.3(a). 

If there are two discontinuities, as in the case of Figure 8.3 (b), there will 

be reflections at each discontinuity with currents traveling back and forth, and 

thus oscillating, in the intermediate conductor. The amplitudes of these 

oscillatory currents will diminish as energy is transmitted from the intermediate 

conductor to the conductors on either end. 

The geometry of Figure 8.3(b) approximates that of a lightning current 

entering an aircraft, since the surge impedance of the lightning channel is higher 

than the surge impedance of the aircraft. When account is taken of all the 

reflections and transmissions at the various transition points, the result is that 
Figure 8.2 Response of the structure to the lightning current. 

(a) To a direct flash 

(b) To a nearby flash 

the current waveforms will appear as shown in Figure 8.3(c). If the entering 

lightning current has a fast front and a slower decay, the current at the entrance 

of the low-impedance section (in the aircraft) will exhibit an overshoot. The 

current at the center of the aircraft will have an oscillatory component 

superimposed upon a waveform fundamentally like that of the input current, 

and the current leaving the aircraft will have a slower rise time than that of the 

incident lightning curren f. 

A way in which one can tally the various reflected and transmitted 

currents is through the use of the lattice diagram shown in Figure 8.4. A current 

wave of unit amplitude enters from the left, sees the transition point, and is 

partly transmitted and partly reflected. Since Z2 is less than Zl, the exiting 

(leaving) current is of higher amplitude than the entering current. In the limit, a 

current wave traveling on a conductor and encountering a short circuit (Z2 = 0) 

would double at that discontinuity. The transmitted current then passes along 

the low-impedance co~~ductor and after a time delay of At encounters the 

discontinuity at the exit end. Here part of the current on conductor 2 is again 

reflected. Since the impedance of conductor 3 is greater than the impedance of 

conductor 2, the transmitted current is less than the incident current. A liniiting 

condition of this would consist of a current wave traveling on a conductor and 

then encountering an open circuit. The current transmitted into the open circuit 

is of course zero. 
(4 

Figure 8.3 Surge propagation at transition points. 

(a) A single transition point 

(b) Two transition points 

(c) Waveshapes 

The current amplitude at any point is then the algebraic sum of the 

reflected and transmitted components at that point. For example, the current 

leaving the intermediate conductor is 0.556 at 1At after the current enters the 

left-hand portion of it, 0.803 after 2At, and 0.913 after 3At, etc. The exit current 

thus rises more slowly than does the incident current. At the center of the 

low-impedance intermediate conductor, the current is oscillatory. 

If instead of a step function a current wave of finite rise time encounters 

the transition, there will be a less pronounced oscillation of current in the 

intermediate conductor and less difference between the waveforms of the 

entering and exiting currents. 

257 
Figure 8.4 Lattice diagram showing amplitude of reflections 

An example of this effect was demonstrated on a simple model of the 

Space Shuttle (Reference 8.1). A twedimensional outline of the Shuttle was cut 

from metal foil and the outline split in the center so that the two halves could be 

slightly separated. The model was suspended in the air and connected to a fine 
wire; a pulse current was passed along the wire, through the model, and along 

the rest of the fine conductor wire to a termination resistor. The fine wire 

represented the channel of the lightning arc. The input and output currents and 

the current at the center of the model were measured with small transformers. 

Typical results are shown in Figure 8.5. One significant point about the 

results was that fast-rising incident currents excited a higher degree of oscillation 

of the aircraft than did slower rising currents. The second significant point 

demonstrated was that in most cases the waveform of the lightning current 

passing through the center of the aircraft was sufficiently similar to the 

waveform of the basic lightning current that the oscillatory component 

superimposed as a result of the change in impedance between the arc channel 

and the aircraft was not significantly large. 

In the case of a nearby lightning flash, the electric field from the flash will 

excite a dipole oscillation of the aircraft. In terms of Figure 8.4, the effect could 

be viewed as one in which the impedance of conductors 1 and 3 was infinite, 

leading to complete reflection and no transmission at the entry and exit points 

on a low-impedance aircraft. A dipole kind of oscillation excited on an aircraft 

structure would be the typical response considered if one were evaluating the 

effect on aircraft of the electromagnetic fields produced by nuclear explosions. 

The period of oscillation would be proportional to the length of the aircraft; 

thus, large aircraft would tend to ring at lower frequencies and higher amplitudes 

than would shorter aircraft. While the response of aircraft and missile systems to 

the rapidly changing electromagnetic fields produced - by nuclear explosions has 

been extensively studied, there have not been any c~rres~ondl~~-studies of the 

response of aircraft to the electric fields associated with the passage of a nearby 

lightning flash. Accordingly, in this record we will not treat further the case of a 

nearby lightning flash, except to add the cautionary statement that it has not 

been proven finally that the indirect effects associated with a nearby lightning 

flash are necessarily lower than those associated with a lightning flash that 

contacts the aircraft. In the case of a nearby lightning flash it would appear that 

the aircraft structure would be subjected to currents of much lower amplitude 

but currents much more oscillatory in nature. These lower amplitude-higher 

frequency currents might lead to more upsets of sensitive electronic circuits than - - 

would currents of higher amplitude and lower frequency. 

8.3 The Internal Response of the Aircraft Structure 

A metallic aircraft is often viewed as a ~arada~ screen, a concept from 

electrostatics which implies that the electrical environment inside the aircraft is 

separate and distinct from the environment outside. To some extent this is true 

for the electrical environment inside the structure: the environment is not nearly 

as harsh as is the external environment. There are, however, some important 

mechanisms by which electrical energy couples to the interior of the aircraft. 

The basic coupling mechanisms are shown on Figure 8.6. The first of these 

relates to the electric field produced alo~lg the inner surface of the aircraft. This 
- 2.2 ps 

(EQUIVALENT) 

Figzire 8.5 Dipole oscillation of an aircraft excited by the passage of lightning 

current. On each oscillogram 

(a) Fast rise of current d is the current entering the aircraft 

(b) Average rise of current e is the current at the center of the aircraft 

(c) Slow rise of current f is the current leaving the aircraft 
coupling mechanism might be defined as a resistive voltage drop. In some cases a 

definite resistance will be involved, though frequently the resistance will be of a 

distributed nature and pwbably frequency- or time-dependent. 

EXTERNAL MAGNETIC FIELD Hex, 

INTERNAL MAGNETIC FIELD 

Eext EXTERNAL ELECTRIC FIELD 

INTERNAL ELECTRIC FIELD 

Figure 8.6 Coupling mechanisms. 

(a) Resistive 

(b) Magnetic fields 

(c) Electric fields 

The second coupling mechanism involves magnetic fields in the interior 

volume of the aircraft. The most common and important type of magnetic field 

is that drawn through apertures from the outside of the aircraft to the interior, 

26 1 
as shown in Figures 8.6 and 8.7. This is frequently called the apertwe fieM. 

There will also be magnetic fields produced by the diffusion of lightning currents 

to the inside surfaces of the aircraft skins. These are referred to loosely as the 

diffusion fields. The diffusion fields are also related to the frequencydependent properties of the resistively generated electric field. Because some of 

the concepts involved in the study of the diffusion fields are central to an 

understanding of other effects, particularly with respect to the response of 

shielded wires, they will be discussed in detail before fields of other origins are 

considered. 

Figure 8.7 Aperture-type magnetic field coupling. 

(a) External field patterns 

(b) Internal field patterns 

The third type of coupling involves electric fields passing directly through 

apertures, such as windows or canopies, to the interior of the aircraft. In metal - 

aircraft this coupling is entirely through apertures, since virtually any thickness 

of metal provides comparatively good shielding. Aperture-type electric field 

coupling is shown in Figures 8.6(c) and 8.8. 

262 
E ext 

Figure 8.8 Aperture-type electric field coupling. 

(a) External field patterns 

(b) Internal field patterns 

The host easily understood mechanism by which the passage of lightning 

current gives rise to voltages on aircraft electrical circuits is that in which the 

current, flowing through joint resistances, produces a voltage by the elementary 

IR voltage drop. Such a case is shown in Figure 8.9. Here lightning is shown 

contacting a wing tip navigation light. The lightning current flowing through the 

resistance of the mechanical mounting structure of the lamp housing produces a 

voltage across that resistance. The voltage drop across the resistance will have the 

same waveform as that of the lightning current. The voltage at some remote 

point, however, may not have the same waveform, since the distributed 

inductance and capacitance of the wire supplying power to the filament of the 

light will be set into oscillation by the suddenly developed voltage. The result is 

that at remote points there probably will be an oscillation superimposed upon 

the basic IR voltage. 

Figure 8.10 shows two other examples of cases in which resistive voltages 

might be encountered. The first would be at the pylons for mounting external 
NAVIGATION LIGHT 

MOUNTED ON THE WING 

(b) I time 

I 

(4 

Figure 8.9 Resistive voltages. 

(a) Physical structure 

(b) Current waveshape 

(c) v1 

(dl V2 

stores, shown in 8.1qa). If lightning current were to contact such external 

stores, it would have to flow through the pylons to enter the aircraft. The 

pylons, not generally designed as current-carrying members and being points 

where the lightning current would be concentrated, might have a high voltage 

developed across them. Another example might be the structural bolts attaching 

a large segment of the airframe, such as the vertical stabilizer shown in Figure - 

8. I qb). 

The effects of joint resistance on circuits are strongly influenced by the 

manner in which circuits are grounded, as shown in Figure 8.1 1. Current flowing 
LIGHTNING FL 

Figure 8.1 0 Other examples of resistance. 

(a) The pylons for external stores 

(b) Joints in structural members 

Figure 8.1 1 Effects of grounding. 

(a) Structural return, V1 = IR 

(b) Single-point ground, V2 = low 

(c) Single-point ground, V3 = IR 

across the joint resistance, R, produces a driving voltage: V = IR. Since the 

circuit across which V1 is measured employs the structure as a ground-return 

path, the circuit couples all of this voltage; thus V1 would be high. A circuit 

employing a single-point ground does not include this resistive drop; hence V2 

would be low. The use of a single-point ground, however, does not eliminate the 

voltage, since in this latter case the voltage at the source end of the circuit, V3, 

would be high. 

These elementary descriptions of joint resistance should not be relied upon 

to predict coupling into circuits extending throughout the entire aircraft. The 

more massive the joint and the lower the dc resistance, the greater will be the 

dependence of resistance on the waveform and frequency content of the 

lightning current, and the greater will be the proportionate effects of changing 
magnetic fields. While these effects will be discussed in more detail in other 

sections, one common oversimplification, shown in Figure 8.12, should be 

pointed out here. If the total end-teend resistance of the aircraft were 2.5 ma 

and a lightning current of 200 000 A were flowing through the aircraft, the 

end-to-end voltage on any circuit could not be depended upon to be less than 

500 V, the product of the lightning current and the dc resistance. 

Figure 8.12 An oversimplified model. 

Maximum voltage is not determined only by total end-to-end resistance. 
REFERENCE 

8.1 F. A. Fisher, Analysis of Lightning Current Waveforms through the Space 

Shuttle, Aircraft Lightning Protection Note 75-1, National Aeronautics 

and Space Administration, Lyndon B. Johnson Space Center, Houston, 

Texas (January 17, 1975). 

CHAPTER 9 

THE EXTERNAL MAGNETIC FIELD ENVIRONMENT 

9.1 Elementary Considerations 

If a long conductor is carrying a current, I, and the return path is far 

removed, the average field intensity, Hav, at a distance, r, from the conductor, as 

shown in Figure 9.l(a), is 

If instead of a solid wire the current is carried on a hollow tube of radius r, 

as shown in Figure 9.l(b), the field intensity, H, at radius r is again 

and at the surface of the tube, where r = ro, the field intensity is 

I H=- 

2nro (9.3) 

In the interior of the tube, r < ro, the field intensity is zero, a concept that 

will be treated in more detail later. 

Since the circumference, P, of the tube is 

P = 2nro (9.4) 

it follows that the field intensity at the surface of the tube is also equal to the 

total current divided by the circumference. 

In all the cases the units of field intensity are amperes per meter if the 

radii are measured in meters. 

If the conductor is not cylindrical, as shown in Figure 9.l(c), the field 

intensity at different points on the surface will be different. However, the 

average field intensity will still be equal to the total current divided by the 

circumference: 

The actual field intensity will be greater than average at points where the radius 

of curvature is less than average; it wd1 be less than average at points where the 

radius of curvature is greater than average, as shown in Figure 9.2. For example, 

the circumference of the fuselage of a typical fighter aircraft just forward of its 

wing is about 5.5 m. Assuming a lightning stroke current of 30 000 A to flow 

through the fuselage, the average field intensity at the surface would be 
Since there are no points of very sharp radius, the field intensity around the 

fuselage would probably not vary greatly from the average value. 

(4 

Figure 9.1 Magnetic fields around current-carrying conductors. 

(a) Current-carrying filament 

(b) Tubular conductor 

(c) Irregular conductor 

270 
ti = HA, 

Figure 9.2 Field intensity vs radius of curvature. 

The situation along a wing carrying lightning current is considerably 

different in that the leading and trailing edges have radii of curvature much 

less than the average. Field intensity along the leading and trailing edges would 

be quite high compared to the field intensity along the top and bottom surfaces, 

for example. 

Figure 9.3 shows, in general, how the magnetic field strengths would vary 

with position on an aircraft if a lightning flash enters through the nose pitot 

boom and leaves through the vertical stabilizer. The field intensity would be 

highest around the pitot boom, lowest around the midsection of the fuselage, 

and high again around the vertical stabilizer. In the vicinity of the nose 

equipment bays, the field would be of greater than average intensity. Since the 

field intensity is inversely proportional to the radius of curvature, it then follows 

that the field intensity outside the fuselage of a large transport aircraft would be 

considerably less than that outside the small fighter aircraft shown on Figure 

9.3. 

Since both the average current density, Jav, and average field intensity, 

Ha,, are equal to the total current divided by the circumference, or 
H IS HIGHEST ON 

LEADING AND TRAILING 

EDGES 

Figure 9.3 Variation of magnetic field strength with aircraft radius of curvature. 

it follows that the tangential field intensity at the surface of a conducting object 

is equal to the current density at that point. This is in fact true, at least for 

transient currents. The relation is not true for dc currents or transients 

sufficiently slow that appreciable magnetic fields penetrate the skin. The 

orientation of the H field vector is always at right angles to the direction of the 

current vector. While small gaps in the structure (Figure 9.4) direct the current 

around the gap, the magnetic field is virtually unaffected, except directly on the 

surface and on a length scale that is small compared to the dimensions of the gap 

interrupting the current flow. 

9.2 Determination of the External Current and Magnetic Flux 

Density 

Both the resistive voltage drops inside a structure and the amounts of 

magnetic or electric field which penetrate through apertures depend upon the 

external field distribution, most importantly upon the magnetic field distribution. This section will discuss some of the ways in which this field distribution 

might be determined. 

The current density at the surface of a conductor is seldom determined by 

the dc resistance of the conductor; rather, it is primarily controlled by magnetic 

effects. The magnetic distribution of current density can be calculated in simple 

geometries. Around the periphery of a cylinder, for example, the current density 

is uniform, assuming the return path for that current is far removed from the 

cylinder: greater than ten times the diameter of the cylinder. 
SLOT INTERRUPTS 

CURRENT FLOW 

STREAMLINES 

Figure 9.4 Current flow and magnetic field around structural gaps. 

(a) Current entering typical wing 

(b) Resultant magnetic field virtually unaffected 

by slot interrupting flow of current 
9.2.1 Elliptical Conductors 

Another geometry of considerable practical importance is the ellipse, 

d shown in Figure 9.5. At the center (X = 0, Y = +-) the magnetic field intensity 

is given by 2 

b and at the edge (X = +-, Y = 0) 2 

At intermediate points the magnetic field intensity or current density is given by 

the expression (References 9.1 and 9.2) 

I 1 H surface = - 

(9.1 1) 

0 

METERS 

Figure 9.5 Magnetic field intensity at the surface of an elliptical conductor. 
This current distribution does not hold for dc currents where the current density 

over the surface is determined by the dc resistance. If the conductor were an 

elliptical cylinder of uniform thickness, the current density would be uniform. 

The region over which a transition takes place between the uniform distribution 

of current governed by resistance and the nonuniform distribution governed by 

magnetic fields (Equation 9.1 1) is indicated as follows: 

where 

The classical skin depth, 6, is 

where 

b = width 

d = depth 

p = resistivity 

p = permeability 

o = angular frequency 

As an example of the frequency range over which the transition takes 

place, consider an elliptic cylinder made from aluminum @ = 2.69 x am) 

with width (b) 1 m, depth (d) 0.25 m, and a wall thickness of 1 mm. The skin 

depth and value of the term 

is given in Table 9.1. 

Table 9.1 SKIN DEPTH AS A FUNCTION OF FREQUENCY 

Frequency Skin depth 

f 6 
Even at 10 Hz the current crowds to the edges of the ellipse and has a current 

density given by Equation 9.1 I. At 1 kHz the skin depth is about 0.25 cm,and 

at 10 kHz it is less than 1 mm. 

Thus, the region over which the current density changes from its uniform 

dc value to the limiting ac distribution is 1 to 10 Hz. The current density 

remains at that value up to indefinitely high frequencies, probably until the 

width of the cylinder becomes on the order of a tenth of a wavelength. The 

current density through the wall thickness of the cylinder remains constant over 

the range 0 to 3 kHz (approx.). 

In many cases the external current density may be determined with 

accuracy sufficient for practical purposes by approximating the surface under 

consideration by an ellipse or ellipses. If such an approximation does not give 

sufficient accuracy, there are other techniques that may be used. 

9.2.2 Numerical Solution of Fields 

Analytically the solution of the field around a currentcarrying conductor 

may be determined by a solution of Laplace's equation: 

where 

v2@ = 0 (9.14) 

@ = the potential. 

In rectangular coordinates Laplace's equation becomes 

For some geometries Laplace's equation can be solved analytically: an 

elliptical geometry is one of those cases The equations defining the field 

distribution will not be presented here, but Figure 9.6 shows an example of the 

field around an elliptical conductor. This field may be viewed as the electric field 

around an isolated ellipsoid, the figure showing the field in one plane passing 

through the major axis of the ellipsoid or the field around an elliptic cylinder of 

infinite length. Alternatively, it may be viewed as the magnetic field around an 

infinitely long elliptic cylinder carrying current into the plane of the figure. If 

the field is viewed as an electric field, the flow lines represent the path along 

which displacement currents flow, and if it is viewed as a magnetic field, the 

equipotential lines define the paths along which magnetic flux lines flow. The 

return path for either the voltage maintaining the charge or the return path for 

the current is assumed to be sufficiently far away from the conductor that it 

does not influence the field around the indicated region. 

The indicated flow lines divide the region into 44 sectors. At the surface of 

the conductor, the magnetic field strength is inversely proportional to the 
spacing between the flow lines. Since the average field strength around the 

surface is 

I H=- (9.16) 

P 

it follows that the field strength at the surface between any two flow lines is 

- 1 

"surface (9.1 7) 

FLOW LINES 

EQUIPOTENTIAL LINES 

Figure 9.6 The field around an elliptical conductor. 

In only the simplest geometries is it possible to calculate the field 

analytically. Usually one must resort to some numerical or graphical method of 

determining the field. Numerically, the field may be determined by a numerical 

solution of Laplace's equation. Figure 9.7 shows a conductor at potential P 

surrounded by a return conductor, a circle in this case, at potential 0. To this 

geometly is fitted a rectangular grid, shown here as a very coarse grid. 

Initially, all the grid points that lie on the conductor would be assigned a 

277 
POTENTIAL = 0 

Figure 9.7 A rectangular grid for evaluation of Laplace's equation. 

potential P, and all the grid points that lie on the return path would be assigned 

a potential zero. Laplace's equation in two dimensions can be shown to be 

approximately 

From this it follows that Laplace's equation is satisfied if the numeric values at 

four points surrounding a central point have values that satisfy the equation 
A determination of the field around the conductor then involves assigning 

field values at all of the points between the conductor and its return path, and 

adjusting the value of these points until Equation 9.19 is satisfied everywhere 

within the grid. The literature (Reference 9.3 and 9.4) indicates a number of the 

numerical techniques by which the potentials at the points may be adjusted to 

their final values. While the process is tedious, it is not completely impractical to 

do by hand. Usually the process is done by computer routines that solve the 

field equations. In addition to tabulating the numerical values of the field at the 

grid points, frequently such computer routines allow one to plot the flow and 

equipotential lines. 

Any solution of Laplace's equation will also allow the pattern of coupling 

through an aperture, as well as the external field pattern, to be calculated. Most 

computer routines deal either with two-dimensional problems or problems with 

rotational symmetry. In a twedirnensional case, the aperture would run the 

length of the conductor being studied. This is not necessarily a disadvantage, 

since some important geometries and apertures are basically two-dimensional. 

An important aperture is that which may exist on the trailing edge of a wing 

when the flaps have been extended. Since it is a convenient region to reach, 

wiring is often placed in this region. Electrically it is a poor place, since the 

aperture is near a region where the magnetic fields external to the wing are high. 

Another important set of apertures that may be approximated as a continual 

opening is that formed by the windows in the fuselage of a transport aircraft. 

In principle, the field around three-dimensional objects may also be solved 

numerically by extending Equation 9.18 to include 8 points on a cube 

surrounding a central point. Solutions of the fields around three-dimensional 

objects not involving rotational symmetry are seldom attempted because of the 

great increase in the amount of computer storage locations required to define 

the field satisfactorily in three dimensions. 

9.2.3 Hand Plotting of Fields 

The fields around any geometry may also be determined graphically by a 

cut-and-try process in which flow lines and equipotential lines are drawn on the 

geometry and adjusted until repeated subdivision always yields smalI squares and 

the flow and equipotential lines intersect at all points at right angles. 

Cut-and-try field plotting, of which Figure 9.8 is an imperfect example, 

may with care, patience, and liberal use of a soft pencil and eraser yieId a field 

pattern of any desired accuracy. 

9.2.4 CaIculation Using Wire Grids 

Another approach to determining current density, shown in Figure 9.9, is 

based on the premise that a two-dimensional geometry can be represented as an 

array of parallel wires. If the current in each wire is known, the average current 

density along the surface defied by any two wires will be the average of the 

current on the two wires divided by the spacing between the wires. 
Figure 9.8 Cut-and-try field plotting 

If the wires are all of infinite length, so that no end effects need be 

considered, and are all connected together at their ends, the manner in which 

current divides among the wires may be calculated with the aid of the simple 

computer program MAGFLD, the elements of which are discussed below, and 

the program given in Appendix 1. The equations governing the current 

distribution are given below. 

Let the location of the wires be defined in terms of the rectangular 

coordinates xi and yi, and Iet the radius of the conductors be rl. The 

self-inductance per unit length of each wire is 

henries per meter. In this equation R is defined as the distance from the 

conductor, or from the group of conductors, to an arbitrary return path. The 

numerical accuracy of the current distribution to be calculated does not depend 

critically upon the value assigned to R, but it should be of the order of 10 to 20 

times the greatest dimension of the structure being modeled. 

Between any two conductors, ij, there will be a mutual inductance: 

henries per meter. 
The spacing between conductors, ri,, can be determined from the 

coordinates of the conductors. If a group of conductors, each carrying a current, 

(Figure 9.10) is considered, and if the self- and mutual inductances of and 

between each conductor are known, the voltages across the self-inductance of 

each conductor will be as follows: 

I 

CONDUCTOR i 

\ 

(4 / 

Figure 9.9 A structure defined as an array of wires. 

(a) The array 

(b) Coordinates defining location 

(c) Definition of the return path for current 

Vl/u = Llil-Mi2i2-MI3i3 .... -Mlnin (9.22) 

V2/u = -M 21 i 1 t L2i2 - MZ3i3 . . . . -M2,,in (9.23) 

V31u = -Mj ,i, - Mj2i2 - Ljij . . . . -Mjnin (9.24) 

VV/a = -M nl i 1 -Mn2i2 - Mn3i3. . . . Lni n (9.25) 

Equations 9.22 through 9.25 may be placed in matrix notation as follows: 
Figure 9.10 Mutually coupled inductances. 

or, in more compact notation 
In Equations 9.26 and 9.27 the normalized angular frequency o has been 

set equal to unity. 

Multiplying by the inverse of the M matrix, I M 1-I , gives the following: 

where ml I, ml2, ml3 are the elements of the inverse of the M matrix. 

If all of the voltages are the same and equal to V, as is the case if all of the 

inductances are connected in parallel, the absolute current in each element is 

- in - (mnl + mn2 + mn3 . . . . + mnn)V (9.34) 

The total current that flows, which is proportional to the impressed voltage, is 

ir = (il + i2 + i3 + .... in)V (9.3 5) 

The fraction of the total current that flows in each circuit is 
9.2.5 Measurements Using Wire Grids 

An alternate experimental method of determining current distribution is 

shown in Figure 9.1 1. It would be entirely feasible to model the surface of an 

aircraft by a grid of wires spaced over a supporting framework or placed on the 

surface of a complete model made from nonconductive materials. If a 

high-frequency current were passed through the wire grid in a path similar to 

that expected for a lightning flash, it would be possible to measure the current 

in the individual wires by a clamp-on current probe. 

To the author's knowledge such a wire grid model has not yet been used to 

study current distribution resulting from lightning flashes striking aircraft, but 

somewhat similar models have been used to measure the current distribution in 

ground-based launching facilities (References 9.5 and 9.6), for studies of the 

interaction of aircraft and missile systems with the electromagnetic fields 

generated by the explosions of nuclear weapons, and for the study of magnetic 

fields in and around the towers used to support high-voltage transmission lines. 

In principle, the wire grids should follow as closely as possible the expected 

current distribution paths and should be interconnected frequently by other 

wires circling the structural member being modeled. It would probably be best if 

the interconnecting wires tended to form orthogonal squares with the wires 

running along the flow lines, in a manner similar to the orthogonal squares 

formed by flux and equipotential lines shown on Figure 9.6. 

In the studies mentioned for current distribution on ground-based 

launching facilities, current with a frequency on the order of 50 to 100 kHz 

proved satisfactory. Clamp-on current transformers of the requisite size and 

sensitivity are commercially available. 

It may well be that the cost of building such a model and making actual 

measurements of current distribution might be cheaper than the cost of trying to 

study current distribution by field plotting methods, either manual or computer 

aided, or by analytical determination of the distribution of currents in parallel 

filaments. 
WIRES TENDING TO 

JOIN AT RIGHT ANGLES 

OSCILLOSCOPE MONITORING 

CURRENT DENSITY 

RRENT MEASURING 

TRANSFORME 

OSCILLOSCOPE MONITORING 

TOTAL CURRENT 

Figure 9.1 I The wire grid model. 

9.3 Examples of External Magnetic Fields 

The magnetic field within and around the cluster of wires can be 

determined by taking the proper summation of the magnetic field produced by 

each individual wire. The computer program MAGFLD provides such a 

summation. 

Some examples of calculations performed with MAGFLD will now be 

described. They will serve to illustrate some of the points discussed in previous 

sections. The geometry chosen for analysis is the fuselage of the hypothetical 

aircraft shown in Figure 9.12. The aircraft, whose airworthiness is not under 

discussion, has a fuselage of elliptical cross section, two meters along the major 

axis and one meter along the minor axis. The fuselage is considered long enough 

that no end effects need to be considered. A lightning current of 1000 A is 

assumed to enter the nose and to exit through the rear of the fuselage. This 

elliptical fuselage is represented by 48 parallel conductors distributed in the 
manner shown in Figure 9.13. Figure 9.13 also shows the current density, or 

magnetic field strength, at the surface of the cylinder, both as determined from 

the program MAGFLD and analytically according to Equation 9.1 1. 

Figure 9.12 A hypothetical fuselage to be modeled as a wire grid. 

One quadrant of the elliptic cylinder is shown on Figure 9.14. This figure 

shows the magnetic field strength both within and around the wire grid as 

calculated by MAGFLD. The orientation of the magnetic field is shown by the 

direction of the arrows, and the strength both by the length of the arrows and 

by the indicated contour lines. The magnetic field strength inside the grid is 

much smaller than that outside, since on the inside the fields from each of the 

filaments tend to cancel, whereas outside the grid they tend to add. In Figure 

9.14 the fields inside the grid are largely the result of the finite number of wires 

defining the elliptic cylinder. If the cylinder were defined by more conductors, 

the fields inside would be smaller, presumably vanishing if the number of wires 

were to become infinite. Figure 9.15 shows a similar plot with one important 

difference: the current was forced to be equal in each wire. The magnetic field 

pattern produced here would be that determined by the resistive current 

distribution, the pattern that represents the final stage after currents and current 

density have become uniform. The orientation of the field external to the grid 

shows only relatively slight differences from that in Figure 9.14, but those 

differences are sufficient to account for the increased field intensity in the 

interior of the grid. A third example of field distribution, shown in Figure 9.16, 

assumes that on each side of the fuselage there is a recessed cavity. Such a cavity 

would be an approximation of the equipment bays for electronic equipment 

frequently found on military aircraft. Figure 9.16 again assumes the current in 

286 
'* t PERIPHERAL DISTANCE IN METERS 

tLLW llC4 

APPROXIMATING \ 

Y 

;,t Om- 

, I I 

0.2 0.4 0.6 0.8 

-0.1 METERS 

-0.2 -- 

Figure 9.13 Wire grid approximation of the elliptical fuselage. 

each fdament to be controlled by the magnetic distribution. The field patterns 

clearly show the field in the recessed cavity to be less than the field at other 

exterior points on the fuselage. 

The degree to which the MAGFLD program can calculate the magnetic 

field external to the aircraft fuselage, while interesting, may not be of great use 

for aircraft studies, since it is usually only the field intensity at the surface of the 

structure that is of interest. The program, however, is capable of calculating the 

field distribution interior to these simple geometries with sufficient accuracy for 

many purposes. One such purpose is to help determine the effective area through 

which circulating currents flow. For such calculations, as shown on Figure 9.1 5, 

the number of filaments defining the structure can be greatly increased, since 

each of them will carry a uniform current. Determining the current distribution 

on many more fdaments when that current is determined magnetically might 

lead to numeric problems, since the order of the matrix that must be inverted is 

equal to the number of conductors involved. The greatest utility of the 

MAGFLD program would be to calculate the current density or magnetic field 

density at the surface of a structure. Figure 9.17 shows the current distribution 

on the filaments for the geometry previously shown on Figure 9.16. The current 

density here is determined magnetically. 
/ //'/ / //' / / / / /,// I? 

/ --*---*---& 

*-,-ae , / -1 

/ /'r / /// r / / , /--- ,/,--- 

/ /- 

/%','- 

/--- --,.. 0 / / /' ' 

J 

ill I i J/'/ / JH/,*- 0 \ \, ,=-- *' / / 

J 

i r /" / ,&I/ I,*/ \ 1 f, .J'Y \:TI\ \&I \cp. , / ,&q I 5~ i 

6 '/ 1 I JI (,p-/ / 5: Z 1 j \Q/&// \ yiE , +- 

I /I - ,X , , II/ 

LLLL~-.---L- lfl-~LLl-L-- 1 I,' J 0 


0 0.1 0.2 0.3 0.4 0.5 

Figure 9.1 7 Current distribution and magnetic field strength at the surface 

(one quadrant only shown). 
REFERENCES 

9.1 V. Belevitch, "The lateral skin effect in a flat conductor," Philips 

Technical Review, 32,61718 (1971): 221-231. 

9.2 H. Kaden, "~ber den Verlustwiderstand von Hochfrequenzleitern," (On 

the Resistance of High-Frequency Conductors), Archiv fur Elektrotechnik, 

28 (1934): 8 18-25. 

9.3 J. B. Scarborough, Numerical Afuthematical Analysis, 6th ed. (Baltimore: 

Johns Hopkins University Press, 1966). 

9.4 G. E. Forsythe and W. R. Wasow, Finite-Difference Methods for Partial 

Differential Equations (New York: John Wiley and Sons, 1960). 

9.5 F. A. Fisher, J. G. Anderson, and J. H. Hagenguth, "Determination of 

Lightning Response of Transmission Lines by Means of Geometrical 

Models," Power Apparatus and Systems, 78, American Institute of 

Electrical Engineers, New York, New York (February 1960): 1725-36. 

9.6 J. G. Anderson and J. H. Hagenguth, "Magnetic Fields Around a 

Transmission Line Tower," Transactions of the AIEE, Part 3: Power 

Apparatus and Systems, 77, American Institute of Electrical Engineers, 

New York, New York (February 1959): 1644-50. 
CHAPTER 10 

THE INTERNAL MAGNETIC FIELDS CREATED BY DIFFUSION 

10.1 Circular Cylinders 

Consider Figure 10.1 in which a current, I, is entering a circular cylinder. 

The cylinder is considered long compared to other dimensions, so that there are 

no end effects, but short compared to the electrical wavelength of any of the 

frequency components of the current I. The return path for the current is not 

shown, but it is assumed to be sufficiently far away from the cylinder that there 

are no proximity effects. Also shown are two conductors, one external (I) and 

one internal (2) to the cylinder. These are connected to an end cap considered 

sufficiently massive that no electromagnetic fields penetrate the cap. At the 

other end of the cylinder are shown two voltages, V1 measured from conductor 

1 to the external surface of the cylinder, and V2 measured from conductor 2 to 

the inner surfilce of the cylinder. 

10.1.1 Resistance 

The cylinder will have a dc resistance 

where p = resistivity 

Q = length 

A = cross-sectional area 

r = radius 

a = thickness (a<<r) 

If the cylinder has the following dimensions and is made of aluminum of the 

indicated resistivity 

Q = 2m 

r = 15.7cm 

a = 0.28 1 mm (0.01 5") 

p = 2.69x10-~ !2. rn 

then the dc resistance, R, will be 1.43 x 10-4 R. If we consider that the input 

current is 116 A, there will then be developed a voltage 

If the current of 116 A is established and allowed to flow for a time sufficiently 

long that steady state conditions are reached, the voltage developed along the 

cylinder will then be 16.6 mV. This same voltage drop would be measured by a 

conductor external to the cylinder or by one internal to the cylinder. 
P 

A GENERAL POINT 

m Y1 

3 I50 w 

i% 

Z g loo 

Ji 

z 

r 

z50 9 

LU 

Y 

METERS 

Figure 10.1 Magnetic fields around a circular 

(a) Geometry 

(b) Field intensity vs radius (for 

cylinder. 

I = 116 

10.1.2 The External Loop Voltage 

Until such steady state conditions have been established, V1 will not be 

equal to V2, and neither of them will be equal to the steady state resistance 

voltage drop. Consider first voltage V1. External to the cylinder the flow of 

current sets up a magnetic field of intensity. 
where I = current 

r = radius 

H = field strength 

having a pattern as shown in Figure 1 O.l(b). V1 will then be 

and @, the flux passing through the loop ABCD, is represented by the shaded 

area of Figure lO.l(b). The flux, @, would be measured in webers. Remembering 

that 

VI becomes 

If 12 is 31.4 cm andtheindicated current of 116 A rises to crest in an equivalent 

time of 0.25 ps (Figure 10.2), V1 will then rise to an initial voltage of 129 V. As 

steady state conditions are reached and the external magnetic field ceases to 

change with time, V1 will decay to its steady state value of 0.0166 V. 

-4 0.25~~ time 

Figure 10.2 External voltage (not to scale). 

This analysis ignores the skin effect, a phenomenon that causes the 

effective resistance to be higher under transient conditions than under steady 

state conditions. For conductors external to the cylinder, the increased 

resistance resulting from skin effect will be of little consequence compared to 

the much larger voltage induced by the changing magnetic field. 
While conductors external to a current-carrying cylinder, such as an 

aircraft fuselage, are not common, they are not unknown. An example might be 

the cables on a missile or rocket that run between a control assembly in the nose 

and the engine controls at the tail. Of necessity, such cables must run, as shown 

in Figure 10.3, external to the fuel and oxidizer tanks. If the cables are not in a 

shielded cable tunnel, they will be exposed to the external magnetic field. 

If the spacing between the wires and the surfaces of the vehicle is not 

large, Equation 10.5 may be somewhat simplified, since the magnetic flux 

density does not vary greatly with distance away from the surface of the vehicle 

and may be considered uniform. 

10.1.3 The Internal Loop Voltages 

Let us now consider the conditions internal to the cylinder, where 

conditions are very different. First of all, the magnetic field internal to the inner 

surface of the cylinder will be zero. This may be shown as in Figure 10.4. Let 

there be passed two planes parallel to the axis through any internal point, P, the 

planes separated by a small angle, a. The magnitudes of the currents in the two 

filaments are to each other as the ratio of the distances from point P. 

Furthermore, the magnetic intensity effects of the filaments at P have the ratio 

unity 

These two filaments then create no field at any internal point; nor does the tube 

as a whole, as it is composed of filament pairs. 

The second important factor is that the phenomenon of skin effect delays 

the buildup of current on the inner surface of the cylinder. The origin of the 

skin effect phenomenon is shown in Figure 10.5. If a magnetic field line is 

assumed to be suddenly established internal to the wall of the conducting 

cylinder, there will be induced eddy currents circulating around that field line. 

These eddy currents will induce a magnetic field of their own of polarity 

opposite to that set up by the external magnetic field. Only as the eddy currents 

decay will the magnetic field penetrate the wall of the cylinder. 

The current density across the wall thickness at several times is shown in 

Figure 10.5(b). Initially, the current density for a step-function input current is 

entirely confined to the outer surface. At a slightly later time, t 1, the current 

296 

Figure 10.4 Cylindrical tube of current. Internal magnetic field intensity is 

zero. 

density will still be high near the outer surface, but a small amount of current 

will have penetrated to the inner surface. As time increases, the current density 

on the outer surface will fall, and the current density on the finer surface will 

increase, until at tfinal the current density will be uniform across the surface of 

the wall. At any time, the area under the current density curve multiplied by the 

peripheral distance around the cylinder will equal the input current. The resistive 

voltage drop along any surface (including any line internal to the wall of the 

cylinder) will be equal to the product of the current density at that point times 

the resistivity of the material. Accordingly, the voltage drop along the inner 

surface must folldw the curve defining the buildup of current density on the 

inner surface. 

10.1.4 The Characteristic Diffusion-Type Response 

The nature of this response is shown in Figure 10.6 and is governed by the 

following equation : 

Figure 10.6 is plotted in terms of a characteristic time constant, T, sometimes 

called the penetration time constant. 
r EXTERNAL MAGNETIC FIELD LlNE 

/ INTERNAL MAGNETIC FIELD LlNE 

/I r CIRCULATING CURRENT 

dl 1 r MAGNETIC FIELD LlNE PRODUCED BY 

THE CIRCULATING CURRENT 

(b) 

Figure 10.5 Factors governing the internal voltage. 

(a) Geometry and decaying eddy currents 

(b) Current density at different times 

R, the dc resistance of the tube, is 

where p = resistivity of the material 

a = thickness of the wall 

r = radius of the tube 
L, the internal inductance of the tube, is given by 

where p = permeability of the material 

Hence the penetration time constant is seen to be 

pa2 7 =- 

n2p 

time 

Figure 10.6 Diffusion-type response to a step function. 

The response curve, shown in Figure 10.6, is called a drfusion-type 

response and is characteristic of many types of situations involving the 

transmission of energy through a distributed medium. An example would be the 

transfer of heat into a block if a heat flux were suddenly applied to one face of 

the block. Another example would be the transfer of electrical energy through 

an R/C or L/R ladder network. Figure 10.7 shows examples, with Figure 10.7(b) 

often being used as the best illustrative example of the phenomenon of skin 

effect. Several important observations might be made about the shape of the 

response curve shown in Figure 10.6. The first is that the response initially 

changes only slowly and thus has a zero first derivative, unlike a simple 

exponential response, which has a finite first derivative. The second is that the 

response approaches its final value much more slowly than does a simple 

exponential response. The third is that in three tirne constants the response has 
reached a large fraction of its final value. In three tinie constants the 

diffusion-type response reaches 90%', while the exponential response reaches 

95% of its final value. 

R/n (n SECTIONS) 

(n SECTIONS) I ouf 

(b) 

Figure 10.7 Ladder networks displaying a diffusion-ty pe response. 

(a) R/C network 

(b) L/R network 

With respect to Equation 10.13 it should be noted that the penetration 

time constant is directly proportional to the pernieability of the material, 

inversely proportional to the resistivity of the material, and directly proportional 

to the square of thickness of the material. The relative pernieability of structural 

material used in aircraft is almost always very nearly unity. Thickness and 

resistivity can vary over wide ranges. 

For reference, Equation 10.13 is shown plotted in Figure 10.8 as a 

function of material thickness and the resistivity of the material. The resistivities 

of some typical metals are shown in Table 10.1. As an exarnple, if we assume an 

aluminum alloy of resistivity twice that of copper and a skin thickness of 0.040 

inches, the penetration tinie constant would be 3.9 ps. If a step-function current 

were established on the outside of a sheet of such metal, it would take 11.7 ps 

for the current density on the other side to buildup to 90% of its final value. 

Returning to Figure 10.5, we are now in a position to evaluate V2. The 

fundamental definition of the voltage between any two points is 
where P is the potential at any point. There is an infinite variety of paths that 

one (night take going from point A to point D, but the simplest to consider is 

the path A-B-C-D. The path A-B is along the inner surface of the cylinder. 

Integrating along this path, the potential difference obtained is the internal 

voltage drop on the wall of the cylinder. Along path B-C there will be no 

potential difference, since we have assumed that the end cap is sufficiently 

massive that there will be no voltage drops along its inner surface. Along path 

C-D again there will be no potential difference, since we will assume that the 

conductor is not carrying any current. To the resistive drops will then be added 

the voltage produced around the loop by the passage of changing magnetic fields 

through the loop ABCD. As explained earlier, in a circular geometry there is no 

magnetic field internal to the cylinder. Consequently, there will be no 

magnetically induced component of voltage, and V2 will have the shape shown 

on Figure 10.6, where the final value is given by the dc resistance drop along the 

cylinder. 

INCHES 

Figure 10.8 Skin thickness vs penetration time constant. 

302 
Table 10.1 RESISTIVITIES OF TYPICAL METALS 

Conductivity Resistivity 

Resistivity as a fraction of relative to 

Material ohm-meters that of copper copper 

copper 1.68 x 1 oe8 1 .O 1 .O 

aluminum* 2.69 x lo-' 0.62 1.6 

magnesium** 4.46 x 0.38 2.65 

nickel lox lo-8 0.17 6 

Monel 42 x 0.04 25 

stainless steel 70x 0.024 4 2 

Inconel lOOx lo-8 0.01 7 60 

titanium 180 x 0.009 107 

*Aluminum alloys range from 2.8 to 5.6 x S2.m; harder alloys generally 

have higher resistivities. 

**Magnesium alloys containing aluminum and zinc range from 10 to 17 x 

Om. 

10.2 Elliptical Cylinders 

If we consider a cylinder of other than circular shape, conditions may be 

considerably different. Figure 10.9 shows an elliptical cylinder into which a 

step-function current is injected. As previously, we will assume that the cylinder 

is long enough that all end effects may be neglected, that it is short compared to 

the wavelengths of any frequency components of the injected current, and that 

the return path for the current is far enough removed that no proximity effects 

need be considered. 

We will also treat the instantaneous current in the cylinder as being 

composed of the sum of a steady state and a transient component of current. 

The transient component takes the form of a circulating eddy current. For the 

following section it should be kept in mind that the eddy currents described 

represent only the transient component of current and that the total current at 

any point or time is the sum of the two components. 

10.2.1 Eddy Currents and the External Magnetic Field 

Under steady state conditions the current density along the wall of the 

cylinder will be governed by the dc resistance and, if we assume uniform wall 

thickness, will have a uniform current density. The current in the cylinder will 

produce a magnetic field. Most of the field lines will completely encircle the 

cylinder, as shown in Figure 10.9(a), but some, because of the uniform current 

density, will pass through the cylinder. The greater the eccentricity of the 
Figure 10.9 Magnetic fields around an elliptical cylinder. 

(a) Penetrating lines of flux 

(b) Detail showing resolution into components 

(c) Circulating currents induced by 

penetrating lines of flux 

cylinder, the greater will be the number of lines of flux passing through the 

cylinder. Those lines that do pass through the walls of the cylinder will have 

important effects on the ~nternal response of the cylinder prior to the time that 

steady state conditions have been established. One such penetrating line is shown 

in Figure 10.9(b). The vector defining that line may, at the point of entry, be 

resolved into two vector components, one normal to the surface, @,, and 

another tangential to the surface, Cbt. If the field line @ is suddenly established, it 

will induce a circulating, or eddy, current in the conducting sheet which it 

attempts to penetrate. The eddy current will produce a magnetic field of its 
own, and the intensity of the eddy current will be of a nature such that the 

magnetic field produced is exactly that required to cancel the normal 

component of the exterior field. If, as shown in Figure 10.9(c), a number of 

lines of magnetic flux attempt to penetrate the surface of the elliptical cylinder, 

the eddy currents produced by each line of flux combine to produce a 

circulating current. In an elliptical cylinder of this nature there will be four 

regions of circulating current, two on each of the two sides. These circulating 

currents will be of a nature such as to increase the current density at the edges of 

the cylinder and to reduce it along the center. They also cancel any penetrating 

magnetic field, forcing the field around the cylinder to be entirely tangential to 

the surface, at least initially. 

The eddy currents cannot exist forever, since energy will be lost as the 

currents circulate through the resistance of the metal sheet. Accordingly, the 

current density at all points will vary with time, eventually becoming uniformly 

distributed in structures having uniform thickness and made of uniform 

resistivity materials. Figure 10.10 shows the manner in which the current density 

will vary. As the circulating current shown in Figure 10.10(c) decays to zero, 

the current at the edge will decay from its initial high value to the final 

resistively determined value, and the current at the center will increase from its 

initially low value. The current densities will change according to an essentially 

exponential pattern, though the transient increase in surface resistance produced 

by skin effects will prevent the circulating current from following a true 

exponential decay. 

One expression giving the magnitude of this redistribution time constant 

that has appeared in the literature (Reference 10.1) is: 

where A = enclosed area of structure 

P = peripheral distance around the structure 

The thickness of the wall, a, is assumed to be very small compared to other 

dimensions. For a rectangular box of sides height h and width d, Equation 10.15 

becomes 

A physical treatment of the redistribution time constant can also be given. 

If the decay is treated as exponential, the factors governing that decay are the 

characteristic inductance and resistance of the path through which the 

circulating currents flow. These factors are illustrated in Figure 10.1 1. The 

circulating currents may be viewed as flowing in one direction along a strip along 

the edge of the cylinder and back along another strip towards the center of the 

cylinder. The rate at which the circulating currents decrease is governed by the 

inductance and resistance of the pair of hypothetical strips. Such a pair is shown 
jedge 

Jcirculoting 

I c---- - 1,- ,,/ 

Jcenter 

b 

(a) 

I I I I I 

0 l~c 2% 3% 4% 5% 

time 

Solid: exponential-type decay 

dotted: decay pottern considering skin effects 

0 --- 

jCir 0 C 1 Tc 2% time 3% 4% 5% 

Figure 10. I0 Variation of current-density with time. 

(a) Current components defined 

(b) Edge and center currents 

(c) Circulating currents 

at the top of Figure 10.1 l(a). The resistance of the pair of strips will be 

where S = width 

Q = length 

a = thickness 
(b) 

Figure 10.11 Geometrical factors governing the decay of the circulating current. 

(a) The character of the L-R circuit 

(b) The width of the strip involved 

The inductance of a similar geometry (Figure 10.12) is given by Grover 

(Reference 10.2) as 

where K and e are factors relating to the geometrical mean radius of the strips. 

For the limiting case of Figure 10.1 2(b) 

log, K = -.I137 

log, e = 0 

L = 0.555 x 10% 
Figure 10.12 Inductance of parallel strips. 

(a) General case 

(b) Limiting case 

The inductive time constant would then be 

and for alu~~tinurn ofresistivity 2.69 x S2-m 

The time constant governing the decay of the circulating currents is thus 

seen to depend upon the width of the region through which the lines of flux 

would tend to penetrate. Since the penetrating lines of flux would tend to be 

confined to the outer portion of the ellipse, it appears that the effective width 

would be of the order of one-eighth of the peripheral distance around the 

cylinder. 

A numerical example is as follows: if the effective-width strip through the 

circulating currents flow is 0.123 m and the thickness is 0.038 cm (0.015 

inches), the L/R time constant is 480 ps. 

308 
10.2.2 Eddy Currents and the Internal Magnetic Field 

As the circulating currents die out and the external lines of flux penetrate 

the walls of the cylinders, there will be set up an internal magnetic field oriented 

as shown in Figure 10.13. In its latter stages the rate at which the internal field 

builds up will be dependent upon the rate at which the externally induced 

circulating currents die away. For an elliptical cylinder made of aluminum of 

thickness 0.038 cm, having a major axis of 47 cm, a minor axis of 9.4 cm, and a 

peripheral distance of 98.7 cm, the internal field will build up with a time 

constant of the aforementioned 480 MS. 

FINAL FLUX ORIENTATION 

INITIAL FLUX ORIENTATION 

t t 

I 

/ 

/ 

Figure 10.13 The internal magnetic field that arises as a result of flux 

pentration. 

The early time buildup of the magnetic field will be strongly influenced by 

the phenomenon of skin effect, and the initial rate of buildup of magnetic field 

will now be seen to have an important effect on the voltages induced on circuits 

contained within such an elliptical cylinder. Figures 10.5 and 10.6 showed how 

an electric field will be built up along the inner surface of a circular cylinder. In 

a similar manner, an electric field will be built up along the inner surface of the 

elliptical cylinder, but, unlike the circular cylinder, the electric field intensity 

will be different at different points on the internal surface of the elliptical 

cylinder, since the initial current density on the outer surface is not uniform. 

The internal electrical field will be essentially proportional to the current density 

on the exterior surface, but delayed and distorted as a result of the pulse 

penetration time produced by eddy currents (shown in Figure 10.6). The 

internal electric field, Ei, being different at different points on the inner surface, 

will give rise to circulating currents on the inner surface. The currents may again 

be visualized as flowing on paired strips. There will be four pairs of strips, two 

on each of the sides of the cylinder. Each of these strips will have associated 

with it an inductance and resistance which may again be calculated from the 
expressions given earlier. The equivalent circuit governing the rate of buildup of 

internal current then becomes that of Figure 10.14, which may be simplified to 

that of Figure 10.15. If eie were a step function, the buildup of current would 

follow the expression 

'circ 

(b) 

Figure 10.14 Factors governing the early time buildup of the internal magnetic 

field. 

(a) Physical factors 

(b) Equivalent circuit 
eie = on effective driving voltage 

reloted to the initiol nonuniformity of external current 

Figure 10.15 Final equivalent circuit governing increase of internal circulating 

circuit. 

If the internal voltage were not a step function but were governed by some other 

function of time, such as the pulse penetration predicted by Equation 10.9, the 

internal current would be 

The convolution integral of Equation 10.25 need not be evaluated analytically, 

since only the early time rate of change of current is important. So long as the 

current is small, the voltage across the resistive part of the circuit will be small 

and the voltage across the inductive part will be equal to the internal resistance 

drop. Under these conditions the internal current will be 

The internal magnetic field will follow the same time pattern as that of the 

buildup of the internal circulating currents. 

10.2.3 Internal Loop Voltages 

We are now in a position to evaluate the voltages on conductors contained 

in a cylinder of noncircular geometry. Figure 10.16 shows an elliptical cylinder 

with two internal conductors, one adjacent to the surface and one in the center. 

Both are connected at one end to an end cap sufficiently massive that no voltage 

drops will appear along its inner surface. The other ends are open circuited. The 
usual assumptions about the length of the cylinder and the return path for the 

injected current are assumed to apply. Voltages V1 and V2 are shown, both 

being measured between their respective conductors and a point on the inner 

wall of the cylinder. Figure 10.16(b) shows that all of the internal flux will pass 

between conductor 2 and the inner wall of the cylinder, while only a small 

amount will pass between conductor 1 and the inner wall. Correspondingly, a 

large fraction of the internal flux passes through the plane defined by 

conductors 1 and 2. 

A POINT ON THE 

INNER WALL 

CONDUCTOR 1 . CONDUCTOR 2 

I \ SMALL AMOUNT OF INTERNAL FLUX I I 

(b) \ LARGE AMOUNT OF INTERNAL FLUX 

Figure 10.16 Factors governing the internal voltaye, 

(a) The geometry 

(b) Internal flux linkages 

The voltage between any two points is defined again as the line integral of 

the potentials around a closed path. Figure 10.17(a) and (b) shows the sinlplest 

paths to consider. V1 would be the sum of the potentials developed around the 

loop ABCD. If there is no current along conductor 1, the potential along path 

A-B will be zero. The potential drop along the path B-C will be zero because of 

the assumptions regarding the end cap. The potential along the path C-D will 

then be the voltage drop produced by the inner current density times the 
=resistive = 1 

--------- ---- 

(4 

0 time 

Figure 10.1 7 The internal voltages. 

(a) Detail of the edge surface 

(b) Paths of integration 

(c) Components of V1 

(d) Components of V2 

313 
resistivity of the material along the path C-D. To these potentials must be added 

the voltage induced magnetically by the changing magnetic flux passing through 

the loop defined by the points A, B, C, and D. If the spacing of the conductor to 

the wall is made vanishingly small, so that C-B and D-A become zero, there will 

be no magnetic flux; hence the voltage V1 between points A and D will be only 

the resistive voltage drop along the path C-D. As in the cylindrical geometry 

case, the voltage for a step-function current injected into the exterior of the tube 

will build up according to the pattern shown in Figure 10.6. Its magnitude will 

be greater than the dc resistance drop by the ratio to which the initial current 

density on the exterior along the end of the ellipse exceeds the steady state 

current density. 

V2 will again be the sum of a resistive voltage drop and a magnetically 

induced drop, this time along the path E-F-C-D. The resistive component will be 

identical to the resistive component of V1, the resistance drop along the path 

C-D. For V2, however, there will be a nonzero magnetic component of voltage 

produced by the passage of a finite amount of magnetic flux through the finite 

loop EFCD. The magnetically induced component of voltage will be given by 

where K is a proportionality constant relating the flux produced in the loop 

EFCD to the internal current. In Equation 10.26, however, it was shown that 

the internal current was proportional to the integral of the internal resistance 

drop. This leads to the rather unusual observation that the magnetically induced 

component of voltage has, initially at least, the same waveshape as the 

component of voltage produced by the flow of internal current density through 

the resistance of the material. The long term response of the magnetically 

induced voltage will be different from the resistively generated component, 

since, as steady state conditions are reached and the internal magnetic field 

reaches its final value, its rate of change of the internal magnetic field will 

decrease to zero. 

The amount of the magnetically induced voltage will depend upon the 

location of the conductor and upon the degree to which the initial distribution 

of magnetic flux around the outside of the cylinder differs from the final 

distribution of magnetic flux. Since the difference between the initial and the 

final flux patterns is greater for cylinders of high eccentricity than it is for 

cylinders of low eccentricity, it follows that the flatter the cylinder, the greater 

will be the influence of the magnetic component. 

10.2.4 Experimental Verification 

In order to demonstrate the above principles, a test was made in which 

current was injected into an elliptical cylinder and the voltage on several internal 

conductors measured. The physical arrangement of the test setup is shown in 

Figure 10.18. The cylinder had a major axis of 47 cm, a minor axis of 9.4 cm 
ELLIPTICAL CYLINDER 

SCREEN ROOM n 

\ 

V = 20kVm,, SIX EQUALLY SPACED' 

CURRENT RETURN WIRES \I 

TWO SHEETS 

OF ALUMINUM 

WITH EDGES 

OVERLAPPED 

ABOUT 14 mm 

Figure 10.18 Setup of test to demonstrate internal voltages in an elliptical 

cylinder. 

(majorlminor = 5) and was made from two sheets of aluminum bent over a 

wooden framework with the overlapping edges stapled at close intervals to the 

internal wooden framework. The edges overlapped by about 1.4 cm. After 

stapling, the seam was covered with a 5 cm wide sheet of adhesive-backed 

aluminum fod. The thickness of the metal sheet was 0.38 mm (0.015 inches). 

One end of the cylinder was bolted to a metal plate facing the outer wall of a 

screen room. The other end of the cylinder was sealed w~th metal strips and 

connected to a pulse generator. The pulse generator consisted of a capacitor 

bank capable of being charged to a maximum of 20 kV. For most of the tests 

the capacitor was 48 pF and was discharged through 160 R into the end of the 

elliptical cylinder. Current was brought into the end of the cylinder through four 

conductors so arranged as to minimize any local end effects. Current was 

returned to the pulse generator through six equally spaced return wires spaced 

about 18 inches from the center of the cylinder. Multiple return wires were used 

both to minimize the external inductance through which the generator had to 

discharge and to minimize any possible proximity effects. Internal to the 

cylinder were placed eight conductors at different locat~ons, the locations being 

as shown in Figure 10.19. Two of the conductors were taped against the inner 

surface of the metal sheet. One of these, conductor number 1, was located at the 

narrow end of the cylinder, while the other, conductor number 8, was located 

about halfway to the center. The rernaining conductors were all located at 

different points along the major axis of the ellipse. The conductors were number 

22AWG and insulated with the th~ckness of insulation normal to such 

conductors. Each conductor was connected at the current input end to the inner 

surface of the end cap through a resistor of 68 L2, the purpose of the resistor 
being to damp out any internal oscillations that might be developed. The 

conductors were terminated on switches so that measurements could be made to 

ground on any conductor or be made between any two conductors. 

CONDUCTORS 1 AND 8 WERE TAPED 

2 FLUSH AGAINST THE INNER 

SURFACE OF THE METAL SHEET 

SCREEN ROOM 

4 

*y 

ABOUT 

1 

400 mm 

OF RG 58 A/U 

COAXIAL CABLE 

ELLIPTICAL CYLINDER 

Figure 10.1 9 Arrangement of conductors inside elliptical cylinder. 

(a) Detail 

(b) End view showing locations of conductors 

(c) ElectricaI termination of internal conductors (typical) 

The measurement principles are shown in Figure 10.20. Line-ground 

voltage measurements were made with the normal input preamplifiers of an 

oscilloscope. The oscilloscope was located in the screen room directly adjacent 

to the end of the elliptical cylinder and connected to the output jacks on the 

cylinder by only a short length of coaxial cable. Accordingly, the measurenients 

would not be influenced by any loading effects of the measurement cables. 

Voltages between pairs of conductors were made with the oscilloscope 

preamplifier in the A minus B mode. Since the voltage between any pair of 

conductors would depend only upon the magnetically induced component of 

voltage, the resistive component being rejected by the d~fferential amplifier, it 

was then possible to integrate the difference voltage and derive a signal 

proportional to the internal magnetic field. This was done using an operational 

amplifier connected as an integrator, shown in Figure 10.10(c). 

Typical tests results are shown in Figures 10.11 and 10.22. Figure 10.21 

shows, on four different time scales, the injected current and the line-to-ground 

voltage on two different conductors, conductor I located adjacent to the small 

radius end of the ellipse and conductor 7 located near the center. With the 
1 eo = - I (V ,-V ,) dt 

Rc 

'+ = I (Vl-V2) dt 

Figure 10.20 Measurement principles. 

(a) Line-ground measurements 

(b) Line-line measurement 

(c) Magnetic field measurement 

exception of the oscillograms displayed at 500 ps per division, the input current 

can be considered as a step-function current. In both cases the line-ground 

voltage is seen to display the pulse penetration-type response shown in Figure 

10.6. The small bump at the leading end of the voltage on conductor 1 is 

primarily the result of leakage from the surge generator into the screen rooni 

housing the measuring instruments and to coupling from the trigger pulse used 

to fire the surge generator. In both cases the initial rise of voltage follows the 

same pattern and reaches essentially its final value in about 3 ,us, a time in 

accordance with the calculated pulse penetration time constant of 1.7 ps 

discussed previously. At later times the voltage on conductor I is seen to decay 

back toward a lower final value in the manner shown in Figure 10.17. Likewise, 

the voltage on conductor 7 is seen to be nearly flat, rising only slightly, as shown 

in Figure 10.21. It is probably only a fortuitous combination of conductor 

location and the characteristics of the elliptical cylinder that result in the voltage 
INJECTED 

CURRENT 

CONDUCTOR 

VOLTAGES 

Figure 10.21 Variation of conductor voltage with time. 

(Leading edges of waveforms retouched for photographic clarity) 

on conductor 7 beirtg nearly flat. If the injected current were truly a step 

function, the two voltages would eventually become equal to the dc resistance 

drop. The bottom set of oscillograms indicates that the current significantly 

departs from a step-function current pattern at later times, and this is reflected 

in the long-time response of the two line-ground voltages. Just as the voltage on 

conductor 7 lagged behind the voltage on conductor I during the initial phase of 

the response, it lags behind during the final decay of the current. The 

redistribution time constant with which VI decays toward its final value is seen 

to be of the order of 300 ps, in accordance with the calculations predicated on 

the theoretical factors described in Figure 10.10. 
VOLTAGE - CONDUCTOR 1 - CONDUCTOR 7 

Figure 10.22 Difference voltages and total flux. 

(Leading edges of waveforms retouched for photographic clarity) 

Figure 10.22 shows the voltages between conductors (displayed on two 

separate oscilloscopes and hence displayed on two separate oscillograms) and the 

i;igral of the conductor-to-conductor voltage. his latter is of course 

proportional to the magnetic flux that builds up internally to the cylinder. 

Since, at early times, the line-to-ground voltages have the same waveshape but 

different amplitudes, it follows that the line-line voltage will also have that 

waveshape. Unlike the line-ground voltages, the line-line voltage decays toward 

zero. Since the input current was not a step-function current, the oscillograms 

displayed at 500 ps per division show the line-line voltage decaying to zero and 

then reversing as the input current decays and the internal magnetic fields 

attempt to follow the decaying external magnetic fields. The input current did 

not last long enough for the magnetic field to reach a steady state value, but the 

oscillograms do indicate that it seems to crest at about 850 ps. Why the 
indicated magnetic field does not crest at the time the line-line voltage goes 

through zero, 1300 ps, is not known. Perhaps instrumentation errors in the 

integrator are to blame. At any rate, the decay time constant for the line-line 

voltage agrees with the value predicted by Equation 10.23. 

With a somewhat different waveshape of injected current, Figure 10.23 

shows the voltages as a function of the position of the conductor. The voltage on 

conductor 8 is not shown; however, it was virtually identical with that shown for 

conductor 5. 

(5ps/div.) 

Figure 10.23 Voltages as a function of position. 

10.3 Fields Within Cavities 

It frequently happens that in the fuselage of an aircraft there will be i 

cavity that is effectively exposed to the external field on only one face, either 

because the inner walls of the cavity are thick enough to provide more shielding 

from the other parts of the external field or because the cavity is much closer to 

320 
Figure 10.24 A cavity exposed to a field on only one side. 

(a) Cavity and field orientations 

(b) An image in the reflecting surfaces defining the cavity 

(c) Current path defining the loop inductance 
one of the external surfaces than it is to any of the other external surfaces. A 

typical cavity would be a gun bay or an electronic equipment bay located dong 

the fuselage of the aircraft and accessible through access panels. For the moment 

the effects of the access panels will be ignored. The geometry is shown in Figure 

10.2qa). 

The cavity may be viewed as being formed from one face exposed to the 

external magnetic field and the current producing that field, and with its other 

faces formed by metal sheets, which, like the exposed surface, may be assumed 

to extend toward infinity. Mirrored in the reflecting surface defining the back of 

the cavity will be an image of the face of the cavity exposed to the magnetic 

field. The image will be spaced a distance, D, behind the face carrying the 

current, and in the image the reflection of the external current will be of 

opposite polarity. The electric field produced on the inner surface of the face 

exposed to the external current, shown in Figure 10.24(c), will act to force a 

current, Icirc, around the interior of the circuit defining the cavity. The length 

of the loop involved will be the same as the length of the cavity, and the width 

will be twice the depth of the cavity. The internal voltage gradient will be 

proportional to the product of the inner current density and the resistivity of 

the face carrying the external current and, as described earlier, will exhibit the 

characteristic pulse penetration buildup described in Figure 10.6. This voltage 

may be viewed as impressed across a loop or cavity inductance, Z 

The resistance and inductance will both be governed by the effective 

characteristics of the loop defining the cavity. 

Typical current paths and their characteristic impedances are shown in 

Figure 10.25. Some cavities may be viewed as being sufficiently long and narrow 

that they may be defined by parallel strips. Others are basically of rectangular or 

circular shape, or of some simple shape that may be approximated by an 

equivalent circular cylinder. The inductance and resistance of each configuration 

are shown. Each of the inductance equations (Reference 10.3) has a correction 

factor--loge k, F', or K-- that relates to the shape of the enclosure. These 

correction factors are shown in Figure 10.26. 

If the cavity is provided with a removable cover, and if this cover is in the 

external current flow or is exposed to the external magnetic field, the effects are 

as illustrated in Figure 10.27. If the cover is assumed to be of the same material 

and thickness as the rest of the face upon which it is mounted, the principal 

effects relate to the resistance of the fasteners used to hold the cover in place. 

The covers will seldom make good electrical contact to the rest of the surface 

except at the fasteners themselves. Accordingly, the external currents flowing in 

the face will be constricted in the vicinity of the fastener and pass from that face 

onto the cover through the fastener. The major effect of this constriction of 

current flow is to introduce a lumped resistance into the electrical circuit, 

although there is a certain amount of inffuence on the inductance of the circuit 

whenever the current is constricted. It will be seen, then, that the greater the 

322 
L = 0.4 x 1c6f [log, d + 1.5 + log, K] S 

2 PI T R=- a S 

L 0.2 as x 106 d - = R [log,-+1.5+log,K] S 

(0) 

2 Pnr R =- as 

L 0.2 x 106 nar - - 

R K 

P 

Figure 10.25 Typical current paths and characteristic impedances. 

(a) Long sheets 

(b) Rectangular box 

(c) Circular cylinder 

number of fasteners, the less the restriction of current flow and the less 

resistance inserted into the current path. The resistance introduced by the 

fasteners is important because it is frequently much higher than the intrinsic 

resistance of the metal surface and because the resistance is not subjected to the 

323 
(b) 0 0.2 0.4 0.6 0.8 1.0 

s /I 

Figure 10.26 Correction factors for inductance. 

(a) Parallel strips 

(b) Rectangular boxes 

(c) Circular cylinders 

324 
Figure 10.27 Effects of covers and fasteners. 

skin effects that retard the buildup of current density on the inner surface. 

An equivalent circuit of the cavity including the effects of fasteners is 

shown in Figure 10.28. Rf is the equivalent resistance of the fasteners. 

Circulated through this resistance is the undistorted current flowing in the 

exterior face of the fuselage. The sum of R1 and R2 is equal to the intrinsic 

resistance of the loop defining the cavity, the resistances given by the equations 

in Figure 10.25. The external current will develop a voltage across only a portion 

of this resistance, since it flows in only a portion of the loop. Letting RI be the 

resistance through which the external current flows, that resistance may be 

assumed to be subjected to the current as retarded by the pulse penetration time 

constant. The two voltages developed across these resistances then circulate 

current around the entire loop. The rate at which the current builds up will then 

be the rate at which the magnetic field inside the cavity builds up. 
CURRENT WITH 

WAVEFRONT 

RETARDED BY 

PENETRATION 

TIME CONSTAN 

UNDISTORTED 

CURRENT 

Figure 10.28 Equivalent circuit governing buildup of magnetic field inside a 

cavity. 
REFERENCES 

10.1 K. Khalaf-Allah, Time Constant for Magnetic Field Diffusion into a 

Hollow Cylindrical Conductor, UKAEA (United Kingdom Atomic Energy 

Authority) Research Group Report CLM-R 14 1, Culham Laboratory, 

Abingdon, Oxfordshire, England ( 1974). 

10.2 F. W. Grover, Inductance Calculations: Working FormuIas and Tables 

(New York: Dover, 1962). 

10.3 Grover, Inductance Calculations, pp. 15-1 6 et passim. 

CHAPTER 11 

THE INTERNAL FIELDS COUPLED THROUGH APERTURES 

11.1 MagneticFields 

The most important mode by which magnetic fields are coupled from the 

exterior of the aircraft to the interior is coupling through magnetic apertures. 

There are two major reasons for the importance of aperture-coupled fields. The 

first is that some apertures may be quite large, windows being a prime example. 

The second reason is that, unlike fields coupled by diffusion through metal 

surfaces, the waveshape of the interior field is not retarded, but tends to be the 

same for small apertures as that of the external magnetic field. The most 

important consequence of the coupled magnetic fields relates to the voltages 

induced by such fields. A field of given peak intensity is more apt to cause 

trouble if it rises to its peak quickly than if it is delayed and distorted. 

Physically, the basic problem is as shown on Figure 11.1. An external 

magnetic field tangential to a metal surface passes across an aperture. At the 

aperture some of the lines of magnetic force leak into the interior region. 

Analytically, the problem is to be able to determine the field strength at any 

point in the interior volume, preferably to be able to determine both the 

magnitude and the direction of the magnetic field. 

Often the analytical problem is best solved by placing in the plane of the 

aperture an equivalent magnetic dipole, the fields around which may then be 

determined from dipole theory. The strength of the equivalent dipole is 

determined by the strength of the external magnetic field and by the size, shape, 

and orientation of the aperture. A geometry amenable to mathematical analysis 

(Figure 11.2) is that of an elliptical aperture in a conductive sheet of a size large 

enough to separate completely the interior volume in which the fields are to be 

calculated from the exterior volume in which the basic magnetic field exists. In 

the following analysis we will treat only fields tangential to the conductive 

surface, since, as explained in the previous sections, a magnetic field with 

components perpendicular to the conductive surface induces circulating currents 

that initially cancel the penetrating component. 

Mathematically the problem is most directly solved if the external 

magnetic field is resolved into two components, Hx and Hy, lying along an axis 

directed through the major and minor axes of the elliptical aperture. These field 

components give rise to two equivalent magnetic dipoles, M, and My, again 

oriented along the major and the minor axes of the aperture. The strength of 

these dipoles is a function of the strength of the external magnetic field that 

would exist in the plane of the aperture if the aperture were absent. The 

governing expressions (Reference I I .I) are 
EXTERIOR REGION 

FIELD FIELD 

/------------ 

I I 

1 I I I 1 I I 

I I I 

\ 

I 

I 

\ \ 1 I (1. lTex4>@ f I 1 EQUIVALENT / -\J' \ I I 

I DIPOLE I 

I I I 

/--A 

I I 

I I 

(b) 

Figure I I. I The aperture-coupling problem. 

(a) A field across an aperture 

(b) Equivalent dipole producing 

the same internal field 
Figure 11.2 An elliptical aperture. 

(a) Geometry and the components of the 

external magnetic field vector 

(b) The equivalent magnetic dipoles 

The expressions Ke and Ee are elliptical integrals of the first and second kinds, 

respectively. The quantities a\ 1 and a52 are shown in Figure 11.3. Since thc 

dipole strength is proportional to the cube of the major axis length of the 

aperture, it follows that large apertures will couple much more field into an 

inner volume than will small apertures, the strength of the dipole increasing 

faster than the area of the aperture increases. 

The pattern of magnetic fields in the interior volume produced by a dipole 
1211 1 

Figure 11.3 Shape factor for elliptical apertures. 

lying along the X axis would appear as on Figure 11.4; Assuming the dipole to 

lie in the XY plane, the field patterns will be concentric closed loops lying in 

planes passed through the X axis. At any point, P, the total magnetic field will 

be tangential to the corresponding closed loop. The components of the magnetic 

field at point P--Hx, Hy, and Hz--(Reference 1 1.2) will be 
Figure 11.4 Field patterns produced by a dipole lying along the - X axis. 

-------.-. L- J ., 

~~~icall~,-~h~ external &&'gnetic field would include an equivalent dipole 

oriented along the Y axis as well. 'This dipole would likewise produce a magnetic 

field at the point P, the componenjs of which would be 
The total magnetic field strength at P, then, would be the sum of the 

components produced by the dipoles lying along the X and Y axes. 

If = H,, + Hxy X 

The total field strength at point P would be the vector summation of the X, Y, Z 

components of the field. 

The field strength is seen to be inversely proportional to the cube of the 

distance from the aperture to the point in question. Accordingly, the 

aperture-coupled fields will be localized in space near the aperture. Small 

apertures are thus less troublesome than large apertures, both because of the 

decreased magnetic field strength associated with the corresponding dipole and 

hyause of the lesser distance one must be removed from the aperture before the 

field strength becomes negligible compared to the field strength closer to the 

aperture. = 

The above formulation has one major restriction: it is valid only in the 

low-frequency region, low-frequency imply~ng that the aperture is small 

compared to the wavele~gths of the field under consideration. Coupling through 

apertures increases with increasing frequency, as witness the quarter-wave 

resonant slot antenna. The fo~mulation also treats only the internal magnetic 

fields produced by external magnetic fields and does not treat the electric field 

produced by a changing magnetic field or the magnetic field produced by a 

changing electric field. Taylor (Reference 11.1) treats the interaction in more 

detail. This interaction between changing electric ~d .~?agne!ic fields is greater 

at high frequencies and, for the frequency s$c;um associated with lightning, is 

possibly of importance only at those frequenc'les for which the aperture is not 

electrically small. 

The above formulation is also valid on'$ when the point at which the field 

is to be calculated is at a distance from the ap,:rture that is large compared to the 

dimensions of the aperture. If an attempt is rnade to calculate the fields close to 

the aperture, the results are inaccurate. If the distance from the point to the 

aperture goes to zero, the calczated magnitude of fieldsbecomes infinite, 

whereas the field itself can never in fact bec~me larger (barring reflections) than 

the external field. It can be shown, in fact, hat the actual field strength in the 
plane of the aperture, as distinct from the strength of the equivalent dipole, will 

be half the field strength that would exist if the aperture were not there. This 

intractable behavior of the calculated magnitude of the field near the aperture 

arises as a result of a simplifying assumption commonly used in dipole analysis. 

This assumption, illustrated in Figure 11.5(a), is that point P is sufficiently far 

from the center of the dipole that each of the radial distances rl and r2 from the 

ends of the dipole to the point P may be represented by the average radial 

distance, r, and that each of the angles B1 and 02 may be represented by the 

angle 8. 

Figure 11.5 Dipole approximations. 

(a) Elementary dipole 

(b) Dipole of finite size 

With a great increase in numerical complexity, it is possible to develop a 

formulation for the fields at point P that, as shown in Figure 11.5(b), allows the 

dipole to have a finite length. Such a formulation allows one to determine with 

33 5 
better, though not perfect, accuracy the field strength close to the aperture. The 

governing equations (Reference 1 1.3) presented here, without proof, are 

H,=G1 xF1 +G2xF2-G3xF3 (1 1.16) 

H Y =G4xF1 +G5xF2-G6xF4 (1 1.17) 

H,=G7 xF1 +G8xF2 (1 1.18) 

where 
Frequently a calculation made to determine the effects of coupling 

through apertures needs to take into account the presence of reflecting surfaces. 

'Ihe simplest case, shown in Figure 1 1.6, treats one reflecting surface of infinite 

dimensions and parallel to the surface containing the aperture. The total field at 

any point between the two surfaces is the sum of the fields produced by an 

infinite array of images of the original magnetic dipole. The field strength from 

each of the images would again be determined from Equations 11.6 through 

1 1.1 1 or 11.1 6 through 1 1.32. Fortunately, in most cases only a few of the 

images need be considered because of the dependence of field strength on the 

cube of the distance to the point under consideration. In principle, additional 

reflecting surfaces could be included in the formulation to define completely an 

interior volume. Such multiple reflecting surfaces are shown on Figure 11.7. 

Typically, one is concerned, not with the magnitude of the fields 

themselves, but with the effects of the fields on electrical circuits. A typical 

problem, as shown in Figure 11.8, might be determining the voltage produced by 

an aperture-coupled magnetic field between a conductor and a metal surface 

containing the aperture. That voltage would be determined by the total magnetic 

flux passing through the plane defined by the conductor and its projection onto 

the metal surface. This total magnetic field would be the summation, both along 

the length and across the width of the loop so defined, of the magnetic field 

strength normal to the loop. If the field strength were to vary considerably with 

position along the loop, it would be necessary to evaluate the field strength at a 

number of points across the area of the loop and to perform some sort of 

graphical integration to determine the total magnetic flux passing through the 

loop. The numerical calculations, while straightforward in principle, would be of 

sufficient complexity to make the problem insoluble in any practical sense if all 

the calculations had to be made by hand. In order to overcome the numerical 
REFLECTING SURFACE 

IMAGE OF 2 IMAGE OF 1 IMAGE OF 3 

Figure 11.6 Reflecting surface. 

I I I I I I i I I 1 

4 4 + + + + 

SURFACE WITH 

APERTURE 

REFLECTING SURFACE 

A 

I t I 

I ! 

Figure 11.7 Multiple reflecting surfaces. 

problems, a computer program, APERTURE, has been written and is described 

in the literature (Reference 11.4). This program allows one to define an arbitrary 

elliptical aperture in an X-Y plane. The orientation of the aperture and the 

orientation and strength of the external field are defined as input quantities and 

do not necessarily have to be oriented along the X and Y - -- reference - -- - axes. - The 

program allows one to tabulate the interior field strength at whatever points are 

desired, in either rectangular or polar coordinates. The program also allows one 

to determine the total magnetic flux passing any arbitrary four-sided plane 
Figure 11.8 Determination of the voltage produced by aperturecoupling. 

surface. The program calculates the field strength normal to that surface and 

performs the graphical integration suggested in Figure 11.8 and necessary to 

determine the total magnetic flux passing through the plane of the loop, 

If the aperture under consideration is not elliptical, a corresponding 

equivalent elliptical aperture can generally be specified, as shown in Figure 11.9. 

The equivalent elliptical aperture would have the same area and the same 

eccentricity as those of the aperture under study. If the aperture were of an 

irregular shape, its area might have to be determined numerically. 

An example of an aperture-coupled field is shown in Figure 11.10 in 

which the elliptical fuselage described earlier has been approximated by two 

sheets of infinite size, one containing the aperture and another serving as the 

reflecting surface. The figure shows only the top half of the field pattern, the 

339 
KCL IMIJGULAR 

APERTURE 

IRREGULAR 

W APERTURE OF 

AREA A 

\ 

' EQUIVALENT ELLIPTICAL 

APERTURE 

--- 

Figure 11.9 Equivalent apertures. 

(a) Rectangular aperture 

(b) Irregular aperture 

field in the bottom half being symmetrical. The aperture considered was 0.2 m 

high by 0.1 m wide, the long axis of the ellipse being oriented at right angles to 

the plane of the figure. The field strength at the surface of the aperture, 

340 
assuming the aperture not to be there, was 167 Alm, the field strength that 

would be produced by the aforementioned current of 1000 A flowing axially 

along the elliptical fuselage. The field strength is seen to decay very rapidly with 

distance away from the aperture. In this particular case, the reflecting. surface 

was sufficiently far from the aperture under study that the field through most of 

the interior volume would have been virtually the same had the reflecting surface 

been disregarded. Doubling the size of the aperture leads to the field pattern 

shown in Figure 1 1.1 1. Roughly in accordance with simplest theory, doubling 

the dimensions of the aperture leads to an 8-fold increase in the field strength at 

points removed from the aperture. Looked at in another light, if one considers 

the volume in which the magnetic field is of a given strength or greater, doubling 

the lineal dimensions of the aperture will double the volume through which the 

given field exists. 

Figure 11.10 Fields coupled through an aperture. 

Major axis oriented vertically 

(0.2 x 0. l rn aperture) 
Figure 11.1 1 Fields coupled through an aperture. 

Major axis oriented vertically 

(0.4 x 0.2 m aperture) 

1 1.2 Electric Fields 

Unlike magnetic fields, which may penetrate into an interior region either 

through apertures or by diffusion through continuous walls, electrical fields, as a 

practical matter, can couple only through apertures. Any metallic surface 

provides sufficient shielding against electric fields that considerations of direct 

penetrations of the field through a metal surface can be neglected. Electric fields 

of practical consequence can exist only perpendicular to a metallic surface, since 

in a manner analogous to penetrating magnetic fields, electric fields of oblique 

incidence have all but the perpendicular component of the field shorted out by 

the low impedance of the metallic surface. The appropriate geometry for electric 

field coupling is then as shown on Figure 11.12. The electric fields of the 

interior region may be calculated as though they were produced by an electric 

dipole of strength (Reference 1 1.5): 
The quantity a33 has previously been presented in Figure 11.3. Subject to 

the provisions that the dimensions of the aperture be small compared to the 

wavelengths of the highest frequency component in the external electric field 

and that electric fields produced by changing magnetic fields can be neglected, 

the electric field in the interior region becomes 

Figztre 11.12 Geometry appropriate tc electric field coupling. 

343 
These expressions are based upon the further assumption of a pure dipole and, as 

with Equations 11.6 through 11 .I 1 for the magnetic field formulation, would 

indicate infinite field strengths as one approached the plane of the aperture, 

whereas in fact the field in the plane of the aperture would be one-half the field 

strength that would exist at that point if the aperture were absent. A power 

series expansion of the field from a dipole of finite size could be prepared but, as 

of this writing, has not been. The bookkeeping routines necessary to handle 

electric fields are virtually identical with those used for magnetic fields and 

could be incorporated into the APERTURE program, although, as of this 

writing, they have not been so included. 

11.3 Magnetic Field Zones 

From the foregoing it can be seen that the task of mapping the magnetic 

field inside a vehicle and actually calculating the voltage on any particular circuit 

could be a formidable task. A possible solution to the problem lies in recognizing 

that aircraft, or at least aircraft within a particular category, tend to possess 

characteristic zones. On a fighter aircraft, for example, the cockpit can be 

regarded as a magnetically open zone exposed primarily to aperture-coupled 

magnetic fields. Within reasonable limits, all fighter aircraft probably have 

approximately the same magnetic field in the cockpit. Another type of 

equipment zone characteristic of fighters would be equipment bays located in 

the forward section and behind the radome. All such equipment bays tend to be 

alike, the differences, perhaps, relating mostly to the type of fasteners used to 

hold the covers in place. The structure within a wing is a type of magnetic field 

structure fundamentally different from either the cockpit or the forward 

equipment bays. Accordingly, it would seem possible to divide an aircraft into a 

relatively small number of typical zones, to assign a ruling or charactxistic 

magnetic field intensity to those zones, and to provide rather simplified tables of 

nomograms listing the characteristic transient likely to be induced in wiring of a 

given length. 

This concept of dividing an aircraft into shielding zones was first used on 

the Space Shuttle (Reference 11 -6). The zones so defined are shown in Figure 

11.13 (Reference 11.7). The electromagnetic fields assigned to each of these 

zones were initially determined by "seat of the pants" engineering by a group of 

engineers knowledgeable in the field. These magnetic fields were then refined 
during the course of an extensive analytical investigation. The magnetic field 

amplitudes assigned to each of these zones, based on the analytical study, are 

given in Table 11.1 (Reference 11.8). The fields Were divided into two 

components, an A-component referring to fields coupled through apertures and 

a B-component referring to fields coupled by diffusion through metal surfaces. 

The A-component of the field would tend to have the same rapidly changing 

waveshape as the external magnetic field, while the B-component would have a 

much slower waveshape. In the Space Shuttle study the waveforms of the 

different components were taken to be as shown on Figure 11.14 (Reference 

11.9). In each case the field intensity was based on a worst case 200 kA lightning 

A THROUGH D - MOST PROBABLE LlGHTNlNG ARC ENTRY AND EXIT POINTS 

SHUTTLE ZONES 

1 . . FI - IT.UT . - . . nFCK - - - . . A - PAY1 . . . . - CIAn - . . - RAV - . . . 2 CENTER EQUIPMENT BAY 7 AFT EQUIPMENT BAY 

3 LOWER EQUIPMENT BAY 8 ENGINE COMPARTMENT 4 FORWARD TOP 9 VERTICAL STABILIZER 

5 FORWARD CENTER 10 WING 

Figure 11.13 Shielded zones within the Orbiter structure. 

current passing through the Orbiter vehicle. The field amplitudes of Table 11.1 

were the maximum amplitudes calculated for any of the possible lightning 

current entry or exit points. While no particular claim for accuracy can be made 

about any individual point within the Orbiter, the field amplitudes at least seem 

reasonable. 

The concept of dividing an aircraft structure into different magnetic field 

zones and assigning a ruling magnetic field strength to each zone, while 

imperfectly formulated at present, is fundamentally no different from the civil 

engineering practice of designing a structure to withstand a standard (generally 

worst case) wind loading. While the wind loading may differ widely at different 

points on the structure, the task of calculating the wind loading on each and 

every structural member would probably be sufficiently expensive that it would 

offset the savings that one might realize by tailoring each structural member to 

its own specific wind loading. 
Table 11.1 MAGNETIC FIELDS IN DIFFERENT ZONES 

OF THE SPACE SHUTTLE 

Aperture Fields Diffusion Fields 

Zone A-component (Aim) Bcomponent (Alm) 

*Payload 

**No payload 

The task of dividing aircraft into typical magnetic field zones and of 

calculating and assigning the appropriate field strength to each of the zones is 

probably of sufficient complexity that it should be done by specialists in the 

field. 

Analytical studies of field intensity could well be supplemented by 

experimental studies in which currents were circulated through an aircraft and 

the magnetic fields inside measured. Aircraft used for such studies should be in 

realistic condition in regard to the mechanical soundness of the structures, 

particularly relating to the access panels, but otherwise they need not contain 

complete electronic systems. They should not, however, have been cannibalized 

unduly, leaving an excessive number of empty bays where equipment niight 

normally have been expected to help provide some electromagnetic shielding 

capability. The use of older aircraft for such studies might well be realistic in 

that the electromagnetic shielding capability of a structure is in large measure 

determined by the electrical characteristics of the various access panels and 

fasteners. Because these tend to degrade with age, an older aircraft might be a 

more realistic test vehicle than a new and carefully maintained aircraft. 
TABLE 1 1.1 FOR AMPLITUDE 

A - - SEE TABLE 11.1 FOR AMPLITUDE 

Figure 11.14 Waveforms of aperture- and diffusion-coupled magnetic fields. 

(a) Lightning current 

(b) Aperture-coupled field, A-component 

(c) Diffusion-coupled field, B-component 
REFERENCES 

1 I. 1 C. D. Taylor, "Electromagnetic Pulse Penetration Through Small Apertures," ZEEE Transactions on Electromagnetic Capability, EMC-15, 1, 

Institute of Electronic and Electrical Engineers, New York, New York 

(February 1973): 17-26. 

11.2 C. D. Taylor, "Electromagnetic Pulse Penetration Through Small Apertures," Interaction Note 74, Electromagnetic Pulse Interaction Notes, 5, 

Air Force Weapons Laboratory, Kirtland Air Force Base, Albuquerque, 

New Mexico (March 1973). 

11.3 K. J. Maxwell, F. A. Fisher, J. A. Plumer, and P. R. Rogers, Computer 

Programs for Prediction of Lightning Induced Voitages in Aircraft 

Electrical Circuits, Air Force Flight Dynamics Laboratory, Air Force 

Systems Command, Wright-Patterson Air Force Base, Ohio (February 

1975). 

1 1.4 K. J. Maxwell et a/, Computer Programs, Section 3. 

11.5 C. D. Taylor, "Electromagnetic Pulse Penetration." 

1 1.6 Space Shuttle Lightning Protection Criteria Document, JSC-07636, 

National Aeronautics and Space Administration, Lyndon B. Johnson 

Space Center, Houston, Texas (September 1 1, 1973). 

11.7 Space Shuttle Program Lightning Protection Criteria Document 

JSC-07636, Revision A, National Aeronautics and Space Administration, 

Lyndon B. Johnson Space Center, Houston, Texas (November 4, 1975), p. 

F-7. 

11.8 Space Shuttle Program Lightning Protection Criteria Document, Revision 

A, p. F-6. 

1 1.9 Space Shuttle Program Lightning Protection Criteria Document, ~evision 

A, p. F-8. 
CHAPTER 12 

VOLTAGES AND CURRENTS INDUCED ON WIRING 

12.1 Elementary Considerations 

The fields by themselves are of little concern; of primary concern are the 

voltages and currents produced by the changing magnetic fields. In principle, 

these may be calculated from the geometry of the aircraft wiring and knowledge 

of the strength and orientation of the internal magnetic and electric fields. While 

calculation of the voltages and currents on actual wiring in aircraft may never be 

practical, because of the mechanical complexity of the wiring harnesses, 

calculations may be performed for simplified geometries. Such calculations will 

illustrate the scope of the problem and indicate the direction of practices that 

minimize the voltages and currents, and hence minimize the risk of circuit 

damage or upset. 

The simplest geometry to consider is that of a conductor, or group of 

conductors, placed adjacent to a metal surface and exposed to a uniform 

magnetic field oriented to produce maximum voltage in the wiring. The 

geometry is shown in Figure 12.l(a). 

If only common-mode voltages are considered and only cable systems 

short enough that transmission line effects need not be considered, the induced 

voltage will be 

where A = area of the loop involved in meters squared 

po = 4n x in henries per meter (permeability of free space) 

@ = total flux linked in webers 

H = magnetic field intensity in amperes per meter 

t = in seconds 

e = in volts 

Expressed in inch units: * 

where II = length of cable bundle in inches 

h = height above ground plant. in inches 

H = magnetic field intensity in amperes per meter 

t = in seconds 
(~1 0 time 

Figure 12. I Response to a changing magnetic field : open circuit voltage. 

(a) Gra~netry 

(b) M~gnctic field waveshape 

(c) Voltage waveshape 

It luust be emphasized that the voltage so calculated is that existing 

between the entire group of conductors (comprising the cable) and the vehicle 

structure. The voltage will divide between the loads at the ends of the cable 

inversely as the impedance of the loads. For worst case analysis, consider one 

end of the cable grounded with the other end open-circuited. All the voltage so 

developed will appear at the open circuit end of the cable. 

350 
Line-to-line or circuit voltages will be less, generally by a factor of 10 to 

200, or 20 to 46 dB, down from the common-mode voltages because individual 

conductors are usually close together and are often twisted, thus reducing the 

total loop area. 

The maximum cable current is that which flows when both ends of the 

cable are connected to the vehicle structure through a low or zero impedance. 

Such an impedance may be an overall shield grounded at each end, or it may be 

a group of semiconductor circuits, each having low input impedance. In the first 

case, the current will flow on the overall shield with the current on the input 

circuits determined by the shielding properties of the shield. In the second case, 

the current will flow directly through the input semiconductors and their bias 

sources 

The short circuit current that flows (Figure 12.2) may be determined from 

the familiar expression 

whence 

where I = amperes 

L = self-inductance of cable in henries 

e = open circuit induced voltage in volts 

t = time in seconds 

Cable inductance may be estimated from the expression 

4h L = 2 x 1 o-~ loge-- H/m (12.5) d 

where h = height above a ground plane 

d = conductor diameter 

The induced voltage, e, which drives the current, is proportional to the cable 

height, but the cable inductance, which impedes the flow of current, is 

proportional to the logarithm of the cable height. 

12.2 Transmission Line Effects 

12.2.1 Response to Magnetic Fields 

Conductors always have associated with them some distributed capaci- 
(4 0 time 

(dl 0 time 

Figure 12.2 Response to a changing magnetic field: short circuit current. 

(a) Geometry 

(b) Equivalent circuit 

(c) Magnetic field waveshape 

(d) Current waveshape 
tance and inductance, the values of which are determined by the size of the 

conductors and the distance of the conductors from adjacent ground planes and 

other conductors. When these are considered, the effect of a changing magnetic 

field is to produce an oscillatory open circuit voltage. As shown in Figure 12.3, 

these oscillations will be superimposed on an underlying voltage proportional to 

the rate of change of the magnetic field. When the internal magnetic field is of 

complex waveshape (as is the usual case) and not the idealized ramp function 

shown in Figures 12.1 and 12.3, the resulting open circuit voltage may be of a 

very complex nature. A few examples will be shown in Chapter 14. While the 

maximum voltage may be difficult to predict, given the complex nature of the 

superimposed oscillations, the amplitude of the envelope can at least be 

approximated from Equations 12.1 and 12.2. The frequency of the superimposed oscillations tends to be inversely proportional to the conductor length. 

Conductors, such as shields, grounded at one end tend to ring as quarter-wave 

dipoles: for example, a conductor 10 m long tends to ring at 7.5 MHz. Even this 

simple relationship is difficult to apply, since one conductor is seldom free of 

the influence of adjacent conductors. 

Aircraft wiring is grouped into bundles, the bundles usually containing 

both short and long conductors. The assembly, even if exposed to a magnetic 

field of simple waveshape will oscillate in a complex manner. Generally there 

will be one dominant frequency with several other frequencies, usually higher, 

superimposed. Each will have its own characteristic decrement. Almost the only 

reliable generalization is that the cables associated with large aircraft will be 

longer than the cables associated with small aircraft and will characteristically 

oscillate at lower frequencies. On fighter aircraft, measurements of induced 

voltages have shown the characteristic frequencies to be in the range 1 to 10 

MHz. 

Currents measured on bundles of conductors tend to be oscillatory, like 

the voltages, as long as the conductors are part of a wiring group employing a 

single-point ground concept. If the conductors are part of a wiring group 

employing a multiple-point ground concept, the conductor currents tend not to 

be oscillatory but to follow the underlying shape of the internal magnetic field. 

12.2.2 Response to Electric Fields 

When the aircraft is struck by lightning, it is forced to assume the electric 

field potential of the lightning flash at that point, perhaps 100 MV with respect 

to ground. The magnitude of voltage with respect to ground is of no real 

importance, but the electric field intensity at the surface of the aircraft is of 

importance. This electric field intensity depends upon the total amount of 

charge stored on the aircraft surface and the size of the aircraft surface, and 

could, in principle, be extremely high. In practice, the electric field intensity at 

the surface cannot exceed more than about 500 kVIm before intense electrical 

discharges take place from the aircraft surface into the surrounding air. The 

electric field strength will be greatest about points of small radius of curvature. 

The discharges from those points will tend to reduce the electric field strength at 
(b) 

time 

(a) Geometry 

(b) Magnetic field waveshape 

(c) Voltage waveshape 

354 
other points on the aircraft. As a result, the maximum field stress to which most 

points of the aircraft can be raised is probably on the order of 100 kV/m. This 

electric field will be radial with respect to the metal surface of the aircraft. In 

places where there are apertures, the electric field lines will pass through the 

apertures and will produce electric fields on the inner surface of the aircraft. 

The response of a conductor to a suddenly applied electric field is shown 

in Figure 12.4. If a conductor is a height, h, above a metal surface upon which 

an electric field is incident, the conductor will assume the potential of the field 

at that point: 

This holds true only if the conductor is isolated from ground. If the conductor is 

completely grounded, as shown in Figure 12.4(b), the voltage on the conductor 

will be zero. If we were dealing only with static electric fields, any ground on the 

conductor would reduce the conductor voltage to zero. However, here we are 

dealing with conductors having distributed inauctance and capacitance, and 

exposed to a suddenly applied electric field. Accordingly, the electric field 

applied at point 1 will raise the conductor to the voltage e = hEi, and the voltage 

will maintain itself at that level until such time as a voltage wave has had time to 

propagate from point 1 to the ground conductor at the left and back to point 1. 

Thus, at point 1 the conductor experiences a square-wave voltage lasting for 

twice the travel time required for the wave to travel from point 1 to the 

grounded end and back. The electric field applied at point 2 likewise raises the 

conductor at point 2 to e = hEi, and it again remains there until a wave has had 

time to propagate to the grounded end and back. Since less of the conductor is 

involved, this voltage lasts for a shorter period of time. Nearing the limit, the 

field applied at point 4 can raise the conductor to a voltage of e = hEi for only a 

very short period of time. 

As the voltage waves that are produced return to the open end of the 

conductor, they again experience a reflection. The voltage waves reflect back 

and forth between the open and grounded ends, reversing their polarity at the 

grounded end and doubling in amplitude at the open end. Thus, the voltage at 

the open end is oscillatory (Figure 12.5), the amplitude of the oscillation 

decreasing with time as energy is lost in the resistance of the conductor or 

radiated into surrounding objects. 

Associated with these traveling waves of voltage are traveling current 

waves. The degree to which the waves are excited depends not only upon the 

magnitude of the incident electrical field but upon its rate of change. Fields with 

longer rise times are proportionately less effective in exciting oscillations. If the 

conductor of Figure 12.4 is viewed as the shield on a group of wires, the current 

and voltage on the shield will be oscillatory. The commonly accepted practice of 

grounding shields at only one end is based on the premise that the electric fields 

to which the shielded conductors will be exposed are of such low frequency that 

all points of the shield may be held at ground potential with only the one 

ground point. When the rapidly changing electric field associated with lightning 
time (4 

Figure 12.4 Response to a suddenly applied electric field. 

(a) Ungrounded conductor 

(b) Well-grounded conductor 

(c) Conductor grounded at one end 

(d) Waveshape at indicated points 

356 
Figure 12.5 Oscillatory response excited by a suddenly applied electric field. 

is present, frequently this condition is not met. 

12.3 Aids for Calculation of Voltage and Current 

In Section 1 1.3 there was introduced the concept of magnetic field zones, 

according to which in any particular area of the aircraft the total magnetic field 

may be viewed as the direct superposition of two components, one the result of 

coupling through apertures and the other the result of diffusion-type coupling. If 

the assumption of a ruling field is made, there may be derived some simplified 

relationships between the magnetic field and the resulting induced voltages and 

currents. Additional assumptions are the following: 

The conductor is of length Q, diameter d, and spaced a height, h, 

above a ground plane. 

The magnetic field is oriented to produce maximum voltages in the 

conductor. 

One end of the conductor is grounded. 

The magnetic fields are of the shape shown in Figure 11.14. 

These are the same assumptions as those illustrated in Figures 12.1 and 12.2, and 

upon which Equations 12.1 and 12.2 were based. Under these assumptions 

where KI = 0.63 if Q and h are in meters 

K1 = 0.63 x lo4 if Q and h are in centimeters 

K1 = 0.41 x lo3 if Q and h are in inches 

357 
In all cases H is expressed in amperes per meter. The waveshape of the open 

circuit voltages would typically be proportional to the derivative of the H field, 

and hence oscillatory. 

If we assume the conductor to be grounded at each end 

where K2 is given by either Figure 12.6 or 12.7, according to the units used for 

measurement of the conductor. Conductor length does not influence short 

circuit current. The waveshape of the short circuit current tends to be the same 

as that of the incident magnetic field. 

The height, h, of the cable bundle above a ground plane is difficult to 

specify because the ground plane is seldom purely a plane surface and because 

cable bundles are frequently strapped directly to a supporting structure. For 

purposes of this analysis, assume 

That the height, h, is measured to the nearest substantial metallic 

structural member 

That if the cable bundle is laid directly on that member, h equals 

one-half the cable diameter 

That if the cable bundle is elevated above the metallic structural 

member, h equals the clear height above the member plus one-half 

the cable diameter. (If the cable height differs along its length, use 

an average height.) 

h/d 

Figure 12.6 K2 - metric units. 
12.4 Voltage and Currents in the Space Shuttle 

Based upon Equations 12.1 to 12.8, Figure 1 1.14, and Table 1 1 .l, one 

may calculate the voltages and currents that would be developed on typical 

wiring. The results are shown on Table 12.1 (Reference 12.1). 

Figure 12.7 K2 - inch units. 

Table 12.1 OPEN CIRCUIT VOLTAGE AND SHORT CIRCUIT CABLE 

CURRENT IN THE VARIOUS ZONES OF THE SPACE SHUTTLE 

h = l in. (0.0259 m) h = 2 in. (0.0508 m) h = 5 in. (0.1270 m) h = I0 in. (0.254 m) 

Zone Voltage Curent Voltage Current Voltage Current Voltage Current 

(volts) (amperes) (volts) (ampms) (volts) (amperes) (volts) (amperes) 

'All values based on cable length of 157.48 inches (4 rn) and diameter of 1 inch (0.0254 m). For other lengths, 

scale the voltage proportionately. 
REFERENCE 

12.1 Space Shuttle Program Lightning Protection Criteria Document, 

JSC-07636, Revision A, National Aeronautics and Space Administration, 

Lyndon B. Johnson Space Center, Houston, Texas (November 4, 1975) p. 

F-6. 
CHAPTER 13 

EFFECTS OF SHIELDS ON CABLES 

13.1 Introduction 

If electronic equipment is to be operated in a region where there are 

changing electromagnetic fields and if experience or analysis indicates that the 

currents and voltages induced by those fields may be harmful, the most 

straightforward approach toward achieving transient compatibility is to shield 

interconnecting wiring from those fields. In aircraft, shielded conductors are 

most commonly used to control low-level and low-frequency interference 

signals, For many good and legitimate reasons shields on cables are often 

grounded at only one end. If the wiring is subjected to the rapidly changing 

electromagnetic fields associated with lightning, however, shields grounded at 

only one end may not be effective, and in some cases can make the surge 

voltages even larger than they would be if the conductors were not shielded in 

the first place. In this section of the report we shall discuss the ways in which 

shielded cables reduce surge voltages, some of the ways in which common 

practice regarding the use of shields may be at variance with the use required for 

the control of lightning effects, and some of the ways in which this apparent 

conflict of use may be resolved. We shall also discuss how the noise currents 

flowing on shields of cables may be related to the noise signals coupled onto the 

signal conductors. 

13.2 Some Elementary Illustrations 

To illustrate some of these effects let us consider a series of tests 

(Reference 13.1) that were made on a 5m long length of RG-58/U coaxial cable. 

The cable was placed adjacent to a metal ground plane and a magnetic field 

passed between the cable and the ground plane. The magnetic field was not a 

distributed field but was confined to the core of a pulse injection transformer in 

the manner shown in Chapter 17. Measurements were made of the voltage 

between the center conductor and ground at each end of the cable. These 

voltages thus represent the common-mode voltages that would exist. The first set 

of results is shown in Figure 13.1. Tn an unshielded conductor or one in which 

the shield on a conductor is not used, equal and opposite voltages appear at the 

two ends if all of the loading impedances are balanced. The voltages at the two 

ends of the conductor are of equal amplitude but opposite polarity, as would be 

produced if the conductor were considered to have an equivalent voltage 

generator at its center, as shown in Figure 13.l(c). If the load impedances at the 

ends are unbalanced by the addition of a 50 fl resistor at one end (Figure 13.2), 

the total voltage induced around the loop remains unchanged, but most of it 

appears at the end with the highest impedance. With reference to the equivalent 

circuit of Figure 13.2(c), the fact that there is any voltage at V2 implies the 

existence of some capacitive loading as well as the desired resistance load; 

otherwise there would be no voltage across V2. 
INCIDENT MAGNETIC FIELD 

END 1 

(4 

Figure 13.1 Shield not grounded at either end. 

(a) Test conditions 

(b) V1 

(c) v2 

(d) Equivalent circuit 

Grounding the shield at one end, shown in Figure 13.3, does not 

significantly affect the common-mode voltage at the other end. The changing 

field induces a voltage between the open end of the shield and ground. The 

conductor is exposed to the same field, and there is thus the same voltage 

developed between the ends of the conductors as that between the ends of the 

shield. Because of the unbalanced load impedances, all this voltage must appear 

between the conductor and ground at end 1. Leaving aside considerations of 

unequal load impedance, the shield can reduce voltage at one end only by 

increasing it at the other end. This effect is shown in Figure 13.4. The reason the 

shield reduces the line-to-ground voltage at V2 is that the capacitance between 
shield and conductor at the grounded end loads the signal conductor, just as did 

the 50 R resistor on Figure 13.3. 

(4 ,,~,~/~-=,/,//, 4-, I 

I 

/ 

Figure 13.2 Unequal load impedances. 

(a) Test conditions 

(b) Voltages 

(c) Equivalent circuits 

Grounding the shield at both ends, as shown in Figure 13.5, produces an 

entirely different response. If the shield is grounded at both ends, the voltage 

induced by the changing magnetic field in the shield produces a flow of current 

through the shield. 

363 
I SHIELD - + -2-- ,-,-- 

I CONDUCTOR 

+ 

"1 

Figure 13.3 Shield grounded at one end. 

(a) Test conditions 

(b) Voltages 

(c) Equivalent circuit 

where E' = voltage induced between ends of the shield 

364 
I = current on the shield 

L = self-inductance of the shield 

Figure 13.4 Shield grounded at one end. 

(a) Test conditions 

(b) Voltages 

(c) Equivalent circuits 

This shield current reduces the voltage induced between the signal conductor 

and ground. The reduction in voltage can be viewed equally well from two 

different viewpoints. 

The first is that the shield current produces a magnetic field that tends to 

cancel the incident field. From this viewpoint the voltages on the signal 

conductors to ground respond only to the difference between the incident and 
the canceling fields. Alternatively, the reduction in voltage can be viewed as the 

effect of the mutual inductance between the shield and the signal conductor. 

This latter approach is illustrated by the equivalent circuit of Figure 13.6(c). The 

voltage between the ends of the conductor is the sum of that resulting from the 

voltage induced in the signal conductor, E, and that coupled through the mutual 

inductance between the shield and the signal conductor. Since the shield is 

grounded at both ends and able to carry a current, the voltage appearing across 

2 - SHIELD 

I- I 

Figure 13.5 Shield grounded at both ends. 

(a) Test conditions 

(b) Shield current 

(c) Equivalent circuit 

(4 

- + ~i\ 

CONDUCTOR 
1 (NEARLY) (NEARLY) 1 

?///////////////////////////////////////////////////////// 

Figure 13.6 Shield grounded at both ends. 

(a) Test conditions 

(b) Conductor voltages 

(c) Equivalent circuit 

the primary inductance of the equivalent transformer is equal to the voltage E' 

induced between the ends of the shield by the incident magnetic field. Since the 

mutual coupling between the shield and the signal conductor is very nearly 

unity, the voltage induced in the secondary, or signal conductor, side of the 

equivalent transformer is about equal to the voltage originally induced in the 

shield. Accordingly, the voltages appearing at the ends of the conductor are 

much lower than they would be if the shield could not carry current. 
An alternative form of the equivalent circuit of Figure 13.6 is shown in 

Figure 13.7. There are two parts of the circuit, one relating the magnetic field to 

the current flowing on the shield and one relating the current on the shield to 

the voltage developed on the conductor. Both the measurements and the circuit 

indicate that voltage on the center conductor is affected by the relative 

impedance of the loads on the ends. A low impedance at one end pulls the 

voltage at that end down but raises the voltage at the other end. M12 and R 

represent transfer quantities, the values of which depend upon the type of 

shield. 

Figure 13.7 Shield grounded at both ends. 

(a) Test conditions 

(b) Conductor voltages 

(c) Equivalent circuit 
13.3 Transfer Quantities 

If the mutual inductance between the shield and the signal conductor were 

equal to the self-inductance--unity coupling--the flow of current on the 

shield of the cable would not cause any voltage to be developed between the 

shield and the signal conductor. This section of the report will now discuss the 

factors that prevent the coupling between shield and signal conductor from 

being perfect. Consider first a shield in which only resistance effects are treated, 

shown in Figure 13.8. In this condition an external noise current flowing 

through the resistance of the cable shield produces an electric field on the 

internal surface of the shield. The nature of the coupling between this internal 

field and the signal conductor depends upon the connections at the ends of the 

cable and upon the distributed inductance and capacitance of the circuit internal 

to the cable. As a simple case, imagine the left-hand end of the signal conductor 

to be shorted to the cable shield. The internal electric field would then be 

completely coupled to the signal conductor, and the voltage between the signal 

conductor and the shield at the right-hand end of the cable would be equal in 

L CONDUCTOR 

- ------- A--- 

------ ------ 

INTERNAL ELECTRIC FIELD 

INTERNAL ELECTRIC FIELD /(------- 

NOISE CURRENT 

0 0 

time 

Figure 13.8 Coupling resulting from resistance effects. 

(a) Current and field polarities 

(b) Current and field waveshapes 
magnitude to the total internal electric field along the length of the shield. This 

simple type of coupling is valid for low-frequency or slowly changing external 

noise currents. For noise currents sufficiently slow that this model applies, the 

waveshape of the internal electric field is the same as the external noise current, 

being related by Ohm's Law: 

E =IR (1 3.3) 

13.3.1 Physical Phenomena 

Generally the coupling is not as simple as this. If the overall shield is a 

solid-walled cylinder, as shown in Figure 13-9, the internal electric field depends 

upon the product of the resistivity of the shield material and the density of the 

current on the internal surface of the shield. As discussed in Section 10.1, this 

internal current density, Ji, will not, in general, be the same as the density of the 

current on the external surface of the cylinder. Because of the phenomenon of 

skin effect, the current density on the inner surface of the shield will rise more 

slowly than does the external noise current. The rate at which the current 

density on the inner surface increases will be directly proportional to the 

permeability of the field material and to the square of the wall thickness and 

inversely proportional to the resistivity of the wall material. Cables with 

solid-wall shields and cable trays of solid metal with tightly fitting covers will 

typically exhibit this type of behavior. 

A different type of behavior is commonly observed on cables with braided 

shields, shown in Figure 13.10. The shields do not provide a perfect conducting 

cylinder, for they have a number of small holes which permit leakage. The holes 

between the individual bundles of wire forming the shield can be approximated 

as a group of diamonds, the size and orientation of which depend upon the 

weave angle of the braided shield. The long axis of the triangles may be oriented 

along the shield, as at end A of the cable shown in Figure 13.10(a), or, if the 

weave angle is great, the long axis of the hole may be oriented circumferentially 

to the cable, as at end B of the cable. The external noise current produces a 

circumferential magnetic field around the outside of the shield. As a first 

approximation, one may visualize the external field lines looping in and out of 

the holes in the braided shield. There will be more leakage in holes with long 

axes oriented circumferentially to the end of the shield (end B of the cable, large 

weave angle) than there will be if the long axes of the holes are oriented along 

the shield, (end A of the cable, small weave angle). Viewed from the end of the 

cable, as in Figure 13.10(b), the field that leaks into the cables produces a net 

magnetic field circumferentially around the inside surface of the shield. This 

magnetic field produces an internal electric field of an amplitude dependent 

upon the degree of leakage into the shield and upon the rate of change of the 

magnetic field. The rate of change of this internal magnetic fieId is to a first 

approximation proportional to the rate of change of the external noise current. 

The total internal electric field, then, is proporti.onal to the resistance of the 

cable shield, to the rate of change of external noise current, and to the number 

of holes in the cable. 
- ------- ---- 

------ ------ 

--- 

-EXTERNAL CURRENT DENSIN - EXTERNAL ELECTRIC FIELD 

Je Ee 

EXTERNAL NOISE CURRENT / 

I 

r 

/' /\ INTERNAL ELECTRIC FIELD Ei 

/ 

EXTERNAL NOISE CURRENT 

(b) 

0 time - 

Figure 13.9 Coupling via diffusion through a solid wall. 

(a) Current and field polarities 

(b) Current and field waveshapes 

Frequently braided cables exhibit a phenomenon, shown in Figure 13.1 1, 

in which the net polarity of the internal magnetic field is opposite to the 

polarity of the external magnetic field. The result of this phenomenon is that a 

change in external noise current induces an internal electric field of the opposite 

polarity. The reasons for this behavior are not clear to this author, but the 

phenomenon seems to depend upon the thickness of the braided covering of the 

shield and the total diameter of the shield, as well as upon the degree of 

tightness of weave of the shield. ' 

Some illustrative examples of the behavior of braided shields are shown in 

Figures 13.12 and 13,13, The figures show the test geometry and the 

conductor-to-ground voltage produced by currents having waveforms of several 

different times to crest. The 47 R resistors shown provide the termination for 

the signal conductor and prevent the voltage at the other end from being 

obscured by the internal oscillations of the cable. Figure 13.12 shows a 

double-shielded cable with the voltage measured from the inner shield to ground. 

This particular cable exhibits the phenomenon shown in Figure 13.8. In Figure 
EXTERNAL MAGNETIC FIELD LINES 

I 

(a) 

FIELD LINES LEAK 

IN THROUGH HOLE 

INTERNAL ELECTRIC FIELD 

1 -/ \ 

1 \ 

I L, ,, ,,,, jF- EXTERNAL NOISE CURRENT 

0 time - 

(c) 

Figure 13.10 Coupling via magnetic leakage through holes. 

(a) Current and field polarities 

(b) Field leakage through holes: end view 

(c) Current and field waveshapes 

13.1 3 a cable similar to RG-8/U cable is shown. This cable exhibits the negative 

coupling inductance shown in Figure 13.1 1. 

Another source of coupling is via capacitive leakage through the holes in 

the shield, as shown in Figure 13.1 4. If the shield is held at ground potential by 

deliberate grounds, if the signal conductor is held near ground potential by 

external loads, and if the shielded cable is subjected to an external and changing 

electric field, dielectric flux will pass through the holes in the cable from the 

external source and onto the signal conductor. The flow of these dielectric, or 

displacement, currents through the external load impedances produces a voltage 

between the signal conductor and the shield. An alternate source of coupling, 

shown in Figure 13.14(b), involves the voltage on the shield itself. Noise currents 

flowing through the external impedance of the shield may, if the shield is not 

perfectly grounded, produce a voltage between the shield and any external 
--- - -- -- 

-- ------ 

-INTERNAL E FIELD RESULTING FROMRESISTANCE COUPLING 

EXTERNAL NOISE CURRENT I 

EXTERNAL FIELD LINES 

(0) WALL OF CABLE SHIELD 

INTERNAL FIELD LINES 

/----- 

INTERNAL ELECTRIC FIELD 

EXTERNAL NOISE CURRENT 

\ I 

(b) (4 \ ,I 

Figure 13.1 I Magnetic coupling giving effect of a negative coupling inductance. 

(a) Current and field polarities 

(b) Direction of field 

(c) Current and field waveshapes 

ground structure. External impedances between the signal conductor and the 

shield, as well as the inherent capacitance between the shield and the signal 

conductor, force the signal conductor to assume nearly the same potential as the 

shield. Because the signal conductor is then at a potential different from the 

surrounding ground, dielectric flux can pass from the signal conductor through 

the holes in the shield and to ground, the dielectric currents again giving rise to a 

voltage between the signal conductor and the shield. 

In most cases the effects of magnetic leakage are probably more important 

than the effects of capacitive leakage. 

1 3.3.2 Calculation of Transfer Quantities 

The factors that affect the voltage which may appear between a signal 

conductor and the shield of a cable, as distinct from the voltages that may 

appear between the signal conductor and ground, are, then, the resistance of the 

shield, the degree to which the shield allows magnetic fields to leak to the inside, 

and the degree to which the shield allows electric fields to leak to the inside. 

These latter two characteristics may be expressed in terms of an equivalent 
CURRENT VOLTAGE 

CABLE LENGTH = 

20 FEET 

dc RESISTANCE = 

0.024 R 

0.5 ps ldiv 0.5 ps/div 

Figure 13.12 Coupling from overall shield to center shield of a double-shielded 

cable. 

transfer inductance of the shield and an equivalent transfer capacitance. These 

effects may be expressed in terms of the equivalent circuits shown in Figure 

13.15. Figure 13.1 5(a) shows the equivalent circuit by which noise current on 

the shield of the cable may be related to the voltage induced between the signal 

conductor and the shield. In the frequency domain, this may be expressed as 

As will be seen momentarily, the term Zd may frequently be treated simply as 

the dc resistance of the cable shield, shown in Figure 13.1 S(b), in which case the 

relation between current and voltage becomes simply 
CABLE LENGTH = 

28 FEET 

dc RESISTANCE = 

0.0326 R 

CURRENT VOLTAGE 

No.40 I No. 41 

0.5 ps/div 0.5psldiv 

Figure 13.13 Coupling from overall shield to center conductor of a cable 

similar to RG8/U. 

Treatment of capacitive effects yields the equivalent circuit shown in 

Figure 13.15(c), in which the relation between loop current induced on the 

signal conductor and external noise voltage becomes 

Vance (Reference 13.2) gives an excellent dissertation on how the transfer 

inductance M12 and transfer capacitance C12 may be related to the characteristics of the braided shield. Many of the equations and figures that follow are 

reproduced from his work. 

A braided shield would, if rolled out on a plane, typically have the 
SOURCE OF A 

NOISE SOURCE 

Figure 13.14 Coupling via capacitive leakage through holes. 

(a) Coupling from an external source 

(b) Coupling from voltage on the shield 

appearance shown in Figure 13.1 6 (Reference 13.3). The characteristics of the 

shield may be described in terms of the following symbols: 

a - the radius of the shield 

t - the number of carriers, or bands of wire, that make up the shield 

N - the number of ends, or the number of wire strands, in each carrier 

P - the number of picks, or carrier crossings, per unit length 

d - the diameter of the individual wires 

Some important quantities that may be defined in terms of the above 

characteristics are 

4naP Weave angle: a =tan-'- (13.7) C 
I SIGNAL CONDUCTOR 

- SHIELD 

-In 

(0) 

1 

zd M12 

SIGNAL CONDUCTOR 

I vs 

SHIELD 

I/ SIGNAL CONDUCTOR 

SHIELD 

Figure 13.15 Equivalent circuits. 

(a) Transfer impedance 

(b) Transfer impedance simplified 

(c) Transfer admittance 

PNd Nd NdC Fill: F === 

sin a W 4na cos a 

Optical coverage: K = 2F - F~ (1 3.9) 

2 Volume of metal: v = n adF (13.10) 

The optical coverage, K, relates to the size of the small holes not covered 

by the carriers. The greater the optical coverage, the smaller the total area of the 

holes with respect to the total area covered by the shield and, other things being 

equal, the better the electrical performance of the shield. The weave angle 

affects the orientation of the diamond-shaped holes. Since holes oriented along 

the axis of the cable are less effective in coupling energy into the interior than 

are holes oriented circumferentially around the cable, it follows that, other 
Figure 13. I6 Properties of a typical braided shield. 

(a) Braid pattern developed on a plane 

(b) Properties of a typical braided shield 

things being equal, a shield with a small weave angle will give better electrical 

performance than a shield with a large weave angle. 

Presented here without proof, the elements of the equivalent circuit of 

Figure 13.15(a) are 

3 78 
Zd = 

4 (1 + j) dl6 

nd2 NCo cos a sinh (1 + j) d/6 

npo MI2 = - (I-K)~/~ e2 (a < 45") (1 3.12) 6C E(e) - (1 -e2) K (e) 

In the above expressions 

a = conductivity of the wires 

po= 4nx10-7 

K(e) and E(e) = complete elliptic integrals of the 

first and second kinds 

For many analyses Zd may be taken, with sufficient accuracy, as the dc 

resistance of the shield. This dc resistance is also slightly dependent on the 

coverage. 

4 - 1 %" 2 2 (13.17) nd2 NC~ cosa n ad o F cos a 

For the specific cable with construction as indicated, Figure 13.17 

(Reference 13.4) shows how the transfer impedance varies with frequency for 

different degrees of optical coverage. Several observations might be made about 

this figure. The first is that at low frequencies, below about 500 kHz, the 

transfer impedance is basically just the dc resistance of the cable shield. The 

second is that for frequencies above about 2 MHz the transfer impedance is 

primarily reactive, being represented by the transfer inductance of the cable 

shield. The third is that for the most common types of shields (those with 

optical coverage of less than 90%) the dip in transfer impedance below the dc 

resistance (the dip being a result of skin effects) is not great. It is for this reason 

that the equivalent circuit of Figure 13.15(b), in which the frequency-dependent 

properties of the shield resistance are ignored, is adequate for most engineering 

analyses. 
The transfer admittance of the braided-wire shield, governing the degree to 

which the shielded cable may be affected by external electric field, is 

Yt =jC12 (13.18) 

C1 is the capacitance per unit length between the internal conductors and the 

shield, C2 is the capacitance per unit length between the shield and its external 

current return path, and e is the permittivity of the insulation inside the cable. 

The preceding equations indicate that both the mutual inductance and the 

mutual capacitance depend on the weave angle of the shield. For a constant 

coverage and for a specific size cable, the variations of M and C with weave angle 

are as shown in Figure 13.18 (Reference 13.5). The mutual inductance per unit 

length increases rapidly with increasing weave angle, since, as explained earlier, 

apertures with their major axes oriented parallel to the lines of magnetic flux are 

FREOUENCY - Hz 

Figure 13.1 7 Transfer impedance of a braided-wire shield. 

100 - I 1 l Ill1 - 

- 

- 

- 

- 0 = 30° 

a= lcm 

T I 1 IlllC - 

d 

- 

- 

- 

K = 85 4% C = 42 - d = 63 m~lr 

N = 10 - K 89 5%, C = 46 

- 

- 

- 

- 

K = 97 8%, C = 58 

1 o5 lo7 
10 20 30 40 50 60 70 80 

WEAVE ANGLE, cr - degrees 

Figure 13.18 Variation of mutual inductance and mutual capacitance with 

weave angle for constant coverage. 

more effective in allowing flux to leak inside than are apertures of the same area 

but with their major axes oriented at right angles to the liues of flux. The mutual 

capacitance varies only slightly with weave angle, since the major factor 

governing the mutual capacitance is simply the projected area of the apertures. 

To simplify the task of estimating M and C from the preceding equations, 

the functions of the eccentricity of the apertures are plotted in Figure 13.19 

(Reference 13.6) and the variation in (I-K)~/~ plotted in Figure 13.20 

(Reference 13.7). 

Tape-wound shields, as shown in Figure 13.21, are often used where 

flexibility of the shielded cable is required. The shield may be formed either 

from a narrow metal sheet spiraled around the core or from a carrier of fine 

wires, again spiraled around a core. Flexible armor and flexible conduit, which is 

normally used primarily for mechanical protection, may also be analyzed as 

tape-wound shields. The tape-wound, or spiral-wound, shield is a rather poor 

shield for preventing coupling of current from the shield onto the internal 

conductor because the shield tends to behave as a solenoid wound about the 

internal conductors. There is thus a rather large mutual coupling term relating 

the internal voltage to the shield current. This is particularly true if the tape is 
WEAVE ANGLE, t2 - degrees 

Figure 13.19 Variation of the eccentricity functions for braided-wire shields. 

wound without any overlap because in this case the entire cable current is forced 

to spiral around the core. Vance analyzes a 1 cm radius cable wound with 

.25 mm thick, I cm wide copper tape, the tape applied without overlap. The 

dimensions of the cable are thus similar to those for the braided-wire shield 

shown in Figure 13.18. The mutual inductance of the tape-wound shield was 

calcuIated as 3.9 pH/m of the order lo4 greater than the mutual inductance 

obtained for a braided-wire shield. 

The thin-waned tubular shield consists of a metal tube of uniform cross 

section and uniform wall thickness. Coupling through the shield can occur only 

by diffusion of the electromagnetic fields through the walls of the tube. The 

transfer impedance of thin-walled tubes such as these, as derived by Schelkunoff, 

(Reference 13.8) is 

1 (l+j)t/6 za =- 

2nrut sinh (1 + j) t/6 

where r is the radius of the shield, a is its wall thickness, u is the conductivity of 

the shield, and 8 is the skin depth in the shield given by 

It is assumed that the wall thickness, a, is small compared to the radius, r, of the 

tube and that the radius, r, is small compared to the smallest wavelength of 
FILL. F 

0.3 - 0.4 0.5 0.6 0.7 0.9 

COVERAGE. K - percent 

0.8 0.9 1 .O FILL. F 

Figure 13.20 Variation of (1-~)3/2 with shield coverage K. 

interest. It is also assumed that the shield is made from a good conductor (metal) 

so that the displacement current in the shield material is negligible compared to 

the conduction current. 
Figure 13*21 Tape-wound shield. 

The maximum transfer impedance is found at dc or low frequencies, or 

frequencies low enough that the thickness of the shield is small compared to the 

skin depth. That impedance is the dc resistance 

The transfer impedance decreases with increasing frequency. Accordingly, 

the internal electric field will, in response to a step-function current, increase 

according to the pulse penetration time constant predicted by Equation 10.9 

and shown in Figure 10.6 of Section 10.1. 

In the absence of any end effects or openings through which magnetic 

fields can leak, the mutual inductance term M1 2 of Equation 13.4 will be zero. 

Cable trays are most often used for mechanical protection of wires, but if 

viewed as shields, they may also provide electrical protection. The characteristics 

of a cable tray that make for good electrical protection are the same as those of 

any other shield: 

It must be able to carry current along its axis. 

It should be of low-resistance material. 

It should completely surround the conductors, implying that the 

tray should be fitted with a conducting cover. 

It should have a minimum number of openings through which 

magnetic fields may leak, 

It should have as few joints as possible, and such pints should be 

made in such a manner as to provide minimum resistance and 

leakage of magnetic fields. 

The transfer characteristics of the tray by itself would be about the same 

as those of the solid tubular shields discussed previously. This comparison 

assumes the tray to be of solid metal and fitted with a well-sealed cover. Since 

covers never form perfect seals, the pulse penetration time constant would 

probably not be as long as predicted by Equation 10.9. Even with an imperfect 

cover, the transfer impedance ZT would not be greater than the dc resistance. 

Trays are most commonly built in short sections and joined by splices or 

transition sections. Such sections frequently provide for thermal expansion and 

contraction and are, at any rate, seldom designed either to provide good 

electrical continuity or to protect against magnetic leakage. When joints are 

considered, the transfer characteristic of the tray is frequently found to depend 

almost entirely on the treatment of the joints. The transfer characteristics can 

seldom be determined by calculation; they must be determined experimentally. 
13.4 Multiple Conductors and Shields 

If there are multiple conductors in the core, each conductor will develop 

nearly the same voltage between that conductor and the shield. This holds true 

whether the conductor is located adjacent to the shield or in the center of the 

bundle of conductors comprising the core. Accordingly, the voltage between any 

pair of conductors in the core will be small. Present analytical tools do not seen1 

of sufficient accuracy to predict with any assurance the magnitude of line-line 

voltages; they are best determined by actual measurement. Line-line voltages are 

much more strongly influenced by load impedances to which the conductors are 

connected than by the position of the conductors within an overall shield. 

13.4.1 Physical Phenomena 

One very common situation involving multiple conductors within an 

overall shield is a configuration in which a shielded conductor is itself contained 

in an overall shield. Figure 13.22 shows such a condition and shows some of the 

basic considerations with which one must deal in order to predict the response 

of this two-cable network to a noise current flowing on the outer shield. The 

total circuit response is a combination of the internal field of the outer shield 

and the internal field of the inner shield plus the coupling from the internal field 

of the inner shield to the signal conductor contained within the inner shield. 

These internal fields are created by noise current flowing on the outer shield plus 

noise current flowing on the inner shield. The noise current on the inner shield 

depends upon the internal field of the outer shield and on the impedance 

presented to signals traveling on the outer surface of the inner shield and 

returning along the inner surface of the outer shield. 

The most important factor concerning this latter impedance is whether or 

not the internal shield is grounded at both ends. Figure 13.32(b) shows the 

internal shield connected to the outer shield at the left-hand end and a switch 

that would allow connection to be made between the shields at the right-hand 

end. It is furthermore assumed that the impedance of any connection between 

shields is zero, a situation that definitely may not be true in practice. 

If the inner shield is not grounded at the right-hand end, the inteinal field 

of the outer shield (Ei-os) produces a voltage difference, Vis.os, at the right-hand 

end of the cable. Neglecting the distributed capacitances between the inner 

shield and the outer shield, the noise current on the inner shield, Iis, is then zero. 

If the inner shield is connected to the outer shield at both ends and the switch, 

S, is closed, the voltage difference between the two ends of the cable is reduced 

to zero. The internal field of the outer shield, however, will induce a circulating 

current, Iis, between the inner shield and the outer shield. The magnitude of this 

current is governed by the strength of the field Ei-,, and the impedance around 

the loop between the inner shield with return on the outer shield. This 

impedance, at least for many cables, is predominantly inductive. The waveshape 

of the current on the inner shield, Iis, is then proportional to the time integral of 

the voltage developed along the inner surface of the outer shield. The current on 
SIGNAL CONDUCTOR 

CURRENT ON INNER SHIELD 

Iis --------------- A 

INTERNAL FIELD OF INNER SHIELD 

------- - Ei-is 

INNER SHIELD 

INTERNAL FIELD 

- 

OF OVERALL SHIELD 

10s 

CURRENT ON OVERALL SHIELD 

SIGNAL CONDUCTOR 

INNER SHIELD is-0s 

tts 

OVERALL SHIELD xis v is-0s 

w b 

(b) 

Figure 13.22 Double-shielded cable. 

(a) Current and field polarities 

(b) Current on inner shield and forced 

through Lis-os 

the inner shield thus typically has a slower rise time and a longer duration than 

does the field along the inner surface of the outer shield. This current, Iis, on the 

internal shield then produces an electric field along the surface of the inner 

shield, Ei-is. 

13.4.2 Examples of Response of Multiple-Shielded Cables 

It will be apparent that there are a number of combinations of shield 

connections and termination impedances which are possible, all of them 

affecting the overall response of the cable system. Some of these effects have 

been demonstrated during an extension of the tests described in Figure 13.1 

through 13.6. The piece of RG-58/U coaxial cable used for those tests was 

modified by pulling over the outer jacket another length of flexible copper 

braid. The test connections and results are shown on Figure 13.23. The 

equivalent circuit has three separate components--separate in the sense that 

each, for all practical purposes, may be treated independently. The first part of 

the circuit determines the voltage induced along the inner surface of the outer 
shield. This is governed by the outer current flowing through R1 and MI, the 

transfer factors of the outer shield. In the second stage of the problem, this inner 

shield voltage is impressed across the inductance existing between the inner and 

outer shields in order to determine the current flowing on the inner shield. In 

the third state, the current on the inner shield is passed through R2 and M2, the 

transfer parameters of the inner shield. 

The resultant voltages are then the voltages that appear between line and 

ground on the central signal conductor. The amplitudes and shapes of these 

(4 

Figure 13.23 Multiple-shielded cable. 

(a) Test conditions 

(b) Conductor voltages 

(c) Equivalent circuits 
voltages are also shown in Figure 13.23. More effects have been illustrated by a 

series of measurements on a rather complex multiconductor cable. The overall 

makeup of the cable itself is shown in Figure 13.34. It consists of eight shielded 

quads surrounding a shielded center complex of nine conductors. The entire 

group of conductors is contained within an overall shield. The overall shield and 

the shields on the individual complexes are insulated from each other. 

Figure 13.25 shows the waveshape of the current injected onto the overall 

shield of the cable. This current was transformer-coupled to the shield, allowing 

the overall shield to be grounded at both ends. 

Figure 13.26 shows voltages measured on three of the different shielded 

INSULATING 

JACKET 

4 CONDUCTORS 

- 

Figure 13.24 Construction of a multiconductor cable. 
18A PEAK 

Figure 13.25 Current on overall shield. 

complexes within the overall shield. Measurements were made on the shields of 

two of the eight shielded quads and on the shield of the center complex. With 

high-impedance terminations at each end of the cable, the voltage from each of 

the individual shields to ground was the same. Voltages on individual conductors 

within the shielded groups would be the same as the voltage on the shield of that 

group. The position of a conductor within an overall shield thus has only a 

secondarder effect on the coupling from the internal electric field of the shield 

to the signal conductor. 

Figure 13.27 shows shield voltages measured with a low-impedance load 

connecting the center shield to ground. A load impedance of 47 52 approximately matches the surge impedance of the transmission line existing between 

any of the shields on the internal groups and overall shield. It prevents the 

development of oscillations which would otherwise be superimposed on an 

underlying waveform, like those shown in Figure 13.26. Note that the voltage on 

the center shield is now increased to nearly twice what it was with a 

high-impedance load on the ends of the shields. Note, furthermore, that the 

voltage on the other shields remains nearly the same as it was with all shields at 

the left open circuited. The slight increase of voltage noted on the unconnected 

shields results from capacitive coupling at the left end between the unconnected 

shields and the one center shield that is held nearly at ground potential by the 

47 52 load. Oscillogram No. 11 on Figure 13.28 shows the nature of these 

superimposed oscillations if the shield on the center complex is grounded 

directly. The voltage, V, shown measured at the right-hand end of the cable was 

in fact measured on one of the conductors within the center complex. The same 

voltage would be measured on any of the other conductors or on the shield 

itself. 

If the ground on the center shield is moved to the right-hand side, the 

389 
1A0. 7 

CENTER SHIELD 

No. 15 

- 0.5 ps/div--- 

Figure 13.26 Signal voltages vs source connections. 

voltage between a signal conductor and ground at the right-hand side is then 

reduced to zero because the capacitive coupling between the center conductor 

and center shield tries to hold the center conductor at the potential of the center 

shield, which, of course, is zero. A voltage of magnitude and waveshape as shown 

on oscillogram No. 11, but of opposite polarity, would then appear at the 

left-hand end of the cable. 

If both ends of the center shield are connected to ground, the voltage on 

the signal conductor is then held at zero at each end of the cable. Note that in 

this case a ground loop is formed between the two shields without the harmful 

effects commonly attributed to ground loops. 

Some additional considerations regarding the effect of connections to the 

signal conductors and the shield around the signal conductors are shown in 
No. 22 

-0.5 ps/div---, I 

Figure 13.2 7 Signal shield voltages: low source impedance on one shield. 

Figure 13.29. Compare oscillogram No. 17 of this figure to oscillogram No. 11 in 

Figure 13.28. Connecting the signal conductor to ground via a low impedance at 

the same end as that at which the shield around that signal conductor is 

connected to ground does not change the voltage appearing at the other end of 

the cable on that signal conductor. If the ground on that shield is moved to the 

right-hand of the cable, however, the voltage on the signal conductor at the 

right-hand end of the cable remains virtually unchanged. This was not the case 

when there was no low-impedance termination on the signal conductor at the 

left-hand end of the cable. Compare oscillogram No. 16 of Figure 13.29 with 

oscillogram No. 14 of Figure 13.28. Finally, note that a ground on the internal 

shield at each end keeps the voltage on the signal conductor low even with a 

low-impedance connection to ground on the other end of the signal conductor. 
CENTER SHIELD 

No. 14 

2 

L3 

\ 

> 

m 

0 

- 0.5 ps/div--+ 

Figure 13.28 Signal voltage vs grounding of signal shield: high source irnpedance - high receiving impedance. 

Compare oxillogram No. 18 in Figure 13.29 to oscillogram No. 12 in Figure 

13.28. 

The significant feature of these measurements is that rather complex 

phenomena can occur on multiple-shielded cables when the grounding of the 

inner shield is different from the grounding of the overall shield. Furthermore, 

the voltages on signal conductors can often be significantly lowered if the 

internal shield is grounded at each end, and the corresponding voltages are 

scarcely reduced at all if the inner shield is not grounded at each end. Most 

commonly, of course, an inner shield is grounded at only one end for control of 

low-frequency or low-level interference. Multiple grounding of the internal 

shields and the production of ground loops, then, is not automatically harmful; 

it may, in fact, be helpful. 
~L~CK;~ 

CENTER SHIELD 

0LUE SHIELD 

No. 18 

PLJk 

Fig 

+-- 0.5 )rs/div--- 

ure 13.29 Signal voltage vs grounding of signal shield: low source impedance - high receiving impedance. 

13.5 Transfer Characteristics of Actual Cables 

The transfer impedance of a shielded cable is not a factor commonly 

specified either by the manufacturer or by procurement specifications. Even 

among cables of the same nominal type, this transfer impedance may vary 

considerably between the cables supplied by different manufacturers. The most 

straightforward method of determining the transfer characteristic of a shielded 

cable, particularly if the cable involves multiple shields and multiple load 

impedances on conductors within the shield, is to make actual measurements of 

the conductor voltages that are produced by currents which are circulated 

through the shield. The techniques by which transient currents may be injected 
into the shields of cabIes through coupling transformers are desckibed in Chapter 

17. 

If actual measurements of coupling effects are not available, the transfer 

characteristics of many cables can be estimated from other published literature. 

One such summary giving the parameters of coaxial cable shields is shown in 

Table 13.1 (Reference 13.9). 

1 3.6 Transfer Characteristics of Connectors 

The transfer impedance of a cable connector or splice can be represented 

by 

where & is the resistance measured across the connector and Mi2 is a mutual 

inductance between the external shield circuit and the internal conductors of the 

cable. The value of ZT is usually not calculable, but it can be measured by 

passing current through a cable sample containing the connector and measuring 

the open circuit voltage induced on conductors inside the shield. Some typical 

values of Ro and MI 2 have been measured on cable connectors and are listed in 

Table 13.2. 

The transfer impedance of a connector is a lumped element in the cable 

circuit, in contrast to the distributed nature of the transfer impedance of the 

cable shield. The effect of leakage through the connector can thus be 

represented by a discrete voltage source 

where I. is the shield current. 

Because most connectors that would be used for shielded cables have 

essentially 100% optical coverage, their transfer admittance is usually negligible. 

In addition, most bulkhead or panel-mounting connectors are located at points 

where the shield voltage is a minimum, so that excitation of the internal 

conductors by the transfer admittance is small even if the transfer admittance 

itself is not small. 

13.7 Grounding of Shields 

It should be emphasized that the transfer parameters of Table 13.2 refer 

only to the properties of the connector itself. Additional transfer impedances 

may be produced by the manner in which the cable shield is connected to the 

connector or by the manner in which the cable connector is mounted to a 

bulkhead. Even slight inattention to detail may introduce into the circuit 

transfer impedances far greater than the impedance of the connector or possibly 

far greater than that of the rest of the cable shield. A common treatment of a 

shield at a connector is to insulate the shield with tape and connect it to the 

back shell through a pigtail. Such treatment is shown in Figure 13.30(a). Equally 
Table 13.1 COAXIAL CABLE SHIELD PARAMETERS.* 

Strand Dimaa d (inchas) 

m.w63 

9.- 

9.o.63 

9.9163 

4.0871 

Q.We.3 

9.eea 

9.- 

9.9163 

9.- 

9.- 

9.- 

0.W71 

Came 

9.mia 

9.9963 

9.016J 

01163 

4.9163 

0.W71 

0.9111 

waw 

Ends. Picks, P Angfep 

Carriers, C N (rnch-'I ldsgrterl ---- 

A - Ahernate I - Inner 

A1 - Ftrn altunrte 0 - OutA2 - S.mnd rhunme S - Shield (~nwbted from bra81dl 

Wcqht per 

FWI (Ib) 

0.01 9 

0.9 19 

@.on 

0.022 

0.033 

0.928 

0.933' 

0.039 

0.929 

0.029 

9.036 

9.035 

9.- 

0.089 

0.01 8 

@,a1 5 

@.@I6 

0.81 5 

0.016 

9.933 

0.833 

DC Resinma. RO 

(mR/m) 

6.6 

6.6 

7.5 

7.5 

4 .O 

5.5 

6.4 

5.9 

4.8 

4.4 

4.4 

3.7 

4.3 

14.2 

13.7 

8.6 

8.1 

8.7 

8.1 

4 A 

4 .O 

Braid aauofly consists of flat tape conductors. Strd dlomotar and wnds an 

ostimatd for the equ~valent cross-rutional araa of the tap. conductor; therefore. 

F. K. ond stranding fector are opproximcrte. 
Table 13.1 COAXL4LCABLESI-IIEWPARAMETERS. (CONTINUE

Outs~de Weave DC Resin- 

Cable Type Strand Ofam- D~ameter Ends, Pzkr. P Angle# Coverage, We~ght pe~ Stranding ance, RD I

{RG-) eter d (inches) (~nch) Csrrjerl. C 1 (~nch-') (degrees) XF K (%I Fool (lbrl Factor (rnIZ/m) M
Table 13.1 COAXIAL CABLE SHIELD PARAMETERS. (CONTINUE

Cable Type 

IRG-J 

outwde 

Strand Dlam- D~ameter 

eter d [inches) f~nchl -- 

Wcaw 

Ends. Plckr, P Anglea 

Carriers, C N (inch-' ) (degrees1 Frll. F ----- Gvnage Weight per K (%) Foot (lbs) -- 

DC Res~rt- L

Strand~ng ance. Ro Ind

Factor (mR/m) hA
Table 13.1 COAXIAL CABLE SHIELD PARAMETERS. (CONTINUED

Outslds Weave OC Rn~st- L

Gble Type Strand D~am- D~ameter Ends. Picks, P Anglep Coverage, VVelght per Strand~ng ance. Ro lnd

[RG-) eter d (~Mer] (unch) Cerr,err. C N lunch-' (degrees1 FlII. F K 1%) ~~t (lbsl Factor (mn/ml ylZ

-- 
Table 13.2 RESISTANCE AND MUTUAL INDUCTANCE OF CABLE 

CONNECTORS 

'Too small to measure in presence of 4 inches of copper tube used to 

mount connector. 

Connector 

Multipin 

Aerospace 

connectors 

(Threaded) 

Type N 

Type BNC 

(Bayonet 1 

Anodized 

Open shell 

Split shell 

Identification 

Burndy NA5-15863 

Deutch 38068-10-5PN 

Deutch 38068-1 8-31 SN 

Deutch 38060-22-55SN 

Deutch 38068-14-7SN 

Deutch 38060-1 4-7SN 

Deutch 38060-14-7SN 

Deutch 38068-1 2-1 2SN 

Deutch 38068-1 2-1 2SN 

Deutch 38068- 12- 12SN 

Deutch 38060-1 2-1 2SN 

Deutch 38068-1 2-1 2SN 

AMP 

AMP 

AMP 

UG 21 B/U-UG58A/U 

UG 88C/U-UG1094lU 

MS 24266R-228-55 

MS 3126-22-55 

MS 3100-165-1P 

MS 3106ARO 

(ohms) 

0.0033 

0.1 5 

0.005 

0.023 

0.046 

0.10 

0.023 

0.0033 

0.01 2 

0.01 2 

<O.M)l 

0.014 

0.0067 

0.0067 

0.0033 

' 

0.002 

5 x lo4 

0.5-1 

0.001 

M12 

(H 1 

5.7 X lo-" 

2.5 X lo-'' 

1.6 X lo-'' 

1.1 X 10-lo 

5.0 X 10-I ' 

8.2 X 10-I 

6.7 X lo-" 

3.0 X 10-l' 

1.3 X lo-" 

1.3 X lo-" 

2.5 X 10-l2 

3.5 X lo-" 

1.6 X lo-'' 

1.5X lo-'' 

1.9 X lo-'' 

' 

4-8 X lo-" 

UM < R, 4 20 MHZ 

wM < R, @ 20 MHz 

a 20 X lo-'' 
common practice is to insulate a panel connector from a panel with an insulating 

block and to ground the panel connector either to the panel through a pigtail or, 

more commonly, to an internal ground bus. A very common treatment of a 

shield at a connector, shown in Figure 13.30(c), is for the shield to be connected 

to one of the connector pins and grounded internally through a pigtail, either to 

the panel or to an internal ground bus. 

The coupling introduced by any of these configurations can be studied in 

terms of the self-impedance of the conductor used for the pigtail and the mutual 

inductance between a portion of the pigtail and the conductors in the connector. 

INSULATING 

CABLE CONNECTOR 

CONNECTOR PINS 

\ 0 

PIGTAIL ->-> 

Figure 13.30 Common treatment of shields at connectors. 

(a) Pigtail connection to a backshell 

(b) Pigtail grounding to a panel connector 

(c) Shield carried on a connector pin 
The flow of current through the self-inductance of the pigtail produces a voltage 

This voltage, less that induced in the mutual inductance between the 

pigtail and the conductors in the connector, will add directly to that produced 

by the flow of current along the shield. 

The self-inductance of a straight conductor is 

where P is the length of the conductor and r is its radius, both measured in 

centimeters. The inductance of a typical pigtail will not be significantly different 

even if it is not straight. 

The mutual inductance between the two conductors is 

where the factor Q, shown in Figure 13.31, is a function of the ratio of the 

length of conductors to their spacing. If d < Q, then 

Mutual inductance so defined is not the same as the mutual inductance 

referred to in Equations 13.4 and 13.1 2. This latter is best referred to as a 

transfer inductance, which one would normally want to be small. The 

corresponding quantity for the parallel conductors would be L minus M. One 

would normally want the mutual inductance (Equation 13.27) between two 

conductors to be large. 

A numerical example will illustrate the problems associated with pigtail 

grounds. Assume a shield is carried, as in Figure 13.30(c), through a panel on a 

pigtail and a set of connector pins. Assume the total length of the pigtail-pin 

combination to be 7 cm, the spacing between the pigtail and a signal conductor 

to be, on the average, 0.05 cm, and the pigtail to be made from number 22AWG 

wire having a radius of 0.0323 cm. The self-inductance of the pigtail would then 

be 74.5 nH, the mutual inductance between the two would be 65.0 nH, and the 

difference, the factor equivalent to the transfer inductance for the shielded 

cable, would be 9.5 nH. A rather mediocre shielded braid would have an 

inductance on the order of 1 nH/m, and a large-diameter, tightly woven braid 

might have an inductance on the order of 0.25 nHlm. The amount of voltage 

injected into a signal circuit by the current flowing through this one relatively 

short pigtail, then, would be as much as that coupled onto the signal conductor 

by 10 to 40 m of cable shield braid. This point may be illustrated by Figure 

13.32, an extension of the test series previously discussed in reference to Figures 
Figure 13.31 Mutualinductance. 

13.1 through 13.6. One end of the cable was grounded through a 33-inch length 

of wire, while the other end was connected to the ground plane directly. Voltage 

V1 was increased from 0.4 V (Figure 13.6) to 13.5 V solely because of the 

inductive drop in the ground lead. 

There are two fundamental drawbacks to pigtail grounding. The first, 

shown in Figure 13.33(a), is that the shield current, being constricted to a path 

of small diameter, sets up a more intense magnetic field at the surface of the 

conductor than it would if the conductor were larger. Because of the higher flux 

density, the self-inductance of the conductor is larger. A more intense magnetic 

field of itself would not really be harmful if all the flux so set up would link the 

signal conductors as well, because then the mutual inductance would also be 

high, and the difference, L minus M, would be low. If the shield current is 

confined to a small-diameter path, the field intensity rapidly falls off with 

distance away from the conductor, and hence the mutual inductance will always 

be smaller than the self-inductance. 

If the shield current is carried to ground on a shell concentric with the 

402 
Figure 13.32 Effects of inductance in a ground lead. 

(a) Circuit 

(b) Voltage 

conductors within the shield, shown in Figure 13.33(b), there is much less 

voltage introduced into the conductors for the following reasons: 

The length of the path through which the shield current must flow is 

shorter. 

The field intensity external to the shield is reduced by virtue of the 

inherently larger diameter of the path upon which the current flows. 

The field intensity inside the shield is low--nearly zero. 

The shorter, and larger diameter, current path implies less self-inductance. 

The absence of magnetic flux within the shield implies that the signal conductors 

are exposed to as much flux as is the shield: that M = L, and L minus M = 0. 

403 
(a) INTENSE MAGNETIC FIELD 

AROUND PIGTAIL CAUSES 

HIGH SELF-INDUCTANCE 

SHIELD CURRENT PATH IS 

UNIFORMLY DISTRIBUTED 

AROUND SIGNAL CONDUCTORS 

\r 

(b) MAGNETIC FIELD ENTIRELY 

EXTERNAL TO REGION 

OCCUPIED BY SIGNAL CONDUCTORS 

Figure 13.33 Grounding of shields. 

(a) Pigtails 

(b) 360" peripheral 

404 
REFERENCES 

F. A. Fisher, Effects of q Clzangilzg Afaglletic Field on Shielded 

Conductors, Lightning Protection Note 75-2, Internal General Electric 

Memorandum, High Voltage Laboratory, Corporate Research and Development, General Electric Company, Pittsfield, Massachusetts (2 June 

1975). 

Edward F. Vance, Couplirig to Cahles, DNA Handbook Revision, Chapter 

11, Stanford Research Institute, Menlo Park, California (December 

1974). 

Vance, p. 1 1-1 25. 

Vance, p. 1 1-128. 

Vance, p. 1 1-129. 

Vance, p. 1 1-1 30. 

Vance, p. 11-131. 

S. A. Shelkunoff, "Electromagnetic Theory of Coaxial Transmission 

Lines and Cylindrical Shields," BeN System Tecl~tlical Jourrlal, 13 

(1934): 532-79. 

Vance, pp. 1-1 33 to 11-1 36. 

13.10 Vance, p. 11-166. 

CHAPTER 14 

EXAMPLES OF INDUCED VOLTAGES 

MEASURED ON AIRCRAFT 

14.1 Introduction 

There have been several sets of tests made on aircraft in which simulated 

lightning currents were injected into the aircraft and the resultant voltages and 

currents on the aircraft wiring measured. Space precludes giving in this report a 

con~prehensive summary of the test results, and the user is referred to the 

reports on the individual tests. A few examples of the measurements will be 

@en in the following sections, and an atternpt will be made to indicate the most 

important facts learned during the tests. 

14.2 Wing from F-89J Aircraft 

The first set of tests to be d~scussed (Reference 14.1) was one in which 

high lightning-like currents were injected into one wing of an F-89J aircraft. 

During the test, represented in Figure 14.1, the wing was fastened onto a 

screened instrument enclosure, which may be viewed as representing the fuselage 

of the aircraft. Lightning-like currents of up to 40 000 A were injected into the 

wing or into the external wing tip tank from a high-current surge generator, 

allowed to flow along the wing to the outer wall of the screened instrument 

enclosure, and then to ground. An example of one of the types of current 

injected Into the wing is shown in Figure 14.l(b). In order to obtain maximum 

current, the surge generator was operated in a mode that essentially allowed the 

production of only one cycle of a damped oscillatory current, unlike a typical 

lightning current, which would rise to crest rather fast and decay at a much 

slower rate. The shape of the current wave must be considered when observing 

the waveshape of some of the voltages that will be discussed. In particular, note 

that at about 20 ps there appears a major discontinuity in waveshape. This 

discontinuity in current waveshape is reflected in the induced voltages. 

Within the wing there were a number ofelectrical circuits, such as those to 

navigation lights, fuel gauges, pumps, relays, and switches indicating position of 

flaps. Some of these ran in the leading edge of the wing and were well shielded 

from many electromagnetic effects, while others ran along the trailing edge 

between the main body of the wing and the wing flaps. These latter were most 

exposed to the electromagnetic fields. All of the circuits were relatively 

independent of each other; they were not, as a general rule, bundled together in 

one large cable bundle, a practice that provides maximum coupling from one 

circuit to another and makes analysis difficult. 

The first circuit that will be discussed, shown in Figure 14.2, was a circuit 

supplying power to a position light mounted on the external fuel tank. An 

electrical diagram of the circuit, shown in Figure 14.2(b), shows that the circuit 

consisted of one wire supplying power to the filament of the position light with 

the return circuit for the light being through the wingstructure. Accordingly, if 
the lightning current contacts the external tank, that circuit will be influenced 

by the resistance Rl of the hangers fastening the tank to the wing, by R2, the 

inherent resistance of the wing, and by magnetic flux arising from the flow of 

current. 

INSTRUMENT 

ENCLOSURE---- 

Figure 14.1 High current injection tests on the wing from an F-89J. 

(a) Test arrangement 

(b) Waveshape of injected current 

Typical results for the stroke position shown are given in Figure 14.2(c) 

and 14.2(d). The open circuit voltage is seen to rise rapidly to its crest and to 

decay more rapidly than does the injected curlent shown in Figure 14. I (b). As a 

result, the open circuit voltage was responding primarily to magnetic flux. 

When the conductor was shorted to ground at the instrument enclosure, 

the short circuit current rose to its crest in approximately the same length of 

time as did the injected current and displayed much the same waveshape as the 

injected current. 

Figure 14.2(e) shows an approximate equivalent circuit that might be 

derived. L1 and R1 represent a transfer impedance between the current flowing 

in the wing and the voltage developed on the circuit. L2 and R2 represent the 

inherent inductance and resistance of the wires between the fuselage and the 

light. The transfer inductance and resistance, which it should be emphasized do 

not necessarily represent any clearly definable resistance or inductance of the 

wing, are merely those values which, when operated upon by the external 
POSITION 

SURGE GENERATOR 

(4 I 

Figure 14.2 A wing tip circuit. 

(a) Circuit orientation 

(b) Electrical details of circuit 

(c) Open circuit voltage 

(d) Short circuit current 

(e) Equivalent circuit 

lightning current, produced the observed open circuit voltage. 

A different type of circuit is that shown in Figure 14.3. In this circuit a 

conductor ran through the leading edge of the wing and terminated in an open 

circuit on a pylon mounted underneath the wing. In the electrical detail circuit 

shown in Figure 14.3(b), it can be seen that this circuit would not respond to 
;", PYLON 

I I 

NOT DETERMINED 

Figure 14.3 A Pylon circuit - open at pylon. 

(a) Circuit orientation 

(b) Electrical details of circuit 

(c) Open circuit voltage 

(d) Equivalent circuit 

the voltage developed across the resistance between the tank and the wing. The 

circuit would respond in some measure to some fraction of the wing resistance 

and to some fraction of the magnetic: field set up by the flow of current in the 

wing, but since the circuit was only capacitively coupled to the wing, the total 

coupling impedance should have been, and was, less than that of the circuit 
shown in Figure 14.2. The purpose of the conductor shown in Figure 14.3 was 

to supply power to a relay and to explosive bolts in the pylon used to hold a 

weapon. In Figure 14.3 the pylon was not installed, so there was no load on the 

conductor. Figure 14.4 shows the results when that pylon was installed, when 

the conductor was connected to a relay with a return through the aircraft 

structure, and when the lightning flash was allowed to contact the pylon. The 

combination of a structural return path for the circuit and a lightning flash 

terminating upon the pylon and thus including the resistive drop across Rj, the 

resistance between the pylon and the wing, served to make the voltage much 

greater than it was when the conductor was open circuited. No attempt was 

made to completely analyze from which area the total amount of magnetic flux 

was coming or whether the flux 91, representing that in the pylon, or 42, 

- 1 

(b) *RELAY 

PYLON 

Figure 14.4 A Pylon circuit - loaded at pylon. 

(a) Circuit orientation 

(b) Electrical details of circuit 

(c) Open circuit voltage 
representing the flux within the wing, was the larger. 

A fourth circuit, and one which illustrates the interaction between the 

lightning-developed voltage and the circuit response is shown in Figure 14.5. In 

this circuit a slot antenna excited by a grounded stub and fed from a length of 

75 52 coaxial cable was installed in the leading edge of the wing. The shorted 

stub that excited the slot antenna was the predominant area intercepting the 

magnetic flux produced by the lightning current in the wing. As in the three 

previous examples, the basic voltage developed in this antenna and shown in 

Figures 14.5(c) or 14.5(e) followed the same pattern as that shown for the other 

circuits. The rapid transition on the leading edge of the voltage, however, was 

capable of exciting an oscillation within the coaxial cable feeding the antenna. 

When the antenna circuit was terminated in a resistor matching the surge 

impedance of the cable, the higher frequency ringing oscillation disappeared, 

leaving only the underlying response of the antenna to the magnetic field 

surrounding the wing. 

Several significant things were learned during this test series. The first was 

that the voltages induced in a typical circuit within the wing consisted of a 

magnetically induced component and a component proportional to the 

resistance of the current path. The location at which the lightning flash 

contacted the wing had an important effect on the magnitude of voltages 

developed on different circuits. None of the circuits displayed the easily 

understandable response of the conductors within the elliptical cylinder, and in 

none of the circuits could a clearly definable transfer impedance that was easily 

relatable to the structural geometry of the wing be derived. Nearly aU the 

voltages and currents measured on the circuits could be explained in terms of the 

simple equivalent circuit shown in Figure 14.6, but there did not seem to be any 

easy way in which the magnitudes of the transfer impedance could be related to 

the physical geometry of the wing or to the location of the conductors within 

the wing. It was possible, however, to say that conductors located in the forward 

portion of the wing were better shielded, and consequently had lower transfer 

impedances, than those along the trailing edge. Likewise, it could be observed 

that circuits which did not have any electrical return through the wing structure 

had lower transfer impedances than those circuits which did have a return 

through the wing structure. 

These equivalent circuits may be based upon the self-inductance of the 

wing, the self-inductance of the circuit within the wing, and the mutual impedance between the wing and the internal conductor. Alternatively, they may 

be based upon a transfer impedance, as shown in Figure 14.5(c). In the former 

the self-inductance of the wing would be nearly equal to the mutual inductance 

between the wing and the internal conductors. Their difference would be equal 

to the transfer inductance in the latter approach. Since the values of the transfer 

impedance will be much smaller than the values representing the self-impedance 

of the circuit, the latter leads to the more easily handled formulation. The latter 

approach is also compatible with the concept of mutual inductance or transfer 

impedance discussed for shielded cables in Chapter 13. 
Figure 14.5 An antenna circuit. 

(a) Circuit orientation 

(b) Electrical details of circuit 

(c) Open circuit voltage - RT = 

(d) Open circuit voltage - RT = 75C2 

(e) Open circuit voltage - RT = 00 

(f) Short circuit current - RT = 0 

14.3 F-8 Digital-Fly-By-Wire (DFBW) 

The second set of measurements (Reference 14.2), about which some 

discussion will follow, was that made on an F-8 aircraft fitted with a fly-by-wire 

control system. The fly-by-wire controls, shown in Figure 14.7, consisted of a 

primary digital system, a backup analog system, and a common set of power 

actuators operating the control surfaces. The major components of the control 
Figure 14.6 Equivalent circuits. 

(a) and (b) Circuits based on self- and mutual inductances of 

the wing 

(c) Circuit based on an equivalent transfer admittance: 

Z = Rt + I,Mt 

I 

system were located in three locations: the cockpit, where sensors coupled to 

the control stick provided signals for the control systems; an area behind the 

cockpit, where there was located the digital computer; and a compartment 

behind and below the cockpit on the left side of the aircraft. This latter 

compartment was one that would normally have been occupied by guns; 

accordingly, it will be referred to as the "gun bay." In this gun bay were located 

the interface and control assemblies. 

Several hydraulic actuators were located at each of the major surfaces. 

LIGHTIJING , L 0 

Lw Rw 

I 
Tbese were interconnected to the fly-by-wire control systems through wire 

bundles that ran under the wings. The control systems did not depend upon the 

aircraft structure as a return path; the system was considered to have a 

single-point ground, and that single-point ground was located at a panel in the 

gun bay. By and large, none of the control wiring in the aircraft was shielded. 

In contrast to the tests on the wing of the F-89J aircraft, in which 

high-amplitude currents were injected into the wing from a high-power surge 

generator, the tests on the complete F-8 aircraft were made with what has been 

called the transient analyzer, a portable and relatively low-power surge generator 

capable of injecting currents of a waveform similar to that found in lightning but 

of a much lower, and nondestructive, level. During the tests on the F-8, the 

injected current was of the order of 300 A. Several different current waveshapes 

were employed; the one to which most frequent reference will be made in 

this abbreviated set of test results is that described as the fast waveform, a 

current rising to crest in about 3 ps and decaying to half value in about 16 p. 

These waveforms are shown in Figure 14.8. 

The first set of measurements to which reference will be made was that on 

a set of spare conductors running between an interface box in the gun bay and a 

disconnect panel located near the leading edge of the vertical stabilizer. The 

routings of the circuit and the waveforms are shown in Figures 14.9 and 14.10. 

The voltage measured between the conductor and ground consisted of a 

high-amplitude oscillatory component and a lower amplitude but longer 

duration component. The oscillatory component was excited by magnetic flux 

leaking inside the aircraft, while the longer duration component was produced 

by the flow of current through the structural resistance of the aircraft. Since the 

voltages induced by the leakage of magnetic flux were proportional to the rate 

of change of that flux, it followed that the oscillatory component would have 

been more pronounced for faster currents injected into the aircraft than it was 

MODE AND POWER PANEL 

cENTE19;P STICK 

SYSTEM ELECTR&IC"\ 

DISPLAY AND KEYBOARD 

Figure 14.7 Location of fly-by-wire control system hardware and wiring 

bundles in F-8 aircraft. 
for slower currents. This effect was noted as indicated by the oscillograrns. 

Voltages measured between conductors, shown in Figure 14.10, were much 

smaller than the voltages measured between either of the conductors and the 

airframe. Such results would be expected on a well-balanced circuit. 

SLOW WAVEFORM 

83 A/div. 20 ps/div. 

FAST WAVEFORM 

83 A/div. 20 ps/div. 

VERY FAST WAVEFORM 

2 ps/div. - 83 A/div. 10 ps/div. 83 A/div. 

Figure 14.8 Simulated lightning test waveforms. 

NOTE: " - ps/div." is horizontal time scale and " A/div." is vertical 

amplitude scaIe. 

41 6 
SURGE 

CURRENT 

INJECTED 

FAST WAVE 

0 2V diu No 373 10~s 'dtv 

Figure 14.9 Spare conductor measurements on P22 (Pin 24 to airframe). 

FAST I, WAVEFORM 

PINS 74-75 

1 V/div. No. 368 1 ps/div. 

Figure 14.10 Spare conductor measurement on P22 (line-to-line, Pins 24 to 

25). 
The second set of measurements was made on a circuit running from the 

interface control unit and the gun bay to a wing position indicator switch 

located underneath the leading edge of the wing. The wing on the F-8 aircraft 

could be raised or lowered around a pivot point towards its rear in order to 

change the angle of attack during landing and takeoff. The purpose of the 

switch was to indicate the position of the wing. The voltages induced on that 

switch circuit are shown in Figure 14.1 1. The voltages measured from 

line to airframe were higher than those measured from line to line, but it is 

significant that the line-line voltages, while of a somewhat different waveshape, 

were not much lower than the line-to-ground voltages. The reason for this lay in 

the fact that the load impedances in the interface box in the gun bay were 

different from each other on the two sides of the circuit. One side connected to 

a power supply bus, while the other side probably connected to an emitter 

follower. Another significant feature about these voltages was that they were 

again of an oscillatory nature. They were apparently excited by the leakage of 

magnetic flux inside the aircraft and were not excited by the drop in potential 

along the structural resistance of the aircraft. 

(SWITCH CLOSED 

WING UP) 

ONE LINE-TO-AIRFRAME 

0.2V/div. NO. 361 I ps/div. 0.2V/div. NO. 364 I ps/div. 

Figure 14.11 Voltages induced in wing position indicator switch circuit at open 

plug P22. 

Figures 14.1 2 and 14.1 3 show voltages measured on two different circuits 

going to actuators, one (Figure 14.12) going to the left pitch actuator and the 

other (Figurc 14.13) going to the left roll actuator. In both cases the voltages 

displayed were the output of the driver amplifier used to control the servo valve 

in the actuator. Both of these were differential measurements, line-line voltage 

measurements. The significant feature about these measurements was, again, that 

the characteristic response was oscillatory and apparently excited by the leakage 

of magnetic flux inside the aircraft. There was some dependence of the voltage 
FAST I WAVEFORM 

LEFT PITCH VALVE DRIVE OUTPUT 

(HIGH TO LOW) PINS 15-16 

0.2V/div. NO. 355 1 ps/div. 

Figure 14.12 Left pitch valve drive output (high to low) at plug P22 (system 

battery powered). 

FAST IL WAVEFORM 

I, N - LWT 1,N-T 

O.ZV/div. No. 359 1 ps/div. 0.2V/div. NO. 360 1 ps/div. 

Figure 14.13 Voltages induced in left roll valve drive output circuit (Pins 

44-45) at open plug P22. 

on the path of current followed through the aircraft, but the dependence was 

not large. Both lightning current paths produced about the same peak voltage of 

transient. 

On the F-8, as is typical of most aircraft, the control wires were laced 

together into fairly large bundles. The routing of some typical bundles in the gun 

bay housing the backup and interface electronic control boxes is shown in 

Figure 14.14. While it was not possible to measure the current on individual 
wires within these cable bundles because of limitations of measurement 

technique and because of the large number of wires within the bundles, it was 

possible to measure the total current flowing on the various bundles. This was 

done by clamping around the bundle a current transformer having a split core. 

Typical results of these cable measurements are shown in Figures 14.15 and 

14.16. The bulk cable currents were also found to be oscillatory, just as were the 

voltages on conductors described earlier. Since the flight control wiring did not 

make use of multiple ground points within the aircraft, it follows that none of 

the currents in these cable bundles would exhibit any of the long time response 

characteristic of multiple-grounded conductors. 

No. 209 FAST 

BUNDLES 

LEADING 

TOWARD 

COCKPIT 

Figure 14.14 Cable bundIes within the gun bay. 

Figure 14.17 shows a statistical distribution of the peak amplitude of 

currents in all of the cable bundles upon which measurements were made. The 

distribution is shown both for the actual current amplitudes injected into the 

aircraft and in terms of what those currents would be if the results were scaled 

up to currents representative of actual lightning flashes. In terms of an 

average-amplitude lightning flash, 30 000 A, the total current on most cable 

bundles would have been on the order of 20 to 100 A. 

Measurements were made of the amplitude and waveshape of the magnetic 

field at a number of points in and around the aircraft. One location upon which 

attention was concentrated was the cockpit, since the cockpit is an inherently 

unshielded region and one in which many control wires would be subjected to 

changing magnetic fields. The positions at which fields were measured, the peak 

amplitude of the fields, and the predominant orientation of the fields are shown 
- 

FASTI, CURRENT SLOW I, CURRENT 

_. -ANo. 207 El ..r - ..- No 208 

0 22A d~v O.llA dl". 

0 4A PP 0.1A pp 

All 2ps/div. 

Figure 14.15 Currents on cable bundles leading toward cockpit and left-hand 

instrument panel. 

FASTI, CURRENT SLOW I, CURRENT 

No. 199 No. 200 

0.44A/div. .22A/div. 

1.3A pp 0.1A pp 

No. 203 NOT SHOWN BECAUSE No. 204 

0.44A/div. , SIGNAL IS OBSCURED 

1.8A pp BY BACKGROUND NOISE 

ALL 2ps/div. 

Figure 14,16 Currents on cable bundles leading toward the area behind the 

cockpit. 

42 1 
[AMPERES PEAK TO PEAK) 

Figure 14.17 Distribution of amplitudes of cable bundle cwrents (measured in 

left gun bay). 

in Figure 14.18. The magnetic fields were measured with a probe which had a 

characteristic time constant of about 4 ps. When exposed to fields changing in 

times less than 4 ,us, it would respond to the absolute magnitude of the field 

intensity, and, when exposed to fields changing in times longer than 4 ps, it 

would respond to the rate of change of the magnetic field. A few typical 

measurements of field waveshape are shown in Figure 14.19. The most 

significant feature about these measurements is that there was no orientation of 

the magnetic field probe that resulted in a zero output, which indicated that the 

orientation of the magnetic field was not uniform with respect to time. The field 

produced at any one point was the sum of the field produced by the total flow 

of current through the aircraft and that produced by oscillatory current in the 

various structural members as the current in those structural members changed 

with time. Similar effects can be expected on other aircraft, and, accordingly, 
Figure 14.18 Magnetic field measurements in the cockpit. 

FIELD ABOVE SEAT - POSITION 4 

- - L No. 363 

PITCH 

7.5A/m/div. 

I 

I \ 

I I 

I I 

@ 18.2A/m / 1 L- - 

.- -- - 

@ 7.2A/m-. 

SEAT 

RIGHT 

CONSOLE 

COCKPIT 

LOOKING AFT FROM 

INSTRUMENT PANEL 

FIELD ON SEAT - POSITION 3 

ALL PICTURES 5ps:div. 

Figure 14.19 Magnetic fields near pilot's seat. 

No. 325 

PITCH 

3A/m/div. 

No. 327 

ROLL 

3A/m/div. 

No. 328 

YAW 

3A/m/div. 

there will seldom be any clear-cut manner in which electrical wiring can be 

oriented so as to pick up the minimum value of magnetic field and hence have 

423 
minimum induced voltages. One might best assume that the magnetic field at 

any point will always be oriented in the worst case. 

Some measurements of the magnetic field within the gun bay housing the 

backup control system and the interface electronics package are shown in Figure 

14.20. Clearly, the magnetic field inside the gun bay compartment was of lower 

amplitude than that of the field external to the gun bay. The waveshapes 

of the field are more difficult to understand. First of all, it must be kept in mind 

that the probe was responding to the rate of change in magnetic field after about 

4 ps and was responding to the field itself for times shorter than about 4 ps. 

Accordingly, the oscillograms displaying the field inside the gun bay indicated 

that there was, first of all, a component of magnetic field that rose to its crest 

about as fast as did the outside field. From then on, the field continued to 

increase at a slower rate, the final value of the field not being indicated by the 

oscillograms. The fact that the field continued to rise, but at a slower rate, was 

in agreement with the behavior predicted in Section 10.2.4, where it was 

predicted that the field would increase with a time constant characteristic of the 

No. 388 

VIEW LOOKING FORWARD 

I I CURRENT FLOW FWD TO AFT 

30A 'm/div. 2ps/div. 

EXTERNAL FIELD 

0.15A/m/div. No. 432 Sps/div. 0.3OA/m/div. No. 440 lOps/div. 

PROBE AGAINST INSIDE PROBE IN MIDDLE OF 

OF DOOR-POS. A LEFT GUN BAY-POS. B 

INTERNAL FIELD 

Figure 14.20 Magnetic fields inside gun bay. 

424 
internal inductance and resistance of the cavity in which the fields were to be 

measured. 

Fields inside a battery compartment located aft of the gun bay are shown 

on Figure 14.21. The measurements showed first some oscillatory magnetic 

field, followed by a field which rose to crest at a time much longer than the crest 

time or even the duration of the lightning current that was injected into the 

VIEW LOOKING FORWARD 

RIGHT 

1.5A/midiv. ipsidiv 

(COVER REMOVED) 

No. 451 

O.lSA/m/div. Ips 'div. 

(COVER IN PLACE) 

No. 452 

0.1 SA/rn/div. 20ps/div. 0.15A/m/div. 100 sldiv. 

(COVER IN PLACE) (COVER IN PLA~E) 

Figure 14.21 Magnetic fields inside the battery compartment (probe time 

constant = 4 ps). 

425 
aircraft. Oscillogram No. 452 in Figure 14.21 indicated the rate of change of 

field as failing to zero at about 400 ys. This would indicate that the field itself 

reached its crest value in about 400 ys. The different nature of the response of 

the fields inside the gun bay and the battery compartment can be explained in 

terms of the types of covers and fasteners used on the two compartments. The 

cover over the battery compartment was held in place by fasteners spaced about 

every 3.8 cm, and hence made good contact to the rest of the airframe. The use 

of multiple fasteners resulted in minimum constriction of the current flow and 

minimum resistance to the circulating current within the battery compartment. 

The gun bay covers, on the other hand, since they were originally designed for 

fast access time, had far fewer fasteners. The fasteners in the gun bay, in fact, 

were spaced about every 30 cm. Accordingly, the cover on the gun bay provided 

a much greater constriction of current flow and a much greater resistance than 

did the cover on the battery compartment. Both of these factors were of a 

nature to allow the field within the gun bay to reach its peak value faster than 

did the field within the battery compartment. 

Some of the significant points about the results of tests on this aircraft 

may be summarized here. The first was that the use of a single-point ground 

system most emphatically did not eliminate all transient voltages produced by 

the flow of lightning current through the structure of the aircraft. The second 

was that the characteristic response of the wiring, both for voltage and for 

current, was a damped oscillation with a frequency in the range I to 5 MHL. The 

very abbreviated series of test results just presented does not indicate the fact 

clearly, but the frequency of oscillation depended considerably on the length of 

the circuit involved: the longer the physical length of the wires, the lower the 

oscillatory frequencies. This was not a clear-cut rule, since the response of any 

one circuit was greatly influenced by the high degree of coupling between all of 

the different circuits. The third significant point is that the total current on any 

cable bundle was of the order of 20 to 100 A for an average lightning flash. This 

bulk cable current was again oscillatory with a frequency tending to correspond 

to the length of the cable bundle. The equipment bays in this aircraft, not being 

designed for electromagnetic shielding qualities, allowed significant amounts of 

magnetic flux to develop within those bays. This is particularly true of those 

bays intended for ease of access. As a rule of thumb, it might be expected that 

on aircraft the equipment bays housing electronic equipment might well be 

fitted with covers designed more with ease of access in mind than with magnetic 

shielding qualities in mind. It must thus be expected that those enclosures for 

which the electromagnetic shielding should In principle be the greatest may well 

be those enclosures having the poorest shielding. 

No attempt was made to determine in this aircraft the degree to which any 

circuit voltages could be reduced by the use of shielded conductors. It was 

noted, however, that on those circuits in which a shield was used and in which 

the shield was grounded at more than one point, the current on that shield 

tended more to have the slower double exponential waveshape of the external 

lightning current than to have the high-frequency oscillatory current excited on 

all of the unshielded cable bundles. 

426 
REFERENCES 

14.1 K. J. Lloyd, J. A. Plumer, and L. C. Walko, Measurements and Analysis of 

Lightning-Induced Voltages in Aircraft Electrical Circuits, NASA 

CR-1744, National Aeronautics and Space Administration, Lyndon B, 

Johnson Space Center, Houston, Texas (February 197 1). 

14.2 J. A. Plumer, F. A. Fisher, and L. C. Walko, Lightning Effects on the 

NASA F-8 Digital-Fly-By-Wire Airplane, NASA CR-2524, prepared by the 

High Voltage Laboratory, Environmental Electromagnetics Unit, Corporate Research and Development, General Electric Company, Pittsfield, 

Massachusetts, for the National Aeronautics and Space Administration, 

Lewis Research Center, Cleveland, Ohio (March 1975). 

CHAPTER 15 

DESIGN TO MINIMIZE INDIRECT EFFECTS 

15.1 Some Premises and Goals for Design 

Any design program aimed at minimizing or eliminating the indirect 

effects of lightning needs to be founded on certain premises of which the 

following might be examples. The first premise is that it is desirable to achieve a 

design that prevents the indirect effects from causing irreversible physical 

damage. The second premise is that it is essential to eliminate that interference 

which provides an imminent hazard to the safety of the vehicle or its crew or 

one that presents a severe risk of preventing the completion of the aircraft's 

mission. (It is, of course, desirable to minimize all interference.) For example, 

indirect effects that cause warning lights to appear might be acceptable whereas 

indirect effects that lead to the tripping of circuit breakers would be 

unacceptable, even if it were possible to reset the circuit breakers. Immediately 

after a lightning flash the pilot of the aircraft might have enough things to do 

that he should not be called upon to reset circuit breakers. Further, interference 

that leads to the scrambling of one channel of a redundant digital control system 

is probably acceptable, but interference that causes computers to shut down is 

unacceptable, particularly if the computers are shut down in a disorderly manner 

which results in the internally stored programs becoming scrambled. A third 

premise is that it is more productive to design electronic equipment that can 

accept transient signals on input and output leads than it is to initiate a retrofit 

program to provide protection to an existing system. A fourth premise is that it 

is more practical to design an electronic system around the capabilities of 

existing and proven protective devices or techniques than it is to develop and 

retrofit new and improved protective techniques to an electronic system 

designed without consideration of the transients that might be produced by 

lightning. A fifth premise is that trade-offs must be made between the cost of 

providing electronic equipment capable of withstanding lightning-induced 

transients and the cost of shielding equipment and interconnecting wiring from 

the electromagnetic effects of lightning. A sixth premise is that designers should 

take as much advantage as possible of the inherent shielding that aircraft 

structures are capable of supplying and should avoid placing equipment and 

wiring in locations that are most exposed to the electromagnetic fields produced 

by lightning. 

15.2 Improvement Through Location of Electronic Equipment 

While it is recognized that the designer may not have much choice in the 

matter, it is often possible to make improvements in the resistance to indirect 

effects by locating electronic equipment in regions where the electronlagnetic 

fields produced by lightning current are lowest and by avoiding the placement of 

equipment in the region where the electromagnetic fields are highest. For 

example, since the most important type of coupling from the outside 
electromagnetic environment to the inside of the aircraft is through apertures, it 

follows that equipment should be located as far from major apertures as 

possible. 

Further, since access doors with their imperfectly conducting covers are a 

major source of electromagnetic leakage, it follows that equipment should be 

located as far from such access doors as possible. In practice this may be more 

easily said than done, however, because frequently the purpose of access doors is 

to provide ready access to electronic equipment. 

One main goal should be to locate electronic equipment toward the center 

of the aircraft structure rather than at the extremities, since the electromagnetic 

fields tend to cancel toward the center of any structure. 

Another main goal should be to locate electronic equipment away from 

the outer skin of the vehicle - particularly away from the nose of the aircraft, 

where the radius of curvature of a prime current-carrying path is smallest. 

Furthermore, if possible, electronic equipment should be located in shielded 

compartments. 

15.3 Improvement Through Location of Wiring 

The designer has somewhat more control over the routing of wiring used 

to interconnect equipment than he does over the location of equipment itself. 

Wiring should be located away from apertures and away from regions where the 

radius of curvature of structural members or the outer skin is smallest. Moreover, 

wiring should be located as close to a ground plane or structural member as 

possible. If the structural member is shaped or can be shaped to provide a 

trough, the member will provide more inherent shielding than it will if the wiring 

is placed on the edge of the member. Some examples of typical structural 

members and the best places for bundles of wiring are shown in Figure 15.1. In 

each case the structural niember is assumed to be carrying current along its axis. 

Some basic principles to follow are these: 

1. The closer a conductor is placed to a metallic ground plane, the less 

is the flux that can pass between that conductor and the ground 

plane. 

2. Magnetic fields are concentrated around protruding structural 

members and diverge in inside corners. Hence, conductors located 

atop protruding members will intercept more magnetic flux than 

conductors placed in corners, where the field intensity is weaker. 

3. Fields will be weaker on the interior of a U-shaped member than 

they will be on the edges of that member. 

4. Fields will be lowest inside a closed member. 

Some examples of cable routing are shown on Figure 15.2. Along the 

interior of a structure a cable clamped to stiffeners, as at position 1, will 

effectively be spaced away from the metal skin by the height of the stiffeners. A 

conductor along the outside edge of the U-shaped member, as shown by 

conductor 2 in Figure 15.2(a), may or may not be better placed than conductor 

1 : effectiveness depends on how closely the conductor is attached to the side of 
Figure 15.1 Flux linkages vs conductor position. 

(a) Conductors over a plane 

(b) Conductors near an angle 

(c) Conductors near a channel 

(d) Conductors near a box 

In each case pictured 

Conductor I - highest flux linkages: worst 

Conductor 2 -- intermediate linkages: better 

Conductor 3 - lowest linkages: best 
Figlire 15.2 Conductor routing. 

(a) A fuselage structure 

(b) A wing structure 

In each case pictured 

Conductor 1 - highest flux linkages: worst 

Conductor 4 - lowest flux linkages: best 

the U-shaped member. Conductor 3, placed along the edge where the U-shaped 

member is attached to the stiffeners would probably be in a lower field 

environment than would either conductor I or conductor 7. Conductor 4, 

located on the interior of the U-channel, would be in the lowest field region and 

hence in the most effective position. Similar considerations apply to conductors 

located in structures like wings or stabilizers. 

A conductor located along the outside trailing edge of the wing, as shown 

by conductor 1 in Figure 15.2(b), will pick up much more flux than will any 

conductor located on the inside, probably by several orders of magnitude. Hence 
conductor 2, located on the inside of the trailing edge, would be better placed. 

Conductors 3 and 4, in that order, would be in the regions of lowest magnetic 

flux. A conductor that could be run inside a major structural member, as shown 

by conductor 4, will be exposed to a minimum amount of magnetic flux. 

Conductor 5, located at the forward edge of the wing, would be in a wellshielded region if the forward edge of the wing were metal, but it would be in a 

high field region and therefore vulnerable if the forward edge were a nonmetallic 

covering. However, even if the covering were metallic, conductor 5 would not be 

in as protected a region as that of either conductor 3 or conductor 4. 

Windshield posts, shown in Figure 15.3, tend to concentrate the current 

flowing on the exterior surface of the vehicle, particularly if a flash is swept 

back, contacting the windshield post directly or the eyebrow region above the 

windshield. Since the current is concentrated, the magnetic field intensity inside 

the crew compartment tends to be very high. The situation is aggravated by the 

fact that the windshields, unlike other regions where the field might have to 

diffuse through the metal surfaces, act as large apertures and so allow the 

internal magnetic flux to build to its peak values very rapidly. Instruments and 

wiring on the control panels are thus in a region of inherently high magnetic 

field strength. Conductors that run from overhead control panels (position A) to 

other instruments (position B) are often run along the windshield center posts. 

They are thus in a region of the most concentrated magnetic fields likely to be 

found on an aircraft, and accordingly they may have induced in them the highest 

voltages. 

1 5.4 Improvement Through Shielding 

In order to make an electronic system immune to the effects of lightning, 

it is almost always necessary to make judicious use of shielding on interconnecting wiring. Figure 15.4 shows some of the basic considerations. In Figure 15.4(a) 

an unshielded conductor, being exposed to the full magnetic field inside the 

structure, will have high voltages developed across the high-impedance 

terminations. The presence of a shield grounded at only one end will not 

significantly affect the magnitude of the voltage induced by changing magnetic 

fields, although a shield may protect against changing electric fields. While a 

shield may keep the voltage at the grounded end low, it will allow the voltage on 

the signal conductors to be high at the unshielded end. 

Shielding against magnetic fields requires the shield to be grounded at both 

ends, as shown in Figure 15.4(b), in order that it may carry a circulating current. 

It is the circulating current that cancels the magnetic fields that produce 

common-mode voltages. There is some merit in grounding such a shield at 

multiple points, since frequently the cable will be exposed to a significant 

amount of magnetic field over only a small portion of its total length. If the 

shield is multiple-grounded, the circulating currents will tend to flow along only 

one portion of the cable whereas, if it is grounded at only the two ends, current 

is constrained to flow the entire length of the cable. There is likewise some 

virtue in staggering the spacing between multiple ground points on a cable shield, 
Figure 15.3 Current flow along windshield posts. 

(a) External current flow 

(b) Internal magnetic fields 

since it is theoretically possible that uniform grounding can lead to troublesome 

standing waves if the shield is illuminated by a sustained frequency interference 

source. 

Of the different types of shields, shown in Figure 15.4(c), the solid shield 

inherently provides better shielding than does a braided shield, and a 

spiral-wrapped shield can be far inferior to a braided shield in performance. In 

severe environments braided shields using two overlapping courses of braid may 

give shielding performance approaching that of a solid-walled shield. 

Conduits, shown in Figure 15.4(d), may or may not provide electromagnetic shielding. Conduits in aircraft tend to be used more for mechanical 

protection than for electrical protection of conductors. Only if the conduit is 

electrically connected to the aircraft structure will it be able to carry current and 

thus provide shielding for the conductors within. Conduits for mechanical 

protection frequently are physically mounted in clamps that use rubber gaskets 

to prevent mechanical vibration and wear. Such clamps, of course, insulate the 

conduit and prevent it from having any magnetic shielding capability. Clearly, 

nonmetallic conduits will not provide electromagnetic shielding. 

The requirement that a shield intended for protection against lightning 

effects must be grounded at both ends raises the perennial controversy about 

434 
LOW 

LOW 

/ 

SOLID 

/ LOWEST 

BRAID 2 

SPIRAL g 

HIGHER 

---- 

(dl 

Figure 15.4 Types of shields. 

(a) Unshielded vs shielded 

(b) Multiple-grounded 

(c) Solid- vs braid- vs spiral-wrapped 

(d) Conduits 

single- versus multiple-point grounding of circuits. For many, usually legitimate, 

reasons low-level circuits need to be shielded against low-frequency interference. 

Most commonly, and usually legitimately, the shields intended for such 

low-frequency interference protection are grounded at only one end. But in this 

practice a fundamental concept often overlooked is that the physical length of 

such shields must be short compared to the wavelength of the interfering signals. 

Lightning-produced interference, however, is usually broad band and includes 

significant amounts of energy at quite high frequencies, frequencies higher than 
those the typical low-frequency shields are intended to handle. The dkhotomy 

between the requirements for shielding against high-frequency, lightningproduced interference and those for shielding against everyday, low-frequency 

interference is usually too great for both sets of requirements to be met by the 

use of only one shield system. 

Most commonly both sets of requirements can be met only by having one 

shield system to protect against low-frequency interference and a second shield 

system to protect against lightning-generated interference. The lightning shield 

can usually consist of an overall braided shield on a group of conductors with 

this overall shield being grounded to the aircraft structure at least at the ends. 

Within the overall shield may be placed whatever types of circuits are needed. 

Frequently these circuits will have a shielded conductor of their own. In a 

coordinated shielded system the designers of individual circuits should have the 

option of grounding such inner shields as their own requirements dictate, but 

they should not have the power of dictating the treatment of the overall shield. 

Such an overall shield (OAS) is shown in Figure 15.5. 

The method of grounding this overall shield can have a great impact on its 

effectiveness in protecting against lightning-generated interference. Figures 

15.5(b) through 15.5(e) show several methods of grounding such a shield when 

that shield is placed over a group of conductors being brought into an equipment 

case. For best performance the OAS should be terminated on the back shell of a 

connector specifically designed for such termination. The shield should make a 

360" circumferential connection to the back shell of the connector. The 

connector shell itself should be designed to have low dc resistance to its mating 

panel connector. Most commonly, such low-resistance mating requires the use of 

grounding fingers within the connector shell. Connectors without such grounding fingers frequently have high resistance between the mating shells, since the 

shells are frequently coated with an insulating coating to reduce problems of 

corrosion. The shell of the panel connector should also provide a 360" peripheral 

connection to the metal equipment case. Providing such a connection frequently 

requires that paint or other coatings on the case of the equipment be removed 

and the bare metal exposed. 

In the absence of a 360" connector, an external pigtail is often used for 

grounding the OAS, as shown in Figure 15.5(c). Such pigtails are definitely 

inferior to the 360" connector because they force an interfering current on the 

shield to be concentrated through the pigtail, and hence to provide a much 

greater degree of magnetic coupling to the core conductors than does the 

distributed current flow on a properly designed connector back shell. If such a 

pigtail is used, it should be as short as possible and should terminate on the 

outside of the equipment case. A pigtail of only a very few inches may introduce 

inore leakage from the shield onto the inner conductors than does a several-foot 

section of the shield itself. 

The practice of grounding an overall shield to the inside surface of an 

equipment case through a pigtail and a set of contacts in the connector is less 

effective than the use of an external pigtail, partly because the length of the 
OVERALL SHIELD (OAS) 

r 1 

INNER SHIELD 

(a) 

CONDUCTORS 

(b) 

(4 

OAS 

INTERNAL PIGTAIL 

(dl 

OAS 

(el 

Figure 15.5 Types of grounding for shields. 

pigtail is inherently longer and partly because it brings currents directly to the 

inside of the case. Such grounding of an overall shield should be avoided 

wherever possible and, in particular, must be avoided whenever the overall shield 

runs through a region where it will intercept a significant amount of energy from 

the external electromagnetic field. 

In no case should an overall shield be connected to a signal ground bus. 
15.5 Improvement Through Circuit Design 

One of the most important considerations in the control of lightningrelated interference through proper circuit design lies in the fundamental 

observation that a device with a broad band width can intercept more noise 

energy than can a narrow band-width device. Some of the considerations that 

derive from this observation are shown in Figure 15.6. The noise produced by 

lightning has a broad frequency spectrum. Considering for the moment only the 

spectrum of the lightning current, the observation is frequently made that most 

of the energy associated with the lightning current is contained in the 

low-frequency region, below 10 or 20 kHz. Before any sense of security is 

derived from that observation, it should be remembered that equipment is 

damaged or caused to malfunction in accordance with the total amount of 

energy intercepted. In a lightning flash there may be plenty of energy in the 

megahertz and multimegahertz region to cause interference. The energy that is 

available for damage or interference may well be concentrated in certain 

frequency bands by the characteristic response of the aircraft or the wlring 

within the aircraft. 

Without reference to any specific frequency regions, however, the energy 

spectrum of the lightning-generated interference on electrical wiring within an 

aircraft will still be a broad spectrum. A receptor with a broad pass band, shown 

in Figure 15.6(a), will inherently collect more energy than will a receptor with 

narrow pass band, shown in Figure 15.6(b). The narrower the pass band, the 

better. In this respect analog circuits have an inherent advantage over digital 

circuits, since a narrow-pass band digital circuit is almost a contradiction in 

terms. If possible, circuits should not have a pass band that includes dc, shown in 

Figure 15.6(c), because, when dc is excluded, the circuits will inherently be able 

to reject more of the energy associated with the flow of current through 

resistance of the structure. 

The studies of types of interference produced in aircraft by the flow of 

lightning current have shown that the lightning energy excites oscillatory 

frequencies on aircraft wiring, particularly ~f the wiring is based on a single-point 

ground concept. Those characteristic frequencies have tended to be in the range 

of several hundred kilohertz to a few megahertz. If at all possible, the pass bands 

of electronic equipment should not include these frequencies, as does the 

hypothetical pass band shown on Figure 15.6(d). Higher or lower pass bands 

would inherently be better than the one shown. As an extreme example, shown 

in Figure 15.6(e), fiber optic signal transmission operating in the infrared reglon 

avoids the frequency spectrum associated with lightning-generated interference 

almost completely. 

Basic considerations about circuit design and signal transmission are shown 

in Figure 15.7. First, as shown on Figure 15.7(a), signal circuits should avoid the 

use of the aircraft structure as a return path. If the structure is used as a return 

path, the resistively generated voltage drops will be included in the path between 

transmitting and receiving devices. On the other hand, signal transmission over a 

twisted-pair circuit with signal grounds isolated from the aircraft structure tend 
I PASSBAND 01 RECEPTOR 

ENERGY SPECTRUM 

OF LIGHTNING 

INTERFERENCE 

(a frequency 

f IR. 

Figlrre 15.6 Frequency considerations. 

to couple lower volt:rges in the signal path. It rnust not be forgotten, however, 

that the use of twisted-pair t~.ansmission lines does not eliminate the commonmode voltage to which electronic systems may be subjected. Comtnon-mode 

voltages applied to the unbalanced transmission path, as in Figure 15.7(b), can 

lead to line-line voltages which may at times be as high as the common-mode 

voltage. 

Differential transmission and reception devices, shown in Figure 15.7(c), 

can offer a many-fold improvement in the ability to reject the con~mon-mode 

voltages produced by lightning. 

In general it is preferable that wiring interconnecting two different pieces 

of electronic equipment not interface directly with the junctions of semiconductors, as shown in Figures 15.7(a) and 15.7(c). Even modest amounts of 

resistance connected between the junctions and the interfacing wires, shown in 
Figure 15.7(d), can greatly improve the ability of sen3iconductors to resist the 

transient voltages and currents. Chapter IG on coniponent damage mechanisms 

gives cxamples of the degree of improven~ent that may be obtained through the 

use of serics resistors. Transmission through balanced transmission lines and 

transformers, coupled with input protection for semiconductors, probably 

provides the greatest amount of protection against the transients induced on 

control wiring. 

TRANSMISSION RECEPTION 

I, PRODUCES AN ERROR SIGNAL 

(0) 

Figure 15.7 Considerations regarding circuit design. 

440 
15.6 Circuit Protection Through Use of Protective Devices 

Circuit protective devices can sometimes be used to limit the amount of 

electrical energy that a wire can couple into a piece of electronic equipment. 

While one can seldom eliminate interference through the use of circuit protective 

devices, circuit protective devices judiciously used can virtually eliminate 

physical damage to electronic devices. "Judicious use" usually means that 

protective devices must be incorporated into a piece of equipment at the time it 

is built, not added after trouble has been experienced. 

There are two basic types of overvoltage or transient protection devices 

(TPD): those which, on sensing an overvoltage, switch to a low-impedance state 

and thus cause the impressed voltage across them to collapse to a low value; and 

those which, on sensing an overvoltage, tend (by virtue of their nonlinear 

voltage-current relation) to maintain the voltage at that level but do not tend to 

collapse the voltage. Examples of the first type are spark gaps or arcing dielectric 

devices. Examples of the second type are Zener diodes and varistors. There are 

also devices which, on sensing an overvoltage, interrupt the power flow to the 

load. If this interruption is accomplished by electromechanical means, they 

should not be considered transient protection devices because they are 

inherently slow to respond. 

Switching devices inherently offer greater surge-power handling capability 

than do the Zener or varistor types of devices. The instantaneous power 

dissipated in a transient protective device is the product of the surge current 

flowing through the device and the voltage across the device. For a constant 

surge current, a switching device like a spark gap, across which the voltage is low 

while in the conducting state, will have less power released in it than a device 

like a Zener diode, across which the surge voltage remains high. For a given 

surge-power handling capability, a spark gap will thus be smaller physically than 

a Zener diode or varistor device. 

Another fundamental difference between switching devices (spark gaps) 

and nonswitching devices (Zener diodes or varistors) relates to their recovery 

characteristics after the surge has passed. If a line is protected by a spark gap and 

if that line is connected to a source of energy (a power bus, for example), that 

energy source must be disconnected from the line before the spark gap can 

switch back from its low-impedance conducting state to its high-impedance 

nonconducting state. Generally this requires opening a circuit breaker on the 

line. A Zener diode or varistor effectively ceases to conduct as soon as the 

voltage returns to its normal value. Operation of remote circuit breakers is not 

required. 

All types of overvoltage protection devices inherently operate by reflecting 

a portion of the surge energy to its source and by diverting the rest into another 

path, all with the intention of dissipating the surge energy in the resistance of 

the ground and interconnecting leads. The alternative to reflecting the energy is 

to absorb the surge energy in an unprotected load. Reflection and diversion of 

the surge energy are not without their hazards. Some are the following: 
The reflected energy can possibly appear on other unprotected 

circuits. 

Multiple reflections may cause the transient to last longer than it 

would otherwise. 

The spectral density of the energy may be changed, either high or 

low frequencies being enhanced. Interference problems on other 

circuits may well be increased even though the risk of damage to the 

protected circuit is reduced. 

Most commonly the appropriate type of transient protective device to be 

used depends on the amount of surge energy to be dealt with. Generally, this 

energy decreases the further away one gets frotn the ~troke. The surge energy to 

be expected can also be related crudely to the normal operating power of the 

circuit involved. One would normally expect lower surge levels on low-voltage 

signal circuits than on medium-power control circuits, and even lower levels than 

those on main-power distribution buses. Thus, one might logically use Zener 

diodes on individual circuit boards, vari~tors on terminal boards, and spark gaps 

on leads running to prime entry and exit polnts. 

There are basically four generic types of transient protective devices 

applicable to lightning hardening of aeronautical subsystems. These types are 

gas-filled spark gaps, Zener diodes, varistors, and arcing dielectrics. The latter are 

currently in development, but the others are readily available commercially. 

Each rype has both advantages and disadvantages. 

15.6.1 Spark Gaps 

Spark gaps are generally composed of two metal electrodes separated by a 

dielectric and held at a fixed distance from edch other. The gap may be sealed in 

a container. Electrodes may be spherical, but in sophisticated devices they are 

not. Sparkover voltage is determined by dielectric composition, density, and 

electrode geometry. Sparkover voltage is also dependent on voltage waveshape. 

If the voltage wave is increasing rapidly, the sparkover voltage will be higher than 

it is on a slowly rising wave. Commercially available spark gaps frequently 

contain minute amounts of tritium or other radioactive elements to reduce the 

dependence of sparkover voltage on voltage waveshape. 

The advantages of spark gaps are as follows: 

They are simple and reliable. 

They have very low-voltage drop during the conducting state. When 

the gap is carrying maximum current, the voltage across the gap is 

typically I0 to 20 V. If more current tries to flow, the arc channel 

increases in diameter and holds the same arc-drop. 4 low arc-drop 

indicates relatively low power absorption during the conducting 

phase. 

They have large power-handling capability. Gas-filled gaps have the 

highest peak current-handling capabilities of any transient protection 

device, and almost any gap can handle the maximum surge currents 

induced by lightning. 
They have high impedance and low capacitance. The low-shunt 

capacity and leakage current characteristics of gas-filled spark gaps 

minimize insertion problems for operating frequencies below 1 GHz. 

They provide bilateral operation, having the same characteristics on 

either polarity. 

The disadvantages of spark gaps include the following: 

They have relatively high sparkover voltage. 

Simple gaps do not extinguish follow current. This is a most 

important point to consider if they are to be used on a power 

circuit. The arc must be extinguished by removing the voltage 

(circuit breaker of fuse) or by inserting resistance rapidly into the 

circuit by an additional element, such as a silicon carbide or zinc 

oxide varistor. Through suitable designing, gaps can be made 

self-extinguishing for applied voltages up to about 100 V. Such 

self-extinguishing properties require the use of the magnetic blowout principle. 

They may have a large dependence of sparkover voltage on the 

waveshape of the voltage. Specifications relating to the impulse ratio 

or volt-time effect should be carefully considered. 

Since spark gaps reflect more energy than they absorb, external 

resistive components may be required to minimize ringing. Discrete 

linear or nonlinear resistors are sometimes used to assist in 

extinguishing a gap. This component may serve the dual purpose of 

limiting current and dissipating power. 

15.6.2 Zener Diodes 

This category includes all single-junction semiconductor devices such as 

rectifiers, in addition to Zener diodes. While other semiconductor devices, such 

as PNPN devices and bipolar transistors, may have appIication as surge arrestors, 

they will not be covered here because of the limited pertinent data available. 

Zener diodes are basically polarized devices which exhibit an avalanche 

breakdown when the applied voltage in the reverse bias direction exceeds the 

device's specified breakdown, or Zener voltage of the device. Operated in an 

opposed series configuration, diodes can be used as effective suppression devices. 

Since Zener diodes are designed to operate in the breakdown mode, they usually 

can perform more effectively as terminal protection devices than can signal 

diodcs. While the energy-handling capabilities of Zener diodes are modest when 

compared with those of spark gaps, they are very well adapted for protection of 

individual components or circuit boards. 

The advantages of Zener diodes include the following: 

They are of small size. 

They are easily mounted. 

They have low "firing" voltage. 

They have low dynamic impedance when conducting. 
They are self-extinguishing. When applied voltage drops below the 

Zener level, they cease conduction. 

They exhibit low volt-time turnup, or impulse ratio. 

The disadvantages of Zener diodes include the following: 

They may be expensive. 

They are not bilateral. To protect against both polarities, two diodes 

in series back-t~back configuration are necessary. 

Diodes have relatively high-junction capacitance; therefore, they 

may cause significant signal loss at operating frequencies above 1 

MHz. Special diode assemblies may extend the useful frequency to 

approximately 50 MHz. 

They do not switch state between a conducting and a nonconducting 

mode. The voltage across the diode does not switch to a low value 

when conducting but remains at the Zener voltage. This characteristic accounts for their ability to cease conduction when the voltage 

falls below the Zener level, but it has a disadvantage thermally. 

During conduction, the power absorbed by the diode is the product 

of the current through the diode and the voltage across the diode. 

The power absorbed for constant current, thus, is directly proportional to the diode voltage. 

Partially offsetting this disadvantage, however, is the phenomenon that surge energy absorbed in the diode is energy that cannot 

be reflected back into the system to cause trouble elsewhere. 

They provide lower energy capabilities than do spark gaps. Since the 

Zener action takes place across a narrow P-N junction, the mass of 

the protecting junction is small and hence cannot store much energy. 

As a result, diode networks cannot be used where extremely high 

transient current or energy is predicted. For most hardening 

applications, this is not a serious limitation, since the induced 

surge-current levels are in the I to 100 A range at those locations 

where Zener diodes are most likely to be used. 

They are not available for voltage below about 5 V. 

They are not normally available for voltages above a few hundred 

volts. 

15.6.3 Forward-Conducting Diodes 

In a forwardconducting state, a diode conducts little current below about 

0.3 V for germanium and 0.6 V for silicon. They can, as a result, be placed 

directly across a low-voltage line and afford substantial protection. 

The advantages of forward-conducting diodes include the following: 

They are of small size. 

They are not costly. 

They provide protection at very low-voltage levels. 

* They have excellent surge-current ratings. 
The disadvantages of forward-conducting diodes include the following: 

They are not bilateral. For protection of both polarities, two diodes 

in parallel must be used. Some vendors supply dipolar diodes for 

protection purposes. 

Conduction may occur on normal signals with attendant signalclipping and frequency-multiplication effects. Diodes must be used 

in series to raise voltage levels. 

They have relatively high capacitance. 

15.6.4 Nonlinear Resistors 

This category includes nonlinear resistors (varistors) that may be characterized by the expression 

I = KV~ 

where Nand K are device constants dependent on the varistor material. 

Varistors may be constructed of silicon carbide, selenium, or metal oxide. 

This section will concentrate on the metal oxide varistor (Reference 15.1). 

The advantages of metal oxide-based nonlinear resistors include the 

following: 

They are bilateral devices. 

They are of small size. 

They are easily mounted. One common configuration is very similar 

in appearance to a disk ceramic capacitor. 

They are self-extinguishing. When applied voltage drops below the 

voltage for which the device is rated, they conduct very little 

current. 

They have an inherently fast responsellow-impulse ratio. 

They provide high power-handling capability. The current- and 

energy-handling capability is second only to certain types of spark 

gaps. This device gives a higher ratio of energy absorbed to energy 

reflected than conventional gaps give. Morever, the energy is 

absorbed throughout the bulk of the material and is not concentrated in a narrow P-N junction. 

The disadvantages of nonlinear resistors include the following: 

They have low impedance and high capacitance. The zinc oxide 

varistor is characterized by high-shunt capacity, limiting its use to 

frequencies below 1 MHz. While it has a much higher idling current 

than have either gaps or diodes, the standby power dissipation is in 

the milliwatt region and is usually not a significant limitation. 

They are not suitable for operating voltages below 20 to 30 V. 

Operating volt~ge is proportional to material thickness. Good surge 

protection of low-voltage circuits would require an impractically 

thin piece of material. 
15.6.5 Surge Protecting Connectors 

At least two lines of activity have been pursued with the aim of 

incorporating surge protective devices directly in electrical connections. One 

approach (Reference 15.2) incorporated spark gaps between the connector pins 

and the connector shell. A high dielectric material, rutile, was used to stimulate 

internal flashover and to achieve a protective level of about 1100 to 1300 V with 

an impulse ratio nearly unity, even with voltage rates of change of 10 kV/p. 

This design withstood extremely large currents (100 to 250 kA). The work on 

this device was performed at Sandia Laboratories and supported by the U.S. 

Atomic Energy Commission. 

Another line of approach uses zinc oxide varistor material to hold the pins 

in a connector. This work is being done by General Electric Company for the 

U.S. Army Harry Diamond Laboratories. Neither of these protective connector 

designs is yet available commercially. 

15.7 Improvement Through Transient Coordination 

Transient coordination is a concept which, when reduced to its simplest 

terms, implies that targets relative to transients will be assigned both to those 

that design electronic equipment ("black boxes") and those that design wiring to 

interconnect those black boxes. The task that is to be assigned the designers of 

black boxes is to produce equipment that will be able to withstand transients on 

all of the input and output wiring. The targets that form a part of the task will 

be specifications describing the type of transients that the equipment must 

withstand and to which it will be subjected as part of an acceptance or proof 

test. The task that is to be assigned to the groups designing interconnecting 

wiring is that no external threat, such as lightning or switching of inductive 

devices, shall produce on the wiring transients larger than those which the black 

box was designed to withstand. The target numbers to which the wiring designer 

must work will be maximum amplitudes of current, voltage, and surge energy. 

The assignment of such tasks implies that there must be a referee who 

assigns the appropriate target numbers and oversees the work to ensure that both 

parties fulfill the tasks assigned. This referee may be called the system integrator. 

The transient coordinator philosophy is illustrated in Figure 15.8. The aims of 

transient coordination would be the following: 

1. Insure that the actual transient level produced by lightning (or any 

other source of transient) will be less than that associated with the 

transient control level number assigned to the cable designer. The 

cable designer's job would be to analyze the electromagnetic threat 

that lightning would present and to use whatever techniques of 

circuit routing or shielding would be necessary to ensure that the 

actual transients produced by lightning did not exceed the values 

specified for that particular type of circuit. 

2. The transient design level controlling the type of circuit or circuit 

protection techniques used, and assigned to the avionics designer, 
would be higher than the transient control level by a margin 

reflecting how important it was that lightning did not in fact 

interfere with the piece of equipment under design. A margin is 

necessary because any single lightning flash might produce an actual 

transient level higher than the assigned transient control level, which 

would have been derived for a predicted average in spite of the cabIe 

designer's good intentions. Prediction of actual transient levels is an 

imperfect art. 

The job of the avionics designer wouId be to ensure that the 

vulnerability and susceptibility levels of the equipment that he is to 

supply would be higher than the assigned transient design level. The 

vulnerability level is that level of transient which, if applied to the 

input or output circuit under question, would cause the equipment 

to be permanently damaged. The susceptibility level is defined as 

AVIONICS DESIGNER 

VULNERABILITY B I 

SUSCEPTIBILITY L 1 

CABLE DESIGNER 

THE CABLE DESIGNER'S 

JOB IS TO ENSURE THAT 

ACTUAL TRANSIENTS DO 

NOT EXCEED A WORST 

CASE TRANSIENT CONTROL LEVEL 

TRANSIENT DESIGN LLEL 

TDL 

I 

(TRANSIENT CONTROL THE JOB OF THE AVIONICS 

LEVEL) - TCL DESIGNER IS TO ENSURE 

THAT HIS VULNERABILITY. OR 

SUSCEPTIBILITY LEVEL, IS 

GREATEK THAN THE 

TRANSIENT DESIGN LEVEL 

b 1 

ACTUAL TRANSIENT 

LEVEL 

Figure 15.8 The transient coordination philosophy. 

SYSTEM INTEGRATOR 
that level of transient that would result in interference with or 

malfunction of the equipment. The vulnerability level by definition, 

then, would have to be at least as high as the susceptibility level. 

There are several ways in which the levels might be set. In the first, the 

system integrator would set the desired transient level, then set the required 

margin, which in turn would set the transient control level. Whatever the 

rationale by which the system integrator sets the transient design level, that level 

would become a part of the purchase specifications and would, presumably, not 

be subject to variation by the vendor of the avionics. As an alternative, the 

avionics designer might determine by suitable testing the vulnerability and 

susceptibility levels of his equipment and provide a guarantee as to the level of 

transients that his equipment could withstand. That level would then be the 

transient design level. After the system integrator had set the desired safety 

margin, the appropriate transient control level for the cable designer would have 

been established, One approach to the setting of margins appears in the Space 

Shuttle Lightning Criteria Documei~t (Reference 15.3). 

The numbers that would be assigned to the transient design level probably 

should be expressed in terms of the maximum voltage appropriate to a 

high-impedence circuit (open circuit voltage) or the maximum current appropriate to a low-impedance circuit (short circuit current). In order for the 

transient coordination philosophy to have most impact, there should be a 

limited number of levels. One set of levels that has been proposed (Reference 

15.4) is shown in Table 15.1. With each level there is associated an open circuit 

voltage and a short circuit current; the two are related by a standard 

transient-source impedance, shown in Figure 15.9 (Reference 15.5) 

Figure 15.9 Short circuit current (ISC) resulting from a transient source with 

VOC open circuit voltage and 50n/50pHz source impedance. 
Table 15.1 PROPOSED TRANSIENT CONTROL LEVELS 

- - 

Proposed Transient Open Circuit Short Circuit 

Control Level Voltage Level Current Level 

Number (volts) (amperes) 

An alternative set of levels for which some voltages are numerically equal 

to the voltages in existing specifications is presented in Table 15.2. 

Table 15.2 ALTERNATIVE TRANSIENT CONTROL LEVELS 

Proposed Transient Open Circuit Short Circuit 

Control Level Voltage Level Current Level 

Number (volts) (amperes) 
REFERENCES 

15.1 Transient Voltage Suppression Manual, Semiconductor Products Department, General Electric Company, Electronics Park, Syracuse, New York 

(1976). 

15.2 J. A. Cooper and L. J. Allen, The Lightning Arrester Connector: A New 

Concept in System Electrical Protection, IEEE International Electromagnetic Compatibility Symposium Record, IEEE 72CH0638-EMC, Institute 

of Electronic and Electrical Engineers, New York, New York (1972). 

15.3 Space Shuttle Program Lightning Protection Criteria Document, 

JSC-07636, Revision .A, National Aeronautics and Space Administration, 

Lyndon B. Johnson Space Center, Houston, Texas (November 4, 1975), 

pp. D-2 to D-5. 

15.4 F. A. Fisher and F. D. Martzloff, "Transient Control Levels: A Proposal 

for Insulation Coordination in Low-Voltage Systems," IEEE Transactions 

on Power Apparatus and Systems, PAS-95, 1, Institute of Electronic and 

Electrical Engineers, New York, New York (JanuaryIFebruary 1976): 

120-29. 

15.5 Fisher and Martzloff, "Transient Control Levels," p. 128. 
CHAPTER 16 

COMPONENT DAMAGE ANALYSIS 

16.1 Introduction 

The energy coupled into a system by lightning induces large current pulses 

into the cabling and wiring of the system and resulting voltage pulses across 

loads. Determining whether or not these voltages and currents cause upset or 

damage to active or passive components requires a knowledge of the failure 

thresholds of devices. These failure thresholds were investigated extensively 

during studies of nuclear electromagnetic pulse (NEMP) effects on electronic 

equipment. 

Assessment of the vulnerability of a system to a transient environment 

leads to the observation that failure of the system to function properly may 

result from a degradation in performance of a single active or passive component 

device when it is subjected to a transient voltage pulse. Systems vulnerability 

evaluation therefore requires that one know the failure thresholds for devices 

subjected to transients. Simulation of failure resulting from transients at the 

component level can be conveniently performed in the laboratory by means of 

high-voltage pulse generators. Information from such simulation can then be 

correlated with the original threat (EMP or lightning) by analytical techniques. 

Semiconductor electronic components are generally more vulnerable under pulse 

conditions than are nonsemiconductor components; therefore, most of the 

discussion of this section will be devoted to semiconductors. 

16.2 Theoretical Models 

The number of 'semiconductor devices that have been experimentally 

tested (mostly diodes and transistors) is only a small sample of the total number 

of devices available. However, the scope of the experimental results has been 

extended by development of theoretical models which relate significant changes 

in the properties of a semiconductor PN junction to high temperatures generated 

in the junction region during application of a high-voltage pulse. Theoretical 

models based on thermal analysis of the junction region yield a mathematical 

relation between junction temperature and power dissipation in the junction 

region. This re!ation can be used to define a constant characterizing the 

operation in a given time domain of each device. 

16.3 Empirical Models 

In addition to the theoretical correlations just described, the experimental 

data have been used in the development of empirical relations which are 

obtained from two models of semiconductor junction devices--the junction 

capacitance model and the thermal resistance model. These models provide a 

framework from which the power failure threshold of an untested device can be 
estimated from the quantities listed in a data sheet description prepared by 

manufacturers for a diode or a transistor. 

16.4 Limitations 

Certain limits to the accuracy and application of the theoretical and 

empirical models exist and must be noted in estimating component device 

vulnerability. The models have been verified only for diodes and bipolar 

transistors. Other devices, such as FETS and unijunction transistors, have been 

tested in insufficient numbers for conclusions to be drawn as to their 

conformance with the model. Also, the results and models which are presented 

apply strictly only to an isolated device--that is, not to a device in a circuit. In 

the case of a multiple-terminal device, they apply only to the two terminals 

connected for test, with any other terminals open. 

The assumptions made about junction heating and transfer of heat in the 

derivation of the models limit their applicability to the region of pulse durations 

of approximately 0.1 to 20 ps. For longer times appreciable heat transfer may 

take place away from the junction area during the pulse input. For short pulses 

the power levels are so high (1 to 10 kW) that very large currents flow; 

consequently, the joule heating in the bulk material is appreciable. The 

transition behavior between these three regions of pulse duration (regions that 

will be more preciseIy described later in this section) is not well defined and may 

vary from one device type to another. Examination of available data indicates 

the transition region generally occurs between 100 ns and 1 MS. 

Still another limitation is fundamental to the work summarized in this 

section: it applies only to junctjon burnout. Other modes of device failure, such 

as metallization burnout and internal arcing, are not treated. Based on the results 

obtained in studies of junction burnout, it would seem that other effects, such 

as metallization burnout, occur at higher power input levels than those input 

levels sufficient to damage the junction. 

16.5 Failure Mechanisms--Semiconductors 

The two principal breakdown modes for semiconductor PN junctions are 

the following: 

I. Surface damage around the junction as a result of arcing 

2. Internal damage to the junction region as a result of elevated 

temperatures. 

Surfuce damage refers to the establishment of a high-leakage path around the 

junction which effectively eliminates any junction action. The junction itself is 

not necessarily destroyed, since, if it were possible to etch the conducting 

material away from the surface, the device might be able to return to its normal 

operating state. This is not practical, of course, in an operational semiconductor. 

It is equally likely that the formation of any surface leakage path would be the 

result of excessive heat formation in the bulk of the material, and this would 

typically be an irreversible phenomenon. It is very difficult to predict 
theoretically the conditions which will lead to surface damage because they 

depend upon many variables, such as the geometrical design and the details of 

the crystal structure of the surface. The theoretical prediction of surface arcing 

under pulse conditions is not practical (Reference 16.1). It should be cautioned 

that surface damage, in the general case, may occur in devices at power levels 

which are orders of magnitude below those sustainable by devices in which bulk 

damage occurs (Reference 16.2). 

Bulk damage, which results in a permanent change in the characteristic 

electrical parameters of the junction, indicates some physical change in the 

structure of the semiconductor crystal in the region of the junction. The most 

significant change is melting of the junction as a result of high temperatures. 

Other types of change may involve the impurity concentrations, the formation 

of alloys of the crystal materials, or a large increase in the number of lattice 

imperfections, either crystal dislocations or point defects. 

The simplest structure to analyze is a diode. Consider Figure 16.1. In this 

diode a current is assumed to flow as a result of some outside stimulus and, in 

doing so, to produce a voltage across the diode. This voltage may be the forward 

bias voltage (0.5 to 1.5 V), or it may be the reverse breakdown voltage if the 

outside stimulus has biased the diode in the reverse direction. Assume that I is a 

square wave and that V is not a function of time, as it might be if V depended 

upon the junction temperature. The instantaneous power dissipated in the diode 

is then also a square pulse of magnitude 

and the total energy produced is 

I 

Assume that the power level is sufficient that the device fails at the end of the 

pulse. Both experimental and theoretical analyses indicate that the power (or 

energy) required to cause failure depends on pulse width: the narrower the 

pulse, the greater the power required to cause failure. Over a broad range of times, 

typically between 0.1 ps and 100 ps, the power required to cause failure is 

inversely proportional to the square root of time. For very short pulse durations 

the power required to cause failure is inversely proportional to the first power of 

time, and for very long pulse durations the power required to cause failure is a 

constant threat. These relations may be expressed by the following equations: 

P = Constant t > 100 ps (16.5) 

where To generally lies between 10 ns and 1 ps. 

Figure 16.2 (Reference 16.3) shows an example for a 10 W diode. The most 
Figure 16.1 Voltage and current through a diode junction. 

(a) Forward bias case 

(b) Reverse bias case, Vg > VBD 

important region is the center region, where 

Junctions are less susceptible to burnout when operated with forward bias: 

first, because the power produced by a given current is lower when flowing 

through the low forward bias voltage than when developed across the higher 

reverse bias breakdown voltage, VBD, and, second, because the current is more 

uniformly distributed across the junction in the forward direction. Accordingly, 

it requires more power in the forward direction to cause failure. An example is 

shown in Figure 16.3 (Reference 16.4). The 2N2222 transistor is a 0.5 W NPN 

silicon high-speed switch. 

454 
Figure 16.2 Expected time dependence of pulse power failure: threshold for an 

example I0 W diode. 

10 oo0 

o - No Failure 

- 

,. loo0 

e 

E 

5 

2 

LY 

an 2 loo 

10 

.O1 0.1 1 10 100 

TIME (Microseconds) 

Figure 16.3 Experimental data points for failure of the base-emitter: junction 

of a 2N-2222 transistor for forward and reverse polarity voltage pulses. 

16.6 Damage Constants 

From curves, such as those of Figure 16.3, the value of a damage constant, 

K (or C), can be determined and tabulated for a range of devices. The 

appropriate threshold damage curve can then be reproduced as desired from a 

single damage characterization number, the damage constant K. It is convenient 
to express K in kW . psM, since the value of K in these units then becomes 

numerically equal to the power necessary for failure, dissipated by a square pulse 

of 1 ps duration. If this point is located on a log-log graph (Figure 16.2 or Figure 

16.31, then a curve of slope, -1/2, drawn through this point reconstructs the 

curve fit to the data for a particular device, and the power for failure at other 

pulse durations can be read directly from such a graph. Ideally, the K factor 

should be known for both the forward bias and the reverse bias conditions. Generally, only the K factor for the reverse bias condition is known. This limitation 

gives conservative answers, since K for the reverse bias condition is almost always 

lower than is K for the forward bias condition. 

The magnitude of the damage constant depends upon the type of junction 

under consideration: broadly speaking, it is larger for large junctions and smaller 

for small junctions. Figures 16.4 and 16.5 (Reference 16.5) show the range of 

the damage constant for typical diodes and transistors. 

16.6.1 Experimental Determination of K Factor 

Experimentally the K factor is determined by injecting power pulses into 

the semiconductor junction, starting at low levels and increasing the levels until 

either failure or significant degradation of the junction occurs. Devices would 

normally be pulsed in both the forward and reverse directions. 

The K factors and breakdown voltages for a number of semiconductors are 

given in Table 16.1 and 16.2 (Reference 16.6). In the case of transistors, the K 

factor listed generally refers to the base-emitter junction, since this is generally 

the junction most susceptible to burnout. In all the cases the K factor refers to 

the reverse bias direction. 

16.6.2 K Factor as Determined from Junction Area 

If the K factor is not measured, it may be estimated by one or more of 

three methods. The most accurate of the indirect methods involves a knowledge 

of the area of the junction. If the area is known, the K factor may be estimated 

from the following relations: 

Diodes - K = 0.56 A 

Transistors - K = 0.47 A 

K in kW psM 

For transistors, the junction area to be used is that of the base-emitter region. 

This is generally the weaker junction (lower breakdown voltage), and it is that 

for which the experimental average value for K(A) was obtained. 

This method is of course limited by the availability of information on 

junction area, but where such information is available, the method yields damage 
Rectifier Diodes 

Reference Diodes 

Point Contact Diodes 

Microwave Diodes 

0.001 0.01 0.1 1 10 100 

Damage Constant, K 

Figure 16.4 Range of pulse power damage constants for representative diodes. 

High Power Transistors 

Germanium Transistors 

Switching Transistors 

Low Power Transistors 

0.01 0.1 1 10 100 

Damage Constant, K 

Figure 16.5 Range of pulse power damage constants for representative 

transistors. 

457 
constants accurate to within a factor of two. For planar devices, the junction 

area can often be measured directly on the silicon chip. 

16.6.3 K Factor as Determined from Junction Capacitance 

The next most reliable method of detertiiining the damage constant is 

from a knowledge of the capacitance (Cj) and breakdown voltage (VBD) of the 

junction. For three different categories of devices, the relations are 

Category 1 - insufficient data 

Category 2 - K = 4.97 x 1 oJ Cj VBD 0.57 

Category 3 - K = 1.66 x lo4 Cj VBD 0.992 

The different categories are 

Category 1 - Germanium diodes and germanium transistors 

Category 2 - Silicon diodes, all silicon transistor structures except planar 

and mesa 

Category 3 - Silicon planar and mesa transistors 

If the junction is a transistor base-emitter junction, the capacitance used should 

be taken at a reverse bias of approximately 1 V. If it is a collector-base junction 

or a diode junction, the value should be taken at the reverse bias of 

approximately 5 to 10 V. 

16.6.4 K Factor as Determined from Thermal Resistance 

A third, but the least reliable, way of estimating the damage constant is 

from a knowledge of the thermal resistance of the junction, either the thermal 

resistance from junction to case (Bjc) or from junction to ambient (Bja). The 

empirical expressions for each category for each of the thermal resistances are 

Category 1 - insufficient data for an accurate curve fit 

Category 2 - K = 707 Bjc (Bjc > 10.0) (16.12) 

Kz4.11 x lo4 Oja (16.13) 

Category 3 - K = 707 8jc (Bjc > 10.0) (16.14) 

K = 2.74 x 10' Oja '2.55 (16.15) 

Normally Bjc and Bja are not given in the transistor data sheets but, rather, must 

be calculated from the maximum operating junction temperature, (Tj[max]), 

the total power dissipation (Pd), case temperature (T,), and ambient temperature (Tamb). 
Generally, at least one of these thermal resistances may be determined from the 

manufacturer's data sheet. 

The accuracy of the damage constant as determined from either the 

junction capacitance or the junction thermal resistance is somewhat limited. 

Table 16.3 (Reference 16.8) gives some estimate of the accuracy within which 

the damage constant can be calculated. 

16.6.5 Oscillatory Waveforms 

Equation 16.6 is based on the assumption that the applied voltage, 

current, and power waves are rectangular in shape. While actual transients are 

very seldom rectangular, it is possible to derive equivalent rectangular pulses for 

more common types of transients. One such type of transient typically 

encountered is the damped oscillatory wave. Based on multiple pulse studies by 

Wunsch and others (References 16.9 and 16. lo), it can be assumed that device 

damage wilI occur, if at all, during the first half cycle of a damped sine wave. 

Therefore, the lower amplitude cycles may be neglected. Two cases may be 

considered, one in which one of the half cycles of the transient does not exceed 

the reverse breakdown voltage of the junction, Figure 16.6 (Reference 16.1 I), 

and one in which the reverse breakdown voltage is exceeded, Figure 16.7 

(Reference 16.12). In either case, one of the half cycles will bias the junction in 

a forward direction. Treating first the case in which the reverse breakdown 

voltage is not exceeded (Figure 16.6), the rectangular wave of the same peak 

amplitude, V,, and producing the same probability of damage as the sine wave, 

has a duration rP 

where 

of the sine wave. If the transient does exceed the reverse breakdown voltage, the 

duration of the equivalent transient depends upon the fraction of the time that 

the oscillatory transient does exceed the reverse breakdown voltage. The 

duration of the equivalent rectangular wave is given by the expression 

I ("e) )' 

Tp = Ts 

. cos -I (+) 

This expression is shown plotted in Figures 16.8 (Reference 16.13) and 16.9 

(Reference 16.1 4). For oscillatory transients whose initial amplitude considerably exceeds the reverse breakdown voltage of the junction, Equation 16.19 
Figrtre 16.6 Device waveforms for Vg < VBD, 

(a) Generator voltage 

(b) Diode voltage 

(c) Junction current 

(d) Junction power 
0 

yy 111 

1'1 

'I Ill 

Figure 16.7 Device waveforms for Vg > VBD. 

(a) Generator voltage 

(b) Diode voltage 

(c) Junction current 

(d) Junction power 
0.24 

0.22 

0.20 

0.18 

0.16 

0.14 

T + 0.12 

0.10 

0.08 

0.06 

0.04 

0.02 

0 

- 

VBD 

Figure 16.8 Plot of 7 17 versus Vo/VBD. P s 

approaches a limiting value of 0.2, and thus becomes identical with the forward 

bias case, Equation 16.1 8. 

A limited amount of data relating voltage and current durations to the 

breakdown of integrated circuits is available. Figures 16.10 (Reference 16.15), 

16.1 1 (Reference 16.16), and 16.12 (Reference 16.1 7) show the results of 

measurements on SN55 107 line receivers, SN55 109 line drivers, and CD4050 AE 

hex buffers. 

VO /VBD 

Figure 16.9 Plot of r P 17 s versus Vo/VBD for values of V,/VBD less than 2. 
CURRENT 1 

PULSE DURATION - MICROSECONDS 

Figure 16.1 0 Damage thresholds of SN 55 107 line receivers. 

16.7 Failure Mechanisms--Capacitors 

Capacitors fail by a mechanism different from that of semiconductors. The 

mechanism of capacitor failure depends upon the type of dielectric. Capacitors 

with solid dielectrics--paper, Mylar, or ceramics--will, when subjected to 

nonrepetitive transients, either fail by puncture of the dielectric or not fail at all. 

Typically, a capacitor can withstand short-duration transient voltages several 

times greater than the dc rating of the insulation. The pulse-breakdown rating of 

the dielectric, however, is not a constant ratio to the dc voltage rating, nor is it 

normally part of any manufacturer's specification. Accordingly, it is safest to 

consider that such a capacitor is in danger of failure if the pulse voltage exceeds 

the dc rating of the capacitor. 

Electrolytic capacitors, on the other hand, do not experience abrupt 

failure when exposed to short-duration transients. If the voltage across the 

capacitor exceeds the voltage used to form the dielectric film, the dielectric film 

begins to conduct. After the pulse has disappeared, the dielectric returns to 

nearly its normal state. During the transient period the dielectric film can carry 

substantial transient current without permanent or catastrophic degradation. 

Transients, however, may lead to increased leakage currents. An example of data 

that is available relates to a series of tests made on tantalum eIectrolytic 

capacitors of value 0.47 pF, 0.047 pF, and 0.0047 pF with a dc voltage rating of 

350 V (Reference 16.18). The data indicate that failure (defined as a substantial 

increase in the leakage current at voltages of less than 350 V) can generally be 
associated with the time during which internal conduction occurs. For these 

components, conduction was initiated at 3 to 4 times the voltage rating. Leakage 

current increased continuously with time of conduction, from initial values of a 

few nanoamperes through milliamperes. The value of the capacitance determines 

how quickly the voltage across the capacitor reaches the breakdown voltage 

range, 90 to 140 V, which then relates to the time of conduction and the extent 

of damage. Figure 16.13 (Reference 16.1 9) shows the data for the nine 0.0047 

pF capacitors tested. For a particular pulse duration of 5 fils, an increase in 

leakage current is expected; for pulse voltages of 100 to 150 V and for pulses of 

150 to 900 V, an increase in leakage current to milliamperes is possible. It is 

cautioned that this behavior may not be readily extended to capacitors of 

different materials or construction. The capacitor data remains insufficient to 

draw general conclusions as to system ilr~plications at this time. 

7 I I I \Il" - 

- 

- 

- 

- 

16.8 Failure Mechanisms--0 ther Components 

A limited amount of pulse test data is available for various nonsemicon- 

- 

- 

b MAX - 

- 

- - 

- VOLTAGE - 

- - 

- - 

- - 

I MIN 

- - 

- - - - 

- - 

- - 

- - 

- - 

- - 

MAX 

- CURRENT - ' MIN 

I I I 11111 I 1 IIIIII I I l lllll. 

0.01 0.1 1 10 

PULSE DURATION - MICROSECONDS 

Figure 16.1 I Damage thresholds of SN 5 5 109 line drivers. 

I 11~llIl I I 111111- - - 

- 

- 

- 
PULSE DURATION - MICROSECONDS 

Figure 16.12 Damage thresholds of CD 4050 AE hex buffers. 

ductor electronic circuit components. These data were mostly obtained by 

testing with a square-wave pulse input of 1 to 10 ps duration and up to 1 kV 

peak. As would be expected, not all such components are invulnerable to pulses 

of this shape. The test results for several kinds of components are presented in 

Table 16.4 (Reference 16.20). The test conditions consisted of an 8 ps, 1 kV 

pulse applied 10 times to each device. In the case of multiterminal components, 

several pairs of terminals were tested in this manner. 

16.9 Examples of Use of Damage Constants 

Some examples of how the preceding material may be used to determine 

whether or not a given transient will cause damage to semiconductors follow. 

The first circuit chosen for analysis (Figure 16.14) (Reference 16.2 1) is a 

simple remote-controlled relay. Across the terminals of the relay coil there is a 

diode which would be exposed to the same transients as those to which the coil 

is exposed. The analysis approach that will be taken is first to calculate the 

current level that would cause the diode to fail and then to see whether or not 
Pulse Height (V) 

Figure 16.13 Representation of pulse test data for Sprague 0.0047 pF tantalum 

electrolytic capacitors. 

r 1 RELAY 

Figure 16.14 Simple remote-controlled relay. 

the transient voltage source could supply that current. It will be assumed that 

the oscillatory pulse is a transient of 1 MHz frequency or 1 ps period. At this 

frequency the inductive reactance of the relay coil would be sufficiently large 

that the relay could be neglected. The current required to cause failure at time, 

t, would be 

~t-% Pf =- IF = - (16.20) 

VBD VBD 
For a 1N540 diode, the reverse breakdown voltage, VBD, is 400 V and the 

damage constant, K, is 0.93 (See Table 16.1). If a 200 ns pulse is used to 

approximate at 1 MHz damped sine wave, the failure current for the diode 

would be 

I~=5.2 A 

Assume now that the impedance of the source from which the voltage transient 

generated is 10 52. The voltage required to produce a current of 5.2 A through 

the diode would be 

- R V~ransient - 'BD + '~ransient Source 

Therefore, a 452 V pulse, 200 ns wide or a 1 MHz damped sine wave having a 

peak amplitude of 452 V would cause the diode to fail. 

The second circuit chosen for analysis is the simple phase-splitter amplifier 

shown in Figure 16.1 5 (Reference 16.22). The first step in determining the input 

current required for damage is to simplify the circuit. Again assume that the 

voltage source producing the transient is a damped sine wave of 1 MHz 

frequency. At such a frequency the reactances of capacitors C1 and C2 will be so 

small that they may be neglected. Lkewise the 12 V power supply line can be 

considered to be at ac ground potential. The resultant circuit after simplification 

is shown in Figure 16.16 (Reference 16.23). The circuit can be further simplified 

by determining the equivalent resistances for the base and for collector circuits. 

The base-emitter junction and the base-collector junction can also be replaced by 

their diode equivalents to represent operation in the breakdown regions. This 

simplified circuit is shown in Figure 16.17 (Reference 16.24). Also shown in 

Figure 16.1 7 are the breakdown voltages and damage constants for the 2N706B. 

Note that for this transistor a damage constant for the collector-base junction is 

available, though not listed in Table 16.2. 

The circuit is now simplified to the point where it lends itself easily to 

hand analysis. The next step is to determine which junction will fail and what 

the failure mode is. The passive components are generally able to withstand 

higher energies for short-duration pulses than can transistors. Therefore, the 

transistor is the element to consider for damage. Failure is also assumed to occur 

in the reverse biased direction. 

Using the Wunsch damage model (P = Kt-%), a calculation is made to see 

whether the emitter-base junction or the collector-base junction would fail first. 
I I 

L ----- I 

Figure 16.15 Phase-splitter circuit. 

Figure 16.1 6 Simplified phase-splitter circuit. 

This calculation shows that the emitter-base junction is the more susceptible. 

The current required to fail the emitter-base junction would be 

IjF = !h = 3.4 A 

V BD 

The voltage from the base to ground is 
Figure 16.1 7 Further simplification of phase-splitter circuit. 

The current through the collector-base junction is 

The power dissipated in the collector-base junction is 

which is below its failure-threshold power. 

The total current into the circuit is then 

'Transient = 'j~ + 'CB +kE=9.4 A 

R~~ I 

and the IT^^^^^^^^ voltage required to cause failure is 

Therefore (assuming a 100 S1 source impedance) a 2.5 kV pulse, 200 ns 

wide, will cause the transistor to fail. 
Table 16.1 DIODE DAMAGE DATA 

I N8 1 

I N82A 

I N9 1 

IN128 

IN191 

IN198 

IN248A 

l N249 

IN2490 

IN250 

IN2506 

IN251 

IN253 

IN254 

IN260 

,003 

.0007 

-0055 

.005 

.005 

.024 

40. 

40. 

40. 

40. 

80. 

.03 

86. 

3.5 

,0027 

10. 

5. 

100. 

40. 

90. 

80. 

50. 

100. 

100. 

200. 

200. 

40. 

95. 

190. 

30. 

SP 

DX 

CM 

CM 

S P 

S P 

SfJ 

SP 

SP 

S P 

SP, DE 

S P 

DX 

SP 

CM 
Table 16.1 DIODE DAMAGE DATA (CONT.) 

Device Number 

IN270 

IN276 

l N277 

1N295, A 

IN320 

IN332 

1 N333 

IN335 

IN337 

l N338 

I N34 1 

IN342 

IN346 

l N429 

1 N457 

1 N458 

l N459 

I N459A 

I N46 1 

l N462 

l N466 

IN467 

IN468 

I N470 

IN474A 

I N482A 

Reference 

CM 

CM 

DX 

S P 

S P 

SP 

S P 

S P 

S P 

DE 

SP 

S P 

S P 

DX 

DX 

S P 

DX 

DX 

S P 

S P 

SP 

S P 

SP 

SP 

CM 

DX 

K 

.022 

.0055 

.027 

.005 

1.2 

3.5 

1.5 

1.5 

1.5 

18.3 

3.5 

1.5 

1.5 

.6 

.12 

.5 

.59 

.96 

.24 

.05 

.78 

.78 

.78 

.78 

.219 

.96 

"BD 

100. 

100. 

125. 

40. 

500. 

400. 

400. 

300. 

200. 

100. 

400. 

400. 

200. 

6.2 

70. 

150. 

200. 

200. 

35. 

80. 

3.5 

4.1 

4.9 

7.1 

5.8 

36. 
- - 

Table 16.1 DIODE DAMAGE DATA (CONT.) 

- 

Device Number 

lN483, A 

IN483, B 

IN484A 

IN4849 

1 N485 

IN486, B 

lN487, Z 

l N488 

l N536 

l N537 

lN538, M 

l N539 

IN540 

l N547 

I N560 

IN561 

l N562 

IN619 

IN622 

l N625 

I N625A 

I N643 

I N643A 

1 N645 

l N646 -- - -- 

K 

.3 

.3 

.45 

.3 

.3 

.29 

.3 

.3 

1. 

.51 

1. 

1. 

.93 

12.1 

.625 

.625 

1.8 

.36 

.347 

,164 

.045 

.44 

.1 

2.8 

2.29 - - - 

v~~ 

70. 

80. 

130. 

130. 

180. 

225. 

300. 

380. 

50. 

100. 

200. 

300. 

400. 

600. 

800. 

1000. 

800. 

10. 

150. 

30. 

20. 

200. 

200. 

225. 

300. - 

Reference 

SP 

S P 

DX 

SP 

CM 

SP 

SP 

SP 

DE 

DX 

S P 

S P 

DX 

DX 

CM 

CM 

S P 

SP 

CM 

CM 

CM 

S P 

DX 

S P 

DX 
Table 16.1 DIODE DAMAGE DATA (CONT.) 

Device Number 

l N647 

1 N648 

IN649 

1 N658 

I N660 

I N66 1 

l N662 

l N663 

l N676 

l N689 

IN691 

IN692 

IN702, A 

I N 703A 

IN704, A 

IN705, A 

IN706 

IN709, A 

IN710 

IN71 1A 

IN712 

IN714A 

IN715A 

IN718A 

lN719A 

IN721, A 

K 

2.8 

2.8 

2.9 

.92 

-44 

.41 

-29 

.44 

.27 

1.1 

.418 

.5 

1. 

1. 

1. 

.91 

,288 

.78 

.78 

2.1 

.78 

.78 

.78 

.1 

.1 

.35 

"BD 

400. 

500. 

600. 

120. 

100. 

200. 

100. 

100. 

100. 

600. 

80. 

100. 

2.6 

3.5 

4.1 

4.8 

5.8 

6.2 

6.8 

7.5 

8.2 

10. 

11. 

15. 

16. 

20. 

Reference 

S P 

SP 

DX 

DX 

S P 

DX 

SP 

S P 

SP 

S P 

S P 

S P 

SP, OX 

SP 

S P 

S P 

CM 

S P 

S P 

DX 

SP 

S P 

S P 

S P 

SP 

S P 
Table 16.1 DIODE DAMAGE DATA (CONT.) 

Device Number 

IN 725A 

1 N729 

1N746, A 

IN747, A 

I N748A 

IN749 

I N750A 

lN751, A 

1N752, A 

lN753, A 

lN754, A 

lN755, A 

1N756, A 

lN757, A 

1N758, A 

lN759, A 

IN761 

IN762 

IN763 

lN763-2 

I N766A 

IN767 

1 N 769A 

IN769-3 

IN816, W 

IN817 

K 

.349 

.06 

1.1 

1.1 

1.1 

1.1 

1.1 

1.1 

1.1 

1.2 

.63 

.63 

.63 

.63 

.63 

.63 

1.8 

1.8 

1.8 

3. 

1.8 

1.8 

1.8 

2. 

1.5 

.46 

"B 0 

30. 

43. 

3.3 

3.6 

3.9 

4.3 

4.7 

5.1 

5.6 

6.2 

6.8 

7.5 

8.2 

9.1 

10. 

12. 

4.9 

5.8 

7.1 

7.0 

12.8 

15.8 

23.5 

26. 

26. 

200. 

Reference 

CM 

S P 

S P 

S P 

SP 

S P 

S P 

SP 

SP 

SP, DX 

S P 

S P 

SP 

S P 

SP 

SP 

S P 

SP 

SP 

DX 

SP 

SP 

S P 

DX 

DX 

SP 
Table 16.1 DIODE DAMAGE DATA (CONT.) 
Table 16.1 DIODE DAMAGE DATA (CONT.) - - -- - 

IN1 126A 

IN1 184 

IN1199, A 

IN 1200 

IN1201 

IN1202 

IN1204A 

IN1206 

IN1217 

IN12228 

IN1317A 

IN1319A 

INi342A 

IN1348A 

IN 1367 

14. 

31.5 

15. 

62.32 

62.32 

21. 

46.106 

62.32 

5.8 

2.563 

.19 

.19 

38.4 

1.827 

34. 

500. 

100. 

SO. 

100. 

150. 

200. 

400. 

600. 

400. 

19. 

28. 

100. 

200. 

47. 

SP 

CM 

S P 

CM 

CM 

S P 

CM 

CM 

S P 

CM 

S P 

SP 

DE 

CM 

SP 
Table 16.1 DIODE DAMAGE DATA (CONT.) 

i 

Device Number 

IN1583 

IN1585 

IN1614 

IN1615 

IN1693 

IN1695 

IN1731 

IN1733A 

I N 1 770A 

IN 1773A 

IN 1780A 

IN1783 

IN1818RA 

1N1823C, A 

IN1828C 

IN1834 

IN1835A 

IN1836C 

IN1904 

IN1909 

1 N2037 

IN2154 

IN2158 

IN2164 

l N2483 

IN2610 

K 

11.391 

3.5 

.38 

.666 

3.2 

3.2 

3.2 

11.3 

14.2 

1.9 

1.9 

21.3 

4.3 

4.3 

4.3 

33.8 

4.3 

4.3 

28. 

6.8 

.05 

20. 

21.5 

2.3 

2. i 

2.6 

"BD 

200. 

400. 

200. 

480. 

200. 

400. 

1500. 

3000. 

9.1 

12. 

24. 

33. 

16. 

27. 

43. 

75. 

82. 

91. 

100. 

200. 

12.8 

50. 

400. 

9.4 

400. 

100 

Reference 

C M 

D E 

S P 

CM 

S P 

S P 

CM 

DE 

D E 

S P 

S P 

D E 

S P 

S P 

SP 

CM 

S P 

S P 

SP 

SP 

S P 

S P 

CM 

S P 

SP 

SP 
- - 

Table 16.1 DIODE DAMAGE DATA (CONT.) 

Reference 

S P 

S P 

SP 

SP 

S P 

S P 

CM 

CM 

CM 

S P 

CM 

S P 

S P 

SP 

SP 

DX 

CM 

SP 

S P 

S P 

S P 

SP 

SP 

SP 

S P 

SP 

V~~ 

200. 

400. 

600. 

400. 

150. 

200. 

10. 

20. 

30. 

33. 

39. 

160. 

200. 

400. 

600. 

1. 

.74 

6.8 

12. 

15. 

20. 

22. 

24. 

25. 

27. 

30. 

Device Number 

IN261 1 

IN2613 

IN2615 

IN2789 

IN2795 

IN2796 

IN 2808 

IN2818 

IN28238 

l N2824 

IN2826B 

IN2844B 

IN28468 

IN2862 

1 N 2864 

I N2929A 

IN2930 

1 N2970B 

IN29768 

IN29790 

lN2984, B 

1N2985, 0, RB 

IN29860 

IN29870 

1 N2988B 

l N2989B 

K 

2.6 

2.6 

2.6 

40. 

40. 

40. 

249. 

249. 

249. 

156. 

249. 

15. 

15. 

2.8 

2.8 

.073 

,196 

15.0 

15. 

15. 

15. 

15. 

15. 

15. 

15. 

15. 
Table 16.1 DIODE DAMAGE DATA (CONT.) 

Device Number 

IN2991B 

lN2995, B 

I N 2997B 

IN3001B 

IN3008B 

lN3015B 

IN3016B 

IN301 78 

IN3019B 

IN30228 

IN30248 

IN30258 

IN30268 

IN30278 

lN3028, B 

I N3029B 

IN3030B 

IN3031B 

IN30338 

IN3035B 

IN30378 

IN3040B 

IN3041, B 

IN3051 B 

IN3064 

I N 3070 

K 

15. 

15. 

15. 

15. 

15. 

33.84 

19.5 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

1.9 

.02 

.365 

"BD 

36. 

47. 

51. 

68. 

120. 

200. 

6.8 

7.5 

9.1 

12. 

13. 

16. 

18. 

20. 

22. 

24. 

27. 

30. 

36. 

43. 

51. 

68. 

75. 

200. 

75. 

200. 

Reference 

S P 

SP 

S P 

SP 

S P 

CM 

DE 

S P 

SP 

SP 

S P 

S P 

SP 

S P 

SP 

SP 

S P 

SP 

SP 

SP 

S P 

SP 

SP 

SP 

SP 

CM 
-- 

Table 16.1 DIODE DAMAGE DATA (CONT.) 

Device Number 

IN3157 

IN3189 

IN3190 

l N3560 

IN3561 

IN3582A 

I N3600 

IN3821 

IN3828A 

IN3893 

1 N3976 

l N424 1 

l N4245 

IN4249 

IN4312 

IN4370A 

IN4816 

IN4817 

IN4820 

IN4823 

l N4989 

AM2 

04330 

FD300 

S G22 

SLDIOEC 

- 

P-P 

"BD 

8 -4 

200. 

600. 

.475 

,475 

11.7 

50. 

3.3 

6.2 

400. 

200. 

6. 

200. 

1000. 

150. 

2.4 

50. 

100. 

400. 

' OOV 

200. 

50. 

125. 

10,000. 

K 

-625 

10. 

4.1 

.038 

.038 

.35 

.18 

1.947 

1.95 

6.41 

132. 

33.84 

2.4 

2.4 

.I16 

.625 

6.8 

6.8 

10. 

.208 

14.33 

1.4 

.001 

.18 

.23 

f 

Reference 

CM 

S P 

CM 

CM 

CM 

SP 

S P 

CM 

CM 

CM 

CM 

CM 

S P 

SP 

CM 

CM 

DE 

D E 

D E 

CM 

CM 

S P 

SP 

SP 

SP 

CM d 
Table 16.1 DIODE DAMAGE DATA (CONCLUDED) 

SP - SAP-I Computer Listing from SAP-I Conlputer Code ilfut~ual (unpublished), U.S. Air Force Weapons Laboratory, 1972. 

DX - Experimental data from DASA (Defense Atomic Support Agency) 

Handbook (See Reference 16.3) 

DE - Estimated data from DASA Handbook (See Reference 16.3) 

CM - Computed data (Reference 16.7) 

Device Number 

SV1035 

SV2092 

SV2183 

TM7 

TM21 

TM27 

TM84 

TM 124 

UT 242 

K 

1.71 

2.6 

2.6 

20. 

18. 

20. 

11. 

11. 

2.6 

v~~ 

26. 

70. 

200. 

200. 

800. 

1200. 

200. 

Reference 

CM 

S P 

S P 

S P 

S P 

SP 

S P 

S P 

SP 
Table 16.2 TRANSISTOR DAMAGE DATA 

3evice 

Number 

2N43, A 

2N117 

2N118 

2N 128 

2N158 

2N176 

2N189 

2N 190 

2N243 

2N244 

2N263 

2N264 

2N274 

2N279A 

2 N 297A 

2N329,A 

2N332 

2N333 

2N335, A 

2N336 

2N337 

2N338 

2N339 

2N34 1 

K 

.28 

.15 

.15 

.017 

.499 

.46 

.17 

.58 

.05 

.05 

-38 

.36 

.0076 

.047 

.499 

.21 

.45 

.32 

.55 

.55 

.12 

.12 

2. 

1. 

BV 

E 60 

5. 

1. 

1. 

10. 

30. 

1. 

1. 

1. 

.5 

40. 

20. 

1. 

1. 

1. 

(4 .-2N335A) 

BV 

CBO 

45. 

45. 

45. 

10. 

60. 

40. 

25. 

25. 

60. 

60. 

45. 

45. 

35. 

45. 

60. 

50. 

45. 

45. 

45. 

BV 

CEO ----. 

30. 

45. 

45. 

4.5 

60. 

30. 

25. 

25. 

60. 

60. 

30. 

30. 

40. 

30. 

40. 

30. 

30. 

30. 

45. 

1. 

1. 

1. 

1. 

1. 

Reference 

Source 

SP 

SP 

SP 

SP 

CM 

CM 

SP 

OX 

S P 

S P 

S P 

SP 

CM 

CM 

CM 

SP 

S P 

S P 

SP 

30. 

30. 

30. 

55. 

85. 

45. 

45. 

45. 

55. 

125. 

S P 

S P 

SP 

SP 

SP 
Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 
Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 

BV 

C BO 

60. 

40. 

40. 

20. 

80. 

45. 

60. 

100. 

TOO. 

-- 

- - 

-- 

-- 

60. 

60. 

120. 

25. 

25. 

40. 

60. 

60. 

75. 

25. 

60. 

80. 

BV 

CEO 

60. 

20. 

30. 

15. 

60. 

30. 

60. 

100. 

100. 

50.(A-C) 

200. (A-C ) 

300.(A-C) 

800.(A-C) 

40. 

40. 

80. 

25. 

20. 

15. 

40. 

40. 

32. 

20. 

40. 

60. 

- 

Device 

Number 

2N551 

2N576A 

2N587 

2N595 

2N618 

2N652A 

2N656 

2N657 

2N657A 

2N682 

2N685 

2N687 

2N690 

2N696 

2N697 

2N699 

2 N 703 

2N7G6,B 

2 N 708 

2N717 

2N718 

2N718A 

2N726 

2N730 

2 N 736 

Reference 

Source 

S P 

DX 

S P 

CM 

DX 

CM 

DX 

DX 

DX 

CM 

DX 

DX 

CM 

CM 

SP 

DX 

SP 

DX 

DX 

SP 

S P 

S P 

CM 

CM 

DX J 

K 

1.6 

.023 

.14 

,012 

.88 

.I18 

.2 

-66 

1.07 

.33 

1.4 

11.7 

3.1 

1 .O 

.2 

.25 

.08 

-0075 

.03 

.13 

.13 

.35 

.021 

,165 

.1 

BV 

E BO 

6. 

15. 

40. 

40. 

30. 

8. 

8. 

8. 

- - 

-- 

-- 

-- 

5. 

5. 

5. 

5. 

3. 

(5.-2N706B 

5. 

5. 

5. 

7. 

5. 

5. 

5. 
Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 
-- 

Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 
Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 

Device 

Number 

2N 1602 

2N1613 

2N1615 

2N 1642 

2N1700 

2N1701 

2N1711 

2N1722 

2N1751 

2N 1753 

2N1772A 

2N1776A 

2N 177A 

2N1871A 

2N 1890 

2N 1893 

2N1916W 

2N2015 

2N2035 

2N2060 

2N2102 

2N2156 

2N2218A 

2N2219 

2N2219A 

K 

.40 

.27 

.553 

.13 

4.134 

4.5 

.36 

54.5 

1.05 

.039 

.651 

1.584 

C-G C-A 

2.0 .46 

1.1 

.27 

.4 

2.22 

26.462 

3.633 

.21 

.77 

.471 

.264 

.3 

.264 

BV 

E BO 

-- 

7. 

8. 

30. 

6. 

6. 

7. 

10. 

2.5 

.5 

-- 

-- 

-- 

-- 

7. 

7. 

-- 

10. 

10. 

7. 

7. 

25. 

6. 

5. 

6. 

BV 

C BO 

-- 

75. 

100. 

30. 

60. 

60. 

75. 

175. 

80. 

30. 

- - 

-- 

-- 

- - 

100. 

120. 

- - 

100. 

80. 

100. 

120. 

45. 

75. 

60. 

75. 

B V 

CE 0 

200. 

50. 

100. 

6. 

40. 

40. 

50. 

80. 

80. 

18. 

100. 

300. 

400. 

60. 

60. 

80. 

400. 

50. 

60. 

60. 

65. 

30. 

40. 

30. 

40. 

Reference 

Source 

DX 

SP 

CM 

DX 

CM 

SP 

S P 

DE 

CM 

CM 

CM 

CM 

DX 

CM 

S P 

DX 

CM 

CM 

CM 

S P 

SP 

CM 

CM 

SP 

CM 
Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 

BV 

CEO 

30. 

40. 

60. 

45. 

100. 

15. 

-- 

30. 

15. 

80. 

60. 

100. 

-- 

20. 

15. 

12. 

60. 

40. 

40. 

60. 

40.(60.-A) 

60. 

10. 

20. 

20. 

I 

Device 

Number 

2N2222 

2N2222A 

2N2223,A 

2N2270 

2N2346 

2N2369A 

2N2417 

2N2432 

2N2481 

2N2509 

2N2516 

2N2563 

2N2646 

2N2708 

2N2857 

2N2894,A 

BV 

E BO 

5. 

6. 

7. 

7. 

- - 

4.5 

30 .* 

15. 

5. 

7. 

8. 

20. 

30.* 

3. 

2.5 

4. 

(4.5-2N2894A) 

Reference 

Source 

DX 

S P 

S P 

S P 

DX 

SP 

CM 

CM 

CM 

CM 

CM 

S P 

S P 

CM 

CM 

S P 

CM 

CM 

DX 

CM 

DX 

DX 

S P 

CM 

CM 

K 

.1 

.1 

.21 

.5 

3.2 

.03 

.549 

.I89 

.099 

.I26 

.209 

.55 

.72 

.018 

.018 

.03 

BV 

C BO 

60. 

75. 

100. 

60. 

-- 

40. 

35.* 

30. 

40. 

125. 

80. 

100. 

35. * 

35. 

30. 

12. 

60. 

60. 

60. 

60. 

60. 

60. 

15. 

40. 

25. 

2N2904A 

2N2905 

2N2906 

2N2906A 

2N2907,A 

2N2920 

2 N 2996 

2N3014 

2N3050 

.221 

.22 1 

.044 

.221 

.1 

.04 

.01 

.02 

.O 1 

5. 

5. 

5. 

5. 

5. 

6. 

.3 

5. 

5. 
Table 16.2 TRANSISTOR DAMAGE DATA (CONT.) 

# 

Device 

Number 

2P.13053 

2N3054 

2N3055 

2N3118 

2N3217 

2N3235 

2N3240 

2N3251 

2N3308 

2N3384 

2N3436 

2N3440 

2N3585 

2N3708 

2N3777 

2N3785 

2N3819 

2N3823 

2N3902 

2N3907 

2N4037 

LN75497 

LN75638 

Mi517331 

T1482 

K 

.72 1 

3.633 

20.084 

.53 

.I26 

20. 

1.5 

.I43 

.12 

.094 

.488 

1.75 

5.278 

.507 

2. 

.012 

.22 

.228 

43.35 

.I65 

.045 

1.9 

2.3 

.1 

.21 

BV 

E BO 

5. 

7. 

7. 

4. 

15. 

7. 

8. 

5. 

3. 

-- 

- - 

7. 

6. 

6. 

8. 

.5 

25. 

30. 

5. 

6. 

7. 

5. 

Reference 

Source 

CM 

CM 

CM 

SP 

CM 

S P 

S P 

CM 

SP 

CM 

CM 

S P 

CM 

CM 

S P 

SP 

CM 

CM 

CM 

CM 

CM 

DX 

DX 

DX 

S P 

BV 

CBO 

60. 

90. 

100. 

85. 

15. 

65. 

160. 

50. 

30. 

30.t 

5O.t 

300. 

440. 

30. 

100. 

50. 

25 .t 

30.t 

400. 

60. 

60. 

20. 

BV 

CE 0 

40. 

60. 

70. 

60. 

10. 

55. 

160. 

40. 

25. 

-- 

- - 

250. 

300. 

30. 

100. 

12. 

- - 

- - 

400. 

45. 

40. 

20. 
Table 16.2 TRANSISTOR DAMAGE DATA (CONCLUDED) 

SP - SAP-I Computer Listing from SAP-I Co~np~trer Code Illar~ual (unpublished), U.S. Air Force Weapons Laboratory, 1972. 

DX - Experimental data from DASA (Defense Atomic Support Agency) 

Hondbook (See Reference 16.3) 

DE - Estimated data from DASA Handbook (See Reference 16.3) 

CM - Computed data (Reference 16.7) 

Table 16.3 ACCURACY OF K FACTOR AS DETERMINED 

BY INDIRECT METHODS (CATEGORY 2 SEMICONDUCTORS) 

Kamb = K as determined from 0. 

J a 

'case = K as determined from 0. 

JC 

Kcj = K as determined from junction capacitance 
Table 16.4 NONSEMICONDUCTOR DAMAGE TEST RESULTS 

'Arcing present continually for 700 V pulses. 

Device Type 

Capacitor 

Capacitor 

Capacitor 

Capacitor 

Coil 

Coil 

Filter 

Filter 

Potentiometer 

Potentiometer 

Potentiometer 

, Relay 

Relay* 

Relay 

Relay 

Relay 

Transformer 

Transformer 

Transformer 

Transformer 

Transformer 

Vacuum Tube 

Vacuum Tube 

Vacuum Tube 

Manufacturer 

Manufacturer's 

Part Number Roperties 

Cornell-Dublier 

Cornell-Dublier 

Sprague 

WES CAP 

Collins 

Colhs 

Test Resu

10 pF 

4700 pF, 500 Vdc 

1 pF, 200 Vdc 

0.022 fl, 600 Vdc 

220 pH 

} 2 pH lo3 Hz 

Notch400and1200Hz 

Bandpass 400 Hz ) 

250 !2 

nmn l 25 0 

Armature 

Armature 

Magnetic Reed 

Magnetic Latching 

Armature 

Audio Frequency \ 

Power, Isolation 

Audio Frequency 

Power. Isolation 

Power. Isolation, Stepdown 

Med v Twin 

t UHF High p 

Med ~1 Twin 

Cl OOK 

CK62 Series 

96P Series 

KF223KM 

240-2524-00 

542-09 16-002 

Bundy 

Var o 

Computer lnsts. 

Ohrnite 

Ohrnite 

Babcock 

C. P. Clare 

Hathaway 

Potter Brurnfield 

Struthers Dunn 

Dektronics 

Dektronics 

Freed 

Varo 

Varo 

No change in c

or leakage res

No change in 

or resistance 

No change in 

frequency res

No changes 

Resistance in

625 C& 7 629 

21-0526-00 

954-0429400 

M18-178105 

51917-1 

5 1927-3 

RP11573-GZ 

A5245-1 

63862 

FLBQOO2 

FC4-365 

D78Z223 

M8Z225 

667-0386-00 

950-1 623-200 

999-0 197-700 

6BX7 

5876 

6BC4 

Resistance in

418 St>425 

Resistance in

Resistance de

No changes 

No change in 

or voltages, n

during pulse 

No change in 

characteristic

Transconduc

decrease, 35% 
REFERENCES 

16.1 R. L. Davies and F. E. Gentry, "Control of Electric Field at the Surface 

on PN Junctions," IEEE Transactions on Electron Devices, ED-11, 

Institute of Electronic and Electrical Engineers, New York, New York 

(July 1964): pp. 313-23. 

16.2 H. S. Velorie and M. P. Prince, "High-Voltage Conductivity-Modulated 

Silicon Rectifier," me Bell System Technical Journal, July 1957, pp. 

975-1 004. 

16.3 DNA EMP (Electromagnetic Pulse) Handbook, Vol. 2, Section 13: 

Analysis and Testing (Confidential), DNA 21 14H-2, Defense Nuclear 

Agency, Washington, D.C. (November 1971). 

16.4 DNA BlZlP Handbook, p. 13-34. 

16.5 DNA EhfP Handbook, p. 13-53. 

16.6 EMP Susceptibility Threshold Handbook, (Appendix D), prepared by the 

Boeing Company for the U.S. Air Force Weapons Laboratory, Kirtland 

Air Force Base, Albuquerque, New Mexico (July 21, 1972). 

16.7 EC-135 Pretest EMP Analysis, I, BDM/A-701-705, prepared by Braddock, Dunn and McDonald, Inc., for the Air Force Special Weapons 

Center, Albuquerque, New Mexico (March 1971). 

16.8 DNA EMP Handbook, p. 13-47. 

16.9 J. B. Singletary and D. C. Wunsch, BDM Final Report, Semiconductor 

Danzage Study, Phase II, Report BDMJA-66-70-TR, prepared by Braddock, Dunn and McDonald, Inc., for the U.S. Army Mobility Equipment 

Research and Development Center, Fort Belvoir, Virginia (June 1970). 

16.10 D. C. Wunsch and L. Marzitelli, BDM Final Report, Semiconductor and 

Nonsemiconductor Damage Study, I, Report BDM-375-69-F-0168, prepared by Braddock, Dunn and McDonald, Inc., for the U.S. Army 

Mobility Equipment Research and Development Center, Fort Belvoir, 

Virginia (April 1 969). 

16.1 1 EMP Susceptibility Threshold Handbook, p. 2 12. 

16.12 EMP Susceptibility Threshold Ilar~dbook, p. 2 17. 

16.1 3 EMP Susceptibility Threshold Handbook, p. 223. 

16.14 EMP Susceptibility Threshold Handbook, p. 224. 

16.15 E. Keuren, R. Hendrickson, and R. Magyarics, "Circuit Failure Due to 

Transient-Induced Stresses, first Symposiun' on Electromagnetic Compatibility, Montreux, Switzerland, May 20-22, 19 75, IEEE EMC Conference Record 75 CH10124, Institute of Electronic and Electrical 

Engineers, New York, New York (1975), pp. 500-505. 

16.16 Keuren, Hendrickson, and Magy arics, "Circuit Failure," p. 504. 

16.1 7 Keuren, Hendrickson, and Magyarics, "Circuit Failure." 

16.18 DNA BitIP Handbook, pp. 13-1 9 to 13-1 01. 

16.19 DNA EhfP Handbook, p. 13-1 01. 

16.20 DNA EAJP Ffandbook, p. 13-1 00. 
16.21 EMP Susceptibility Threshold Handbook, p. 144. 

16.22 EMP Susceptibility Threshold Handbook, p. 146. 

16.23 EMP Susceptibility Threshold Handbook, p. 148. 

16.24 EMP Susceptibility Threshold Handbook, p. 148. 

16.24 EMP Susceptibility Threshold Handbook, p. 148. 

16.25 EIectromagnetic Susceptibility of Semiconductor Components, AFWLTR-74-280, Air Force Weapons Laboratory, Kirtland Air Force Base, 

Albuquerque, New Mexico (1974). 

16.26 B. D. Faraudo and L. C. Martin, Review of Factors for Application in 

Component Damage Analysis, Protection Engineering and Management 

Note PEM-52, Lawrence Livermore Laboratory, University of California, 

Livermore, California (September 1976). 

16.27 D. M. Tasca, Theoretical and Experimental Studies of Semiconductor 

Device Degradation Due to High Po~ver Electrical Transients, 735D4289, 

Corporate Research and Development, General Electric Company, 

Schenectady, New York (December 1973). 

16.28 Determination of Upset and Damage Circuit Thresholds, PREMPT 

Program TN-25, the Boeing company, Seattle, Washington (September 

15, 1975). 

CHAPTER 17 

TEST TECHNIQUES FOR EVALUATION OF INDIRECT EFFECTS 

17.1 Introduction 

There are as yet few official standard tests or even generally recognized 

test techniques for evaluation of how well electrical and electronic equipment 

withstands the indirect effects of lightning. The purpose of this section is to 

discuss some of the existing test techniques, to discuss some of the philosophy 

behind some of the existing or evolving standards, and to suggest possible 

avenues of improvement in test techniques. 

The lightning current to which an aircraft may be exposed cannot be 

predicted beforehand except in statistical terms. Aircraft are thus designed to 

some type of model lightning flash--one which duplicates the essential 

characteristics of lightning, even though the chances of any particular lightning 

current actually having all those characteristics is vanishingly small. A design 

model that has strongly influenced test techniques is the severe composite 

lightning model developed for the Space Shuttle Program (Reference 17.1). This 

model, with some revisions, seems to be emerging as a standard for aircraft 

design, testing, and certification. The model, as incorporated in present test 

practice (Reference 17.2), is shown as Figure 1.36. 

As regards indirect effects, only the first and possibly the second return 

stroke are likely to be of consequence. The wave fronts of the initial return 

stroke and the restrike are not specified in Figure 17.1 (Reference 17.2), since 

they are not significant for the evaluation of direct effects. In theSpaceShzcttle 

Program Lightning Protection Criteria Document and in the earlier MIL 

specification Bonding, Electrical and Lightning Protection for Aerospace 

Systems (Reference 17.3), the current is taken to rise linearly to crest in 2 ps. 

Dividing the peak amplitude of 200 000 A by a front time of 2 ps yields an 

average rate of change of 100 kA/ps, a figure frequently quoted with respect to 

indirect effects. 

Ideally, for proof tests to determine indirect effects, a simulated lightning 

current having this waveshape (or at least the waveshape of the initial stroke 

would be circulated through the aircraft and the electronic equipment observed 

to insure that no harmful effects would be caused by the current. In practice this 

is not done, since available surge generators do not have the stored energy 

required to circulate this amount of current through the inductance that would 

be associated with a complete aircraft. Compromises with perfection have to be 

made, and these take the following three forms. 

Tests may be made on a complete aircraft, during which tests 

currents of high amplitude are circulated through the aircraft and 

the effects noted. In such tests the waveshape of the current may 

differ radically from the waveshapes of actual lightning currents. 

Tests of this nature might be called proof tests and are discussed in 

Section 17.3.1. 
Tests may be made on a complete aircraft, during which tests 

currents of low amplitude but correct waveshape are circulated 

thlough the aircraft. Such tests are sometimes called frutzsietlt 

ut2u!ysis testr and are discussed in Section 17.3.2. 

Tests may be made upon individual pieces of electronic equipment 

or upon interconnected electronic subsystems. During such tests it is 

assumed, explicitly or implicitly, that the waveshapes of the current 

or voltage surges injected into the pieces of equipment or subsystems 

are representative of those produced by natural lightning. Some of 

the aspects of such tests are discussed in Section 17.5. 

17.2 Waveform Definitions 

Before discussing test techniques in detail or discussing the advantages and 

disadvantages of different types of tests, it would be well to discuss general types 

of waveforms and the terms used to describe these waveforms. 

Waveforms of either voltage or current tend to fall into one of the 

categories shown in Figure 17.1. The waveforms shown in Figure 17.1 are 

described in two ways: in terms of the natural type of waveform produced (1) 

by physically reali~able test circuits and (2) by their triangular approximation. 

When any test waveform is described in terms of its triangular approximations, it 

must be understood that there is no intent that an actual test be made with 

triangular waveforms, since such waveforms can never be produced with 

physically realizable hardware. Any test that is to be done should be performed 

with natural waveforms, of which a specified triangular wave is only a 

convenient approximation. 

Perhaps the most common waveform encountered during indirect effects 

testing is the damped oscillatory waveform of Figure 17.l(a). The major 

characteristics of importance for such a waveform are the peak amplitude, A, 

and the period, T, or alternatively, the frequency, f. Since an oscillatory 

waveform may be either a basically cosinusoidal waveform (one starting at the 

peak amplitude A) or a sinusoidal waveform (one starting at zero), the initial 

time to crest, t 1, should generally be specified in addition to the period, or 

natural, frequency. In addition, specifications may be required as to the 

decrement of the wave, or the time required for the oscillations to damp out to a 

negligible amplitude. 

The second most commonly encountered waveform is the exponential 

wave of Figure 17.l(b). This waveform describes the characteristics of a 

capacitor discharging through a resistor. The most common specifications relate 

to the peak amplitude, A, the time to crest, ti, and the duration of the tail, or 

decay, t2. The front time is described most commonly in terms of its triangular 

approximation, t',, the time required for a tangent line through the major 

portion of the leading edge to intercept the projection of the peak amplitude. 

Surge generators frequently have enough inherent inductance that the leading 

edge is fairly linear over a reasonable length of time, in which case the best fit 

tangent line may be drawn by inspection. To cover those situations in which the 
Figure 17.1 Description of waveforms. 

(a) Damped oscillatory wave 

(b) Exponential wave 

(c) Exponential wave with backswing 

(d) Generalized triangular wave 
tangent line is not obvious by inspection, specifications frequently require that 

the straight line defining the front be that passing through the 10% and 90% 

points. Sometimes specifications require that the straight line be drawn through 

the 30% and 9Wo points. The duration of the tail, t2, is most frequently defined 

as the time required for the wave to decay to one-half its initial amplitude. Some 

alternate specifications may define the tail in terms of the e-folding time, the 

time required for the wave to decay to l/e, or approximately 37% of its initial 

amplitude. 

With some types of test circuits the characteristic wave produced is that of 

an exponential wave with a relatively small backswing, as shown in Figure 

17.l(c). Technically, such a wave is a damped oscillatory wave with a high 

degree of damping. Specifications made with regard to the peak amplitude of 

such a wave, A,, and to the time to crest, t, or t'l, will be the same as those for 

the exponential wave. Specifications relating to the decay time may refer to 

either the time required for the wave to decay to one-half (or alternatively 37%) 

of its initial amplitude or the time required for the wave to decay to zero, tj. 

Specifications which relate to an exponential wave with backswing frequently 

require that the maximum amplitude of the backswing, A2, be less than some 

specified fraction of the initial amplitude. The duration of the backswing is 

frequently unspecified. In summary, then, the characteristics required to specify 

the generalized triangular wave of Figure 17.l(d) are the peak amplitude, A, the 

time to crest, tl, and a measure of the duration of the wave. This latter may be 

either the time required to decay to half value, t2, or the time required to decay 

to zero, tj. If tl is small with respect to either t2 or tj, t2 may be assumed to be 

equal to one-half of t3. 

1 7.3 Tests on Complete Vehicles 

Tests on complete vehicles may be intended either as full-scale proof tests 

of the aircraft and all its internal electrical and electronic systems, or they may 

be transient analysis tests to determine only how the wiring responds to the 

electromagnetic fields produced by the simulated lightning current. 

17.3.1 Proof Tests Techniques and Waveforms 

Proof tests on a complete vehicle are intended primarily to identify system 

incompatibility problems--systems which may operate properly by themselves 

but which may fail when the complete vehicle is subjected to lightning currents. 

For example, redundant systems may provide no redundancy when they are 

simultaneously affected by the same induced voltages. Typical test setups are 

shown in Figure 17.2 (Reference 17.4). 

In a test of this nature a high-power surge generator is placed adjacent to 

the aircraft; current is passed into one extremity of the aircraft and returned to 

the surge generator through an external wire or wires--preferably wires 

sufficiently well insulated that the test current is confined to the test circuit and 

does not enter the ground network of the facility where the tests are made. 
CURRENT GENERATOR 

TEST VEHICLE 

GROUND RETURN WIRES 

CONNECTION INSULATED FROM GROUND 

WHERE INDUCED VOLTAGE 

MEASUREMENTS ARE MADE 

GROUND WHEELS ON RETURN WIRES 

CONNECTION ELECTRICAL INSULATION INSULATED FROM GROUND 

Figure 17.2 Typical setups for complete vehicle tests. 

(a) Vehicle under test grounded 

(b) Low side of current generator grounded 

There should be a ground on the system. If the low side of the surge generator 

can be insulated from ground, it is generally preferable that the vehicle under 

test be grounded as shown in Figure 17.2(a). If this is not possible, the test 

vehicle must be insulated from ground and the ground made at the low side of 

the surge generator, as shown in Figure 17.2(b). 

Since present surge generators are not sufficiently powerful to circulate 

the full-design lightning current through the aircraft, it is accepted practice to 

perform the proof test with two or more damped oscillatory currents having as 

waveforms one which provides the fast rate of rise characteristic of a natural 

lightning stroke wave front and the other a long-duration period characteristic of 

natural lightning-stroke duration. The two different components are necessary, 
since the mechanism of coupling electron~agnetic fields into the interior of a 

vehicle is different at different frequencies. The high-frequency waveform is 

intended primarily for evaluation of aperture effects and has a minimum 

amplitude of 10 kA and a fundamental frequency of 100 kHz. The lowfrequency waveform is primarily for evaluation of diffusion effects and has a 

minimum amplitude of 20 kA and a fundamental frequency of 2 kHz. As a 

result of the inductance of the test circuit, the waveforms are typically of a 

sinusoidal nature, and no independent definition of the front time, tl, need be 

made. 

During the tests the responses of the electrical equipment should be 

measured. Since neither of the test current waves is of the full amplitude 

associated with lightning, the test results must be extrapolated to the full 

lightning-threat level, 200 kA. For the high-frequency test current wave, the 

average rate of rise of the first quarter cycle of the test waveform should be 

determined by measurement or calculation and the test results extrapolated to 

100 kA/ps. Measurements made with the low-frequency current wave must be 

extrapolated from the measured peak amplitude of the current wave to the 200 

kA level. 

The use of these oscillatory currents is one of the necessary compromises 

with perfection previously mentioned. These waves are used for proof tests 

simply because test facilities capable of subjecting a complete aircraft to a 

full-threat lightning current do not exist. For purposes of simulation the 

oscillatory waves have both merits and drawbacks. These are summarized below. 

The fast oscillatory wave has the following merits: 

Over a narrow band of frequencies it simulates one important aspect 

of indirect effects--the aperture coupling of electromagnetic fields. 

High amplitude is available from relatively low-energy-surge generators, since little energy is lost in waveshaping resistance. 

It can be nearly a full-threat-level test because of the higher 

frequency coupling effects. 

The fast oscillatory wave has the following drawbacks: 

It overemphasizes (and deliberately so) one of the component 

frequencies of lightning currents. 

It underemphasizes low-frequency coupling effects, particularly the 

effects of vehicle resistance and the diffusion penetration of 

magnetic fields. 

It is fundamentally different in waveshape from the waveshapes of 

natural lightning currents. 

It is potentially damaging to avionic equipment. In this respect, 

however, it must be remembered that if avionic equipment can be 

damaged by a simulated lightning proof test, it can also be damaged 

by an actual lightning flash. 

Bulky and expensive test equipment is required to make the test. 

Some of the drawbacks of the fast oscillatory test wave are offset by its 

companion, the slow oscillatory wave. The slow oscillatory wave has the 

following merits: 
It simulates over a narrow band of frequencies one important aspect 

of indirect effects--the resistive coupling of electromagnetic fields. 

High amplitude is available from relatively low-energy-surge generators, since little energy is lost in waveshaping resistance. 

It can approach a full-threat-level test of lower frequency resistive 

coupling effects, at least for lightning strokes of low to average 

amplitude. 

The slow oscillatory wave has drawbacks parallel to those of the 

high-frequency test wave: 

It overemphasizes (and deliberately so) one of the component 

frequencies of lightning currents. 

It may not correctly simulate diffusion penetration of magnetic 

fields. One can take refuge, however, in the observation that 

imperfect simulation is better than no simulation. 

It does not constitute a full-threat-level test for the higher amplitude 

lightning flashes. Data obtained at low amplitudes require extrapolation to give effects that would be present at full lightningcurrent amplitude. 

It is fundamentally different in waveshape from the waveshapes of 

natural lightning currents. 

It is potentially damaging to avionic equipment. The observation 

made previously about the high-frequency test waveform applies 

equally well to the same drawback in the low-frequency test. 

It too requires bulky test equipment. 

17.3.2 Transient Analysis Tests 

A major virtue of the fast oscillatory test wave is that it can simulate fairly 

well one of the important aspects of natural lightning--the electromagnetic 

effect, associated with the initial rate of change of lightning current. Because of 

the emphasis that is placed on rate of change, it is often believed that only rate 

of change of current is important. Likewise, it is often believed that a 

nondamaging test for indirect effects analysis can be made by applying a current 

that initially rises at the full-threat-level rate, 100 kA/ps, but not allowing that 

current to flow for very long or not allowing it to increase to a high amplitude. 

Such concepts are fallacious. 

Consider Figure 17.3 (Reference 17.5) in which two different current 

waves are applied to an equivalent circuit of the type shown in Figure 17.3(a). 

One current rises at the rate of 100 kA/,us to a value of 10 kA, with tl equal to 

0.1 ps, and the second rises at the rate of 100 kA/w to a value of 100 kA, with 

t 1 equal to 1 ps. The responses of the circuit, shown in Figure 17.4 (Reference 

17.6), are fundamentally different. 

The most nearly correct method of making a nondestructive test of 

indirect effects is to keep the basic waveshape the same, reducing only the 

amplitudes. Such scaling is shown in Figure 17.5 (Reference 17.7). Frequently, 

concern is expressed about the fact that reducing amplitude but keeping front 

50 1 
.I ,us 1 ,us 

Figure 17.3 Why duration of rate of rise is as important as rate of rise. 

(a) Circuit analyzed 

(b) Current waveshapes 

time the same reduces the rate of change of current. The initial rate of change is 

indeed reduced, but it is reduced in exactly the same ratio as that of the peak 

amplitude. 

Such considerations of scaling have led to the development and use of the 

transient analysis technique in which currents of waveshape similar to those 

produced by natural lightning, but of lower amplitude, are circulated through 

the aircraft. During such a test many measurements would be made of the 

transients induced on the wiring of the aircraft. These measurements, when 

scaled up to full-threat level, provide a measure of the transients to which 

lightning would subject the individual pieces of electronic equipment. An 

assessment of whether or not such transients might present a hazard to the 

equipment might involve additional tests on the individual pieces of equipment. 

Some of the considerations in such tests are given in Section 17.5. 
Figure 17.4 Responses of circuit. 

(a) Response to a 10 kA current 

(b) Response to a I00 kA current 

The techniques of making transient analysis tests have been described in 

other literature (References 17.8, 17.9, and 17.10) and will not be discussed 

here. A summary of their advantages and disadvantages follows. 

The transient analysis type of test has, among others, the following 

advantages: 

The tests may be made with current surges of waveshape nearly 

representative of those of natural lightning. 

The waveshapes have the correct ratio of high- and low-frequency 

components. 

The test correctly simulates frequency-dependent resistance and 

diffusion-coupled electromagnetic effects. 

The test is nondamaging to avionic equipment if it is made at low 

current levels. 

Relatively small and inexpensive equipment is required to make the 

tests. 

The transient analysis type of test has, among others, the following 

drawbacks: 

Because a considerable amount of energy is lost in resistive circuit 

elements, a given size test generator produces less current than it 

would if it were used to produce damped oscillatory current waves. 

Data obtained at low amplitudes must be extrapolated to full-threat 
I LOWER CURRENTS OF SAME WAVESHAPE 

I KEEPING INITIAL RATE OF RISE 

- 

I CONSTANT GIVES A DIFFERENT WAVESHAPE AND 

/ HENCE A DIFFERENT TYPE OF RESPONSE 

(b) lf2 

Figure 17.5 Considerations regarding amplitude scaling. 

(a) Correct scaling 

(b) Incorrect scaling 

level. Such extrapolation requires an assumption of linearity--that 

is, 10 A produces 10 times the response of 1 A. While such linearity 

can usually be relied upon, there are some nonlinear effects involved 

in the indirect effects of lightning. 

It is not a proof test and hence has no pass or fail criteria. 

17.4 Tests on External Electrical Hardware 

The test techniques described so far have dealt with determining how the 

wiring and equipment within the aircraft respond to the passage of lightning 

current through the aircraft. There are, however, some types of equipment 

containing electrical hardware, or containing wiring that connects to electrical 

hardware, that are frequently mounted on the exterior of the aircraft. Some of 

these are likely to be struck by the lightning flash. A pitot boom containing a 

deicing heater and the electrical wiring to supply power to that heater is an 

excellent example. An externally mounted navigation light is another example. 
Tests are frequently required to determine what voltages and currents might be 

induced on the electrical wiring by a lightning stroke to such hardware. These 

tests are needed because the currents and voltages may be conducted to other 

electrical systems in the vehicle. The basic principle by which one of such tests is 

made is shown in Figure 17.6 (Reference 17.1 1). The test object should be 

mounted on a shielded test chamber so that access to the electrical connectors of 

the object can be obtained in an area relatively free from the electromagnetic 

fields produced by the surge generator. The test object should be mounted to 

the shielded enclosure in a manner similar to the way in which it is mounted on 

the aircraft, since normal bonding impedances may contribute to the voltages 

induced on circuits. If the shielded enclosure is large enough, the measurement 

oscilloscope may be contained within it. If not, a suitably shielded instrument 

cable may be used to transfer the induced voltage signal from the shielded 

enclosure to the oscilloscope. Measurements of the voltage induced on the 

electrical circuits within the test objects should be made on both a line-ground 

and a line-line mode. Since the direct contact of the lightning flash with the test 

object frequently produces quite high voltages, dividers or attenuators external 

to the oscilloscope may be needed. If used, these should be checked to ensure 

that the signals displayed on the oscilloscope are not the result of either spurious 

pickup on the attenuator or overload effects in the amplifiers used in the 

oscilloscope. 

The appropriate waveforms of the surge currents to be injected into the 

test object depend on where the object will be located on the aircraft and 

whether or not it will be likely to intercept the initial lightning current stroke or 

whether it is likely only to intercept a stroke blown back along the surface of 

the vehicle by the swept-stroke mechanism. The literature on test requirements 

discusses in considerable detail the waveforms appropriate to the different zones 

on the aircraft, and for any such test the reader is referred to that literature 

(Reference 17.1 2). 

17.5 Tests on Internal Equipment 

The second major category of tests involves tests to determine how the 

avionic equipment within the aircraft responds to the transients induced on the 

wiring by the flow of lightning current. Equipment for and specifications 

relating to these tests are in an embryonic state of development at the present 

time. In the following sections we will attempt to discuss some of the available 

test techniques, the goals they are intended to accomplish, and some of the 

factors to which attention must be given during the conduct of such tests. 

1 7.5.1 Waveforms and Amplitudes 

Tests and measurements have shown that the transient voltages and 

currents to which avionic equipment is subjected are most commonly of a 

damped oscillatory nature. Voltages are almost always of this nature and 

currents usually are as well, although the currents flowing on the shields of 

shielded conductors have more of a tendency to be of a unidirectional nature 
SHIELDED ENCLOSURE r------------- --' 

I 

I DIVIDER 

OR 

ATTENUATOR OSCILLOSCOPE 

I 

LIGHTNING 

Figure 17.6 Typical electrical hardware indirect effects, test and measurement 

circuit. 

like that of the external lightning current than do the induced voltages. 

Some examples of transients found in aircraft during transient analysis 

tests are shown in Figure 17.7 (Reference 17.1 3). Whether the measurement is 

of an open circuit voltage or of a short circuit current is indicated in this figure. 

Figure 17.7 Some examples of lightning-induced transients found in aircraft. 
Information relating to amplitude and to the conditions under which the 

transient was measured has deliberateIy been omitted, since the intent of the 

figure is merely to illustrate the broad diversity of waveforms that might be 

encountered. 

The question of what waveforms to specify for transients used during tests 

related to indirect effects is another subject for which there are as yet only 

evolving standards. Some of those standards, or evolving standards, that have 

been used outside as well as inside the aircraft community are summarized in 

Tables 17.1 (Reference 17.14) and 17.2 (Reference 17.1 5). Specifications 

relating waveforms that are basically unidirectional will be discussed first. 

TABLE 17.1 SPECIFICATIONS RELATING TO 

UNIDIRECTIONAL TRANSIENTS 

Peak Crest Decay Amplitude of 

Amplitude Time Time Backswing 

A1 T1 T2 A2 

Space Shuttle - Component 1 

10 A 2 Ps 100 ps 

50 V 2 1s 100 ps 

Space Shuttle - Component 2 

5 A 300 ps 600 ps 

0.5 V 300 ps 600 ps 

Standard ANSI Test Waves 

(depends on 4 Ps 10 ps 

application) 1.2 ps 50 ps to half value 

Fisher and Martzloff 

(depends on 5 I.ls not specified 

application) 0.25 p 5 Ps 

not specified 

A2 < 0.5 A1 

The waveforms of longest standing are the American National Standards 

Institute (ANSI) waves relating to lightning effects on high-voltage apparatus 

(Reference 1 7.1 6). These call for current waves of 4 ps to crest and 10 ps to half 

value, and voltage waves of 1.2 ps to crest and 50 ps to half amplitude. Although 
other waveforms are specified for other purposes, these are the most common 

waveforms. 

TABLE 17.2 SPECIFICATIONS RELATING TO 

OSCILLATORY TRANSIENTS 

Peak Damping 

Amplitude Period Frequency Definition 

Space Shuttle - Component 1 

10 A 8~ 125 kHz 

50 V 8~ 125 kHz 

Space Shuttle - Component 2 

5 A 1200 /.IS 833 Hz 

ANSI SWC Test 

N A N A NA 

2.5-3.0 kV 0.67-1 /.ti 1 .O-1.5 MHz 

Martzloff and Howell 

NA NA N A 

0.8 kV 10 ps I00 kHz 

A2 > 0.25 A] 

on fifth cycle 

A2 > 0.25 A, 

on fifth cycle 

A;! >0.25 A1 

on fifth cycle 

A2 > 0.25 A1 

on fifth cycle 

Envelope decay 

to 50% in no 

less than 6 ps 

Some of the older American literature refers to a voltage wave having a 

time to crest of 1.5 /.ti and a decay time to half amplitude of 40 ps. For all 

practical purposes, this is the same as the 1.2 x 50 ps waveshape, the 1.2 x 50 ps 

wave representing a fusion of slightly different American and European 

standards. The 1.2 x 50 ps figure is often viewed as representing the waveshape 

of the lightning current. This view, however, is incorrect; it represents the 

voltage produced at high-voltage equipment, such as transformers, by the effects 

of a lightning stroke at some remote point. The 4 x 10 current wave 
represents the lightning-related current passing through a lightning arrester after 

it has been caused to spark over internally by a voltage surge of different 

waveshape. Both the voltage and current waveshapes represent transients that 

may be produced by basically simple laboratory test equipment. They do not 

necessarily represent the shape of the current or voltage transients produced by 

natural lightning. 

At one stage in the development of the Space Shuttle Criteria Document, 

there was an allowance made for two basically unidirectional transient test 

waves. While those test waves had some deficiencies and have been largely 

superseded, discussion of them is still appropriate, since that discussion will 

illustrate some of the problems inherent in waveform specifications. 

The first of these specifications dealt with a transient rising to crest in 2 ps 

and decaying to zero in 100 ps. The intent of this test wave was to dupIicate in 

some manner the effects produced by magnetic flux leaking through apertures. 

The waveform had its basis in the ANSI test waves. These waveforms permitted a 

backswing, a feature characteristic of transformer-coupled surges. The backswing 

amplitude was required to be less than 25% of the initial amplitude, but the 

duration of the backswing was left uncontrolled. The test waveform was not 

intended to be interpreted as requiring a backswing. An overdamped wave was 

perfectly satisfactory. If an overdamped waveform was used, there would be no 

clearly defined time to zero. In such cases the decay time was intended to be 

taken as one-half the indicated value (50 ps instead of 100 ps) and measured to 

the time at which the wave had decayed to 50% of its initial amplitude. The 

waveform was thus similar to the standard ANSI test waveform derived for tests 

on high-voltage apparatus. The front time of 2 ps reflected the 2 ps front time of 

the basic lightning design current. The transients themselves were specified as 

having amplitudes of 50 V open circuit and 10 A short circuit. 

The second specification related to a long-duration transient representing 

the effects produced by magnetic flux diffusing through the walls of cavities. 

Such flux would have rise and decay times much longer than those of the 

lightning current. The specification called for a short circuit current of 5 A and 

an open circuit voltage of 0.5 V, both taking 300 ps to reach crest and another 

300 ps for decay to zero. The specification of equal times to crest and from crest 

back to zero is incompatible with the response of real physical elements. In 

practice, any waveform with a rise time of 300 pi would have a decay time 

longer than 600 ,us. 

One common deficiency of the above specifications was that both the 

short circuit current and open circuit voltage had the same waveshape. In 

practice the open circuit voltage would be of a duration shorter than that of the 

short circuit current. 

In the original paper in which the concept of transient control levels was 

first presented (Reference 17.17), Fisher and Martzloff proposed a test wave 

that was primarily unidirectional. The open circuit voltage was characterized by 

a fast rise to crest and then a decay to zero in 5 ps, or greater. To allow for 

transformer coupling of the transient, a backswing was allowed after the 

transient had decayed to zero. The character of the backswing was not 
specified--only that its amplitude should be less than 50% of the initial 

amplitude. 

The rationale behind this waveform concept included the following several 

aspects. 

It should be in some measure proportional to the derivative of the 

magnetic field produced by a lightning current. 

The duration of the transient should be long enough that possible 

failures of semiconductors would not be strongly affected by the 

waveshape of the transient, since, with transients of duration shorter 

than about a microsecond, the failure levels of semiconductors are 

strongly affected by waveshape. 

The duration of the transient should be roughly comparable to the 

duration of clock cycles in digital equipment. 

The transient should include a rapidly changing phase to excite 

inductively coupled circuit elements. 

The transient should be one that could be produced by and coupled 

to equipment by relatively simple test equipment. 

Directly associated with the open circuit voltage transient was a short 

circuit current transient. The short circuit current was that current which would 

flow from a source the internal impedance of which could be represented as 50 

pH in parallel by 50 S2. The test waves taken together as a set were thus more 

consistent than were those relating to the Space Shuttle. Amplitudes of neither 

current nor voltage were specified, since the test levels would be part of transient 

control level specification. 

The most common type of transient encountered would be oscillatory. 

Some existing specifications relating to osciIlatory transients are summarized in 

Table 17.2. The first two shown deal with the Space Shuttle and were essentially 

oscillatory versions of the unidirectional transients shown in Table 17.1 As 

presented at that time there was no specification on the initial front time; 

however, the maximum front time would be governed by the oscillatory 

frequency. The decrement of the transient was defined by requiring that the 

amplitude of the fifth half-cycle be more than 25% of the amplitude of the 

initial amplitude. 

The low-frequency component 2 was ultimately dropped from further 

consideration, since there was little evidence from transient analysis tests of such 

low-frequency oscillations being excited by diffusion-coupled magnetic flux. 

A specification of somewhat longer standing is the ANSI (SWC) Surge 

Withstand Capability Test (Reference 17.18). This test was derived from 

considerations of the transients found on control equipment in high-voltage 

substations. The specification makes no reference to short circuit currents. The 

open circuit voltages are specified as being between 2.5 and 3 kV, with the 

oscillatory frequency being from 1.0 to 1.5 MHz. Decrement is specified by 

requiring that the envelope decay to 50% in no less than 6 ,us. 

The test was intended for simulation of the transients found in 

high-voltage electrical substations and to which equipment in such substations 

might be exposed. The voltage range specified is reasonable for such apparatus 

510 
but may be high for most electronic equipment, particularly for equipment 

located in shielded locations. The waveshape, while not unreasonable, is 

sometimes difficult to inject into circuits by transformer coupling. This test does 

not treat in any way the short circuit current that would flow on a shorted 

conductor connected to a low-impedance load or flow upon the shield of a 

cable. The specification that the envelope decay to 50% in no less than 6 ps 

implies a lower loss test circuit than may be found in practice. The 

recommended test generator circuit is somewhat uncontrolled, using an 

untriggered spark gap firing virtually at random at the crest of an ac charging 

voltage, though the waveform itself does not imply any particular generator test 

circuit. The specification does call for a 150 R source impedance. The question 

of source impedance deserves serious consideration and 150 52 may not be the 

optimum value. 

This test (and waveform) is one of the few that have been formally 

approved by an industry standardizing group. 

Martzloff and Howell (Reference 17.19) proposed a test wave with a 

frequency of 100 kHz and a voltage range of 0 to 8 kV. The wave was intended 

for the duplication of transients found in residential circuits on 120 V ac lines. 

The damping was specified such that the ratio of successive half-cycles should be 

greater than 0.6. This also may be somewhat high, particularly if the transient is 

to be coupled into equipment through transformers. No specification was made 

about the magnitude or waveshape of the short circuit current. Martzloff and 

Howell show a test circuit capable of injecting the transient onto 120 V ac lines 

where, since the output impedance of the circuit is basically 150 R resistive, the 

shape of the short circuit current would be about the same as that of the open 

circuit voltage. 

This waveform and this test circuit have been widely accepted in some 

fields. One example is their use in relation to ground fault interrupters 

(Reference 17.20). 

In answer to response from readers and users of the original Fisher and 

Martzloff paper (Reference 17.21) on the transient control level philosophy, 

Crouch, Fisher, and Martzloff have proposed a test wave somewhat different 

from the original one (Reference 17.22). The revised voltage wave, shown in 

Figure 17.8, emphasizes the oscillatory nature of the wave, rather than 

deemphasizing it, as did the original test wave. The front time was raised to 0.5 

ps, the course of the wave after crest being defined in terms of its oscillatory 

frequency of 100 kHz and the decrement specified by requiring that the ratio of 

successive half-cycles be greater than 0.6. The voltage wave thus becomes nearly 

identical with that proposed by Martzloff and Howell. 

The proposed waveshape should be viewed in relationship to its companion short circuit current. The current, similar to that from natural transient 

sources, is essentially the integral of the voltage. 

17.5.2 Equipnlent for Direct Injection of Transients 

Surge generators capable of producing the types of transients described 

51 1 
Al (100V) OPEN CIRCUIT 

0 

Figure 17.8 Revised proposals for TCL test waves. 

(a) Open circuit voltage 

(b) Short circuit current 

here generally en~ploy capacitors that are discharged into a wave-shaping circuit. 

That circuit would be resistive if unidirectioi~al transients are to be produced, 

and inductive if oscillatory transients are to be produced. Some representative 

examples of circuits are shown in Figure 17.9 (Reference 17.13). The switching 

gap may be either triggered or untriggered depending on the degree of 

sophistication desired. Components should be chosen and laid out in such a way 

as to minimize undesired residual inductance. Circuit voltages are typically high 

enough to be hazardous; therefore, appropriate safety precsutions should be 

taken in both their construction and their use. Figure 17.10 (Reference 17.24) 

shows an exanlple of the types of transients that are produced by the revised 

TCL test circuit shown in Figure 17.9(b). 

One way in which the generators may be used is to connect the output 

directly to the terminals of the device under test. When generators are used in 

such a manner, one may determine the level of transient that causes a device to 

fail or, alternatively, determine that a given transient does not cause the device 
to fail. The test may be applied in a common mode, as it is in Figure 17.1 1(a) 

(Reference 17.3,5), or in a differential mode, as it is in Figure 17.11(b). 

Depending on the nature of the circuits under test, it may be appropriate to 

couple the transient into the terminals of the device through coupling capacitors. 

Direct injection of transients into the terminals of a device has the 

drawback that the device becomes connected in a manner different from its 

normal connections. As a result, it may not be possible to check the device to 

see whether the transient interferes with the normal operation of the device. One 

way the operation of the device may be checked is to discharge the test 

generator into shields placed over interconnecting cables, if there are such 

shieids, or into the cases housing electronic equipment. In either case, the 

generator would be used to generate a short circuit current, and that current, 

C1 = Mica type 

L1 3 turns no 9 wire on 2% in. diameter form 

IMR 50pH 

dc OUTPUT 

CHARGING 

VOLTAGE 0.25pF 

(b) > - r T A b 

* LOW INDUCTANCE 

**LOW INDUCTANCE & NON POLARIZED 

FRONT TIME = 3R2C 2 

TAIL TIME = 0.7 R1 (C, + C2) 

.-~ 

-"--J--= v w 

Figure 17.9 Representative transient generator circuits. 

(a) SWC test circuit - I .5 MHz oscillatory 

(b) Revised TCL test circuit - 100 kHz oscillatory 

(c) Unidirectional - 1.2 x 50 ps 

c2 

OUTPUT 

0.03pF 

R2 

Rl dc -- 

CHARGING -?- 75R F- 

(4 - A - -P > a 

VOLTAGE 

1 PF 
Figu~e 17.1 0 Output characteristics of TCL generator shown in Figure 17.9(b). 

Current - 40 A/major div. 

Voltage - 1000 Vlrnajor div. 

flowing through whatever transfer impedances might be present, would induce 

transients upon the internal circuitry. Presumably one would have previously 

determined in some manner what would be an appropriate amount of current to 

inject into the shields. 

Direct injection of current is an appropriate test technique if the system 

wder test will be exposed to a severe electromagnetic field threat, since any 

successful system will almost certainly be fitted with shields on the interconnecting wires. 

1 7.5.3 Equipment for Transformer Injection of Transients 

Another powerful test technique by which transients may be injected into 

the electrical wiring interconnecting different pieces of electronic equipment is 

through the use of a pulse-injection transformer and a suitable pulse generator. 

The elements of such a system are shown in Figure 17.1 2. Current that is passed 

from the pulse generator through a primary winding on a magnetic core sets up a 
TERMINALS OF 

ISOLATING REACTORS SYSTEM UNDER TEST 

- 0 

TO OTHER CrYYYI 1 o--- 

EQUIPMENT ----+ 

0 

I /I 

TEST 

(0) GENERATOR REFERENCE GROUND 

TERMINALS OF 

SYSTEM UNDER TEST 

TEST 

GENERATOR 

Figure 17. I1 Direct injection of transients. 

(a) Common mode 

(b) Differential mode 

magnetic field in that core. If the core is placed around a cabIe interconnecting 

two pieces of electronic equipment, the magnetic field contained in that core 

induces current or voltage in the interconnecting wiring in a manner very similar 

to that produced by magnetic flux from an external lightning source. This type 

of equipment has been used for NEMP effects testing on the B-1 (Reference 

17.26) and has also been specified for use in the Space Shuttle Lightning Criteria 

Document (Reference 17.27), though only in principle, not in detail. An 

example of the type of test waves that may be produced by a transformer 

pulse-injection circuit is shown in Figures 17.13 and 17.14 (Reference 17.28). 

The inclusion of more complex waveshaping elements in the pulse generator 

(shown dotted in Figure 17.1 2) would allow other waveshapes to be produced. 

Some considerations regarding the capabilities of such transient injection 

follow. First, it should be remembered that the circuit under test is primarily 

exposed to the changing electromagnetic field confined in the core of the pulse 

injection transformer. If the circuit under test is a high-impedance circuit, the 

natural effect of that changing magnetic field is to induce an open circuit 

voltage, the magnitude and duration of which will depend upon the amplitude 
DEVICE 

UNDFR - . - - - . . TEST 

CURRENT VIEWING 

MAGNETIC TRANSFO 

Y CORE . 

ZENER DIODES TO LIMIT 

v GROUND PLANE OR CURRENT RETURN PATH 

Figure 1 7.12 Transformer pulse injection. 

and rate of change of the magnetic field in the core but which will not be 

significantly affected by the length of the circuit under test. If the circuit is of 

low impedance, however, the magnetic field wlll induce a circulating current in 

the circuit. The shape of this current will be of a duration longer than that of the 

open circuit voltage, and will be nearly the same as the duration of the current 

produced by the pulse gerlerator in the primary of the transformer. The 

magnitude of this current will depend upon the impedance of the circuit under 

test, and th~s impedance will in turn depend upon the height of the circuit above 

a ground plane and upon the diameter of the circuit or bundle of wires in which 

current is injected, since these latter parameters affect the self-inductance of the 

circuit. A complete specification of the test circuit must deal both with the 

maxlmum open circuit voltage that may be produced and with the maximum 

short circuit current that may be produced. The inductance of the circuit under 

test is almost directly proportional to the length of the circuit, but proportional 

only to the logarithm of the diameter of the circuit or its height above the 

ground plane. In the absence of any other specifications, it is suggested that the 

pulse generator and core be such that the specified short ci~cuit current may be 

induced on a conductor 3 m long (about 10 ft) and having a diameter of 0.41 crn 

and spaced 5 cm above a ground plane. Such a conductor may be provided by 

the sh~eld of RG-58 coaxial cable. Furthermore, it is proposed that the open 

circuit voltage be measured when the pulse generator is set at the ainplitude 

producing the specified short circuit current. Specifications have yet to be 
Figure 17.13 Oscillatory current injected. 

(a) ISC 

(b) EOC 

(c) EOC (different time base) 

developed as to what this open circuit voltage should be in terms of the 

corresponding short circuit current, although suggestions are given by Fisher and 

Martzloff and by Howell and Martzloff (Reference 17.29). If the open circuit 

voltage is too high, there are two alternatives: (1) to modify the pulse generator 

so as to reduce the rise time of the current pulse in the primary of the 

transformer, or (2) to use a Zener diode or other clipping device on an auxiliary 

winding on the transformer core. After the required short circuit current was 
VERTICAL - 1 A/div. 

HORIZONTAL - 1 ps/div. 

VERTICAL - 20 V/div. 

HORIZONTAL - 0.2 ps/div. 

Figure 17.14 Unidirectional current injected 

(a) IsC 

(b) EOC 

shown to be developed on the standard conductor (Figure 17.12), that standard 

conductor would be replaced by the actual cable into which transients were to 

be injected. The actual current induced in the cable could be monitored by the 

current-viewing transformer. Recording it as well should probably be one of the 

requirements of the test plan. The actual current, of course, may be lower than 

the standard short circuit current if the actual cable is terminated in impedances 

higher than zero. Conversely, if the actual cable under test is provided with an 

overall shield and that shield is connected to the ground structure at both ends 

and if it has a diameter larger than 0.41 cm, the actual short circuit current may 

be larger than the standard or specified current. 

Whether or not it is appropriate to measure the voltage actually induced 

on the individual circuits would depend upon the characteristic of those circuits. 

Connection of a voltage probe would certainly be more likely to disturb the 

circuit and cause spurious response than would the insertion of a current-viewing 

transformer. 

During the demonstration tests in which the waves shown in Figures 17.13 

and 17.14 were measured, the conductor was a 15-foot section of RG-58 coaxial 

cable lying directly on a metal-covered floor. One end of the shield of this cable 

was connected to the floor, and the other end could be left either open circuited 
or shorted to the floor. Measurements were made only of the open circuit 

voltage or the short circuit current on the shield of the cable. Measurements 

were not made of either the current or the voltage induced on the signal 

conductor within the shield. The short circuit current is seen to rise to crest in 

about 1 ps and to oscillate at a frequency of 123 kHz. The peak current 

obtained with this configuration was about 23 A. 

If the conducter was left open at one end, the magnetic field that was then 

produced in the core of the transformer induced an open circuit voltage of about 

210 V. The voltage rose to crest in about 0.2 ps and oscillated at a frequency of 

91 kHz. The frequency was different because the inductance seen by the 

capacitor in the pulse generator was different, depending on whether or not 

there was a short circuit on the secondary of the transformer. The higher 

frequency oscillations on the front of the open circuit voltage were caused by 

the stray inductance and capacitance of the conductor under test. 

Significant points to observe here are that the open circuit voltage will 

invariably rise to crest faster than does the short circuit current and that the 

open circuit voltage will typically oscillate at a lower frequency than does the 

short circuit current. When the conductor under test is shorted, the inductance 

seen by the pulse generator is lower than it is when the secondary is open. This 

difference in oscillatory frequencies should be either recognized and allowed for 

in test specifications or corrected by adjusting the capacitance of the pulse 

generator used to excite the transformer. 

When a unidirectional current was injected into the transformer (Figure 

17.14), the short circuit current rose to a crest of about 5.5 A in about 0.5 p. 

The rise time was greater and the amplitude less because of the damping resistor 

that was required to be in series with the primary winding. The open circuit 

voltage again rose to crest more rapidly than did the short circuit current. If an 

equal number of turns are used on the primary and secondary windings, there 

will be less current in the secondary winding than in the primary. With the 

transformer core used, the secondary current was about 65% of the primary 

current. The factors that affect the performance of puke-injection transformers 

are discussed in more detail in paragraph 17.5.4. 

As an alternative to transformer coupling, it is possible to discharge the 

pulse generator into a wire run parallel to the cable interconnecting the two 

pieces of equipment. This technique is less efficient in coupling low-frequency 

energy into the cable under test than is the transformer technique, but it may be 

more effective in coupling high-frequency energy. 

17.5.4 Equipment for Generation of Electric and Magnetic Fields 

The test equipment described so far has been aimed at the injection of 

current or voltage pulses into the terminals of electronic equipment. There is 

sometimes a need for subjecting an entire piece of electronic equipment to an 

engulfing electric or magnetic field, principally to check for magnetic field 

leakage of a cabinet. Sometimes these effects are checked with a wire-wrap 

technique in which a coil of wire is wrapped around the item to be tested and 
the coil then excited from a pulse generator. The type of magnetic field so 

produced is perhaps not the best that may be derived for indirect effects testing. 

In the study of nuclear electromagnetic pulse (NEMP) effects, use is often made 

of strip line simulators in which a pulse generator is matched to a large, open 

transmission line. In such simulators attention is given to ensuring that there is a 

suitable transition from the small geometry of the pulse generator to the large 

geometry of the test chamber, with the intent of producing plane wave fields in 

the working chamber. The pulse generators and working chambers are also 

carefully matched to each other to allow the production of electromagnetic 

fields having very fast fronts. 

For testing the electromagnetic effects resulting from lightning, the 

physical design of such chambers can frequently be simplified. The rise times 

involved are not as fast as those in NEMP studies, and the requirement of plane 

wave propagation in the test chamber is not as important. A test chamber 

suitable for many types of equipment might be like that shown in F~gure 17.15. 

It would basically consist of a Helmholtr. coil of rectangular cross section about 

1 m3 in volume. Such a coil, when excited by the pulse generator sketched, 

would produce in its working volume a magnetic field of approximately 1000 

A/m or an electric field intensity of about 30 kV/m. Magnetic fields of this 

nature might frequently be encountered in typical equipment bays of an aircraft 

unless these bays were deliberately built to provide electromagnetic shielding. 

The impedance (E field intensity/H field intensity) would be lower than that of 

free space or lower than the impedance of an NEMP strip line simulator. This 

would be appropriate, since the electromagnetic field impedance inside an 

aircraft equipment bay would also be lower than the impedance of free space. 

The duration of the pulses produced by such a simulator with the generator and 

load resistor constants shown would be about 100 ps. The front times would be 

determined by the inductance of the chamber and the values of any waveshaping 

elements, L and C, employed in the pulse generator. Longer duration magnetic 

fields, typical of those that might be found in an aircraft equipment bay, might 

be produced by short circuiting the 20 R termination resistor shown in the 

figure and, possibly, increasing the value of the capacitance and inductance of 

the pulse generator. 

17.6 Design Notes for Transformer Pulse-I~~jection Equipment 

This section will discuss some of the detailed considerations in the design 

of equipment suitable for pulse-injection equipment. The pulse generator and 

transformers described should not be viewed as representing the only designs 

that might be suitable or even the best that are suitable. They are intended only 

to provide a starting point for those who wish to initiate a test program at 

minimum cost. 

17.6.1 Pulse Generator 

A simple single-polarity current injection-pulse generator could be made as 
Figure 17.15 Possible test chamber for evaluating electric and magnetic field 

effects. 

shown in Figure 17.16 (Reference 17.30). This pulse generator is basically the 

one used to produce the waveforms shown in Figures 17.13 and 17.14. The 

energy storage element is capacitor C, which is charged through D2 to a voltage 

determined by the ratings of transformer T2 and the setting of the variable 

autotransformer TI. Charging of the capacitor would take place on the positive 

half-cycle (b) of the supply voltage. During this time the discharge switch, RYI, 

a relay with mercury-wetted contacts, would be unenergized and resting on 

contact b. During the negative half-cycle (a) RYl will be energized through 

diode Dl and closed to position a, thus discharging the capacitor during the 

half-cycle when there is no voltage being applied through diode D2. Because of 

mechanical inertia, the switch armature does not necessarily operate in phase 

with the ac drive signal, and, consequently, during a portion of the positive 

charging half-cycle, the contact may remain closed to point a. In order to 

prevent excessive power drain from the transformer or excessive power 

dissipation in R3, the series resistor R2 is used to limit the charging current 

drawn from the transformer. 

The capacitor would be discharged into a single-turn primary on a 

transformer core. If a unidirectional pulse were desired, switch S3 would be 

open and resistor R2 inserted in the discharge circuit to provide damping. If S3 

were closed, the capacitor would oscillate through the inductance of the 

transformer. A parts list for the pulse generator is shown in Table 17.3. 

Higher voltages could be obtained by increasing the ratings of T2, D2, and 

C1. If the charging voltage were appreciably higher, the mercury-wetted contact 

relay RYI used for switching the charged capacitor into the transformer would 

52 1 
have to be replaced by a different device, either a higher voltage relay or some 

type of discharge tube. 

TABLE 17.3 PARTS LIST - SINGLE-POLARITY PULSE GENERATOR 

T3 - 

F - 

RYl - 

Yellow (power on) 

Red (pulse on) 

0.1 pF, 600V 

IN4003 (1 A, 200 V) 

IN4007 (1 A, 1000 V) 

6800 R, 1 W 

15KR,1 W 

27 R, 2 W 

SPST (power and relay on) 

SPST (pulse on) 

SPST (unidirectional or oscillatory) 

Variable autotransformer (1 20 V, 2 A) 

120/350 V, 40 mA 

(Any of a variety of plate transformers is satisfactory. 

Half or all of a center-tapped secondary may be used). 

Ferroxcube 1F53C5 U-core (2 required) 

120V, I A 

Mercury-wetted contact relay 

Clare HG 1002 or 

Potter and Brumfield JML 1200-8 1 

RYI MUST CLOSE TO CONTACT (a) ON 

HALF-CYCLE (a) OF THE SUPPLY VOLTAGE. 

CONTACT (b) SHOULD BE THE REST OR 

UNENERGIZED POSITION 

Gnd 

4 J I 

CONDUCTOR 

UNDER TEST 

Figure 1 7.16 Single-polarity current injection-pulse generator. 
17.6.2 Current-Injection Transformers 

Some of the parameters affecting circuit response are shown in Figure 

17.1 7 (Reference 17.31). These include the capacitance of the storage capacitor; 

the inductance of the lead connecting the storage capacitor to the transformer; 

the primary, secondary, and leakage inductances of the transformer; and the 

inductance and resistance of the conductor under test. 

The characteristics of the pulse transformer are among the most important 

properties. These parameters relate to the primary and secondary inductances, 

the flux in the transformer core, and the proportion of flux produced by current 

C = STORAGE CAPACITOR 

Lp = INDUCTANCE OF PULSE GENERATOR AND 

CONNECTING LEAD (0.77pH/ft FOR 500 

COAXIAL CABLE) 

L I = PRIMARY INDUCTANCE OF TRANSFORMER 

L2 = SECONDARY INDUCTANCE OF TRANSFORMER 

Ls = INDUCTANCE OF LEAD UNDER TEST 

Rs = RESISTANCE OF LEAD UNDER TEST 

#, = FLUX DUE TO CURRENT IN PRIMARY 

#,=FLUX DUE TO CURRENT IN SECONDARY 

) ,=LEAKAGE FLUX, FLUX ORIGINATING IN 

PRIMARY, BUT NOT LINKING SECONDARY 

Figure 17.1 7 Parameters determining circuit response. 

523 
in the primary that links the secondary winding. The difference in these two 

fluxes is the leakage flux, and it is the leakage flux that prevents the secondary 

current from being as large as the primary current. 

For a given transformer, the primary and secondary inductances may 

either be measured on an inductance bridge or be calculated by discharging a 

known capacitor through the winding and observing the frequency of oscillation. 

Mutual inductance, which is related to the degree of coupling between windings, 

can be determined by connecting the primary and secondary windings in series, 

first in series so that the two magnetic fields set up are in the same direction 

(series aiding) and then so the magnetic fields are in opposite directions (series 

bucking). 

Knowing the primary, secondary, and mutual inductances, one can 

produce an equivalent circuit of an injection transformer. One such circuit is the 

Pi circuit of Figure 17.18(b) (Reference 17.32). If the number of turns on the 

primary winding is equal to the number on the secondary winding (the usual 

case), the circuit reduces to that of Figure 17.1 8(c). 

The amount of open circuit voltage or short circuit current that can be 

transformed by a given transformer depends upon the magnetization curve of 

the core. A typical curve for a core without any air gap in the magnetic path is 

shown in Figure 17.19(a) (Reference 17.33). The operation of the core depends 

on the type of pulse to be produced. The most difficult operation involves a 

train of pulses, all of the same polarity. If the core is initially demagnetized, the 

first pulse will follow path a on the magnetization curve, whereas successive 

pulses will follow curve b. Since the amount of open circuit voltage that may be 

produced depends upon the change in magnetic field, operation along curve b 

cannot support as high a voltage as would be expected if one considers only path 

a, nor one lasting for as long a period of time. If alternate positive and negative 

pulses are applied, then operation along paths c and d will occur, and a much 

higher output voltage may be produced. If the first pulse carries the magnetic 

field along path a and then lets it relax to point 1, the second pulse will carry the 

magneti~ation along path c and let it relax to point 2. The third pulse would 

then carry it along path d and let it relax back to point I again. 

If there is an air gap in the magnetic path, the B-H curve is flattened out 

and the remnant magnetic flux in the core becomes very small. In such a case, 

the core can support a series of pulses all of the same polarity and amplitude, 

since operation along path e is virtually the same as that along path f. 

The B-H curve of a typical ferrite core suitable for a current injection 

transformer is shown in Figure 17.20 (Reference 17.34). The core dimensions 

are shown in the figure. The core was a ferrite U-core with a removable end 

yoke. Tf the end yoke butts closely against the pole pieces of the U-core, the 

core shows a clearly defined B-H loop. The initial magnetic path, corresponding 

to curve a of Figure 17.1 9(a), is identified as the zeru-milgap. With gaps of 1.8 

or 3.3 mils (.0018 or .0033 inches) on the two-pole faces, the curve flattens out, 

or becomes less steep. ln each case the core begins to saturate when the magnetic 

field reaches 1.8 x 10-5 W, corresponding to about 3500 gauss (lines per square 

centimeter). 

524 


AMPERE-TURNS ' 

Figure 17,20 Saturation characteristics of ferrite current injection transformer. 

The primary or secondary inductance is defined as the slope of the B-H 

curve: 

The inductance is proportional to the square of the number of turns in the 

winding. For the three curves the inductance would be as indicated. 

In the absence of a controlled gap in the magnetic path, the inductance of 

a winding depends largely on the amount of pressure exerted on the pole piece, 

since with more pressure any small residual gap is reduced. When the B-H curve 

shown in Figure 17.20 was obtained by exciting the core with 60 Hz ac, the 

magnetic forces involved held the poles closely together. The result is that the 

inductance that would be calculated from the zero-mil gap curve in Figure 

527 
17.30 would be higher than one that would be encountered if the core were 

excited by a short-duration pulse, one lasting only a few microseconds. With 

short pulses the magnetic field may be high and may exert a significant pressure 

on the pole pieces, but the inertia of the pole piece prevents the poles from 

moving together and reducing the residual air gap. 

The transformers are not constrained to be operated with only one turn on 

CONDUCTOR UNDER 

TEST 

Figure 17.21 Effect of more turns on secondary. 

Less short circuit current 

- More open circuit voltage 

- A decrease in the frequency of the natural 

oscillatory mode of the cable under test 

CONDUCTOR UNDER 

TEST 

\ 

Figure 17.22 Effect of more turns on primary. 

-- More short circuit current 

- Less open circuit voltage 

- A decrease in the frequency of the natural 

oscillatory mode of the pulse generator 

- A longer rise time of the current pulse 
each winding or, for that matter, with the same number of turns on the two 

windings. Figures 17.2 1 and 1 7.22 (Reference 17.35) give quaIitatively some of 

the ways in which turns on one winding or the other affect the results if the 

number of turns on the two windings are not equal. For instance, if the wire 

under test is looped through the core several times, a given excitation on the 

transformer will produce less short circuit current, more open circuit voltage, 

and a decrease in the frequency of the natural oscillatory mode of the cable 

under test. If there are more turns on the primary, there will be more short 

circuit current, less open circuit voltage, and a decrease in the frequency of the 

natural oscillatory mode of the pulse generator. There will also be a longer rise 

time of the current and voltage pulses. 
REFERENCES 

17.1 Space Shuttle Program Lightning Protection Criteria Document, 

JSC-07636, Revision A, National Aeronautics and Space Administration, 

Lyndon B. Johnson Space Center, Houston, Texas (November 4, 1975), 

p. A-5, and Lightning Test JVaveforms and Techniques for Aerospace 

Vehicles and Hardware, Society of Automotive Engineers, Warrendale, 

Pennsylvania (5 February 19761, p. 9. 

17.2 Space Shuttle Program Llgh ming Protection Criteria Document, p . 7-6. 

17.3 MIL-B-508 7-B: Bonding, Electrical and Lightning Protection for Aerospace Systems (August 3 1, 1970). 

17.4 Lightning Test Waveforms and Techniques for Aerospace Vehicles and 

Hardware, Society of Automotive Engineers, Warrendale, Pennsylvania (5 

February 1976), p. 2 1. 

17.5 F. A. Fisher, Some Observations Regarding Amplitude Scalirzg for 

Lightning Currents, Aircraft Lightning Protection Note 74-1, High 

Voltage Laboratory, Environmental Electromagnetics Unit, Corporate 

Research and Development, General Electric Company, Pittsfield, 

Massachusetts (14 June 1974). 

17.6 Fisher, Some Observations Regarding Amplitude Scaling, (n.p.1. 

17.7 Fisher, Some Observations Regarding Amplitude Scaling, (n.p.). 

17.8 L. C. Walko, A Test Technique for Measuring Lightning-Induced Voltages 

on Aircraft Electrical Currents, NASA CR-2348, National Aeronautics 

and Space Administration, Lewis Research Center, Cleveland, Ohio 

(February 1974). 

17.9 J. A. Plumer, F. A. Fisher, and L. C. Walko, Lightning Effects on the 

NASA F-8 Digital Fly -By-Wire Airplane, NASA-2 524, National Aeronautics and Space Administration, Flight Research Center, Edwards, 

California (March 1975). 

17.10 J. A. Plumer, Guidelines for Lightnirzg Protection of General Aviation 

Aircraft, FAA-RD-73-98, Federal Aviation Administration, Department 

of Transportation, Washington, D.C. (October 1973). 

17.1 1 Lightning Test Waveforms and Techniques, p. 20. 

17.12 Space Shuttle Program Lightning Protection Criteria Document and 

Lightning Test Waveforms and Techniques. 

17.1 3 F. A. Fisher, Equipment for Indirect Effects Testing, Aircraft Lightning 

Protection Note 75-3, High Voltage Laboratory, Environmental Electromagnetics Unit, Corporate Research and Development, General Electric 

Company, Pittsfield, Massachusetts (4 August 1979, p. 6. 

17.1 4 Space Shuttle Program Lightning Protection Criteria Document; Techniques for Dielectric Tests, ANSI C 68.1-1968/IEEE No. 4 (April 1969), 

the Institute of Electrical and Electronic Engineers, New York, New 

York (1969); F. A. Fisher and F. D. Martzloff, "Transient Control 

Levels: A Proposal for Insulation Coordination in Low-Voltage Systems," 

IEEE Transactiorrs orr Power Apparatus and Systems, PAS-95, I, 

(January/February 1976): 120-29. 
17.15 Space Shuttle Program Lightning Protection Criteria Document; Surge 

Withstand Capability (SWC) Tests, ANSI C 37.90a-1974ITEEE STD 

472-1 974; F. D. Martzloff and E. K. Howell, IJi-Voltage Impulse Testers, 

75 CRD 039, Corporate Research and Development, General Electric 

Company, Schenectady, New York (March 1975). 

17.16 ANSI C 68.1 -1968lIEEE No. 4 (April 1969). 

17.17 Fisher and Martzloff, "Transient Control Levels: A Proposal for 

Insulation Coordination in Low Voltage Systems." 

17.18 ANSI C 37.90a-1974. 

17.1 9 Martzloff and Howell, Ili-Voltage Impulse Testers. 

17.20 Proposed Requirements for Surge Tests on Ground Fault Circuit 

Interrupters, Underwriters Laboratories, Inc., Melville, New York (July 

1975). 

17.21 Fisher and Martzloff, "Transient Control Levels: A Proposal for 

Insulation Coordination." 

17.22 K. E. Crouch, F. A. Fisher, and F. D. Martzloff, Transient Control 

I,evels: A Better Way to Voltage Ratings in Power Converter Applications, TIS 76CRD154, Electronic Power Conditioning and Control 

Laboratory, Corporate Research and Development, General Electric 

Company, Schenectady, New York (July 1976), p. 2. 

17.23 Part (a): ANSl C 37.90a-1974; Part (b): Crouch, Fisher, and Martzloff, 

Transient Control Levels: A Better Way to Voltage Ratings, p. 3. 

17.24 Crouch, Fisher, and Martzloff, Transient Control I,evels: A Better Way to 

Voltage Ratings, p. 3. 

17.25 Adapted from ANSI C 37.90a-1974. 

17.26 EMP Susceptibility Threshold Handbook, Draft Report submitted by the 

Aerospace Group of the Boeing Company for the U.S. Air Force 

Weapons Laboratory, Kirtland Air Force Base, Albuquerque, New 

Mexico (July 2 1, 1972), pp. 18-25. 

17.27 Space Shuttle Program Lightning Protection Criteria Document, 

Appendix D. 

17.28 Fisher, Equipment for Indirect Effects Testing, pp. 20, 22. 

1 7.29 "Transient Control Levels: A Proposal for Insulation Coordination" and 

Hi-Voltage Impulse Testers. 

17.30 Fisher, Equipment for Indirect Effects Testing, p. 37. 

17.3 1 Fisher, Equipment for Indirect Effects Testing, p. 24. 

17.32 Fisher, Equipment for Indirect Effects Testing, p. 26. 

17.33 Fisher, Equipment for Indirect Effects Testing, p. 28. 

17.34 Fisher, Equipment for Indirect Effects Testing, p. 29. 

17.35 Fisher, Equipment for Indirect Effects Testing, p. 34. 

APPENDIX 1 

SR-52 CALCULATIONS - CONDUCTOR 

SR-52 

User Instructions 

PAGELOF 1 
SR-52 CALCULATIONS - CONDUCTOR (CONCLUDED) 

TITLE Canductor Temperatwe Rise and ElorpiIan PAGELOF, SR-52 PROGRAMMER J. A. Plum= DATE 25 October I976 
APPENDIX 2 

SR-52 CALCULATIONS - TUBE 

SR-52 

User Instructions 4~ 
SR-52 CALCULATIONS - TUBE (CONT.) 

TITLE Qoundiq n~ubc Ternpcrat~~c ire md Elo~ation PAGELOF 

DATE I6 March 1976 

SR-52 PROGRAMMER I. A. numcr 
SR-52 CALCULATIONS - TUBE (CONCLUDED) 

TITLE Groundimg Tube Tcrnpcratun Riw and Elongallon PAGE 'OF- SR-52 PROGRAMMER J.A.P1ume' DATE i6 Much lW6 Mq Fwm 

APPENDIX 3 

PROGRAM MA GFLD 

Table A3.1 LISTING OF PROGRAM MAGFLD 

................................................................... 

MAGFLD 10/L6/76 

I lo HEM I20 HEX 130 ;(EM 140 2Ehl 150 REM lo0 dEH I I0 EM 

l a0 rci!4 I90 HE!4 

200 QEM 210 HEM 220 REU 230 REM 232 REhl 234 HEM 236 HEM 23b HC'4 239 AEM 

FA FlSHCd GENERAL ELECTRIC COUPAIIY ENVIllONUENTAL ELECTRO AGNETICS UNIT BLDG 9-209 PITTSFIEL9,MA 0120 l PHONE (41 3 1-494-4380 02 DIAL COYhl 8-236-4370 

THIS PRoGHAM CALCULATES THE MAGNETIC FIELD INTERIIAL OR FXrERNAL TO A GR'jJP OF CURRENT CARRYING CONDUCTORS. ALL COIIDUCTORS AnE ASSJMED LONG EIIOUGH THAT END EFFECTS AdE NESLIGIBLE IT ALSO CALCULATES THE DISTPIBUTIOII OF CURRENT AMONG TH; CONDUCTORS. THE COliWCTORS AHE ALL ASSU'!ED TI1 BE E7UIDISTAKT FRI THE RETdRN PATH. T3E DISTANCE TO TO ME RETURN PATH IS ASSUME7 METCR. THE PROGRAM DOES IlIlT CALCULATE ANY PRoXI).IITY EFFECTS. 

DATA 15 STORED IN A FILE ;IHIC'I IS SIVE:I THC WAh(E DATFIL THE FIR,T RECOdO OF THE FILE SHOULD GIVE %E TOTAL NUMB~R OF CONDUCTORS 

SllCCESSlVE dECORDS OF THIS FILE SHOUL9 SlVE THE LOCATIOII OAND SIZE OF THE COll9UCToRS. THEY SHOULD sf IN THE FOR).! wrollnUCToR 

NUUbER" ( 1 3i20UGl PII 1, "X 3rXQDINATE1' (It( '!ETERS). "Y COOP71:.IAE" (IN I~IETILHSI ANO "CO:iDUCTo? DIA!!ETER8' (1i.l 'IFTERS) . 

240 RE!A A SAMPLc DATA FILE IS S~lo~lil UEL0,r. 241 RE14 THIS FILE DFSCRI:>ES FOUR COII~UCTII\?S EACH fl.Ol :IETEHS QIAUFTER. 242 KEsi AHHAJJGEI' AT-THZ CORiJEPS OF A SOUA~E'I MITER OY A 517E AIJn 243 HEY CEIITEHE I AT X=Y=O. 244 HEY 

242 REU 

240 REM 10W> 4 247 REM 101, 1, .5..5..01 Z4J {?EM 102: 2,.5,-.5,.9l 232 HE% 103"s 3,-.5,--5, -01 234 HE61 104'; 4, .5,-.5,.01 236 ifEl.( 

25d REM 

260 dEM THE LENC'TH OF THE CONDUCTORS IS NOT ENTERET). THEY ARE ASSUME9 m 202 REX BE flF 1,rFI;IITE LEVCTH. 270 2EM 

2dO REM THE PHIKRAU IS PilESENTLY I)IhlENSIoNEO Fo2 A YAXIYUM oF 290 HEM 50 ClIIII)JCTOt4S 

300 REM 

310 ttE84 THE IlUTi-UT IS S0;lMALIZED TO A TOTAL CUddiliT OF ONE AMPERE 320 dE!A FLOtrING TIIRIIUGH THE OlcoiiP OF CONDUCTORS. 

345 FILES DATFIL 3>0 DIY h~(50,50),V(i0,1~,1(50,11.~l50,4).N(50,50) 360 DIM J(SD),K(50) 303 HEAD# I .(I I 370 rlAr HEAD #1.D(I.il,4) 300 YAT V=ZEd(I.II ,I ) 

390 FOR 1-1 TI) . I 400 LET V(1.I )=I 
Table A3.1 LISTING OF PROGRAM MAGFLD (CONT.) 

-- ---- --- .- ------------- - ----------------- - ......................... 

YAT I=ZEH(lil .I I :.!AT N=ZER(NI .:I1 

PHIHT"COi.IDU,'TO;Iil,"X COORI)IIIATE"."Y CO!I;I?INA'TF"."DIA7EER-KflERSn I #### -F.###-AAHAT PhINT '3 

~hl!<T 

LET P2=6.20dld LET E2=2.71,.28 LET gl=300 FllH I=I To :I FOR J=I To ,.I IF I=J TriEN 383 

LET n2=SOd((D(I,21-~(J,21)*2 +(O(I,3)-D(J,3))^21 

LFr M(I,J)=I.OG(RI/HZJ 

GOT!) 590 

LET C(I,J)=:.OC;(RI*Z/D(I,4)) IdEXT J NEXT -.~ I 

,(AT N=IIIV(Y) MAr 1=1t*v LET Il*O 

FOd I= ITOiIl LET Il=I(I.I)+II NEXT I FOR I=i TO ..I Lrr J( I )=I ( I, I )/I IlEXT I PBI.1T 'rCOIID~CTOR" PHI!JT 

FObi I=l TO .I 

PdINT USING 455.1 IiEXT I PHii4T 

I 

. "FUACTIONAL 

PdlliT 

PHIIIT"TI1E iI HST CiiARACTEe OF A.4Y YEYEOAX! I :IPUT DATA IS A CO?iTROL" 

?dIliT"Ci(ARA<'TEbi. iidEW Ail INPUT OUESTIO!d ASKS FOR 'COIITl)OL',FI:TEil" 

Pd1NT"I (O..E) IF YllU NISti TO CollTINUE 4'iIiil THAT TYPE OF CALCUL-" 

PlcI!lT"ATIoii AND EWTER O (ZERO) IF YOU :.iISi To Go ON TO THE" Pi(INT"i4EXT iYPE [IF CALCULAi'I0;I" ?HINT 

HE'A THIS STi.r(TS CALCiJLATION OF FIELD AT A OEFINED POINT Pd1:IT"DO YO.J wAKT A POIIJT UY POIIIT CALCULATIOII OF THE FIELO?" Pr(IriTnYES=I i4o=O" 4 

I~lPiii Al IF AI=0 THE 1022 PBINl 

 EM TYIS STiUTS CALCULATIIlil OF THE FIELD AT A &TINED POINT PRINT X Y Y-X 3-Y H-T!)T AYGLF-OFG" 
Table A3.1 LISTING OF PROGRAM MACFLD (CONT.) 

930 PRINT 

940 REM XI AND YI ARE C(N)RDINATES OF POINT PI 950 PRINTUCONTROL, XI .Yl"i 

960 INPUT O,XI,YI 962 PRINT 970 IF 0-0 THE,, I020 

980 GOSUB 1048 

990 PHINT USING 1021 ,XI ,YI ,B2,B3,D4.B5 ION PHINT l 01 0 GOT0 950 

HEY THIS EiIDS ME CALCULATIoll rtF THE FIELD AT A DEFINED POINT I #. ###*-*- #.###"*"- #. ###**** #. *##**** #.###*-*- ####. ## 

REM THIS STARTS CALCULATION OF THE FIELD AT A SERIES OF POINTS 

PRINT"D0 YOU rlANT A CALCULATION OF THE FIELD AT A SERIES OF POINTS" 

PR1NT"COVErING AN AREA? YES=( .No=O~~f INPUT (I 

IF 0=0 THE11 1280 

PRINTrthHAT ARE START,SToP AND INCREMENT VALUES FOR X"I INPUT S1 ,Si,S3 

Pt4INTndHAT AGE START,STOP AND INCREMENT VALUES FOR Y"f INPUT 54, SJ, 56 PRINT 

DUTMT 

, ,,*... 

PRINTU X Y Y-X H-Y H-TOT AHGLE-DEG" 

FOR XI =S I TO S2 STEP 53 

FOR Y I =S4 TO S5 STEP S6 

GOSUB 1049 PRINT USING 1021 ,XI,Yl,B2,B3,84.B5 NEXT YI NEXT XI REM THIS ENDS THE CALCULATION OF FIELD AT A SERIES OF POINTS 

PdINT 

PRINT 

GOT0 1280 HEM THIS STARTS THE SUBROUTINE FOR CALCULATION OF FIELD AT A POINT LET B2=O LET B3=0 

1060 FOR 

1070 REM 1080 REM 

I090 REM 1100 LET 1110 LET 

1120 LET 1 130 LET 1135 HEM 1 137 REM 1140 LET 1150 LET 1 160 LET 

i3=~-COMPONENT BETWEEN POINT I AND POINT I DITTO FOR Y3 

R3= TOTAL DISTANCE BETWEEN POINT I AND POINT I 

X3=XI-D(I.2) 

Y3=YI-D(I.3) 

R3=SORIX3*X3+Y3*Y3 

BI=J(I)/(P2*R3) 

C2=INCHEMEKTAL,COMPONENT IN X DIRECTION 

C3=INCREMENTAL' COMPONENT IN Y DIRECTION 

C2-B I *Y3/R3 
Table A3.1 LISTING OF PROGRAM MAGFLD (CONT.) 

MAGFLD I 0/2 67 76 

1170 LET B3=83+i3 

lldO NEXT I l I90 REM B2= X COMPONENT OF TOTAL FIELD 1200 HEM B3=Y COMPIINENT OF TOTAL FIELD 

REM 84- TOIAL FIELD HEM B5=ANGLE [IF TOTAL FIELD LET 84=SQR182*!32+B3*83 ) 

dEM THE FOLLIINING STATEMENTS IF B2<>0 THEN 1240 IF B3<0 THtN 1236 LET B5=90 

LET ~5=-90 

GOT0 1260 LET B5=ATN 1 B3/B2 I LET 85=85*>7.29583 BEM IF B2>U ANGLE IS HEW IF B2<J AND 83~0 REU IF BZeO AND 83>0 IF 82>0 THtN 1260 IF 83~0 THkN 1258 

GUARD AGAINST DIVISION BY ZERO IF; LINE 1240 

IN FIRST OR FOURW QUADRANTS AND LINE THE ANGLE IS IN THE THIRD QUADRANT THE ANGLE IS IN THE SECOND OUADRANT 

LET B5=180+85 GOT0 1260 LET 855-183+85 RETURN 

HEM THIS ENDS THE SUBROUTINE FOR CALCULATION IIF THE FIELD AT A POINT PRINTnDO YllU IIANT THE FLUX INTEGRATED OVER A PATH BETNEEN" PRINTnTIIO POINTS. YES- l NO-O"I INPUT A2 IF A2-0 THtN 1400 PRINT 

HEM X4 AND Y4 ARE THE CtUlRDINATES (IF PI REM X5 AND Y5 ARE THE COUIRDINAES [IF P2 PRINT"CONT~0LX-PI .Y-PI ,X-P2.Y-Pa") INPUT O.X4.Y4,X5.Y5 IF 0-1 THEh 1470 PRINT 

PRINTnDO YOU dANT TO GO BACK AND DO MORE POINT BY POINT CALCULATIONS" PRINT1'(ENTER 21, DO YORE CALCULATIONS OF FLUX OVER AN AREA (ENTER I)" PRINT"0R SiOP (ENTER 0) 

INPUT A2 IF A212 THEN 090 IF A2-I THEN 1330 

IF A2-0 THEN 1720 REY THIS SrARTS CALCULATION OF FLUX OVER AN AREA LET - - - X5rX5-Xl 

. . - . . - . . LET Y6=Y5-Y4 LET R6= SOR(X6+X6+Y6*Y6) REY 01 IS 'THE ANGLE BETWEEN THE LINE JOINING PI AND P2 

REY AND THE X AXIS. SIL(OIl=Y6/R6 
Table A3.1 LISTING OF PROGRAM MAGFLD (CONCLUDED) 

1230 liE1.I All0 COS(01 )=X6/d6 

1240 RE.4 THIS SiAI1TS T!IE 1:iTEGHATIflN OVER THE AREA BETWEEN PI Alin P2 1342 LET 03~50 1545 LET 04=R6/1:3 

1320 FOh J=O TO 03 

1260 LET O2=J/O3 1370 LET XI =X4+\)2*X6 1330 LET Y I =Y4+ 12*Y6 

I390 OOSUB 1049 1600 HEM Dl AND D2 A2E THE COYPONEYTS OF H-X ANr) il-Y NORMAL TO THE 1010 dEH LIkE JOIIIING PI AEln P2 

1020 LET DI=-B2<Y6/R6 1630 LET D2=B3* .6/R6 

LET K(J)=DI+D2 INEXT J Lti 87=0 

fOR J=0 TO 03-1 

LET B7=(K(J)+K(J+I NEXT J LET B7=2E-/*P2*B7 1690 PttII.IT"TOTAL FLDX="~B~I"IVEBERS" 1700 PI~INT 171 0 GOT0 1360 1720 END 

I 

0 1 0 ONE OF FOUR CONDUCTORS. 

EACH CARRYING 0.25 A 

I 

I 

Figure A3.1 Pattern of field around the conductors described in lines 246 to 

254 of Table A3.1. 
Table A3.2 SAMPLE OUTPUT FROM PROGRAM MAGFLD 

MAGFLD 09127EDT 10/26/76 

CONDUCTOR X COORDINATE Y CIHIRDINATE DI AMETER-METERS 

CONDUCTOR r-HACTI ONAL CURRENT 

THE FIRST CHARACTER OF ANY KEYBOARD INPUT DATA IS A CONTROL 

CHAHACTER. nHEN AN INPUT QUESTION ASKS FOR 'CONTR0L8,ENTER 

I (ONE) IF YOU WISH TO CONTINUE NITH THAT TYPE OF CALCULATION AND ENTER 0 (ZERO) IF YOU dISH To GO [IN TO THE 

NEXT TYPE OF CALCULATION 

00 YOU NANT A POINT BY POINT CALCULATION OF THE FIELD? 

YES=l NO=Ol l 

X f H-X H-Y H-TOT ANGLE-DEC 

Do YOJ #(ANT A CALCIILATION OF THE FIELD AT A SEHIES OF POINTS 

COVEhING AN ARM? YES=l.NIl=O? I 

dHAT ARE STAdT SLOP AND IhCREh'ENT VALUES FoQ X7 0 1.2 4 

MAT ARE START:STOP AND IilCHEYENT VALUES FOR Y? 0:1.2::4 

H-X 

. I49E-W 

.370E-01 

.124E+00 

. llYE+00 

H-Y -. 186E-08 

. I4I)E-Ob 

. I PEE-06 

.649E-09 

TfiE FLJX I lTEG2ATFD 

YES=I iiO=O? ! 

DO YOU iiAiuT TO 30 bACK AX3 L)o 90BE POIIVT fry POI:IT CALCULATIOPiS 

(EiTcd 21, 30 S!I*E CALCULAT1OI:S OF FLilX OVE? A!l AREA (E?ITEl! I I 

0.2 STOP (EIITER :) 

7 i) 
INDEX 

Access doors, effects on magnetic 

field, 327 

sparking at, 159-164 

Action integral, 82, 188, 190 

Airworthiness regulations, 105-1 13, 150 

Aluminum foil, as diverter strip, 222 

as protective coating, 246 

shielding effectiveness, 241, 243, 244 

Anodizing, effects on conductivity, 

163, 166 

effccts on couplings, 168 

Aperture coupling - Electric fields 

cquations, 343-344 

shape factor, 34 1 

Apcrturc coupling Magnetic fields 

dipole approximations, 329-332, 

335 

cquations, 333-334, 336-337 

equivalcncc of aperture shapes, 340 

examples, 342-343 

patterns of field intensity. 333 

reflecting surfaces, 337, 338 

shape factor, 332 

Arcing (See sparking) 

Attachment of flash (See attacl~ment 

zones), as influenced by diverter 

strips (See diverter strips) 

initial attachment, 48, 49, 53 

on nonmetallic surfaces, 207-209 

to fuel tanks (Scc ,fuel tanks) 

to radomes (Scc radomes) 

Attaclnncnt zones, 50-53, 105-108, 

132-1 34 

defined, 5 1, 106 

dcter~ninctl by scalc models, 106 

field experience, 107, 109 

ligl~tning environment in, 1 10-1 1 1 

swcpt-stroke Lone, 50-53, 106, 107, 

108, 109 

Avoitlancc, of storms, 59, 65-71 

of strikes, 69-70 

I%ontling, jumpers, 204 

rnagnctic force cffccts, 201-203 

of access doors, 163, 164 

of composites, 240 

of fuel caps, 160 

of hinges, 205 

of pipe couplings, 169, 170 

on fuel dump pipes, 135 

via fasteners, 165 

Boys camera, 8,9, 149 

Breakdown, attachment point tests, 

211-213 

of air, 7, 174 

of canopies, 88 

of electrical insulation, 9 1, 93 

of fiberglass, 87, 2 16 

of spark gaps (See protective devices) 

Canopies, puncture of, 88 

scorching of, 90 

strcamcrs inside, 88 

Capacitance, aircraft, 40 

sphere, 39 

Channcl of lightning, initial development, 5-9 

pressure of, 83 

return stroke, 9-14 

temperature of, 83 

Charge, delivered by stroke components, 16, 19, 25, 27 

in leader, 7, 41 

on aircraft, 40, 41 

produccd by engines, 48-49 

to burn through metals, 136, 143, 

152 

Circulating currents, 304, 306, 307, 

31 1, 

Coatings, 224, 245-250 

('omposite materials, boron, 230-233 

dielectric coatings for, 248-249 

electrical modcls of, 231, 232 

cncrgy absorption, 229 

cxamplcs ofdamage, 234, 235, 237 

extent of damage, 236,268, 239 

graphite, 232, 237-238 

invisible damage, 233, 235 

joints, 240 

n~ccl~anisn~ of failure, 87-89,230,233 
protective coatings, 245-247 

shielding properties, 24 1-245 

temperature rise, 230 

Conductivity, at access doors, 163-164 

at joints, 165-167, 240 

of composites, 229-230, 233-238 

of metals, 190, 303 

through couplings and interfaces, 

171-172 

Connectors, grounding of, 394,400,404 

transfer characteristics, 394, 399 

Continuing current, 18 

effects of, 78, 136-154, 186-188 

Corona, 3, 4,44,49 

Coulomb ignition thresholds, 140-143, 

150, 152 

Covers, access (See access doors) 

effects of internal fields, 325 

sparking at, 163 

Current distribution, 280,283, 287, 291 

in ~nternal conductors, 160-171 

time conqtant of redigtribution, 

305-308 

Damage Direct effects, from shock 

waves, 85, 86 

to con~posites, 88, 84, 230-240 

to fiberglass, 83, 87,92, 206-213 

to metal surfaces, 78, 136-154 

to radomes, 83, 85, 87 

to structures, 79-86, 199-201, 

206-213 

Damage - Indirect effects, 98 

constants, 453,455-462,465-492 

examples of, 92, 99-101 

to capacitors, 463-464 

to semiconductors, 45 1-462,465-492 

Dart leader, 18 

Design to minimize indirect cffects, circuit design, 438-440 

in fuel systems, 172-179 

location of equipment, 429-430 

location of wiring, 430-433 

premises for design, 429 

protective devices, 441-446 

shielding, 176-179, 433-437 

transient coordination, 446-448 

Diffusion (See pulse penetration) 

Direct effects, defined, 76, 77 

on fuel systems (See fuel systems and 

fuel tanks) 

on composites (See composites) 

on metallic surfaces (See metallic 

surfaces, fuel systems) 

magnetic force, 78-79, 81 

melting and burnthrough, 78, 80 

pitting, 79 

resistance heating, 82 

Diverters, application of, 2 13-220 

electromagnetic interference from, 

223 

foil, 222-223 

inductance of, 222 

internal, 22 1 

metallic coatings, 224 

on composite materials, 247 

on wing tips, 132, 156 

required size, 220 

segmented, 225-226 

solid bar, 220-221 

tests on, 2 18-2 19 

thermal elongation, 194, 195 

types of, 220-228 

woven wire, 224 

Dwell time (See swept strokes) 

Eddy currents, 299, 303, 309 

Electric field, as influenced by aircraft, 37, 39 

at aircraft surface, 38-41 

at surface of a sphere, 40 

coupling through apertures, 341- 

342 

effects on circuits, 353-357 

external intensity, 353 

on interior of cylinders, 298-302 

on interior of shield, 369 

Electron density, 48 

Elljptical conductors, current distribution, 274-276 

internal magnetic field, 288, 289, 

309-320 

tangential field intensity, 274-276, 

287 

Engincs, lightning cffects on, 75, 96, 

97.98 

Explosion, in fuelvent systems, 127- 

132 

in fuel tanks, 115-1 16, 162 

of conductors, 82, 84, 188-193 

of radomes, 83 
Fasteners, effects on magnetic fields, 

325 

needed to carry lightning currents, 

166 

Flame arresters, configurations, 124, 

125 

location of, 125, 126, 133 

Flame sprayed aluminum, on composites, 247 

shielding properties of, 243, 244 

Flame suppression, 124-126 

Flammability, limits of jet fuels and 

gasolne, 1 15-1 20 

Flash blindness, 7 1, 75 

Flash, lightning, charge formation, 2 

components of, 1, 2, 18 

direction of, 18 

influence of aircraft, 37 

influence of ground, 15 

polarity, 18, 19 

statistics (See statistics, lightning) 

types of, 1, 19 

F-8 aircraft, 41 3-425 

F-89 aircraft, 407-41 3 

Filler caps, protection of, 159-163 

sparking at, 92-95, 159-163 

Fuel quantity probes, 156-158, 160, 

175-179 

Fuel lines, bonding of, 134, 135 

Fuel systems, 89-91 

protection of, 115-180 

wing tip tanks, 93, 95, 153-159, 

162 

Fuel tanks, attachment zones related 

to, 132-134, 153-156 

diverters on (See diverters) 

hot spots from arcs, 136 

inerting of. 120 

integral fuel tanks (See integral 

fuel tanks) 

nonmetallic, 154-1 59 

objects within, 157-158 

puncture of, 154-159 

required skin thickness, 136, 154 

streamers in, 156 

titanium vs aluminum, 140 

trailing edge, 153-154 

wing tip, 93, 95, 155-159 

General aviation aircraft, strike to, 92 

Hinges, 203-205 

Ignition of fuel, effect of altitude, 117, 

119 

effect of temperature, 116-120 

energy required, 90 

flame retardant foams, 9 1 

flammability of fuel, 11 5-1 20 

from vents, 91, 121-134 

fuel probes, 90 

hot spots, 136-153 

resulting from filler caps, 139-162, 

190 

sources of ignition, 120-179 

Indirect effects, defined, 76, 77,99- 

101 

examples of, 92, 98, 100 

Induced voltages, on external loops, 

294, 297 

on fuel probe wiring, 177 

on internal loops in circular cylinders, 293. 296, 298 

on internal loops in elliptical cylinders, 3 11-3 14 

Inductance, of bond straps, 135, 164, 

170 

of diverter strips, 2 14 

in ground conductors, 403-404 

mutual, calculation of, 280, 401- 

402 

self, calculation of, 280, 307, 323- 

324 

transfer (See transfer inductance) 

Interaction, of aircraft with the lightning 

channel, 42-43, 49-50, 85-86, 

140, 144-149,254-259 

Interception of lightning, altitude, 57-58 

in clear air, 1, 59-60 

flight conditions, 59 

as influenced by temperature, 58, 

63-64 

as influenced by time of year, 63 

as influenced by weather, 59-64 

warning symptoms, 65-7 1 

Ionization, at aircraft extremities, 

38-46 

of engine exhaust, 47-49 

of leader, 6 
Jointh, g~~idclities for design, 166-167 

in pipes and couplings, 168-173 

on skins, spars, rib\, 165-1 67 

Laplace's equation, 276, 278 

I.attice diagrams, 258 

Leader, as influenced 1)y aircraft, 37-38 

charge on, 7,9 

dcvcloptitcnt of, 5-6, 8 

dialiietcr of, 7 

electric field strc~~ptlt, 7 

propagation of, 38, 12-43 

resistance of, 9 

step-, 5-9 

velocity of, 9 

protection from te~nperaturc rise in, 

188-193 

Models, for attachment point studies, 

52-53, I08 

for determining estcrnal current 

density, 285-286 

of liglitning flaslies, 31-33, 110 

Nonmetallic materials (See carlopies, 

composites, fiberglass) 

Orbiter (See Space Skiirtle) 

MAGFLD, coniptlter prograni, 280 

Magnetic fields -- euternal, cleliientary 

concepts, 269-280 

esamples of, 277, 285, 291 

intlucncc of radius of curvature, 271 

methods of determining, 274, 284 

Magnetic fields internal, apcrturecoupled, 261-262, 329-344 

diffusion-coupled, 293-326 

esperiment and nicasurcment. 3 14- 

320, 41 3-425 

in cavities, 310-326 

pulsc pcnctration titlie constant, 

298-303 

redistribution time constant, 305 

Magnetic fields, mcasurcrnent of, 420- 

4 26 

Magnetic field zones, 341-344, 359 

Magnetic forces, at joints, 200 

betaeen conductors. 196 

esamples, 8 1, 197 

on bond straps, 201-203 

on riletal surfaces, 195-203 

Magnetos, 75 

Metals, properties of, 150-152 

Metallic surfaces, holes burned in, 

136-140, 187 

magnetic force effects, 195-203 

pitting of, 79, 138 

protection from holes, 141, 149- 

153, 186-188 

P-sta tic (See precipilatiotr static) 

Precipitation static, 42-43 

a7 influrnced by btatic dischargers, 

45 

frequency spectrum, 45 

Properties of metals (See rndals, 

properties on 

Propellers, 96-97 

Protcctive devices fur wiring, forward 

conducting diodes, 444 

nonlinear resistors, 442 

varistors, 442 

Zener diodes, 443 

Radar returns, 65, 68 

Radomcs, 87, 89, 207-217, 225-228 

attachment point tests, 208-2 12 

da~iiage to, 83, 87, 208 

Rain gradient, 66 

Reflections, of apertures, 338 

of pressure waves in fuel vent systems, 118-1 29 

of traveling waves, 254-260 

resulting from protective devices, 44 1 

Requirements, airworthiness related 

to liglitning, 11 1-1 13 

Resistance, 261, 264-266 

of bond straps, 164 
of composites (See conductivity of 

composites) 

of hinges, 205 

of metals (See metals, properties on 

Resistivity, influence of, on pulse penetration time, 298-302 

on redistribution tirne, 308 

on temperature rise, 188-190, 193, 

230-231 

Resistivity of typical metals, 303 

Restrike, 18 

Return stroke, current, 1 1 

development of, 9-14 

resistance, 12 

surge impedance, 14 

velocity of propagation, 10 

voltage gradient, 12 

St. Elmo's fire, 4, 44, 69 

Segmented diverter strip (See diverters) 

Shielded conductors, braided shields, 

effects of grounding, 355, 362- 

368, 385-393, 437 

effects of load impedancc, 362-369, 

387, 390-393 

multiple shields, 385-393 

response to current, 363-368 

transfer admittance, 380-381 

transfer inductance, 368-379 

Shielded conductors, solid shields, 298- 

303,382 

Shielded conductors, spiral shields, 

381-382 

Shock wave, 83, 86 

Single point grounding, 765 

Skin depth, 775 

Space Shuttle, 33, 34, 258, 344, 356, 

507,508 

Sparking, at access doors, 163-164 

at couplings, 168-173 

at fuel filler caps, 161-162 

at joints, 165-168 

voltage required for, 170, 173, 175 

Static discharge, 37, 45, 46 

forewarning by precipitation static, 

45-46 

Statistics - Aircraft, frequency of, 

strikes to, 71-72 

Statistics - Characteristics of lightning, 

amplitude of return stroke, 21 

charge in continuing current, 14 

charge per flash, 25 

duration, 23 

frequency of occurrence, 27-3 1 

front time, 15-16, 21 

interval between strokes, 24 

rate of rise, 22 

return strokes per flash, 24 

tail time, 15-17, 23 

Statistics, strikes to aircraft, altitude, 58 

flight condition, 59 

frequency of occurrence, 71-72 

temperature, 64 

time of year, 63 

weather, 60, 64 

Stoichiometric mixture, 123 

Streamers, discharge, 5 

from helmets, 88 

near fuel vents, 132-1 36 

on nonmetallic surfaces, 207-2 13 

Swept strokes, as affected by environment, 110 

dwell time, 143-147 

influence of air velocity, 146-147 

influence of coatings, 232, 140, 147 

mechanism, 49 

on rivets, 166-167 

ways to reduce dwell time, 153 

zones, 106-107 

Synoptic meteorological conditions, 59 

Tests, of diverters, 2 18-2 19 

to observe sparking, 17 1 

to verify protection, 1 10-1 f 3, 180- 

181,446-450, 495-529 

Test equipment, electric and magnetic 

fields, 5 19 

surge withstand capability (SWC), 

508, 513,515 

transformer injection of transient, 

514-529 

transient control level, 447, 448, 

509-5 14 

Tests for indirect effects, full-scale on 

complete vehicles, 499-501 

transient analysis tests on complete 

vehicles, 499-504 

transient control level, 446-450 

waveform definitions, 505-5 11 

Tests for direct effects, airworthiness 

certification, 11 1-1 13 

diverter placement, 2 13-2 19 
fiberglass structures, 207-2 19 

fuel system protection, 180-181 

full-scale attachment, 207-212 

swept-stroke, 146-147 

Thermal expansion, 194-195 

Thunderstorms, cells, 2-5, 58, 65-68 

frequency of occurrence, 27 

Time lag curve, 209, 2 10, 2 1 1 

Transients, analysis technique, 4 13-426, 

501-504 

control levels, 446-450,5 1 1 

specifications, 507-5 11 

tests, 5 11-529 

Triboelectric charging, 43 

Triggering of lightning, by an aircraft, 

37,47 

by Apollo 12, 48 

by electromagnetic radiation, 49 

by jet engine exhaust, 47-49 

Turbulence, 64, 66 

Varistors (See protective devices for 

wiring) 

Vcnt outlets (See also ignition of fuel, 

fuel systems), Atlantic Research 

study, 123 

effects of air flow, 121, 126-1 27 

effects of turbulence, 129 

explosive ignitions, 127 

location of, 132-1 33 

Lockheed -- LTRl study, 126 

velocity of flame propagation, 123, 

I25 

velocity of flames, 127, 130-131 

Wire fabric, cffects of coatings, 246 

mechanism of operation, 234, 246 

Wiring, aid to calculation of current and 

voltage, 357-359 

current induced upon, 351 -352 

influence of location, 179,430-433 

in fuel tanks, 172, 179 

response to electric fields, 35 3-357 

response to magnetic fields, 349-353 

transmission line effects, 351-353 

voltages induced upon, 349-350 



For sale by the Nat~onal Techn~cal Inlormalio~i Service Sprinef~eid Virgin~a 22161 

1 Report No 

NASA RP -1008 

-- 

2 Government Accarnon No 3. Recipsenr'r Catalog No 

. .. 

4 Tttie and Subtntle 5 Report Date 

LIGHTNING PROTECTION OF AIRCRAFT October 1977 

6 Performing Organization Code 

7 Authorlrl 

Franklin A. Fisher and J. Anderson Plumer 

9 Performing Organizafron Name and Addrerr 

General Electric Company 

100 Woodlawn Avenue 

Pittsfield, MA 01201 

- - 

8 Performing Organization Report No 

10 Work Unct No 

11 Contract or Grant No 

13 Type of Report and Perlod Covered 

12 Sponsoring Agency Name and Addrerr 

National Aeronautics and Space Administration 

Washington, D. C. 20546 

Reference Publication 

- 

14. Sponsoring Agency 

- 

Code 

15. Supplementary Noter 

16 Abrtiact 

This handbook summarizes the current knowledge concerning the potential lightning effects on 

aircraft and the means that are available to designers and operators to protect against these 

effects. The impetus for writing this book comes from two sources - the increased use of 

nonmetallic materials in the structure of aircraft and the constant trend toward using electronic equipment to handle flight-critical control and navigation functions. NonmetaIlic structures are inherently more likely to be damaged by a lightning strike than are metallic structures. Nonmetallic structures also provide less shielding against the intense electromagnetic 

fields of lightning than do metallic structures. These fields have demonstrated an ability to 

damage or cause upset of electronic equipment. The persons who can best use information on 

aircraft protection from lightning are the aircraft designers and operators, but generally they 

are not among those who produced this information. Moreover, they are often unaware of its 

existence, and they seldom have the background to distill from it the important facts that can 

and should be applied to achieve safer designs. The purpose of this book is to present the 

most important parts of this body of knowledge in a manner most useful to the designer and 

the operator. 

17 Key Words ISuggested by Authorlrl) 

Lightning; Thunderstorms; Electronic 

equipment; Aircraft equipment; Aircraft 

instruments; Aircraft safety; Aircraft 

structures 

18. Distribution Statement 

Unclassified - unlimited 

STAR Category 03 

19. Security Clarsil iof th~s report1 

Unclassified 

22 Price' 

A2 4 

20 Security Clarr\f (of this pagel 

Unclassified 

21 No of Pager 

560