NASA Contractor Report 3329 

NASA 

CR 

3329 C.1 

Improved Test Methods-- 

Lightning-Induced Voltages in Aircra 

K. E. Crouch and J. A. Plumer 

CONTRACT NAS4-26 13 

SEPTEMBER 1980 
NASA Contractor Report 3329 

TECH LIBRARY KAFB, NM 

1llIIllIllIIIlllIlll

00bll952 

Improved Test Methods for Determining 

Lightning-Induced Voltages in Aircraft 

Prepared for 

Dryden Flight Research Center 

under Contract NAS4-2613 

NASA 

Nalional Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1990 

CONTENTS 

INTRODUCTION .................. 

OBJECTIVES ................... 

BACKGROUND ................... 

EXPERIMENTAL INVESTIGATION ........... 

Full Scale F-8 Test Setup ......... 

F-8 DFBW airplane ........... 

Aircraft instrumentation ........ 

Basic test configuration ........ 

Test circuit instrumentation ...... 

One-Tenth Scale Model F-8 Test Setup 

Relative geometric scale models .... 

Model construction ........... 

Model test circuit ........... 

Test Technique Evaluations .......... 

Return conductors ............. 

Transmission lines ............. 

Traveling wave transit times ........ 

Grounded versus ungrounded aircraft 

Aircraft resistively terminated versus short 

circuited to return conductors ....... 

Improved Test Circuit Evaluations ....... 

Lumped parameter ladder network ...... 

Inductance and capacitance selection .... 

Number of segments ............. 

Leader attachment simulation ...... 

Return stroke simulation ........ 

Induced voltage relationships ..... 

Streamer initiation effects ...... 

Page 

1 

2 

3 

9 

9 

9 

9 

15 

17 

23 

23 

29 

29 

33 

33 

39 

45 

52 

52 

56 

59 

59 

60 

62 

65 

68 

75 

iii 
CONTENTS - Continued 

Extrapolation of induced voltages 

Further investigation of transit times 

Simulation of longer lengths of leader channel . . . 

Additional cause-effects evaluations using the model. 

RECOMMENDATIONS . . 

CONCLUSIONS 

REFERENCES 

Page 

75 

77 

84 

86 

90 

96 

98 

iv 
I .

 

AWG 

BCS 

C 

CT 

d 

DFBW 

DFCS 

e 

f 

h 

i 

I 

iL 

L 

R 

LTA 

MJ?C 

R, RF> RR’ Rs’ Rt 

Zo’ ZL’ 'A/C 

SYMBOLS 

American wire gauge 

backup control system 

capacitance, farads 

current transformer 

diameter of a conductor, meters 

digital fly by wire 

digital flight control system 

voltage, volts 

frequency, hertz 

distance between 

current, amperes 

conductors, m 

peak current, amperes 

simulated lightning current, amperes 

inductance, henries 

length, meters 

lightning transient analysis 

mode and power control 

resistance, ohms 

time, seconds 

period of oscillation, seconds 

impedance, ohms 

surge impedances 

(also characteristic impedances), ohms 

V 
SYMBOLS - continued 

electrical conductivity, mhos/m 

permittivity, F/m 

permeability, H/m 

traveling wave transit times, seconds 

vi 
INTRODUCTION 

The increasing use of solid state electronics to perform 

flight control and other critical functions in aircraft has led 

to increased concern about the reliability of such electronics 

under adverse environments. Among the potentially hazardous 

environments are the surge voltages induced in aircraft electrical wiring by lightning strikes. Electronic microcircuits 

operate at very low power levels and cannot tolerate surge 

voltages of the magnitude that can be induced by lightning. 

To learn more about lightning-induced voltages and how to 

control them, a series of NASA-sponsored programs was conducted 

during the years 1967-72. This research led to a better understanding of how lightning causes interference in aircraft electrical circuits and development of a nondestructive method, 

known as the lightning transient analysis (LTA) test, for determining the magnitudes of possible lightning-induced voltages in 

various electrical circuits. 

In response to a concern about fly-by-wire system vulnerability in particular, an LTA test was performed on the NASADryden Flight Research Center (DFRC) F-8 digital fly-by-wire 

(DFBw) aircraft in 1974. The aircraft was subjected to electric 

currents which were similar in waveform to lightning return 

strokes but greatly reduced in amplitude. The test data were 

then extrapolated to correspond with severe lightning strike currents. 

The F-8 data was widely studied and prompted positive efforts 

to design lightning protection into aircraft then being designed 

with fly-by-wire control systems, including the USAF/General 

Dynamics F-16 and the NASA space shuttle. 

However, due to the lack of an apparent relationship between 

the driving test current and the voltages it induced in some electrical circuits, concern was expressed about the validity of the 

test method and the influence of the surrounding facility on results. Questions have also arisen about the validity of linear 

extrapolation of test data to predict induced voltages under fullscale lightning stroke conditions, and about whether the test 

current being driven through the aircraft is in fact the primary 

cause of most induced voltages. 

The purposes of this program, therefore, were to obtain a 

better understanding of tne several cause-effect mechanisms inherent in LTA testing, develop ways to minimize unwanted facility 
effects, and define improvements that enable a better simulation 

of natural lightning. The program was conducted in three experimental phases and utilized the NASA F-8 airplane as a test bed. 

Besides incorporating a typical flight-critical electronic system, 

the F-8 has been utilized in prior LTA tests and a large amount 

of data produced by the original LTA technique was available for 

baseline reference. The first and third experimental phases of 

this program were conducted on the F-8 at Dryden using transportable test equipment and instrumentation. The second experimental 

phase, which utilized a one-tenth scale model of the F-8, was conducted at Lightning Technologies, Inc. The program was conducted 

during the period February, 1979 through February, 1980. 

OBJECTIVES 

The primary objectives of this program were to obtain a better understanding of the basic mechanicms that related induced 

voltages appearing in aircraft electrical circuits to various test 

circuit parameters, and to utilize this knowledge to define improvements in the test method, such that overall validity is 

increased. 

Of particular interest were the following relationships: 

l The cause(s) of higher frequency oscillations that 

occur within a microsecond or so after current is 

applied to the aircraft. 

0 The effects of test circuit return conductors (not 

present in natural lightning but a necessity in most 

tests), and the nature of the transmission line thus 

formed. 

l The nature and effects of traveling waves in the airframe. 

The following improvements in simulation were also desired: 

l Bring the test current and voltage into a proper 

relationship. 

l Bring the traveling wave currents into proper 

relationship with the return-stroke currents. 

l Enable simulation of lightning leader effects as well 

as the return stroke. 

l Minimize unwanted facility effects. 

2 
BACKGROUND 

Concern for the effects of lightning-induced voltages on aircraft electronic systems has been increasing with the advent of 

fly-by-wire flight controls and other systems which utilize electronics in flight-critical functions. Lightning strikes have 

already demonstrated an ability, on occasion, to disrupt aircraft 

electronics, and several trends in electronics and airframe design 

may aggravate the situation further unless designers are made 

aware of the lightning effects environment their equipment must 

survive in. 

To learn more about lightning-induced voltages in aircraft 

electrical circuits and how to control them, a series of NASAsponsored research programs was conducted during the years 1967-72. 

This research led to a better understanding of how lightning 

causes interference in aircraft electrical circuits (refs. 1 and 

2) and development of a nondestructive method of simulating a 

lightning strike to the aircraft for the purpose of determining 

the susceptibility of particular circuits (ref. 3). This has become widely known as the LTA test. 

Since the aircraft must be available before the susceptibility of its electrical circuits can be determined by the LTA test, 

a need exists for analytical techniques with which to predict this 

susceptibility during the design phase. Contractual efforts to 

develop such tools have been underway since 1972 (ref. 4) but 

these have borne little fruit thus far. The difficulty of solving 

this problem analytically is recognized, and proven techniques are 

not close at hand. 

In response to the concern about fly-by-wire system vulnerability, an LTA test was performed on the NASA F-8 aircraft in 1974 

(ref. 5). The aircraft was subjected to electric currents which 

were similar in waveform to lightning return strokes but greatly 

reduced in amplitude. The test data were extrapolated to correspond with average (30,000 A) and severe (200,000 A) lightning 

stroke currents. The extrapolated data is summarized in Table I. 

The F-8 data was widely studied, and prompted positive 

efforts to design lightning protection into the aircraft then 

being designed with fly-by-wire control systems: the USAF/General 

Dynamics F-16 (ref. 6) and the NASA space shuttle (ref. 7). 

Nevertheless, the magnitude of the voltages measured in the F-8 

and the apparent absence of a logical relationship to the injected 

test current has caused some skepticism regarding the validity of 

3 
the basic LTA test and the authenticity of the lightning strike 

amplitude and waveform simulated in the test. 

TABLE I - RANGE OF PEAK INDUCED VOLTAGES IN NASA F-8 LTA TEST 

(Scaled to 200 kA) 

[From ref.51 

INDUCED VOLTAGE AMPLITUDE 

INTERFACE (Peak Volts) 

MS. MAX. 

Stick Trim and MPC Inputs to DFCS (525) 233 900 

Stick Transducer Inputs to DFCS (P4) 40 87 

DFCS Control Outputs (52) 233 400 

BCS Control Inputs (512) 222 422 

Mode and Power Control (515) 833 1132 

Mode and Power Control (514) 213 732 

Power Dist. Bay (+28VDC Bus) 160 200 

DFCS Ground to A/C Ground 0 666 

The occurrence of return strokes, such as the one represented 

by the 2 x 50 microsecond waveform (time to reach crest x time to 

fall to l/2 crest) used in the LTA test has been confirmed (ref. 8). 

Much less is known about the nature of streamer and leader currents occurring in the airframe, and no attempt was made during 

the LTA test development to simulate these currents. Whereas 

their amplitudes are almost certain to be less than return strokes, 

their rise and decay times may be considerably shorter, producing 

substantial rates of rise. 

Flight research programs (.ref, 9) are getting under way to 

learn more about these currents and other characteristics of lightning by flying instrumented aircraft near thunderstorms, but a 

number of years will be required to accumulate meaningful data. 

In the meantime, the trend toward further miniaturization of 

electronics anduse of nonconducting materials in airframes has increased the concern over induced voltages. This necessitates that 

questions of LTA test validity be resolved and the test improved, 

if necessary, so that it can be used with greater confidence. 

4 

..-_. .a,-.. .,.a-, , 
Most of the questions about the validity of the LTA test have 

arisen because there appears to be little relationship between the 

test current and some of the voltages it induces, particularly in 

parallel-pair, single-point grounded circuits such as are employed 

in flight control systems. 

A typical LTA test circuit is illustrated in Figure 1 and examples of voltages induced by the test current in two aircraft 

electrical circuits are shown in Figure 2. 

Figure 1 - Typical Lightning Transient 

Analysis Test Circuit. 

Figure 2 - Relationships between Test Current and 

Induced Voltages. 

5 
The voltage measured in oscillogram (b) displays a structural voltage rise (iLR,) component that is in phase with the 

simulated lightning current, plus a magnetically induced component (d@/dt) which is proportional to its derivative. This is 

indicated by the shorter rise time and undershoot of the induced 

voltage waveform. 

The voltage measured in oscillogrsm (c), however, oscillates 

at a high frequency about the zero-axis and bears no evident 

relationship to the lightning current. Circuits which display 

voltages such as shown in (b), where the familiar relationship 

exists, nearly always have a relatively large loop area, such as 

occurs when a single wire is used to transmit a signal (or power) 

and the airframe is utilized to return it. 

But when a separate wire is used for the return and routed 

next to the first one, the area between the two is small and the 

familiar derivative voltage is greatly reduced, There remains, 

however, a high frequency component of short duration occurring 

immediately after t = o, as shown in (c). When extrapolated to 

full-scale levels, these high frequency voltages sometimes exceed 1,000 volts - a level that cannot be tolerated by many 

electronic components. 

Some aircraft designers, in disbelief, have rationalized 

that these high frequency voltages are peculiar to the LTA test 

only and should not be extrapolated upward by the same factor as 

the familiar voltages are. Other designers have incorporated 

shielded cables or surge suppressors to deal with these voltages 

on the assumption that they are authentic. Such measures, of 

course, add to the cost and weight of the aircraft. Thus a 

better understanding of the origin of the high frequency voltages 

that occur when test current first enters the airframe is required. 

Since the frequencies of some of the oscillatory voltages 

appeared related to the length of the airframe, it has been 

thought that traveling wave currents in the airframe may be the 

source of the high frequency voltages. 

Traveling wave currents may be expected to reflect back and 

forth in the airframe at the speed of light if, for example, a 

mismatch exists between the surge impedances of the lightning 

channel and the airframe, as illustrated in Figure 3. 

6 
iL 

MiAfC== 

ZL ,450 R 1 

I 

2*/C = 50 R 

I 

ZL =450 R 

I 

I 

REFRACTED 2ZL 

CURRENT -ZAI~ +ZL‘1'8iL 

2zAlC 

I "A/C + 'L 

=D.2lL 

I 

REFLECTED "L- ZA/C 

CURRENT .- 

zA,C~+ZL'"'8 'L 

ZAIC-ZL 

=ZA,C+ZL'-a8'L 

I I 

THESE REFLECTIONS CAUSE AN OSCILLATING CURRENT IN THE AIRFRAME: 

2T 1T 

Figure 3 - Traveling Wave Currents. 

The surge impedance of any conducting path is defined(ref.lO) 

as the ratio of the instantaneous surge voltage to the instantaneous surge current at a single point along the conducting path, 

as follows: 

z. = + (1) 

where, 

e = conductor voltage with respect to its 

return path, V 

i= conductor current, A 

It will be a constant for each conductor and return path. It can 

be shown that the surge impedance may also be expressed as: 

(2) 

7 

--- 
where, 

L= inductance per unit length of conductor, H 

C = capacitance per unit length of conductor, C 

If the surge impedances of the lightning,channel and the airframe are not the same (that is, if the ratios of voltage to current are not the same), some of the energy in the lightning 

channel may be reflected back into the channel and may not immediately enter the aircraft. Similarly, energy that has been 

deposited on or in the airframe may not leave it at once, but may 

instead reflect back and forth as shown on Figure 3. 

In Figure 3, the surge impedances of the lightning channel 

(ZL) and the aircraft (ZA C) are assumed to be 450 ohms and 50 ohms, 

respectively. If the lig h tning channel and aircraft can be represented as transmission lines, and if the reflection and refraction 

coefficients (ref. 10) for a lossless line are applicable, then 

the traveling wave current in the aircraft would appear as shown 

in Figure 3. The velocity, (v) of the traveling waves in the 

aircraft would be that of light, 3 x 10' meters per second, and 

the period (t) of oscillation would be proportional to the length 

(a) of the aircraft, or 

and the frequency (f) would be 

f = + hertz (4) 

and voltages induced by such traveling wave currents should be of 

the same frequency. In LTA tests to date, some voltages of these 

frequencies have appeared, but they are so cluttered with components of other frequencies as to make positive identification 

difficult. 

This experimental investigation focused on the transmissionline aspects of the LTA test setup, These were aspects that had 

not been addressed thoroughly enough in any of the previous LTA 

programs, and included return conductor effects, aircraft surge 

impedance, transit time and ringing frequency, and transmissionline aspects of the test current generators. To enable a large 

number of variations to be made in test circuit parameters, a 

simple, geometric scale model of the F-8 airframe and return 

conductors was constructed. The geometric model technique has 

been utilized successfully in related work (ref. 11) and enables 

variations in test circuit and return conductor configurations to 

be made far more quickly than if done on a full scale setup. 

8 
The following paragraphs describe the basic full scale and 

model test arrangements, and present the results obtained. 

EXPERIMENTAL INVESTIGATION 

Full Scale F-8 Test Setup 

F-8 DFBW-airplane.- The NASAF-8 DFBW airplane was utilized as 

the test aircraft in this program. The aircraft is a single engine, 

all metal fighter-type aircraft with a high wing that can be raised 

a few degrees to obtain a higher angle of attack during landing and 

takeoff maneuvers. The NASA F-8, the dimensions of which are shown 

in Figure 4, has been utilized by NASA as a test bed aircraft to 

evaluate various aspects of digital flight control technology. 

Whereas the ultimate goal of this work is to enable a better 

understanding of flight-control systems vulnerability in the lightning environment, this particular program was directed toward test 

technology development rather than assessment of the NASA F-8 DFBW 

system vulnerability. Thus, the DFBW system was not operating during this program. In fact, the flight control computer package 

was removed from the aircraft to provide room for a battery operated oscilloscope. 

Aircraft instrumentation. -Induced voltages were measured in 

one flight control circuit only, so that the effects of many variations in test parameters could be studied on a common basis. The 

circuit chosen was a twisted two-wire pair that extended aft from 

plug AP3 in the flight control computer pallet area a distance of 

about 8 meters to plug PlOO at the base of the vertical fin as 

shown in Figure 5. 

The circuit was part of a multi-conductor bundle of DFBW circuits, and is typical of circuits that extend between the computer 

and secondary actuators located near control surfaces. It was 

basically unshielded, except that portions of it passed through 

several sections of metal conduit in this exposed, area. 

Both wires of the twisted pair were grounded to the airframe 

at PlOO, and terminated to ground with 50 ohms in the pallet area. 

The induced voltages were recorded during the first full-scale 

test series by a battery-powered 200 MHz oscilloscope and during 

the second full-scale series with a 100 MHz oscilloscope with a 

wide-band differential preamplifier, also powered from a battery. 

Most of the oscillograms reproduced in this report were produced 

by the 100 MHz oscilloscope. The pallet area where the oscilloscopes were located is pictured in Figure 6, and the location of 

the aft terminal of the instrumented circuits is shown in Figure 7. 

9 
l- --- , ., --- ,;

5.7m r 

., .,.'J 

I " 

3Am. 

*I- .7.m 

-- a' 470 

.4m L 

I.5m --I 

Figure 4 - Overall Dimensions of the NASA 

F-8 Digital Fly-by-Wire Airplane. 

10 
Pallet / Plug AP3, pins 10 and 11 

Wing 

# Sections of 

50QTerminations 

I 

conductor wounded 

\ to airframg at 

Battery operated oscilloscope 

preamplifier and time base. 

plug PlOO. 

F-8 DFBW airplane 

I 

Forward Aft 

Figure 5 - Instrumented Twisted-pair Circuit in NASA F-8 

Digital Fly-By-Wire Airplane. 
Figure 6 - Location of Battery-Powered Oscilloscope 

to Measure Induced Voltages. 
Figure 7 - Location of Aft Termination of Instrumented Circuit. 
. . 

For most of the tests, the aircraft oscilloscope was triggered by the voltage induced in a Rogowski coil around the nose 

boom as shown on Figure 8. A coaxial cable conducted this voltage 

to the oscilloscope trigger input. 

Nose Cone 

Figure 8 - Rogowski Coil to Provide Aircraft 

Oscilloscope Trigger Signal. 

The aircraft oscilloscope had a differential preamplifier so 

that the line-to-line voltage between conductors could be measured, as well as the voltage between either conductor and airframe 

ground. The oscilloscope was connected to the circuit conductors 

via a pair of short coaxial cables fitted with alligator clip 

leads. 50 ohm line-to-ground terminations were used to terminate 

each conductor. The alligator clip leads were short and twisted, 

and the oscilloscope was grounded to a shelf within the pallet 

area to minimize extraneous noise in the measurement system. 

14 
At intervals throughout the tests, measurements were made of 

the noise that appeared in the oscilloscope and the alligator clip 

leads when test current was applied to the aircraft. A typical 

noise measurement is shown on Figure 9. 

0.2v 635 0.5 us 

Figure 9 - Noise in Induced Voltage Measurement System 

with Channels A and B Connected to Airframe 

Ground in the Pallet Area. Voltage Shown Is 

the Differential Voltage between Channel A 

and Channel B. 

Most of the line-to-line induced,voltage measurements were 

made with the same vertical sensitivity in Figure 9, and the 

signals ranged from one quarter of a centimeter to over one full 

centimeter of deflection, such that the noise level was nearly 

always less than 10% of the induced voltage signal. 

Basic test configuration. - The aircraft was positioned in a 

metal-walled hangar with its main and nose gear wheels parked upon 

insulating sheets of clear polycarbonate resin, 6.lmsquare and 

9.5mmthick. These insulating pads enabled the aircraft to be 

electrically charged up to 50,000 volts (50 kV) with respect to 

building grounds. 

Since many of the tests would require return conductors, an 

array of four conductors, supported by wooden stanchions, was 

erected along the aircraft. In the first series of full scale 

tests, each conductor consisted of a group of three steel wires 

spaced approximately 0.4mapart. The wires were covered with 

aluminum foil to see if this made any significant difference in 

test results (it did not). Figure 10 illustrates the location of 

the four return foils or groups of wires. 

15 

The four conductors were positioned approximately 2 meters 

from the aircraft in upper and lower and left and right quadrants 

to encourage cancellation of return current proximity effects of 

the aircraft. 

The stanchions supporting the return conductors were made of 

wood with spacings and dimensions as shown on Figure 11. The four 

foils or groups of wires were brought together to facilitate a 

common connection point at each end of the aircraft. 

Tests were begun using an "original" LTA test circuit in 

which electric charge is stored in a single generator capacitor 

cc> - A waveshaping resistor (R) and inductor (L) are connected 

in series between this capacitor and the aircraft to provide a 

unidirectional current pulse which rises to its peak in 2 microseconds decays to 50% of its peak value in 50 microseconds. This 

is the familiar 2 x 50 '~.l.s waveform. 

The capacitance consisted of one or more 0.5 PF, 50 kV capacitors connected in the manner necessary to provide the desired 

capacitance. The resistance was a bi-filar wound noninductive 

card-type resistor, and the inductor was a coil of AWG No. 10 insulated wire, wound on a 5 cm diameter plastic tube. Values of 

each element were varied during the tests and are given in individual circuit diagrams that follow throughout this report. 

The generator capacitors and waveshaping elements were positioned between the nose of the aircraft and the forward ends of 

the return conductors. It was determined early in the program 

that each end of the return conductor array should be grounded to 

the building ground through terminating resistors. The terminating resistors were suspended at the apexes of the return conductor arrays and grounded to the building via aluminum conductors 

as shown in Figure 12. 

The capacitors were charged from a 50 kV DC power supply located on a table outside of the return conductor array. The capacitors were discharged into the waveshaping circuitry and aircraft 

by means of a flapper switch mounted on one of the capacitor bushings, as shown on Figure 13. 

Test circuit instrumentation. - The electrical parameters 

measured during the program were the current entering or leaving 

the aircraft, or at other places in the test circuit, and the 

voltage between the airplane and the adjacent external return 

conductors at the nose and tail of the aircraft, and occasionally 

at other points in the circuit. 

17 
1.8 m 

t-3.4 m+ 8m >I3.7 m 

-8rnA 

Figure 11 - Dimensions and Location of Wooden Return Conductor Stanchions. 
S. 

Figure 12 - Location of Generator Capacitors and Return Conductor 

Terminating Resistor at Forward End of Aircraft. 
Figure 13 - Flapper Switch to Discharge Capacitors 

into Aircraft Test Circuit. 
The test current was measured with a 1OO:l impulse current 

transformer (CT) loaded with a 0.1 ohm burden resistor, so that 

the ratio between test current and signal voltage at the oscilloscope was as follows: 

. 

escape = 

=pg (O.lQ) (5) 

. 

'test 

= 1000 

,A test current of 250 amperes produced 0.25 volt at the oscilloscope - a level easily measurable. The CT was built and calibrated at the General Electric Corporate Research and Development 

Center and has a frequency response of 1 Hz to 100 MHz. It is 

shown at the aft end of the aircraft termination resistor in 

Figure 14. The CT output voltage was transmitted to a measurement 

oscilloscope located outside of the return conductor array via a 

33 meter RG-58U coaxial cable. To minimize noise induced in this 

cable, it was enclosed within a braided shield and laid upon a 

0.5 m wide aluminum foil strip that extended between the return 

line termination resistor and the oscilloscope enclosure. 

Voltages were measured between the aircraft and return lines 

with a 1OO:l resistance voltage divider comprised of two legs of 

five 1000 ohm carbon resistors arranged side-by-side in a plastic 

enclosure. The two 5000 ohm resistors were each connected to conductors of a 33 m, RG-22 twinaxial cable, which was terminated 

with 50 ohm line to ground resistors at the oscilloscope. To obtain higher voltage divider ratios, 5x or 10x terminations were 

added at the oscilloscope. For example, to measure traveling wave 

voltages between the aircraft and the return conductors when the 

generator capacitors were charged to 25 kV, the 10x termination 

was utilized so that: 

e etest etest 

scope = 

-- 

(lOO)(lO) = 1000 (6) 

or, if the test voltage was 25 kV, the oscilloscope voltage was 

e = 25x103V 

scope = 25 volts 1000 
Figure 14 - Impulse Current Transformers at Aft Termination, 

Showing Aircraft and Return Line Termination 

Resistors. 
Test circuit current and voltage were recorded by a 400 megahertz dual beam oscilloscope with two wide band differential 

preamplifiers. This oscilloscope was powered from 115 VAC building power and housed within a five-sided aluminum box to minimize 

electromagnetic field coupling between the return conductors and 

the oscilloscope. This oscilloscope is shown in Figure 15. 

Most of the time, this oscilloscope was also triggered externally by a voltage induced in a Rogowski coil located at the nose 

boom or other location close to the generator discharge flapper 

switch. This was done so that voltage and current measurements 

could be made at various locations along the test circuit and 

related time-wise to a common zero time reference. A Rogowski coil 

was chosen to provide the trigger voltage because it responds to 

the derivative of the test current and thus rises faster than the 

test current. The length of the trigger voltage, test current and 

test voltage measurement cables were made the same (33m)to eliminate signal arrival time differences at the oscilloscope. 

The response characteristics of the test circuit and current 

measurement systems was evaluated by applying square wave voltage 

and current signals to the CT and voltage divider and observing 

the response at the oscilloscope, which was separated from the 

sensors by the 33msignal cables. The results of these evaluations are shown on Figure 16. 

Occasionally, several measurements were made of radiated 

fields or other parameters also associated with the test circuit. 

These are described at appropriate points in the following text. 

One-Tenth Scale Model F-8 Test Setup 

As outlined in Objectives, a major purpose of this program 

was the evaluation of the relationships between various test aircraft. Since alteration of the test circuit and return conductor 

arrangement surrounding the F-8 airplane is a time-consuming operation, it was decided to use a relative geometric scale model to 

enable the large number of test circuit changes to be made and 

evaluated within the program. The following paragraphs describe 

the basic geometric scale model technique and the F-8 model test 

setup employed in this program. 

Relative geometric scale models. - When studying the electromagnetic response of a system to lightning, it is often difficult to visualize and express mathematically all of the factors 

affecting the response. This is especially true when the system 

is physically large and the distributed nature of the electromagnetic fields produced by lightning conduction in the structure must 

23 
-- 

Figure 15 - 400 MHz, Dual Beam Oscilloscope within Shielded 

Enclosure for Measurement of Test Circuit Current 

and Voltages. 

24 
0.1 

PF 

5000 R ea. 

Preamplifiers 

Measurement Circuit 

12 ns Response 

>lOO mV/div 10 ns/div 

Figure 16 - Response Measurement of Differential Voltage Divider. 

25 
be considered as well as the distributed nature of the electrical 

circuits of interest that are contained within the airframe. 

Often the mathematical expressions which describe these distributed fields are quite complex, and the data necessary to define the 

expressions are not readily available. In such cases, it may be 

easier to measure directly the desired system response than to 

derive and employ mathematical expressions. 

If the system of interest is physically large, as, for example, a Boeing 747 aircraft or an Apollo launching tower, the 

injection of currents approaching the characteristics of natural 

lightning strokes in either pulse rise time or amplitude is not 

possible with laboratory facilities available today or in the foreseeable future. Instead, a geometric scale model of the system 

can often be tested to obtain some, at least, of the desired data. 

Sometimes the characteristics of the geometric scale model are 

such that quantitative measurements can be made. 

The geometric scale model technique is not to be confused 

with computer models, about which much activity has recently centered. A digital computer model involves a digital computer 

providing numerical solutions for a mathematical model of a system. 

Other common models are electrical circuit models such as used in 

analog computers where lumped electrical parameters of inductance, 

capacitance and resistance are used to represent distributed parameters or other physical quantities which obey analogous laws of 

behavior. Neither of these models, of course, bears any physical 

resemblance to the system it represents. A geometric scale model, 

on the other hand, looks like and has a dimensional relationship 

to the full scale system of interest. The use of geometric scale 

models is much more prevalent in the fields of civil, mechanical 

and aerodynamic engineering than in electrical engineering, where 

geometric model techniques have only been used for antenna radiation pattern studies. 

In a relative geometric model, the impedance, voltage between 

points, current distributions and magnetic and electric fields 

in the region around the model are simulated. 

Relative geometric models are described in terms of scale 

factor, which is the relationship between the model property 

and the full scale property. Scale factor involves the fundamental physical dimensions of length, mass, time and charge. Only 

some of these factors can be chosen independently, usually 

three. Table II gives scaling relationships that apply in terms 

of the length (R), time (t) and impedance (Z) parameters. The 

scale factors of all other quantities are derived in terms of 

these three. Any three parameters in Table II can be chosen arbitrarily and the rest of the parameters defined in terms of those 

three. 

26 
TABLE II - Scale Factors in a Relative Geometric Model 

Scale Factor (S) fn Terms of length CL), 

guantity 

time (t) and Impedance (2) of Full Scale 

(F) and Model (M) Systems 

length CL) $4 

SL - 4 

time (t) 

reaistnnce (R) ‘R - ‘2 

capacitance (C) 

S 

SC - f 

inductance (L) 

impedance (2) 

conductivity (o) So - A- L 7. 

permittivity (E) St 

s E -3-r L z 

permeability (u) 

frequency (f) Sf - $ 

Relative geometric models were first used to study lightning 

effects on power system transmission lines (ref. 12). In those 

studies, the scale factors were chosen such that the impedances 

of the model were equal to the impedances of the full scale lines. 

Thus, the impedance scale factor was set equal to unity. Since 

the model work was to be done in the air, the scale factor for 

permittivity (E) was also set equal to unity. Since only one more 

scale factor could be chosen arbitrarily, length (a) was selected 

to reduce the model to a manageable physical size. The three 

arbitrarily selected scale factors were then: 

sz = 1 

SE = 1 

% = l/10 

These choices result in length and time scale factors being 

the same. If the scale factor (SR) for length is chosen to be 

27 
l/10, then the time scale factor (S&e) is also l/10, from Table II. 

With a length scale factor chosen to be l/10, all other scale 

factors will be as shown in Table III. 

TABLE III - Scale Factors for a OneTenth Scale F-8 Model 

Quantity Scale Factor 

TJ 10 

SE: 1 

1 

sf 10 

All of the requirements of Table III were met in the F-8 

model except the requirement that the material conductivity of 

the model should be 10 times that of the full scale airplane. 

Materials with conductivity 10 times that of aluminum or copper 

do not exist. 

Material conductivity requirements set one of the important 

dividing lines between what a practical relative geometric scale 

model can and cannot be used for. If, in general, the effects 

under study are influenced mainly by the electromagnetic fields 

outside of the model materials, then the model may be used for 

quantitative studies. In the l/10 relative scale model of the 

F-8 aircraft which was tested in this program, the peak of the 

oscillatory current through the model with a 200 volt capacitive 

generator was about 5 amperes. The model circuit resistance was 

less than 0.5 ohms so it seems obvious that an electromagnetic 

property (inductance) was controlling the circuit response and the 

material resistance was of very little importance. Therefore, 

the model can yield quantitative data on the response of the F-8 

to lightning-like current pulses. 

28 

-... - .-. .._ 
Studies involving electromagnetic fields internal to a conductor enclosure that were produced by currents diffusing through 

the wall of that structure are quite another matter. The diffusion and shielding properties of materials are quite heavily 

dependent on the electrical conductivity. Therefore, the response 

of systems contained within the structure to fields that penetrate 

the structure material cannot yield quantitative results. However, 

the qualitative measurements are still useful and give insights 

into the trends that will occur in the full size system. Of 

course, aperture coupling, in which fields are admitted to the 

interior through holes, is not affected by material conductivity 

and will, therefore, be accurately represented in the model. 

The above discussions present some of the practical background 

necessary to understand the concepts and practices involved in the 

application of relative geometric scale models. For a more rigorous derivation of the laws of similitude as applied to models, the 

reader is referred to reference 12, where the original work of P.A. 

Abetti is reproduced. Examples of the application of relative 

geometric scale models to aerospace lightning studies can be found 

in references 13, 14 and 15. 

Model construction. - The model was constructed of 0.41 mm 

(0.016") aluminum sheet riveted together. The fuselage was formed 

as a simple cylinder 14.6 cm (5 3/4") in diameter, and the wing, elevators;and rudder were formed as flat surfaces. The nose of the 

aircraft was simulated with a funnel. A dimensional drawing of 

model is given on Figure 17 and a photo of the model in a typical 

test configuration is shown on Figure 18. A twisted pair of wires 

was installed inside the model fuselage and two major apertures 

(the tail opening and a 7.6 cm x 10.2 cm (3" x 4") opening to 

represent the cockpit) were placed in the model. Numerous 

incidental apertures remained between the rivets and seams. 

Model test circuit. - The test circuit utilized for the model 

included scaled versions of those used during the full scale tests 

of the F-8. The same return conductor configurations were utilized as were present in the full scale test setup, as shown in 

Figure 19. The test voltages and currents applied to the model 

were much lower than those applied in the full scale tests to 

allow use of a mercury wetted relay switch in place of the spark 

gap used in the full scale tests. Repetitive switching at repetition rates low compared to the duration of the signals being measured will produce a steady display on the measurement oscilloscope 

and will produce a very clear trace even at fast sweep rates. 

The mercury relay was driven from a 60 Hz source, providing sixty 

operations per second. The time between switch closings was 

17 milliseconds, and the duration of the applied pulses did not 

exceed 20 microseconds. A schematic diagram of the model generator circuit is given in Figure 19. 

29 
w 

0 

A 

&------ 1.65m / 

- j i. i.,‘ _ I 

57. 8 cm 

Figure 17 - Design Drawing of l/10 Scale Model of the F-8 Aircraft. 
Figure 18 - One-Tenth Scale Model of F-8 DFBW 

Airplane and Test Circuit. 
5600 i-i 

4 T g=0.15~F 

(when present) 

to Aircraft 

from 

Return 

Conductors 

Figure 19 - Model Generator Capacitor Charging 

and Switching Circuit. 

During the full scale LTA tests, the aircraft and return conductors were shown to have a loop inductance of 18 PH. According 

to the inductance scale factor of Table III, the model test circuit should have had an inductance one tenth as great, or 1.8 vH; 

whereas the actual inductance as determined from ringing-frequency tests, was 1.7 ~.IH. This is an average difference of less 

than 6%, a phenomenal similarity considering the simplicity of 

the F-8 model, 

To simulate the conductive structural steel in the floor of 

the hangar where the full scale F-8 was tested, an aluminum foil 

ground plane was placed under the model. Traveling wave oscillations were found to exist between the return conductors on the 

model, just as they had in the full scale test setup. Thus, they 

were resistively terminated with 115Rresistances. This returnline traveling wave phenomenon also correlated very closely with 

the response observed on the full scale F-8. 

Measurements of the surge impedance of the transmission line 

formed by the aircraft and the return conductors yielded a value 

of 100 ohms on the model as compared to 124 ohms for the fullscale F-8. These and other comparisons indicated that it was 

possible to obtain useful quantitative data about the response of 

the F-8 aircraft to various lightning current generator and test 

circuit configurations. Consequently, the model was used to 

determine the generator configuration and waveshaping elements 

required to produce the traditional 2 x 50 us current waveshape 

with and without the aircraft terminated in a resistance equal to 

its surge impedance. Evaluations of the test circuit response to 

a lumped parameter ladder network as a means of simulating a 

lightning channel were also first carried out on the model and 

32 
the design was established for a B-segment ladder network to be 

evaluated on the'full-scale F-8 aircraft. In the test results 

that are described in the following paragraphs both model and 

full scale results are presented, 

Test Technique Evaluations 

Return conductors. 

is struck by lightning, 

- When an aircraft in free space (flight) 

it becomes part of the stroke channel and 

conducts the lightning currents carried in that portion of the 

channel. The lightning current circuit is very large and the 

return paths of the lightning current are very diffuse. In most 

cases, the lightning current return path for that portion of the 

channel containing the aircraft is far enough removed that it has 

no effect on the aircraft. 

Unfortunately, when tests are conducted on an aircraft using 

simulated lightning currents, two major differences are present. 

First, the aircraft is not in free space, It will be on the 

ground, and usually in a building which greatly alters its electromagnetic environment. 

(moist soils, concrete, 

These mediums of high conductivity 

building steel) are relatively close to 

the aircraft. Second, copper or aluminum conductors will normally 

be used to return the test currents to the current generator. Practical circuit limitations require that these return conductors be 

relatively close to the aircraft. 

The first measurements of lightning induced voltages using 

laboratory simulated lightning currents (ref. 1) were conducted 

using single conductor return paths which were usually laid upon 

the floor beneath the aircraft. 

Consideration was given to the influence of the return conductor during later testing (ref. 5) where two or more return lines 

were laid on the floor beneath the aircraft and routed around its 

perimeter. It was generally felt that these arrangements resulted 

in fuselage current distributions closer to actual inflight conditions, but no tests to verify this were performed. In almost all 

of the tests, the aircraft was maintained at building ground potential while the return conductors and generator were electrically 

insulated from ground. Also, the test generators employed were 

generally configured to provide a unidirectional simulated lightning current pulse and no consideration was given to the distributed transmission line represented by the aircraft and the adjacent 

return conductors. 

33 
More recently, investigators have begun to consider this distributed nature and have configured the return lines to be coaxially distributed around the aircraft in an attempt to get a 

fuselage current distribution the same as the free space situation (ref. 16). Each improvement in return conductor positioning 

around the fuselage has improved the current distribution in the 

aircraft which in turn has brought the magnetic fields surrounding the aircraft closer to the actual free space situation. However these changes have also affected the capacitive and electric 

field environment by increasing the distributed capacitance of 

the aircraft. 

When the aircraft is charged to some voltage level, an electric field exists around the aircraft and electric flux lines are 

established. Basically, the flux lines can be thought of as lines 

drawn between units (electrons or protons) of charge on the aircraft to charges of opposite polarity in the area surrounding the 

aircraft. If the return conductor is solid and completely encircles the aircraft then all electric flux lines will terminate on 

the return. In this case when a traveling wave is introduced 

on to the aircraft, the wave will propagate between the aircraft 

and the return conductors and will be contained within them. 

However, in most situations, it is not possible to build a 

solid coaxial return around the aircraft and a portion of the electric flux lines will terminate on other conductors outside of the 

return path. Basically, this represents the capacitance between 

the aircraft and the building floor or wall. When a traveling 

wave is introduced into this system, propagation will occur between the aircraft and the return conductors as well as the other 

conductors. The traveling wave currents that propagate between 

the aircraft and return conductors are returned to the source 

(generator capacitor) along the return conductors, but the waves 

propagating between the aircraft and the other conductors must 

find conductive paths back to the generator. When the aircraft 

and its return lines are isolated from building ground, the currents must be capacitively carried back to the return conductors, 

so traveling waves will exist between the airframe and the return 

conductors, the airframe and the building, and between the building and return conductors. r 

Similarly, if the aircraft is grounded, as has been the case 

in previous LTA tests, the return conductors rise in voltage, with 

respect to surrounding ground. So now waves propagate between 

the return conductors and aircraft, between the return conductors 

and building ground, and to complete the return path a wave must 

propagate between the building ground and the fuselage. How the 

electromagnetic waves which are propagating between the aircraft 

and building and between the return lines and building behave depends on how the transmission line that is formed between these 

34 
elements is terminated. If these lines are shorted or open circuited, reflected waves will occur. 

During the first test series on the full scale F-8 aircraft, 

measurements of the voltages between the return conductors and 

the hangar grounds verified the existence of these traveling 

waves. Initially, the return conductors were connected directly 

to the building ground at the nose end, and the tail end was left 

floating. Later, both ends of the return conductor configuration 

were connected to building ground through resistors of approximately 100 ohms. Figure 20 shows examples of the voltages measured between the aircraft return conductors and the building 

ground. By adding termination resistors to the ends of the 

transmission line formed between the return conductors and the 

building grounds, the duration of the oscillations was limited to 

one cycle and the magnitude was reduced by a factor of about 5. 

Various schemes were tried in an attempt to completely eliminate the remaining single cycle pulse that appeared between the 

return conductors and the building ground. The lowest voltage 

was obtained by laying a wide aluminum foil on the floor beneath 

the aircraft and connecting it to each end of the return conductor 

array. Unfortunately, this scheme diverted about 50% of the return currents to the foil and thus defeated the purpose of the 

coaxial return conductor arrangement. 

The arrangement and type of return conductors utilized during 

the LTA tests was shown to have an effect on the induced voltage 

results, as indicated by tests on the full scale aircraft in which 

induced voltages in the aircraft electrical circuit were measured 

to compare a return array of four groups of three parallel wires 

in place of the four 0.5 m wide aluminum foils shown in Figure 10, 

and to compare a return conductor array of four positions around 

the aircraft with only two on the lower sides of the aircraft. 

The results of these comparisons are shown in Figures 21 and 22. 

Essentially, no difference in induced voltage response was 

detected when parallel wire returns were utilized in place of the 

aluminum foils, as indicated by the induced voltage comparison of 

Figure 21. This was an important point since it is much simpler 

to install and support an array of wires than the wide aluminum 

foils. At various times during the full scale tests, both copper 

and steel wires were used, with no discernible difference in induced voltage response. The steel wires added approximately 2 ohms 

of resistance to the circuit, which was of no consequence since 

much higher resistances are already necessary for waveshaping 

purposes. 

It should be noted, of course, that one wire cannot replace 

one foil due to the differences in capacitance between the wire or 

35 
RF 

(n> 

RR 

(n> 

Model 

20 v 306 0.1 us 

117 117 

20 v 310 0.1 US 

Full Scale 

0 a3 

2.5 V 136 1.0 US 

144 132 

2.5 V 141 1.0 ?Js 

Figure 20 - Typical Return Line-to-Building Voltages 

Before and After Termination. 

36 
Four Foil Returns 

Four 3-wire Groups 

in same locations as foils 

2v 179 1.0 l.ls 

Figure 21 - Comparison of Induced Voltages with 

Foil and Wire Return Conductors. 

37 
F-8 

\ 7 / 2v 713 0.5Lls 

/ 

Return Conductors (2) 

Figure 22 - Comparison of Induced Voltages 

with a Balanced and Unbalanced 

Return Conductor Arrangement. 

38 
foil and the aircraft. In the first full scale test series three 

wires spaced 0.25 m apart (0.5 m total width) effectively gave 

identical performance to the 0.5 m wide foils as shown in Figure 21. 

During the second test series, two wires spaced 18" apart were 

used, and no discernible difference was again noted. 

Comparisons of induced voltages measured with all four return 

conductors in place versus only the lower returns are shown in 

Figure 22. Similar comparisons were conducted during the full 

scale test series, and in both cases, induced voltage levels were 

reduced by a factor of close to 27% when the top two conductors 

were removed. The reason for this is undoubtedly that the fuselage 

current becomes unbalanced and less dense on the upper portion of 

the fuselage when the upper return conductors are removed. The 

reduction in current density produces a corresponding reduction in 

magnetic flux density along the top of the fuselage, where the 

instrumented conductor is located. 

Transmission lines. - As discussed previously, it is probable 

that during the first few microseconds after lightning contact 

with the aircraft the airframe will behave as a short transmission line. That is, voltage and current waves will originate at 

the point of attachment and propagate along the airframe, possibly 

reflecting back and forth. The same phenomenon will occur during 

simulated lightning testing, although as discussed earlier, the 

aircraft environment will differ considerably from the inflight 

condition. 

In order to understand how these traveling waves relate to 

the aircraft induced voltage measurements, characteristics of 

these waves must be measured and identified, It was possible to 

measure the traveling wave currents entering and leaving the aircraft with the current transformer shown in Figure 14, but no such 

instrument exists with which to directly measure the current on 

the airframe. Attemptsto measure the magnetic flux produced by 

the airframe current with d$/dt coils were only partially successful, due to the difficulty of isolating the magnetically induced 

signal from that also capacitively coupled from the fuselage to 

the probe. Measurements of the voltage waves at various points 

along the fuselage were possible using the resistive voltage divider described earlier, and all data presented here are from those 

measurements. 

As was shown in Figure 3, the surge impedance (Zp,c) of an 

aircraft may be important in determining the magnitude of traveling waves that may occur along it when it is struck by lightning. 

The surge impedance (Z,) of any transmission line (also known 

as its characteristic impedance) was defined as: 

zO = /x7z (ohlJ=) 

39 
where, 

L = inductance per unit length (H/m) 

and, 

where, 

C = capacitance per unit length (F/m) 

Vt = ZoIt (8) 

Vt = traveling wave voltage 

It = traveling wave current 

If a transmission line is terminated by a resistance equal to 

Z then the traveling wave passes along it once and is fully absgrbed by the termination resistance and no reflections occur. If 

the line is terminated in any other resistance or imp>edance, reflections or refractions occur and the transmission line will exhibit 

a damped oscillation. The rules for reflections and refractions 

on the transmission lines are illustrated in Figure 3 and described 

more fully in several texts (ref. 10). In general, voltage waves 

reflect positive at terminations with resistances higher than Zo 

and negative at terminations with values lower than Z,. In other 

words, terminating a transmission line in a resistance greater than 

Z. will result in an increase in the total circuit voltage after a 

reflection, and terminating in a resistance less than Z, will result in a reduction of the total circuit voltage after the reflection. Current waves behave in the opposite manner, reflecting 

negative at high termination impedances and positive at low impedances. 

The surge impedance (Zo) of the aircraft and its return conductors was determined by inserting various values of resistance 

between the aircraft tail and the return conductors and measuring 

the voltage appearing across this resistance when the generator 

capacitor is switched on to the other end of the line as shown in 

Figure 23. Current through this resistance was also measured most 

of the time. For the full scale tests, it was necessary that the 

termination resistors be capable of dissipating all of the energy 

stored in the generator capacitor. The maximum generator capacitance of 2.0 ~.IF charged to 50 kV stores an energy of 2500 wattseconds. Assuming that all of the energy is delivered and dissipated within less than one second, a resistor thermal capacity of 

about 2500 watt-seconds is required. 2500 watt noninductive 

resistors ranging in value from 30 ohms to 300 ohms (the probable 

range) were not readily available, so suitable resistors were 

assembled from 150 2-watt carbon resistors mounted on a fiberglass 

40 
Full Scale With 

Wire Returns Resistively 

Terminated to Building 

Full Scale With 

Foil Reeurns Resistively 

Terminated to Building 

Rt * 20 

Test No. 38 

Rt = 104 R 

Test NO. 205 

e 

m 

RI: = 88 

Rt = 124 

R, < Z, 

Rt = 117 n Rt = 152 n 

Figure 23 - Voltage across Various Termination Resistances. 

41 
perforated board and wired 15 in parallel and 10 in series. Two 

watt resistors have thermal ratings of approx?mately 20 watt-seconds each, for a total of 3000 watt-seconds. Two of these resistor boards can be seen in Figure 14. 

Resistor values of 45, 54, 67, 82, 88, 104, 124, 132, 143, 

160, 181, 212 and 226 ohms were installed successively in the Rt 

position in the circuit of Figure 23 and the value of resistance 

observed to give the least reflection was considered to be Z .I 

Unfortunately, since other transmission lines are present in'the 

system, aircraft to building and return conductors to building, 

which each have different values of Z,, it was not possible to 

obtain an exact terminated wave and some oscillations were always 

present. 

Determinations of Z. were made for the full scale F-8 test 

circuit as well as the l/10 scale model. 

Figure 23 compares the results of four such tests. The first 

tests were conducted prior to terminating the return conductors to 

the building ground. The oscillations between the building ground 

and airframe are apparent in these measurements. During a later 

test series, the wire return conductors were replaced with aluminum foils. These foil returns were also resistively terminated to 

the building with 140 ohm resistors at each end which removed most 

of the return-to-building oscillations. 

Changing from wire return conductors to aluminum foil lowered 

the surge impedance by about 20% from over 100 ohms to about 90 

ohms. This was no doubt due to the higher capacitance between the 

airframe and the foils versus the wires. During the second full 

scale test series, wire returns were again used and the surge impedance was again over 100 ohms. The relative geometric scale 

model which also utilized wire returns resistively terminated to 

the buildings, compared very closely with the full scale tests of 

the same arrangement. 

It was suggested earlier that when an aircraft in flight is 

struck by lightning, the airframe currents will be oscillatory 

due to traveling wave reflections. The cause of this phenomenon 

has been related to a surge impedance mismatch between the aircraft and the lightning stroke channel. Estimates of the lightning channel surge impedance have ranged from 400 to 600 ohms 

(ref. 10). The surge impedance determinations made on the aircraft test circuits ranged from approximately 80 ohms to 125 ohms. 

This would suggest a significant mismatch and the formation of 

reflected traveling waves in the airframe as illustrated in 

Figure 3. However, the aircraft surrounded by return conductors 

is obviously not in "free space" as it is in flight. 

42 
To find out what the aircraft surge impedance might be if no 

return conductors were nearby, the one-tenth scale model was 

suspended in air and measurements of Z, were made. Since connecting a termination resistor between the tail of the model and 

ground would compromise the set-up by moving ground close to the 

tail, a series resistor was inserted between the mercury-wetted 

switch and the nose of the aircraft. If this series resistor is 

equal to Z 

one half B' 

then the voltage at the nose (e) will initially be 

o the capacitor charge voltage (E) and reach E after 

the reflection at the open tail returns to the nose. The test 

setup and resultsofthese tests are shown in Figure 24. With the 

return conductors around the model, Z values of 88 to 105 ohms 

were measured as indicated in Figure 93. With the return conductors removed and the model suspended 0.7 m off the ground plane, 

a value of 250 ohms was determined. An elevation of 1.3 m increased the surge impedance to about 400 ohms, but further increases in elevation above ground made no difference because 

most of the coupling was between the front of the model and the 

vertical foil that was necessary to connect the low side of the 

capacitor to ground. All foils were then removed and the capacitor was grounded via a thin wire to the power system ground 

and earth. This time the final measurements indicated that the 

surge impedance had again risen and was about 730 ohms. 

The main point of this experimental exercise was to illustrate 

that the transmission line surge impedance of an aircraft in 

flight is probably higher than the 88 to 124 ohm levels determined 

with the nearby return wires. If the aircraft in flight is 

treated as a classical transmission line problem, with the aircraft fuselage as one conductor and the earth as the other, a calculation of the characteristic impedance can be made. These calculations ignore the aircraft wings and consider the earth as a 

single return conductor while in fact it is a conducting plane. 

The inductance and capacitance of the aircraft are then given by: 

P 

L= $ In ($4 (H/m) 

where: 

C 2nco 

=a 

(F/m) 

d 

pO = permeability of free space - 41~10'~ 

&O = permittivity of free space = 

1 

36~rlO' 

(9) 

(10) 

V/M) 

(H/M) 

h= distance between conductors 

d = diameter of the fuselage 

43 
h 

Aluminum Foil Ground Plane 0.3m Wide 

Rs < Z Rs QJ zo 

355.0 

RS = 150 52 R, = 400 n 

h=l. 3m 

356.0 

- ~ 

356.4 

h=2m 

with all 

foil 

removed 

364.0 

: 

Figure 24 - Surge Impedance of the F-8 Model as 

a Function of Distance Above Ground. 

44 
Since many aircraft lightning strikes occur at an altitude of 

about 3 km, h was chosen to be 3 km. The F-8 fuselage (d) is 

approximately 2m. Using these numbers in equations (9) and (10) 

gives: 

L = 1.5 H/m 

C = 7.6 pF/m 

Then, from eq. (2): 

= 444n 

This is near the range of surge impedances often attributed to a 

lightning flash channel which is thought to be 400 to 600 ohms. 

It therefore seems quite probable that the field factors which 

determine the value of surge impedance of a lightning stroke 

channel are also controlling the surge impedance of the airframe 

in flight. 

Traveling wave transit times. 

ports traveling waves whichKin 

- If an aircraft fuselage supfact reflected at each end of 

the aircraft, the frequency of oscillations will be related to the 

length of the aircraft and the velocity of wave propagation. The 

electromagnetic waves propagate in the air medium around the fuseIage, and traveling wave currents flow on the fuselage itself. 

From Maxwell's equations, it can be shown that the velocity of 

propagation is a function of the medium in which these waves travel and can be calculated from: 

1 

V= 

G 

(11) 

The permittivity (E) and permeability (u) of air is very 

close to that of free space so use of these values (u. and so) 

will produce very small errors, thus: 

v = /$!g = 3xlO'm/s=c (12) 

where, 

c = velocity of light 

A traveling voltage wave introduced at the nose of an aircraft will require a transit time (T) to reach the tail. Since 

the aircraft is in air, the transit time should be equal to: 

R 

T =- 

C cd (13) 

45 
where, 

R = length of the aircraft (m) 

In the LTA test circuit, with the tail of the aircraft open 

circuited and the nose connected to the return conductors through 

the capacitor, applied voltage and current waves should reflect 

and oscillate as illustrated in Figure 26. 

Every four transit times, the observed voltage waveform will 

repeat, giving an observed oscillation which has a period (t) 

which is equal to four transit times. 

During the first full scale test series, traveling wave voltages were measured on the F-8 aircraft. A typical measurement is 

shown in Figure 25, triggered at the point of measurement so that 

the time between generator switching and arrival of the wave at 

the measurement point is not shown on the oscillogram. Stray 

electrical losses, as well as the non-uniform geometry of the aircraft eliminate the squareness of the traveling wave, but the 

basic frequency of the reflected wave is still very evident. 

Figure 25 - Typical F-8 Ringing Frequency Oscillogram. 

Voltage measured between aft end of aircraft and return conductors after 10 kV 

capacitor was switched onto aircraft. 

This measurement gave a frequency of 1.74 MHz. Other measurements were slightly lower, so an average value of 1.7 MHz was 

obtained. This corresponds to a transit time of 145x10-" seconds, assuming that four transit times elapse in the period of 

one cycle of voltage, as indicated on Figure 26. 

At the speed of light, a distance of 43.5m would be traveled 

in this time. The F-8 airframe is 17.lm long and the distance 

from one apex of the return conductor array to the other was 22m. 

46 
(2)Return Conductor 

(2)Return Conductors 

t 

. 

i 

Figure 26 - Voltage and Current Traveling Waves on an 

Idealized Test Circuit; as Observed at the 

Center of the Fuselage. 

47 
Assuming that the larger apex-to-apex distance applies, the transit time determined from the measurement of Figure 25 is still 

twice as long as the actual distance would predict. The discrepancy between the measured transit time and calculated time is 2:l. 

To verify these observations, tests were conducted on the 

l/10 scale model of the F-8. The model test circuit from apex to 

apex of the return conductor was 2.25m. In these tests, typical 

results of which are shown in Figure 27, average voltage waveform 

ringing frequencies of 15 to 18 MHz were measured. 

5V/div 346 O.OZps/div 

Figure 27 - Typical Relative Geometric Scale Model 

Ringing Frequency Oscillogram. 

In the oscillogram of Figure 27, an average frequency of 

16.5 MHz was obtained which corresponds to a transit time of 

15.2 x lo-'s, or a distance of 4.55m; again twice the length of 

the circuit. 

To investigate the phenomena further, and to see if the wings, 

etc. on the model were somehow involved, the model was removed 

from the circuit and replaced first by a 13 cm diameter galvanized 

steel tube and then by a 1.3 cm copper pipe. The test results, 

shown by the data in Table IV indicate that the transit time was 

reduced as the conductor got smaller, but was still 23% longer 

than calculated. 

Measurements of ringing frequency and transit time made during the second test series on the full scale F-8 yielded results 

essentially identical to those obtained during the first series. 

In all of the above tests, the return conductors were not 

uniformly spaced around the center conductor. They start and end 

in conical fashion and spread at the fuselage center. The effects 

of the introduction of bends in the return path in this configuration on transit time is not known. Consequently, tests were 

48 
TABLE IV - Ringing Frequency and Transit Time 

Measurements on Various Diameter 

Cylindrical Conductors in Place on 

Aircraft in One-tenth Scale Model 

LTA Circuit 

Conductor 

Voltage 

Frequency Transit Time zO 

1.3 cm tube 27.3 MHz 9.2 ns 280 

13 cm tube 21.0 MHz 11.9 ns 120 

l/10 scale 

F-8 Model 16.5 MHz 15.2 ns 82 

Circuit Total Length = 2.25m, T = 7.5 ns 

performed on a configuration which is shown in Figure 28. Here, 

return and center conductors are parallel and the ends are flat. 

Four configurations were tested with this fixture, as follows: 

(1) AWG no. 20 center conductor and a single AWG no. 20 return 

wire spaced 0.3 m apart; (2) same as (1) but eight AWG no. 20 

return wires spaced in a 0.61 m circle around the center conductor; 

(3) same as (2) but a 0.13 m dia. tube used as a center conductor; 

and (4) same as (3) but the eight return conductors were moved 

closer to the center tube, forming a 0.15 m dia. circle around the 

0.13 m dia. tube. 

The test results from these four configurations are given in 

Table V. The total length of the configurations of Table V was 

5.64 m, which corresponds to a transit time at the speed of light 

of 18.8 nanoseconds. As in the LTA circuit tests of Table IV the 

large tube transit time is longer than the thin wire transit time 

and both are longer than the calculated transit time. In these 

flat-ended tests the differences are smaller than in the previous 

tests. For example, the wire line transit time is 15% longer than 

the calculated time whereas in the LTA circuit test the thin tube 

transit time was 23% longer than the calculated time. Also, the 

large tube transit time is 15% longer than the wire for the flat 

.&- ended-configuration while in the LTA circuit, the larger tube had a 

transit time 29% longer than the small tube. 

It is also interesting to note that the surge impedance, determined by dividing the voltage traveling wave magnitude by the current traveling wave magnitude, for a two wire system of configuration (1) was the same for the semicoaxial system of configuration 

(2) - Moving,the return wires closer to the tube did reduce the 

surge impedance. 

49 
Figure 28 - Transit Time Test Circuit 

Configurations. 

50 
TABLE V - Ringing Frequency and Transit Time 

Determinations in 2 Coaxial Test 

Configurations 

Frequency Transit Time zO 

-- OfJJz) (nd (n> 

(1) ---+*0.3mOf- 11.4 21.9 %400 I 

1 I I I 

I 0 ’ 

(2) 

(3) 

i* 1 I 0 0 0 0 11.7 

1 0 I 0 I a 

I I 

t I. ‘ 0 

I ’ 0 0 0 10.0 

I 

21.4 

25.0 

24.9 

%400 

s400 

lLlO0 

Using these same experimental techniques, measurements of 

traveling wave transit times were made on a length of RG58U coaxial cable, for which transit times have been established. The 

cable was 15.5 m long and has a published propagation velocity of 

2x10' m/s due to the cable insulation dielectric. A measured 

transit time of 83 ns was determined as compared to a calculated 

value of 78 ns. The difference between these two values is less 

than 7%. Since the exact velocity of propagation is a function of 

the insulation dielectric constant which could vary slightly from 

one cable sample to another, the value obtained is sufficiently 

close to the calculated value to be considered identical for 

practical purposes.' Thus, the transit time appears to be configuration dependent and most closely approaches the speed of light 

when the center conductor in a coaxial arrangement is thin. This 

51 
transit time "discrepancy" was an area of continuing concern in 

this investigation and additional tests were made to help evaluate this phenomenon. The results of these additional tests are 

reported in later paragraphs of this report. 

Grounded versus ungrounded aircraft. - In most LTA tests conducted prior to this program the aircraft has been grounded at the 

point on the fuselage where the induced voltage measurement is 

being made. This allows the signal to be conducted outside of the 

aircraft via a twinaxial cable and recorded with an external oscilloscope - a necessity in places where there is no room for an 

oscilloscope within the aircraft. Fiber-optic or other dielectric 

waveguide techniques offer a means of making this measurement 

while leaving the aircraft ungrounded. Thus, a comparison test 

was made to determine if a difference exists in the induced voltages measured within an ungrounded and a grounded airframe. 

The original RLC LTA circuit was utilized for this evaluation, 

with the aft end of the aircraft shorted to the return conductors. 

In each case, the induced voltages were measured by the oscilloscope within the pallet area, even though the oscilloscope could 

have been placed outside of the aircraft and connected to the 

instrumented circuit via a cable in the grounded airframe test. 

The original 2 x 50 ~.ls LTA test waveform was applied and the induced voltages with the aircraft ungrounded and grounded are shown 

in Figures 29 and 30, respectively. A comparison of the induced 

voltage oscillograms shows them to be nearly identical.. 

It should be noted that both ends of the return conductors 

were resistively terminated in this case,an improvement over the 

original circuits in which no attempt was made to dampen spurious 

oscillations existing between the return conductors and the "outside world". 

Aircraft resistively terminated versus short circuited to 

return conductors. - The one-tenth scale model was utilized to 

determine the circuit element needed to pass a 2 x 50 us current 

through the aircraft. Basically, this required that the total 

circuit inductance (L) be increased to accommodate an increase in 

the circuit resistance needed to terminate the aircraft to the 

return lines with its 124 ohm surge impedance while providing the 

same 2 us rise time, since rise time is proportional to L/R. The 

value of the generator capacitance was also decreased to retain 

the fall time of 50 vs (to half peak value) with the higher circuit resistance, since this parameter is proportional to the RG 

product. 

The resistivity terminated circuit is shown on Figure 31 together with the corresponding induced voltages. For the full 

scale generator charging voltage of 25 kV, the resistively 

52 
Model 2.75 A 

L = 2.15 pH 

C = 0.15 PF 

R,= 43 9 

E - 130 v 

0.5 A 13A O.lPS 

Full 

Scale 

L = 18 yH 

C = 1.5 vF 

R,= 45 R 

E = 25 kV 

100 A 687 0.5 us 

Figure 29 - 2x50 us Return Stroke into Aircraft 

Terminated with Short Circuit with 

Aircraft Ungrounded. 

53 
Induced Voltage: 

Line to ground 

691 0.5 us 

No model tests were made of this configuration. 

Full Scale 

L 18 pH 

C = 1.5 pF 

RS = 4552 

E = 25 kV 

I 

-I- - - 

Figure 30 - 2x50 ~J.S Return Stroke into Aircraft 

Terminated with Short Circuit with 

Aircraft Ungrounded. 

54 
Induced Voltage: 

Model 

L = 7.1 pH 

C = 0.04 I-IF 

Rt= 100 R 

E= - 110 v 

1.08 A - 

0.2 A 14A 0.1 I.ls 

Full .~ Scale 

198 A - 

L = 90 pH 

C = 0.5 ~-IF 

Rt= 124 R 

E = 25 kV 

40A 680 1.0 ps 

Figure 31- 2x50 us Return Stroke into Aircraft 

Terminated with Rt = Zo. 

55 
terminated circuit provided only half the current provided by the 

short-circuited circuit of Figures 29 and 30. Taking this into 

account, it is seen that the line-to-ground induced voltage is 

about the same in either case (that is, about 2 volts per 250 amperes of peak current) but the line-to-line voltage is much lower 

with the aircraft terminated. This is probably because the higher 

frequency traveling wave oscillations that appear on the current 

wavefront are eliminated in the resistively terminated case. This 

result suggests that the line-to-ground voltage is more nearly 

proportional to the peak amplitude of the test current, whereas 

the line-to-ground voltages appear more closely related to the fast 

rate of rise and/or high frequency components of the test current. 

Further comments on these relationships will appear later in this 

report. 

Improved Test Circuit Evaluations 

One of the objectives of this program was the evaluation of 

possible improvements in the LTA test circuits. Selection of 

candidate circuits was to be based on the results of cause-effect 

studies conducted during the first part of the program. During 

the investigations reported thus far, several important causeeffect relationships were identified. Among these are the following: 

1. It is necessary to have return conductors to enable 

the return-stroke portion of the lightning flash to 

be simulated and these conductors must be distributed around the aircraft in order to minimize proximity effects on the distribution of current in the 

aircraft. 

2. The airframe and its return conductors behave like 

a transmission line although the velocity of traveling wave propagation in this transmission line is 

somewhat slower than that of waves in a "pure" airinsulated transmission line with a narrow center 

conductor. As with other transmission lines, the 

magnitude of the traveling voltage and current surges 

are related by the surge impedance of the line, and 

the surge impedance, in turn, is determined by the 

geometry of the line. In other words, it is determined by the size of the aircraft and its return 

conductors, and by their spacing. 

3. The surge impedance of the F-8 aircraft and four 

aluminum foil or wire return conductors is about 

100 ohms. 

56 
4. The surge impedance of an aircraft in flight is 

greater than this, perhaps between 400 and 800 ohms, 

and therefore may not be appreciably different than 

that of the lightning flash channel itself. 

5. Reflected traveling waves may be eliminated by terminating the aircraft to its return conductors in a 

resistance equal to the transmission line surge 

impedance (in the F-8, 100 ohms) but an initial wave 

propagating once from the generator end of the aircraft to the opposite end will nearly always occur 

in the test, and very likely occurs in flight. 

It therefore appears necessary that improved test circuits 

be capable of generating both the traveling wave phenomenon and 

the return stroke current. Since little is presently known 

about the magnitude of traveling waves produced on aircraft in 

flight, it may not be possible to accurately simulate these 

phenomena. It is necessary, however, that the traveling wave 

phenomenon be controllable. 

In actuality, the traveling wave phenomenon would probably 

occur first, when the lightning leader first approaches or contacts the aircraft. A period of relative quiet then probably 

ensues, until the leader has propagated to the earth (or another 

opposite charge center) and the return stroke has been initiated. 

The return stroke is formed by the rapid discharge of the leader 

into the earth, from the ground up. 

It may require a few milliseconds for the leader to reach 

the earth, and then the return stroke will propagate up from the 

earth at about 1 x lo8 m/s (about l/3 the speed of light). These 

two different phenomena are illustrated in Figure 32. 

Thus, the basic phenomena that should probably be simulated 

are leader currents and return stroke currents. When each of 

these currents reaches the aircraft, traveling wave currents are 

also likely to be present. 

Since traveling wave phenomena are involved in the return 

stroke as well as the leader attachment phase, it seems most 

appropriate to simulate the lightning channel as a transmission 

line rather than with lumped RLC components as in the original 

LTA configuration. Clearly, natural variations in lightning phenomena and aircraft size imply that a range of lightning channel 

and aircraft surge impedances are possible, so the test technique 

should enable these parameters to be known and controlled, and 

changes in them to be made. Configuration of the lightning 

generator to approximate a transmission line was therefore decided 

upon for evaluation. 

57 
Leader Currents Return-Stroke Current 

- 

Figure 32 - Lightning Strike Phases to be 

Simulated in the LTA test. 

Lumped parameter ladder network. - A transmission-line representation of the lightning chai!inZlXas Long been considered appropriate (ref. lo), and mathematical analyses of such models have 

recently been accomplished by Kim, DuBro and Tessler (ref. 17), 

Little (ref. 18) and Strawe (ref. 19). In each case the lightning 

channel has been represented as a lumped parameter ladder network --- 

as shown in Figure 33. 

Cloud 

resistance 

+ 

Element of 

channel inductance 

+ 

Element of 

channel resistance 

4 

Cloud Element of 

capacitance copocitonce channel copocitance 

Earth 

resistonce 

Figure 33 - Lightning Channel Represented as a Lumped 

Parameter Ladder Network (from ref. 18). 

Each of the referenced analyses utilized different numbers 

of ladder elements and different values of R, L and C in each 

element. Values of inductance ranged from 1400 to 6100 pF per 

meter of channel length. The results of these analyses are 

58 
reported completely in the references and will not be repeated 

here, but in each case they show that: 

1. The amplitude and rise time of the return stroke 

diminish as it propagates upward from the earth. 

2. The wavefront of the return stroke currents is 

concave, and reaches its maximum rate of rise 

shortly before it reaches its peak amplitude, 

rather than at the beginning of current flow as 

in the double-exponential current produced by 

the original LTA circuit. 

Inductance and capacitance selection. - For the test circuits to be evaluated in this nrogram. it was necessary to select 

appropriate values of L and C.L The rationale for thisdselection 

was as follows: 

The surge impedance (ZL) of a lightning channel has been 

estimated by Bewley (ref. 10) to range between 400 and 600 ohms, 

and most other researchers appear in agreement with this. As 

reported earlier, the surge impedance (ZA/ 

E 

), of the aircraft 

with respect to the return lines is about 00 ohms. It will be 

important to have the surge impedance of the lightning channel 

match that of the aircraft and return line configuration on 

some occasions, so a value of 100 ohms was selected for initial 

evaluations of the ladder network and the first test circuit 

evaluations were made using the l/10 scale model. 

Assuming that the return stroke travels through the channel 

with a velocity one-third the speed of light, then equations (14) 

and (15) should be satisfied as follows: 

= 100 ohms (14) 

and 

1x10* m/s (15) 

combining (14) 

and (15): 

L= ++ loon 

lx108m/s = 1 uH/m (16) 

and 

ccl= 

vz (1x10; (100) = l~lO-~'F/m (17) 

= 100 pF/m 

59 
Simulation of a lightning channel 3000 m long would thus 

require capacitors totaling: 

C total = (3000m)(lOO x 1O;-12 F/m) (18) 

= 0.3 VF 

and a total inductance of: 

L total = (3OOOm)(lxlO 6 H/m) (19) 

= 3 nH 

Number of segments.- The analysts referred to earlier utilized eight or more ladder segments to provide what was considered 

to be an adequate representation of a lightning stroke waveform, 

but it would be desirable to reduce the number of capacitors and 

inductors that must be assembled in a test circuit. At the same 

time, a closer approximation of an actual transmission line is 

achieved if each- element is made to represent a smaller fraction of the total line. Therefore it was-decided to represent a 

shorter portion of the total lightning channel with a lumped parameter ladder network representing six segments of 25 meters each. 

This would reduce the duration of the return stroke passing through 

the aircraft, but should not affect its rate of rise or peak amplitude; nor should such a reduction in length of channel simulated 

significantly affect the nature of the short-duration traveling 

waves in the aircraft. 

For the full scale test circuit, this required that for each 

25-meter segment, 

L = (25m)(luH/m) = 25 PH (20) 

and, 

C = (25m)(lOOpF/m) = .0025 PF (21) 

For the l/10 scale model evaluation, the values of L and C 

were divided by the scale factor of 10 in accordance with the 

rules of Table III. 

A comparison of the lumped parameters selected for this program with those utilized in the analyses referred to earlier is 

presented in Figure 34. It will be noted that the surge impedance of the experimental line is considerably greater than that 

of either of the analyzed lines. Also, the capacitance per meter 

of the experimental line is considerably lower than that of either 

of the analytical models, Considering the remoteness of a typical 

lightning flash from earth or any other conductor, the 100 pF/m 

60 
Parameter Kim, et al. Little 

Length of Channel 

Being Represented 3000m 3000 m 

No. of Ladder 8 8 

Elements 

Length of Each 

Element 375 m 

Parameters 

R 0.01 Q/m 1.0 Q/m 

L 0.04 vH/m 2.0 pH/m 

C 3000 pF/m 1400-6100 pF/m 

L zL= d- c 13 n 38-18 Q 

F-8 Test in 

This Program 

150 m 

6 

25 m 

0 1 

1.0 uH/m 

100 pF/m 

100 n 

Figure 34 - Comparison of Lumped Parameters Utilized in This 

Program with Those Utilized in Analytical Models. 
estimate seems more appropriate than the higher values utilized 

by Kim et al.and Little; however, no attempt was made in the 

analysis to confirm or disprove either of these values by calculation. 

At first, the lumped parameter ladder network was tested in 

the one-tenth scale model configuration. A series resistance of 

one ohm per segment was included initially, but measured currents 

were indistinguishable from those occurring in the ladder circuit 

without the resistance present, so testing was continued without 

the series inductance. 

Leader attachment simulation. - Figures 35 and 36 illustrate 

the model and full scale test results, with the lumped parameter 

ladder network used to represent a leader first attaching to the 

nose of the aircraft. These figures show the current entering 

the airframe (and in some cases, airframe voltages also) and the 

line-to-ground and line-to-line induced voltages in the instrumented circuit. 

In Figure 35, only the approaching leader is simulated. The 

airframe currents produced in the model are very similar to those 

that entered the full scale F-8, and in each case, related to the 

line charging voltage andthe 100 ohm surge impedance, as expected. 

Since the aircraft was unterminated, the current was reflected back and forth in the airframe. The reflections, which can 

be seen in the current oscillograms of Figure 35, are lower in 

amplitude than the initial surge due to losses. 

In the full scale tests, the voltage (e) between the nose of 

the aircraft and an adjacent return wire was also measured. During the first pulse the voltage oscillates in phase with the current, as expected, and is then out of phase. The fuselage remains 

at a DC voltage level determined by the redistribution of charge 

from the 75msimulated leader. 

The induced voltages shown on Figure 35 peak during the 

first half microsecond which is the time the airframe current is 

rising and changing most rapidly, indicating the voltages are 

induced by magnetic flux passing through apertures such as the 

wing-to-fuselage seam and the cooling air louvres in the pallet 

area covers. The line-to-ground voltage includes a lower amplitude oscillatory component which appears to be in phase with the 

successive current oscillations, indicating that it may be related to the structural resistive voltage along the length of the 

fuselage. This voltage would appear in the line-to-ground measurement but not in the line-to-line measurement which observes 

only the voltage in the loop between the two wires of the circuit. 

62 
Induced Voltage: Line to ground Line to line 

2.0 v 666 0.5 us 0.5 v 665 

Model 

L = 2.5 JJH 

C = 220 pF 

E= 80V 

I = 0.8 A 

Full 

Scale 

e Nose-to-Return 

L = 25 ~-IH 

; 1 ;;O(&PF 

i 

I = 250 A 

100 A 582 1.0 us 5 kV 588 

0-l w Figure 35 - Simulated Leader Attachment to Unterminated Aircraft. 
Model 

L= 2.5 ~.IH 

C = 220 pF 

E= 100 v 

I=lA 

Full 

Scale 

L= 25 PH 

V 662 0.5vs 0.5 v 661 0.5 us 

0.5 A 510 0.1 I.ls 

i i i 

0.5 A 511 0.1 ps 0.5 A 512 0.1 us 

100 A 601 1.0 I.ls 

500 v 

100 A 600 1.0 us 

500 v 

100 A 617 1.0 us 

500 v 

Figure 36 - Leader Attachment with Aircraft Terminated. 
Figure 36 illustrates a similar case as that of Figure 35, 

except that an uncharged ladder network is attached to the tail 

of the aircraft. This has the effect of terminating the aircraft 

in its surge impedance of 100 ohms, and of providing a path for 

the leader to continue onward from the aircraft. Whether this is 

a better simulation of the actual leader attachment situation or 

not is not known. If an aircraft is intercepted by a leader, continued propagation of the leader will occur, but it takes place 

in a series of steps, with a pause of many microseconds between 

each step. 

As shown in this figure, the peak amplitude of the airframe 

current is about the same as that of Figure 35 when the aircraft 

was unterminated, but the duration of the current surge is longer 

and reflections are essentially eliminated. This is due to the 

100 ohm termination of the aft end of the aircraft. It will also 

be noted that the rate of rise of the airframe current is greatest 

at the nose of the aircraft where the switch is located. It is 

somewhat less at the tail, and even slower at the earth end of 

the simulated channel. The airframe voltage is very similar in 

waveform to the current, due to the absence of reflections, and 

there is excellent similarity between the model and full scale 

current waveforms. 

In this circuit, line-to-ground induced voltage shows a 

greater IR component than was evident when only the attaching 

leader channel was present. This is because the current remained 

at its peak level for a longer period of time, permitting more 

current to diffuse to the interior of the aircraft skin and cause 

a more pronounced resistive voltage rise, which is superimposed 

on the magnetically induced component. Only the magnetically induced component appears in the line-to-line measurement, of course. 

The majority of the voltage still appears to be related to di/dt 

though. 

Return stroke simulation. - By moving the switch to the 

ground end ofthe ladder network and beginning with the entire 

channel charged, the return stroke can be simulated. By closing 

the switch, the channel begins to discharge at the ground end 

first, creating a current wave that propagates up the channel in 

the manner of a return stroke. This situation is shown in Figure 37. This time, the current rate of rise is greatest at the 

ground end of the channel because the switch is located there. 

The amplitude remains approximately the same but the rate of rise 

diminishes as the wave progresses further up the channel. 

Since the current rate of rise is now much slower at the 

aircraft than was the case in the leader attachment simulation of 

Figures 35 and 36, the magnetically induced voltage is also much 

lower than that measured during leader attachment. The resistive 

65 
z Induced Voltage: 

Line to line 

- 1.8V 

2.0 v 631 0.5 ps 0.2 v 638 

Model 

L= 2.5 PH 

C = 220 pF 

E = 170 V 

I = 1.5 A 

Full 

Scale 

L = 25 UH 

C = 2000 pF 

E = 25 kV 

I = 218 A 

100 A 569 1.0 ps 

Return Stroke Through Aircraft 

50 A 
voltage, which appears only in the line-to-ground measurement, is 

now higher than before but not more than l/2 of the total lineto-ground induced voltage and is approximately the same amplitude 

as the resistive voltage induced.by.the leader current in the 

instrumented circuit of Figure .36. This is because the resistive 

.voltage is directly proportional to the airframe current and the 

current in these circuits is, in turn, proportional to the charging voltage divided by the 100 ohm surge impedance, and the charging voltage and surge impedance were the same for all of these 

tests. 

This is not to imply that the transmission line charging 

voltage should be the same for a leader simulation as for a return stroke simulation. Additional discussion of this matter 

appears in a later paragraph; and in any event, additional information is needed on leader and return stroke channel characteristics 

to clarify this aspect. 

Study of the full-scale current oscillograms of Figure 37 

shows that the duration of the current wave is proportional to 

the length of the charged transmission line and that the effective 

velocity of propagation is about one-half the speed of light. 

At the ground end, the 

stroke current (until sharp 

638 of Figure 37: 

total time duration (T) of the return 

drop-off begins) is, from oscillogram 

This should be the time necessary for the current wave to propagate all the way up the channel and discharge the entire length. 

For a total simulated channel length (R) of 150 m plus the aircraft length of 17 m, the effective velocity (v) was 

V = T effective a (22) 

= 2.5~10-~s 

167m 

= 1.5~10' m/s 

Considering the imperfections in the simulation, this result compares favorably with the goal of representing a channel with an 

effective velocity of 1x10' m/s. 

The return stroke current duration in the model test was one 

tenth of that in the full scale test, producing an effective velocity along the one tenth scale channel of also about 1.5x10* m/s. 

67 
It should be remembered that the actual velocity of current 

flowing in the lumped parameter ladder network remains at (or 

very near) the speed of light. The lumped parameters in the network make the network appear physically longer than it actually 

is. 

Induced voltage relationshipA, - As mentioned in the foregoing paragraphs, cause-effect relationships appear to exist between the induced voltages and the airframe currents, with lineto-ground voltages appearing to be a combination of structural 

resistive as well as magnetically induced voltages, and the lineto-line measurements indicating only induced effects. A closer 

comparison of these measurements for each of the conventional 

LTA and lumped parameter ladder circuits evaluated is presented 

in Table VI. In this table, oscillograms of the wavefronts of 

the airframe currents generated by each test circuit are shown, 

together with values of the peak current maximum rate of rise 

and peak induced voltages. 

A comparison of the five results shown on Table VI indicates 

two important results: 

appear to be 

2. Line-to-line voltages appear to be 

related to current rate of rise. 

For example, the line-to-ground voltages appearing in the instrumented aircraft circuit during the leader attachment in test 666 

of Figure 35 are about the same as those measured in the returnstroke simulations of test 631 of Figure 37 and 680 of Figure 31, 

when the peak currents were approximately the same; even though 

the rate of rise in the return stroke simulations was much lower. 

Conversely, the line-to-line voltages were significant only during the fast rate-of-rise tests. The data of Table VI have been 

plotted on Figures 38 and 39 to further illustrate this result. 

In each graph, one of the test results is displaced significantly 

from the line approximating all of the others. The reason for 

this is unclear, and may simply be due to an experimental error. 

Unfortunately, the'data were not compared in this manner until 

after the full scale test program was complete and the setup 

dismantled, so the test could not be repeated. 

The relationships- between induced voltages and simulated 

lightning current in the airframe are further illustrated in 

Figure 40, in which the oscilloscope traces of typical line-toline and line-to-ground voltages measured in the instrumented 

circuit are shown together with the test current. Since the 

circuit voltages were recorded with an on-board oscilloscope it 

68 
Table VI -Current Rate-of-Rise vs. Induced Voltage 

Peak Maximum Peak Induced 

Test Current (i) and Voltage (e) Current di/dt L-G 

Config. Entering Aircraft (A) (A/w) vJ> 

Leader 

Attachment 

to i 

Unterminated 

Aircraft 

i Return 

Stroke 

250 2055 -1.16 

(Test 666) 

100A 583 O.lus 

218 393 -1.8 

(Test 631) 

40A 573 0.2ps 
Table VI - Current Rate-of-Rise vs. Induced Voltage (continued) 

Peak Maximum Peak Induced Voltage 

Test Current (i) and Voltage (e) Current di/dt L-G L-L 

Config. Entering Aircraft (A) (A/w) w w> 

Lumped Circuits 

2X5OjJS 

Return 

Stroke 

with 

Aircraft 

Terminated 

inZ, 

2X5OUS 

Return 

Stroke 

with 

Aircraft 

Short 

Circuited 

190 140 

Lost in 

-1.80 noise 

(Test 680) (Test 681) 

i 

50A 679 0.5ps 

500 886 -4.10 -0.24 

(Test 689) (Test 690) 

i 

1OOA 688 0.5jJs 
Line 

to 

Line 

Voltage 

09 

0.4 

0.3 

0.2 

0.1 

(690) 0 / 

0 (66% 

0 1 

Current Rate 

Figure 38 - Line-to-Line Vo 

2 3 

of Rise (kA/ps) 

ltage vs. Current Rate of Rise. 
5 

4 

Line 3 

to 

Ground 

Voltage 

w 2 

1 

0 

0 (689) 

100 200 300 460 

Peak Current Amplitude (A) 

500 

Figure 39 - Line-to-Ground Voltage vs. Peak Current Amplitude. 
Current 

Line-to-Line Voltage 

Line-to-Ground Voltage 

At Ipk, 

Structural Resistance = 

R,= =V 

25oA = 0.004 s-2 

Figure 40 - Relationship Between Induced Voltage 

and Simulated Lightning Current. 

73 
was not possible to positively synchronize the measured voltages 

with the test current that was recorded by oscilloscopes'outside 

of the aircraft. Thus, the time relationship indicated in Figure 40 by the dashed line is an estimate. 

If this time relationship is correct, it is evident that 

the line-to-line voltages are maximum when the current is changing most rapidly, pass through zero when the current is maximum 

and unchanging, and become opposite in polarity when the current 

is changing again in the opposite (decreasing) direction. This 

is a derivative relationship that indicates the voltage is induced by the changing magnetic flux (Q) that accompanies the test 

current according to the familiar relationship: 

e=-- d9 

dt (23) 

On the other hand, the voltage measured between either of 

the aircraft circuit conductors and the airframe (the line-toground voltage) appears to be a combination of magnetically induced and structural resistive voltages, as would be expected. 

This is indicated by the fact that the voltage reaches its peak 

during the time the di/dt is maximum, but stays positive until 

the time when the current is maximum when it is lower than for 

the line-to-line voltages, verifying the existence of an in-phase 

resistive component. 

If the time relationship of Figure 40 is correct and if the 

line-to-ground voltage is in fact a simple combination of magnetically induced and structural resistive components, the voltage at peak current should be due to resistance only, and the 

structural resistance (Rs) responsible for this component can be 

estimated from the relation: 

Rs =?- 

i 

=: l.OV - = 0.004 ohms 250A 

At first glance, a resistance of 4 milliohms along 8m of conventional aluminum fuselage appears somewhat high; but since most of 

this resistance is probably in the fastened joints and seams which 

are heavily treated with anticorrosion (and therefore non-conductive) coatings, this result may indeed be realistic. 

To obtain another data point on the aircraft structural resistance, an ohmmeter measurement was made along the fuselage using 

a kelvin-type bridge. Several readings were made which yielded 

resistance values of 5 to 7 milliohms. 

74 
Streamer initiation effects. - A high frequency oscillation 

appearing as "hash" is superimposed on the induced oscillograms 

of Figures 35 and 36. It was suspected that this resulted from 

high voltage streamers that occur in the switch gap during the 

switching process or from airframe corona effects. 

In an attempt to evaluate this a segmented diverter was 

inserted in series with the switch in the leader attachment simulation of Figure 36. This device provides a longer path for 

streamers to develop prior to complete closure of the gap and 

current flow onto the airframe, thus forcing streamer development 

to begin a finite period of time prior to leader current flow 

from the ladder network onto the aircraft. The result of this 

test is shown on Figure 41, where it is evident that the high 

frequency oscillations now occur well in advance of the leader 

current. When separated from the high frequency oscillations, 

the leader current pulse appears quite smooth. A region of "hash" 

remaining on the decay of the leader current pulse may be due to 

an instability in the arc along the segmented diverter, but the 

smoothness of the front and peak of the leader pulse is unmistakable. This is apparent on oscillogram 620 as well as 622, in 

which the test was repeated with a slower oscillograph sweep to 

enable the entire leader current to be captured. The leader current and airframe voltage appearing at the aft end of the aircraft 

are also shown on Figure 41. 

The 11.1 MHz frequency of the streamer initiation effects 

measured in the instrumented circuit was the highest frequency 

observed during any of the induced voltage measurements. It 

does not appear elevated to any of the other frequencies observed, 

and appears too high in frequency to be due to traveling waves in 

the airframe. These "streamer effects" were always considerably 

lower in amplitude than the voltages induced by simulated leader 

or return stroke currents in any of the test circuits evaluated, 

and time did not permit further evaluation of them in this program. 

Extrapolation of induced voltages. - In the past, voltages 

induced during LTA tests have usually been extrapolated to correspond with full scale current amplitudes of 30 kA to represent an 

"average" lightning stroke and 200 kA to represent a severe stroke 

that is thought to be exceeded in amplitude only 1% of the time. 

The LTA tests have usually been performed with 2x 50 PS current 

waveforms that provide an average rate of rise of 150 A/ps. 

Thus, linear extrapolation according to current amplitude has 

carried with it a linear extrapolation of rate of rise to 150 

u/sls. 

The waveforms produced by the lumped parameter ladder networks have a higher maximum rate of rise than that produced by 

75 
Induced Voltage L-G 

320A23 kVi 

e 

Voltage Induced by 

Streamer Initiation 

Period = 0.09 ps 

f = 11 . 1 flHz 

Induced Voltage L-G 

100 A 619 1 us 

5 kV 

3.15 v - 

1.0 v 622 1.0 us 

Figure 41 - Separation of Voltages Induced by Streamer 

Initiation and Leader Current Entry by 

Use of Delayed Sparkover. 

76 
the original LTA circuit, for a given current amplitude, so extrapolation of voltages induced by the ladder networks according to 

current amplitude may predict unrealistically high voltages. 

Since it has been found that some induced voltages appear to respond to current rate of rise and others to current amplitude, it 

appears proper to extrapolate induced voltages according to the 

current parameter that they are related to. For example, in the 

circuit instrumented in this program, the line-to-ground voltages 

were related to peak current, and the line-to-line induced voltages were related to current rate of rise. Examples of how these 

extrapolation factors may be determined for voltages measured in 

the original LTA circuit as well as the 100 ohm lumped parameter 

ladder network are shown on Tables VII and VIII. In each case, 

the extrapolations are to a severe current amplitude of 200 kA 

and a severe rate of rise of 100 kA/us. Several things are of 

interest from the analyses of Tables VII and VIII. 

1. The predicted line-to-ground voltage is considerably greater than the line-to-line voltage, indicating the protection afforded by independent 

return circuits. 

2. The predicted voltages resulting from the 100 ohm 

ladder network test are quite close to those resulting from the original LTA test. 

3. Predicted line-to-line voltages, which are of primary concern to digital data transfer circuits in 

such systems as fly-by-wire flight controls, would 

be several times higher if they had been extrapolated according to amplitude along with the line-toground data, as was usually done in the past. 

4. The need to apply the correct extrapolation factor 

is therefore apparent, and emphasizes the need, in 

turn, to correctly identify the dominant cause 

factors during the tests. 

Further investigation of transit times. - As described earlier in this report, attempts to derive the traveling wave transit 

time and velocity from the frequency of airframe voltage or current oscillations resulted in indicated velocities lower than the 

speed of light. 

Other attempts to determine the traveling wave transit time 

were made by measuring the elapsed time between appearance of 

the wavefront at the nose and tail of the aircraft. This was 

done for both the current and voltage waves, and was accomplished 

by triggering of the oscilloscope for all measurements with a 

signal from the Rogowski coil on the nose boom. The cable 

77 
TABLE VII -Extrapolation Analysis Using the 

Original LTA Circuit 

(using data from Tests 680/690of 

Table VI and Figure 29) 

1. Line-to-Ground Voltages were most closely related 

to Peak Current, so: 

L-G Extrapolation Factor % 200 kA 

500 A 

'Ir 400 

2. Line-to-Line Voltages were most closely related 

to Maximum di/dt, so: 

L-L Extrapolation Factor Q 100 kA/vs 

0.89 kA/ps 

'L 112 

3. Thus, for the Return Stroke Phase: 

Line to 

Ground 

Line to 

Line 

! 

Test Ext. Predicted 

Data (V) x Factor = Voltage 

400 1640 

I 112 

I 

27 

78 
TABLE VIII - Extranolation Analysis Usine: the L lOOs2 Ladder ketwork - 

(using data from Tests 631/638of 

Table VI and Figure 37) 

1. 

2. 

3. 

Line-to-Ground Voltages were most closely related 

to Peak Current. so: 

L-G Extrapolation Factor % 200 lu4 

218 A 

Line-to-Line Voltages were most closely related 

to Maximum di/dt, so: 

L-L Extrapolation Factor % 100 kA/vs 

0.39 kA/lls 

'L 256 

Thus, for the Return Stroke Phase: 

1 

I Test Ext. Predicted 

Data (V) x Factor = Voltage 

Line to 

Ground 

Line to Line 

1.66 917 1523 

0.04 256 10 

I 

79 
transmitting this signal to the oscilloscope was shorter than that 

bringing the traveling wave measurement from either the nose or 

tail, providing a short delay time that enabled the start of the 

waves always to be seen. The same instrument cable and probes 

were used to make the measurements at the nose and tail. 

The current wave measurements are shown on Figure 42 and the 

voltage waves on Figure 43. 

For the model, a velocity of 2.09~10~ m/s was determined, 

but in the full scale test using the same circuit, a velocity of 

only 1.39x108 m/s is determined, as shown on Figure 42. Note 

that the oscillograph sweeps are the same for the model and full 

scale test currents in this figure. Thus, in the full scale situation the beginning of current rise is less abrupt. For the purpose of establishing more definite times of arrival, the intersect 

of the second graticule line above the zero line was selected. 

This should be valid if the waveforms are identical, which was 

verified by superimposing one upon the other. The reason for the 

apparently slower traveling wave velocity between the model and 

full scale tests is not clear, 

In the tests of Figure 43 the traveling wave voltage waveform 

appearing at the nose and tail were measured and the difference in 

their times of arrival was determined to be 0.07 ps, which resulted in a velocity along the 17 m F-8 aircraft of 2.45x108m/s. 

This was the highest velocity determined from any of the transit 

time measurements, and is about 80% of the velocity of light. The 

reason that the voltage waveform of this circuit apparently propagated along the aircraft faster than the current waves of Figure 

42 is also unclear. 

The influence of generator circuit elements on the frequency 

of airframe voltage oscillations was found to be significant in 

other tests in which the voltage between the unterminated tail of 

the aircraft and the adjacent return lines was measured when series inductance and resistance was inserted between the generator 

capacitor and the airframe, as shown in Figure 44. From these 

results, it is clear that the period of oscillation depends significantly upon the lumped test circuit elements as well as the 

length of the aircraft. In other words, the voltage observed at 

the open ended tail of the aircraft is dependent upon all elements 

in the circuit loop, not just the aircraft. The addition of a 

20 PH inductor, for example, doubles the effective length of the 

circuit and produces a period twice as long as that when no lumped 

inductance is present. Finally, addition of a series resistance 

reduces the reflection magnitude the aircraft appears to become 

slowly charged to the capacitor voltage. 

80 
-1.67m-- 

= 

O.'O78us 

'I = (0.078- 0.070)~s = 0.008+.002 1~s 

- 

v=-= d 1.67m 

T 0.008~10-~s 

= 2.1+0.7x108 m/s - 

581 580 

Full 

Scale 

Switch 

at 

Earth 

(O.lp I 0.lp.s 

0.58~~ 0.46~s 

T = (0.58- 0.46) us = 0.12~s 

d 1.67m v = - = 

-r o.12x10-6s 

= l.39x108 m/s 

Figure 42 - Transit Time Determined from Current 

Wave Entry and Exit Times. 

81 
. . . ,,._ - 

I 

0.0311s 0.05P.s 

q-5,yF / jyjj---t 

1OkV ? 

T = (0.10-0.03) x 10-6s 

= 0.07 x 10-6s 

from which 

d - 17.15m v=-- 

T 0.07x10'"s 

= 2.45~10' m/s 

Figure 43 - Transit Time Determined from Voltage Wave Entry and 

Exit Times. 

l Closer to speed of light than when 

derived from the current waves. 

82 
next to 

Wing Period = 0.54 US 

T = 7 = 0.135us 

-- 

e 

4 foils 

132t 

= 

f = 900 kHZ - 

0.5 13 

-\t---/ 1OkV --p-G--k 

e 

4 foils 

\ > 1.4 L :. 137i-i 1 2

c 

= & 

e 

2.5kV 213 1.01.ls 

Period = 1.1~~ 

T =1.1= 

4 0.275~s 

Figure 44- Airframe Voltage Oscillations vs. Driving Circuit. 

83 
Review of all of the traveling wave transit time evaluations 

conducted during this program indicates a trend toward the speed 

of light as the diameter of the center conductor diminishes. 

Parallel wires or coaxial arrangements of either wires or insulated cables showed velocities closer to those expected, whereas 

when an aircraft or large cylinder becomes the center conductor, 

the velocity diminishes. Discussions with researchers familiar 

with the theory of antennas and the interaction of solid conductors within an electromagnetic field have disclosed that this 

phenomenon should, perhaps, have been expected. It has been shown, 

for example (ref. 21) that the velocity of propagation of waves 

on a thin wire in an electromagnetic pulse field can be less than 

the speed of light, and that larger diameter conductors may experience slower velocities. 

Notwithstanding this, the (supposed) transit time and velocity discrepancies were causes of concern in this investigation! 

and indicate the need for further investigation aimed at determining a physical exp,lanation for this phenomenon. 

Simulation of longer lengths of leader channel. - As described earlier, the lumped parameter ladder network simulated in 

the full scale tests consisted of six 25 m segments which represented a total 150 m length of channel. Simulation of a longer 

length would, of course, require additional capacitors and inductors. In order to evaluate the test current produced by a longer 

length of simulated channel, a 20-segment network representing 

500 m of channel was evaluated on the scale model. Each segment 

had the same length and circuit elements as in the previous model 

tests. The resulting currents are shown in Figure 45. 

In Figure 45, which represents a return-stroke, oscillograms 

470, 468 and 469 show the current entering the nose, exiting the 

tail, and at the ground end of the channel. The duration of the 

current at the ground end is longer than at the aircraft, due to 

the longer length of channel to be discharged. As before the 

fastest rate of change appears at the switch, which in this case 

is located at the ground. Comparison with the same situation in 

a shorter channel of Figure 37 illustrates that the rates of current rise and decay are similar. The one difference is the duration of the current pulse. 

Oscillogram 483 shows the current in each of the 20 segments, 

from the earth to the cloud end. Here again it can be seen that 

the stroke duration is longest at the ground end and shortest at 

the cloud. It must be remembered. of course. that these simulations represent channel discharge'phenomena only and do not include simulation of the intermediate or continuing current discharges that are caused by drainoff of the charge-remaining in 

the original cloud center. These are much slower and lower in 

amplitude currents that do not significantly influence induced 

voltages. 

84 
1 1 1 - 

0.5 A 470 0.1 I.ls 0.5 A 468 0.1 1J.s 0.5 A 469 0.1 us 

Current in each Segment 

Model 

L = 2.5 PH 

C = 220 pF 

Rt= 100 R 

earth 

cloud 

0.2 A 483 0.1 US 

Figure 45 - Return Stroke Current in a 500 m, 20-Segment Network 

(Each Segment Represents 25 m of Lightning Channel.) 
Following the 20-element evaluation, a second test was made 

to see if the 500 m channel could be represented with fewer segments, each representing a longer length of channel. If this 

were satisfactory, a smaller'number of components would be necessary. The results of this evaluation are shown on Figure 46. 

The lumped inductors and capacitors in this network are each four 

times larger than those of the 25 m segments. Thus, each segment 

now represented was 600 m. As shown in Figure 46, the amplitude 

of the stroke current waveforms was equivalent to that of the 

currents in the 25 m segments, but the rate of rise of the currents entering and leaving the aircraft is much less than that 

produced by the shorter segments. Clearly, the larger segments 

act less like a transmission line and are not appropriate for use 

in this test circuit. 

Additional cause-effects evaluations using the model. - To 

obtain another look at the causes of the induced voltages, a series of tests was made on the one-tenth scale model with a-single 

wire installed in approximately the same location as the instrumented circuit in the full-scale F-8. The wire was grounded to 

the model fuselage, so induced voltage measurement is representative of a line-to-ground situation. 

According to the model scaling laws of Table III the conductivity of the model fuselage material should be ten times that of 

the full scale aircraft materials if structural IR rises are to be 

faithfully represented in the model. Such materials are not available so the model was made of aluminm with the same conductivity 

as the full scale F-8. The location and size of apertures in the 

model are also not good replicas of those in the F-8. Thus, the 

amplitude of the induced voltages measured in the model cannot be 

compared directly with those observed in the full scale aircraft, 

and no attempt was made to do this. 

The model induced voltage waveforms should, however, be fairly representative. Measurements were made using the RLC test circuit with the aircraft terminated in an open circuit, resistor 

equal to its surge impedance, and short circuit to the return 

lines. The original LTA RLC circuit and a similar circuit modified to include a resistive termination were also studied. In 

addition to the induced voltages, measurements were made of the 

voltage surge appearing between the fuselage and an adjacent return line, and of the current exiting the tail of the aircraft. 

The test circuit and circuit elements for each test are shown 

together with the measurements in Figure 47. 

The measurements shown in Figure 47 indicate the induced 

voltage is clearly related to the traveling waves on the model 

fuselage. The first case shown is with the aircraft terminated 

in an open circuit. 

86 
F-8 

0.5 A 492 0.1 j.l.5 0.5 A 490 0.1 F.rs 0.5 A 488 

Figure 46 - Return Stroke in a 500 m, 6-Segment Network 

(Each Segment Represents 100m of Lightning Channel.) 
Test Condition -____ 

Open Circuit 

R, = 0 

L, = 0 

Rt = m 

Terminated 

n, = 0 

L, = 0 

R, = 82 R 

Short Circuit 

R, = 0 

L, = 0 

a, = 0 

STU 0.2x5.0 us 

R, = 43 R 

L, = 1.7 IIH 

Rt = 0 

Terminated 0.2x5.0 

R, = 0 

L, = 1.7 llti 

Rt = 83 R 

2v 

m 2v 

m 

-1 v 

0.5 v 

m 

2v 

Note: 0.50 us/div sweep for all oscillograms. 

0.5 A 

Figure 47 - Geometric Scale Model Induced Voltage Measurements. 

88 
When the mercury wetted switch is closed, the capacitor voltage is immediately impressed on the front of the model. This 

voltage wave propagates down the fuselage on the transmission 

line formed by the model and the return lines. The surge impedance (Z,) of this system was determined to be about 100 ohms. 

Therefore, associated with the voltage is a current wave of magnitude equal to the capacitor voltage divided by Zo. The rise time 

of this wave will be fast, approximately equal to the voltage wave 

rise time. When the traveling voltage and current waves reach the 

open end of the fuselage, the voltage wave reflects back toward 

the nose. Ideally, this would occur at the same amplitude (E) as 

it began, and with the same polarity. Thus, it would return at 

double the original amplitude (E + E = 2E) while the current wave 

(I) reverses in polarity and returns, cancelling out the current 

(I - I = 0). 

Back at the nose of the aircraft, the traveling waves encounter the capacitor. The capacitor acts as a short for the transmission line and causes the voltage wave to reverse. However, only 

the portion reflected at the open end (E) reverses, so the voltage 

on the line after the wave again reflects if 2E - E = E. At the 

capacitor, the traveling current wave (-1) is reflected without 

polarity change and returns. So the current on the line after this 

reflection is -1. When these reflections again reach the far end 

of the model fuselage, the voltage and current traveling waves (-E, 

-1) reflect as before with the voltage remaining the same polarity 

(-E) and current changing polarity (I). After these waves start 

back to the nose, 

(E - 

the voltage and current on the line are both zero 

E = 0, -I + I = 0). At the nose the waves reflect as before, 

with voltage reversing and current remaining the same polarity so 

the traveling waves (E, I) and the voltage and current levels on 

the line are the same as at the start. The cycle can then repeat 

and would do so forever if there were no losses on the line. 

Losses due to insulation resistance and stray capacitance and inductance remove energy from the line by absorbing it or diverting 

it into other systems. Thus the reflections shown in the fuselage voltage wave in the first case of Figure 47 are less than the 

original amplitude. 

In the ideal case of a lossless transmission line, the traveling waves would be perfect square waves with durations equal to 

the wave transit times. However, the losses in the fuselage 

system are quite high and the waves never become the ideal square 

waves. Thus, visualization of the wave propagation described 

above is somewhat difficult. To help identify traveling wave 

phenomena in the other cases, traveling wave voltage and current 

measurements were made on a 50 ohm (Z,) coaxial cable system 

where the losses were lower and the square waves more identifiable. The traveling voltage and current waves were measured in 

the cable at three points corresponding to the nose, mid-fuselage 

and tail for three of the termination conditions of Figure 47, and 

89 
..- -._ ..-. _ ____ -. .__ 

the results are shown on Figure 48. One test was also conducted 

with a lumped inductor inserted in series with the input nose end, 

and it can be s,een that the inductance changes the square waves to 

triangle-like waves. In this 50 ohm cable test, the center conductor represents the aircraft and the cable shield represents the 

return conductors. 

By studying Figure 47 and 48, the model fuselage currents can 

be determined and from them the current changes (dI/dt) which 

should be related to the induced voltages in the model. For the 

first condition with the aircraft terminated in an open circuit, 

the.current is continuously oscillating, so there should be an induced voltage that is the derivative of the current with the same 

basic frequency as the current. This derivative relationship is 

apparent in the induced voltage oscillograms of Figure 47, and it 

is also evident in the full scale test of Figure 35, for example. 

In the third case, with the aft end of the model short circuited to the return lines, the voltage measured in Figure 47 

shows that the traveling waves are reflecting back and forth along 

the fuselage similar to the behavior of the cable currents shown 

in Figure 48. A superimposed low frequency oscillation is responsible for the waves not returning to zero, and losses cause the 

distortion of the traveling waves. The oscillation frequency of 

the voltage is about twice that of the open circuit case, since 

now the current traveling wave frequency is twice the open circuit 

case, as shown in Figure 48. 

In the final two cases (4 & 5 of Figure 47), inductors were 

inserted between the capacitor :and the model. In the fourth case, 

the model tail remained shorted, and in the final case it was terminated on an 82-ohm resistor. The rate of rise of the traveling 

current waves is reduced in both cases by the inductor, which resulted in lower amplitude induced voltages, as evident in the final 

two induced voltage oscillograms (Figure 47). 

It should also be noted, in comparing all of the induced voltages of Figure 47, that the first induced voltage peaks in the 

first three cases are nearly the same amplitude, since the wavefront 

of the first traveling wave was the same in each of these three 

cases. 

RECOMMENDATIONS 

A. Any of the circuits evaluated may be utilized for LTA 

testing provided the driving cause-effects relationships are established, and accommodated during data analysis as described in 

Tables VII and VIII. 

90 
Test Conditions 

Inductive Input 

2 pH 

Shoreed 

R=0 

AOrx : All Oscillogrms taken ar 0.1 usldiv. 5 V/div and 1 Ajdiv 

Upper trace is voltage; Lower trace is current. 

Figure 48 - 50 Ohm Cable Transmission Line Voltage 

and Current Traveling Wave Measurements. 

91 
B. The lumped parameter ladder network may be- an improvement 

because: 

1. 

2. 

3. 

4. 

5. 

6. 

C. 

1. 

It is a better simulation of a lightning channel, because 

it represents the channel surge impedance as well as the 

charge per unit length of a typical lightning channel. 

Lightning voltage and current may be controlled and related by the surge impedance of the simulated channel and 

of the aircraft with respect to its return conductors. 

Impedance matching between the simulated lightning channels and the aircraft can be known and controlled. 

Both leader attachment and return stroke phases can be 

represented. 

The ladder network is efficient. Fast rates of rise and 

high currents can be obtained with reasonable DC charging 

voltages, as shown in Figure 49. This figure shows the 

maximum rate-of-rise and peak currents that can be obtained from the 100 ohm ladder network, the original 

lumped RLC LTA circuit, and the lumped RLC circuit with 

the aircraft terminated in its surge impedance (Z,), as 

a function of generator capacitor charging voltage. Both 

of the lumped RLC circuits produce the basic 2 x 50 US 

return stroke waveform, but the unterminated circuit 

produces reflections that act to increase its maximum 

rate of rise. A study of the figure shows that the ladder network produces the best combination of rate of 

rise and peak current. 

It produces a return-stroke waveform with concave front 

more representative of lightning than the double exponential waveform, as shown in Figure 50. This results in 

the maximum rate of rise occurring closer in time to the 

peak current than is the case with the lumped RLC circuit, 

in which the maximum rate of rise is at t = O+. A study 

of return-stroke waveforms recorded in the literature 

(ref. 22) shows that many lightning return strokes have 

concave fronts, as shown on Figure 51. 

Whichever test circuit is used: 

The relationship between induced voltages and airframe 

currents or their rates of change should be determined 

and appropriate extrapolation factors calculated. In 

particular: 

92 
50 

30 

Max. 20 

Rate 

of 

Rise, 

difdt 

(WlJs) 

10 

Test +a, 

ab ' 

Level 8~' 

0 0 4 0 0 / 0 

30 50 100 200 300 

Capacitor Charging Voltage (kV) 
Original RLC Circuit 

1OOA 622 0.5PS 

Figure 50- Wavefront Comparison. 

94 
0 10 20304050 100~s 0 10 20 30 40 50 loops 0 10 20 30 40 50 

kA' 

0 5 10 15 20 25~s 0 5 10152025~s 0 5 10 15 20 25~s 

0 5 10 15 20 25~s 0 5 10 15 20 25~s 0 5 10 15 20 25~s 

-20 km-20 lx----- -207 

-200r -20Ky 1: 2; ?5w 

Figure 51 - Front Waveshapes of Lightning Currents as 

Measured by Berger. 

95 
A. 

B. 

C. 

Voltages dependent on amplitude should be 

extrapolated by a factor: 

200 kA 

ITest 

Voltages dependent on rate of rise should 

be extrapolated by a factor: 

100 kA/PS 

diNttest 

The above factors may be the same if the 

test current has a waveform that is a proportional replica of the lightning current 

waveform that is desired to be simulated. 

2. Traveling waves between return conductors and the surrounding building may affect the test results. These effects 

may be reduced by terminating each end of the aircraft return conductor array to building ground with a resistance 

equal to their surge impedance with respect to the building. This will be approximately 150 ohms in many cases. 

3. At least 4 return conductors should be used, spaced symmetrically around the aircraft. Either wires or foils may 

be utilized. 

D. The geometric scale model technique is a powerful tool for 

evaluating test clrcurts needed to obtarn desired current and voltage waveforms and generator requirements. 

CONCLUSIONS 

This investigation resulted in an improved understanding of 

the cause-effect relationships inherent in the induced voltage process taking place in aircraft electrical circuits when the aircraft 

is subjected to lightning-like current pulses. It also enabled 

some evaluations to be made of several modified LTA test circuits, 

with the objective of identifying one(s) that might be improvements. 

While this investigation did not answer all of the questions surrounding the LTA process, some conclusions are apparent from the 

work reported herein, and some improvements in the LTA test method 

appear possible. These are summarized on the following page: 

96 
Induced voltages. are due primarily to magnetic flux and struccurrents in t e air rame. - This 

each of the induced voltage oscillograms with the airframe current that apparently caused them. 

Voltages must be present to drive the airframe currents, but 

since less is knoti abouiijgmning voltages than about currents, 

it seems best, for the present, to relate LTA test criteria to 

airframecurrents. - Measstning currents that 

actually flow in the aircraft during the leader attachment and 

return-stroke phases of the lightning attachment process must be 

obtained to ascertain the full threat levels that should be simulated or extrapolated to in future LTA tests. Measurement of 

airframe currents, however, is a much simpler task than measurement of airframe voltages, due to the absence of a reference. 

craft is unterminated (floatingl 

Substantial current may flow in the airframe even if the air- 

--"') when subJected to a high voltage 

impulse. - In this case the current is solely related to the applied voltage via the airframe surge impedance. The duration of 

such currents is short, and they will oscillate back and forth 

along the airframe. 

The frequency of oscillatory currents depends on: 

1. The length of the airframe and 

2. The characteristics of the attached lightning 

generator circuit. 

Test circuit generated traveling waves appear to travel at 50% 

to 80% of the speed of light through an airframe. - The actual 

velocity appears to depend on the diameter of the fuselage or wing 

structure, and the configuration of the return conductors. It may 

not be possible to identify a definite relationship between an aircraft test circuit and traveling wave transit times, due to the 

complex geometries involved, but this not necessary since precise 

simulation of induced voltage frequencies is not necessary for 

assessment of component damage effects. 

Lightning current surges may either. - 

1. Travel once through the airframe, from entry to exit point. 

2. Enter the airframe, reflect back and forth, and exit. 

3. Enter and reflect back and forth only. 

What happens in a particular case may depend on the characteristics of the attached lightning channels. 

97 
REFERENCES 

1. Lloyd, K.J., Plumer, J.A. and Walko, L.C.: Measurements and 

Analysis of Lightning-Induced Voltages in Aircraft Electrical 

Circuits. NASA CR-1744, 1971. 

2. Plumer, J.A.: Analysis and Calculation of Lightning-Induced 

Voltages in Aircraft Electrical Circuits. NASA CR-2349, 1974. 

3. Walko, L.C.: A Test Technique for Measuring Lightning-Induced 

Voltages on Aircraft Electrical Circuits. NASA CR-2348, 1974. 

4. Maxwell, K.J., et al. : Computer Programs for Prediction of 

Lightning-Induced Voltages in Aircraft Electrical Circuits. 

AFFDL-TR-75-36, Volume I, 1975. 

5. Plumer, J.A., Fisher, F.A. and Walko, L.C.: Lightning Effects 

on the NASA F-8 Digital Fly-by-Wire Airplane. NASA CR-2524, 

1975. 

6. Plumer, J.A.: YF-16 #l Lightning Transient Analysis Test 

Report. General Dynamics Report 16PR051, 1975. 

7. Space Shuttle Program Lightning Protection Criteria Document. 

JSC-07636, Revision A, 1973. 

8. Cianos, N. and Pierce, E.T.: A Ground-Lightning Environment 

for Engineering Usage. Stanford Research Institute Technical 

Report No. 1, 1972. 

9. Nanevicz, J.E., Bly, R.T. and Adamo, R.C.: Airborne Measurement 

of Electromagnetic Environment near Thunderstorm Cells. Proceedings of the 1977 IEEE International Symposium on Electromagnetic Compatibility, IEEE Publication 77CH 1231-O EMC. 

10. Bewley, L.V.: Travelling Waves on Transmission Systems, 

Second Edition, Dover Publications, Inc. 1963. 

11. Fisher, F.A.: Geometric Model Studies of KSC Launch Facilities. 

General Electric Report SRD-76077. 

12. Fisher, F.A.: Feasibility of Geometric Models for the Determination of Shielding Effectiveness. General Electric Report, 

1966. 

13. Crouch, K.E.: Lightning Current Distribution on a Scale Model 

Space Shuttle Rocket Engine. General Electric Company Corporate Research and Development Report SRD-73-177, 1973. 

98 
REFERENCES - continued 

14. Fisher, F.A,: Geometric Model Studies of KSC Launch Facilities. 

General Electric Company Corporate Research and Development 

Report SRD-76-077, 1976. 

15. Crouch, K.E.: Lightning Protection Evaluation of Launch Complex 36A Using Model Techniques. Lightning Technologies, Inc. 

Report LT-78-07, 1978. 

16. Little, P.F., Hanson, A.W. and Burrows, B.J.C.: Test Techniques 

for Simulating Lightning Strikes to Carbon (Graphite) Fibre 

Composite Structures. Proceedings of the 1979 Federal Aviation 

Administration-Florida Institute of Technology Workshop of 

Grounding and Lightning Technology. 

17. Kim, D.G., Dubro, G.A. and Tessler, L.P.: Transmission Line 

Theory Applied to Aircraft Lightning Interactions, Proceedings 

of 1977 IEEE International Symposium on Electromagnetic Compatibility, 1977. 

18. Little, P.F.: Transmission Line Representation of a Lightning 

Stroke. Journal of Applied Physics, Vol. 11, 1978, pp. 1893- 

1910. 

19. Strawe, D.F.: Non-Linear Modeling of Lightning Return-Strokes. 

Proceedings of the 1979 Federal Aviation AdministrationFlorida Institute of Technology Workshop on Ground and Lightning Technology. 

20. Fisher, F.A. and Plumer, J.A.: Lightning Protection of Aircraft. 

NASA Reference Publication 1008, 1977. 

21. Hess, R.F.: EMP Coupling Analysis Using the Frequency (Transfer Function) Method with the SCEPTRE Computer Program. IEEE 

Transactions on Electromagnetic Compatibility, 1975. 

22. Berger, K.A.: Novel Observations on Lightning Discharges: 

Results of Research on Mount San Salvatore, Franklin Institute, 

1971. 

99 
1. Report No. 

NASA CR-3329 

4. Title and Subtitle 

2. Government Accession No. 3. Recipient’s Catalog No. 

IMPROVED TEST METHODS FOR DETERMINING LIGHTNING-INDUCED 

VOLTAGES IN AIRCRAFT 

5. Report Date 

September 1980 

6. Performing Organization Code 

7. Author(s) 

K. E. Crouch and J. A. Plumer 

8. Performing Organization Report No. 

LT-80-59 

9. Performing Organization Name and Address 

Lightning Technologies, Inc. 

560 Hubbard Avenue 

Pittsfield, Massachusetts 01201 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, D.C. 20546 

15. Supplementary Notes 

10. Work Unit No. 

512-51-14 

11. Contract or Grant No. 

NAS4-2613 

13. Type of Report and Period Covered 

Contractor Report 

14. Sponsoring Agency Code 

H-1126 

NASA Teihnical Monitor: Wilton P. Lock, Dryden Flight Research Center 

Final Report 

16. Abstract 

Tests performed on aircraft at reduced lightning current 

levels to measure the magnitude of voltages induced on the electrical cables have used standard lightning current generators 

which produce double exponential current waves. A lumped parameter transmission line with a surge impedance (Z ) 

that of the aircraft and its return lines was P 

matching 

eva uated as a 

replacement for the earlier current generators. Various test 

circuit parameters were evaluated using a l/10 scale relative 

geometric model. Induced voltage response was evaluated by 

taking measurements on the NASA-Dryden Digital-Fly-by-Wire (DFBW) 

F-8 aircraft. Return conductor arrangements as well as other 

circuit changes were also evaluated, with all induced voltage 

measurements being made on the same circuit for comparison purposes. The lumped parameter transmission line generates a 

concave front current wave with the peak dijdt near the peak of 

the current wave which is more representative of lightning. 

However, the induced voltage measurements when scaled by appropriate scale factors (peak current or di/dt> resulting from both 

techniques yield comparable results. 

7. Key Words (Suggested by Author(s)) 16. Distribution Statement 

Lightning 

Transients 

Induced voltages 

Test techniques 

Unclass .fied - Unlimited 

Subject category 05 

I 

9. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price’ 

Unclassified Unclassified 107 $9.00 

*For sale by the National Technical Information Service, Springfield, VA 22161 

NASA-Lang1 ey, 1980