NASA Contractor Report 174919

Analysis and Tests for Design of an

Electro-Impulse De-Icing System

-CR-174919) ANALYSES AND TESTS FOE N86-27268

DESIGN OF AN ELECIB0-IMPULSE DE-ICING SYSTEM

Interim Beport (Wichita State Univ.) 198 p

HC A09/MF A01 CSCL 01C Unclas

: G3/03 43300

G.W. Zumwalt, R.L Schrag, W.D. Bernhart,

and R.A. Friedberg

Wichita State University

Wichita, Kansas

May 1985

Prepared for

Lewis Research Center

Under Grant NAG 3-284

NASA

National Aeronautics and

Space Administration

A°
TABLE OF CONTENTS

CHAPTER 1 PAGE

THE EIDI DEVELOPMENT PROJECT

I. Introduction 1

II. The Basic Principles 1

III. Prior History 6

IV. Current NASA/WSU/Industry Project 7

V. References . . 9

CHAPTER 2

ELECTRODYNAMIC STUDIES AND TESTS

I. Electro-Impulse Circuit Characteristics

and Design Aspects 11

II. Ballistics Pendulum Studies 23

III. Magnetic Field Diagnostics Experlement 39

IV. Electrodynamics Modeling 60

V. References 75

CHAPTER 3

STRUCTURAL DYNAMIC TESTS AND STUDIES

I. Indroductlon 77

II. Strain Measurements Near EIDI Colls . . . . 81

III. Boeing (BCAC) 767 Leading Edge Slat Strain 84

IV. 2.5 Inch Radius Semi-Cylinder Studies 87

V. Recent Developments Related to EIDI Force Measurements ... 94

VI. References ........ 96

VII. Figures 3-1 to 3-32 97

CHAPTER 4

FABRICATION TECHNIQUES

I. Coll Wire Rolling 129

II. Coll Making 129

III. Coll Mounts 136

IV. Electrical Leads 140

V. Weight Estimates . . . i 140

VI. Con Attachment 141

VII. Icing Tunnel Models . . . 141

VIII. Test Model Instrumentation 143

CHAPTER 5

ICING TUNNEL TESTS

I. Octs. 25 - Nov. 5, 1982 Tests 147

II. April 18-22, 1983 Tests 149

III. August 15-18, 1983 Tests 153

IV. November 7-9, 1983 Tests ,' 157

V. May 14-23, 1984 Tests 161

VI. August 13-17, 1984 Tests 164

VII. Sept. 17-21, 1984 Tests 166

VIII. Nov. 26 - Dec. 4, 1984 Tests 172

IX. References / 175
TABLE OF CONTENTS

(Continued)

CHAPTER 6

FLIGHT TESTS

I. NASA Icing Research Aircraft 177

II. Cessna TV 206 Flights (1984) 180

III. References 187

CHAPTER 7

PLANS AND PROSPECTS . . . . . . 189

CHAPTER 8

APPENDICES

WSU Participants on EIDI Project ..... 191

Industrial Participants . . . . 192

Prospects and Problems for Applying

EIDI to Helicopter Blades 193

ii
CHAPTER 1. THE EIDI DEVELOPMENT PROJECT

I. Introduction

Ice accumulation on aircraft wings In flight has been a danger since the

earliest days of flight. The total accumulation needs not be large to be

fatal. Although 1ce normally Is accrued on frontward facing surfaces only,

giving a few centimeters thickness on the front 2 percent of the wing chord,

this Is enough to cause flow separation and destroy 11ft, particularly 1f the

aircraft slows or maneuvers. Also, drag may Increase enough to exhaust fuel

reserves or destabilize the flight. Helicopter rotors are even more vulnerable to the detrimental effects of 1ce, and engine Inlet dlffusers require

1c1ng protection to a greater degree than lifting surfaces. This need brought

the NASA Lewis Research Center Into aircraft 1c1ng research almost from Its

start.

Although several methods of de-1c1ng or ant1-1dng are available, all have

some undesirable aspects 1n regard to energy requirements. A method suggested

(Ref. 1-1) as early as 1937 offers an alternative which has not been

adequately developed. The electro-magnetic Impulse phenomenon has been used

for metals forming, and holds the promise of 1ce removing with very low

energy, minimal maintenance (no moving parts), great reliability, and weight

and cost competitive with existing methods.

This report summarizes work done under a NASA-Lewis grant to develop the

Electro-Impulse De-Ic1ng (EIDI) system. Wichita State University has been the

grant recipient, charged with researching the phenomena to provide the underlying technology, and also with coordinating the efforts of a team of participating Industries. The program has consisted of basic analyses, laboratory

tests, 1c1ng tunnel tests and flight tests. At the two-and-a-half year point,

the EIDI system has been tested and refined and shown to be a low-energy,

highly reliable de-1c1ng method. This report 1s a status report and an outline of design methods developed to this date.

II. The Basic Principles

The physical form of the Electro-Impulse De-Ic1ng (EIDI) method 1s shown

1n Figure 1-1. Flat-wound colls made of copper ribbon wire are placed just

Inside the leading edge of a wing's skin with a small gap separating skin and

coll. Either one or two colls are placed at a given spanwlse station, depending on the size of the leading edge. Two methods of supporting colls are

shown; support by the front spar or from a beam attached to ribs 1s generally

used, but mounting to the skin Itself 1s sometimes used.

The colls are connected by low resistance, low Inductance cables to a high

voltage capacitor bank, and energy 1s discharged through the coll by a
remote signal to a silicon-controlIed-rect1f1er (M

thyr1storn). Discharge of

the capacitor through the colls creates a rapidly forming and collapsing

electro-magnetic field which Induces eddy currents 1n the metal skin. The

fields resulting from current flow 1n the coll and skin create a repulsive

force of several hundred pounds magnitude, but a duration only a fraction of a

millisecond. A small amplitude* high acceleration movement of the skin acts

to shatter, debond and expel the 1ce. Two or three such "hits" are performed

sequentially, separated by the time required to recharge the capacitors, then

1ce 1s permitted to accumulate until It again approaches an undesirable thickness.

COIL

ALUMINUM

DOUBLER

COIL

TO CAPACITOR

TO SWITCHING

UNIT

SILICON

CONTROLLED

RECTIFIER

Fig. 1-1 Impulse Coils in a

Leading Edge

Figure 1-1 also shows "doublers," unalloyed aluminum discs, slightly

larger than the colls, bonded to the skin opposite the coll. These are used

when the skin thickness 1s less than the minimum required to provide

adequate conductance for the eddy currents. Composite, non-metallic,

leading edges require a similar special treatment. A fundamental study of

the phenomena and parameters for electro-Impulse was undertaken to provide a

basis for such geometric and electrical design choices. Prof. R. L. Schrag

of Wichita State University has led this study which 1s presented 1n Chapter

2. Ref. 1-2 reported Initial results of this work.
TBIOOER

ENERGY

STORAGE :

CAPACITOR

THYRISTOR

OIOOE

CLAMP

IUPUL8E

COIL AIRCRAFT

SKIN

Fig. 1-2 Basic Circuit

In Figure 1-2, the basic circuit 1s Illustrated. An electro-Impulse 1s

Initiated by supplying a trigger pulse to the thyrlstor, allowing the capacitor to discharge through the coll. A typical current waveform 1s shown in

Figure 1-3. Since a thyrlstor has diode properties, the current follows the

first positive loop of the RLC response, after which the thyrlstor re-opens

the circuit. This leaves the capacitor reverse-charged. Such reverse charging reduces capacitor life substantially. For that reason a clamping diode 1s

placed across the capacitor. A typical current and resulting skin displacement are shown 1n Figure 1-3.

a

u

.002

U

<

240

TIME. \1 SEC

SOO

Fig. 1-3 Typical Coil Current and

Skin Displacement
Figure 1-4 1s a flat geometry Illustration of the coil's magnetic field

and Induced eddy currents 1n the electrically conducting skin. The magnetic

field due to the eddy currents 1s not shown, but 1t has a significant

Influence (by self Induction) on the magnitude, time history, and radial

distribution of the eddy currents. In addition, the electromagnetic "skin

effect" phenomenon affects the eddy current distribution across the aluminum

skin thickness. Current densities are greatest on the coll side. A reverse

coupling effect 1s also present. Time-changing eddy currents Induce a voltage

1n the Impulse con, modifying Its current. From a circuit aspect, the consequence 1s a modification of the effective Inductance and resistance of the

coll. The effective Inductance decreases/ and the effective resistance Increases, due to the proximity of the metal sheet.

MAGNETIC

HEtD ^— EDDY

CURRENTS

SKIN

COIL

Fig. 1-4 Coil Magnetic Field Pattern and

Resulting Eddy Currents

When the aircraft skin moves 1n response to the electro-Impulse force the

coll-to-skln gap changes and that modifies the magnitude of the proximity

Influence. In addition, the skin's movement relative to the coil's magnetic

field further modifies (by motional Induction) the electromotive forces that

drive the eddy currents. These Influences due to skin motion are, however

relatively small because of the time delay Involved 1n the motion, and appear

to be negligible when the skin 1s 1ce loaded. The assertion 1s, 1n effect,

that the coll current and the strength of the force Impulse may be calculated

without the need to also analyze the complex structural response. This 1s

discussed In detail In Chapter 2.

Figure 1-5 shows a wing with colls placed spanwlse, separated by about

0.4 meters. These are all supplied by a single power unit. A more effective

modified version 1s shown 1n F1g. 1-6. Energy requirements are small, being

comparable to those typical of landing lights for the same size aircraft. Deicing has been accomplished 1n the 1c1ng wind tunnel and 1n flight for typical

general aviation wings under a wide range of velocities, angles of attack,

1c1ng rates and temperatures.

Just as effective de-1dng by electro-Impulse requires a matching of

electric circuit dynamics and skin electrical properties, so also does 1t
IfjAHSFORjgR

CAPACITOR

PROGRAMED

SWITCHING

Fig. 1-5 Electro-Impulse Coils

Installed in a Wing

TRANSFORMER

CAPACITOR

PROGRAMED

SHIRKING

Fig. 1-6 E.I.D.I. with Series-Connected Coils

("Odds and Evens")

demand properly related structural dynamic matching. The propagation of

skin movement 1n chordwlse and spanwlse directions from the EIDI coll 1s

necessary. It 1s also a complex* three-dimensional* transient structural

dynamic analysis problem; Ref. 1-3. Coll location and pulse duration must

have the proper relations to the structural vlbratlonal mode to be excited.

Failure to do so severely Impairs the effectiveness of de-1c1ng.

The Imposed force cannot be modeled as a simple Impulse* since Its

duration Is comparable to the period of the first vlbratlonal mode.

Similarly* the force cannot be treated as a concentrated "point" load, since

the width of the coll 1s of the same order as the wave lengths of low-order

vlbratlonal modes. Thus* structure-dynamic analysis must Include the

details of the spaclal and temporal results of the electro-dynamic

analysis.
Under the leadership of Professor Walter D. Bernhart, analytical and

experimental studies have been carried out on the structural dynamics of

leading edge structures under EIDI-type Impulsive loading. These are

reported 1n Chapter 3.

The participation of Industry 1n the project provided constant reminder

of the practical aspects of safety* weight* cost and manufacturablHty.

While research and testing were carried on, parallel efforts were made to

develop fabricating and design method for the hardware. In particular* coll

making and mounting were developed under the leadership of Robert FMedberg.

The current state of materials and processes for these* together with

testing methods* are given 1n Chapter 4.

Icing tunnel tests (eight sets at the time of this report writing) have

been essential to the system development. The unique capabilities of the

NASA-Lewis Icing Research Tunnel (IRT) provide testing conditions for confident evaluation of the system 1n flight. Chapter 5 summarizes the IRT tests

and results. Reference 1-4 reports 1n more detail the engine nacelle and

large wing testing.

Chapter 6 gives a brief account of the two flight test programs

completed to date. These were earlier reported 1n References 1-5 and 1-6.

A third 1s 1n progress at the time of writing this Interim report. Finally*

plans and prospects for the project are delineated 1n Chapter 7.

III. Prior History

The use of electro-magnetic Impulse force to remove 1ce was first

suggested by Rudolf Goldschmldt* a German national residing 1n London before

World War II. He was granted a patent (Ref. 1-1) and a series of patent

extensions 1n 1937 through 1939. His patents* now expired* anticipated most

of the applications now being used or considered. However* there 1s no

evidence that Goldschmldt ever attempted to build the devices he Imagined.

During the 1950's and 1960's, electro-Impulse methods were used for

metals forming 1n various Industrial processes* but no record can be found of

the use for de-1c1ng until researchers 1n the USSR either discovered

Goldschmldt's patents or rediscovered this application Independently. In

1965* I.A. Levin 1n the Soviet Ministry of Power and Electrification* seeking

methods for cleaning frozen and sticky materials from surfaces (coal bunkers*

transformer boxes* towers* etc..) published work on electro-Impulse

possibilities. He Immediately received Inquiries from other ministries

(aviation* fisheries, dairy* housing) regarding de-1c1ng of vehicles, buildings* ships* and for cleaning of dry milk from hoppers (Ref. 1-7 and 1-8).

Responding to their requests Involved him 1n bureaucratic territorial

struggles and he was fired. Eventually* however, he was set up 1n his own
laboratory under the State Committee for Meteorology and Environmental

Monitoring, and apparently made some Installations 1n aircraft; the 11-18 has

been cited as having been the first* but confirmation by Western observers 1s

difficult to obtain.

In any case* 1n the early 1970's* Russian representatives were granted

EIOI patents 1n several Western nations* Including the USA* and USSR salesmen

began calling on American and European aircraft companies offering to sell

their design and construction services for an EIDI system. Their lack of

candor discouraged most customers, but Interest was stirred and during the

1970's* work was done on this method 1n France (A1r-Equ1pment division of

DBA), Great Britain (Lucas Aerospace and B.A.C.), and the United States

(Lockheed and McDonnell-Douglas). For various reasons* the development

stopped short of full Implementation. The system stm lacked a well

developed underlying technology and known design parameters.

IV. Current NASA/WSU/Industry Project

A. Initial Feasibility Demonstration

In mid-1982* NASA Lewis Research Center funded a six-month grant to

Wichita State University to work with two small plane makers* Beech and

Cessna* and an aircraft electrical system manufacturer* the Engine Systems

Division of S1mmonds-Prec1s1on» to do a feasibility study resulting 1n an

1c1ng tunnel demonstration 1n Oct.-Nov. 1982. Two wing sections were

tested* m1d-w1ng portions from a Beech Bonanza and a Cessna 206. These

were similar 1n size and flight speeds* but had very different leading edge

structures* one very stiff and small* the other flexible and large 1n extension from the front spar. Good cleaning was accomplished with fairly low

energy expenditure for air speeds from 96 to 230 knots* air temperatures

from 29° to -15°F (-2° to -26°C), angles of attack from 2 to 8 degrees and

liquid moisture contents from 0.6 to 2.4 g/m.

B. The Industry Consortium

The results were encouraging enough to lead NASA to expand the aim to

full development of the method for the whole range of civil aircraft. A

consortium of participating Industries was formed for this purpose. Each

company agreed to contribute some services or equipment to the project* and

1n return became eligible to submit Its own products for de-1c1ng design and

tests by the EIDI method. WSU was charged with doing the needed research

Into the electrodynamlc and structural dynamic phenomena Involved* and

developing manufacturing methods for colls and their mountings. In

addition* WSU coordinated the effort and conducted further tests 1n the

7 .
NASA/Lewis Icing Research Tunnel (IRT). The Industries represented a wide

scope of aircraft sizes* air speeds* skin thicknesses and of attitudes

regarding the Introduction of a new device.

Realizing the ultimate requirement for certification, FAA personnel

were Invited Into the group meeting so that their concerns could be

considered as early as possible.

Consortium members for the 1982-83 period were:

Small Aircraft: Beech Aircraft Co., Wichita, KS

Cessna-Pawnee D1v., Wichita, KS

Business Jet Aircraft: Gates Learjet Corporation,

Wichita, KS

Cessna-Wallace D1v., Wichita, KS

Composite, High Performance: LearFan Ltd., Reno, NV

Transport Aircraft: Boeing Commercial, Seattle, WA

McDonnell-Douglas, Co., Long Beach, CA

Electrical Equipment: S1mmonds-Prec1s1on, Norwich, NY

In late 1983, Rohr Industries of Chula Vista, CA, joined the group with

an Interest 1n de-1c1ng engine Inlets. In 1984, a second aircraft

electrical equipment developer was added, Electro-Delta of White Oak, Texas,

and the first helicopter maker, Sikorsky Aircraft Co., Stratford, CT. also

Joined. At the time of preparation of this report, Bell Helicopter of Ft.

Worth, TX, and the Boelng/Vertol Co. of Philadelphia, PA, have requested

membership.

C. Objectives

The program objectives were, from the start, quite comprehensive:

1. Develop computer models for the structural dynamics of

leading edge portions of wings and engine Inlets to

provide design guidance for coll location, coll size.

Impulse Intervals and coll spanwlse spacing. An alternative approach was also desired, namely the development

of a standard measurement method for existing structures

to extract the structural dynamics characteristics

needed for the design.

2. Develop a computer model for the electro-dynamics and

provide detailed design data for the electro-Impulse

equipment, Including coll design, power, voltage, 1nsultlon, pulse duration, and switching equipment.

8
3. Test wing sections and engine Inlets 1n the IRT to guide

and prove EIDI designs.

4. Consider practical aspects of retro-fitting the EIDI

system to existing aircraft.

5. Devise methods for optimal design of a wing structure

for using the system.

6. Estimate and attempt to minimize the cost of the EIDI

system 1n terms of weight* maintenance and capital

outlay.

7. Details to be considered are:

(a) Limits of application (size, stiffness, etc.)

(b) Standardization of components

(c) Fatigue of skin, mountings, switching gear, bondings

(d) Electro-magnetic interference problems and solutions

(e) Use with composite materials

(f) Integration with present avionics and electrical

systems

8. Conduct flight demonstrations using aircraft from NASA

and participating Industries.

9. Carry out at least the first stages of FAA certification.

V. References

1-1. British Patent Specification No. 505,433 Issued to Rudolf Goldschmldt;

May 5, 1939.

1-2. R.L. Schrag and G.W. Zumwalt, "Electro-Impulse De-1c1ng: Concept and

Electrodynamlc Studies," AIAA 22nd Aerospace Sciences Meeting, Reno,

NV, Jan. 9-12, 1984, Paper No. 84-0021.

1-3, W.D. Bernhart and G.W. Zumwalt, "Electro-Impulse De-1c1ng: Structural

Dynamic Studies, Icing Tunnel Tests and Applications," AIAA 22nd

Aerospace Sciences Meeting, Reno, NV, Jan.'9-12, 1984, AIAA Paper No.

84-0022.

1-4. G.H. Zumwalt, "Icing Tunnel Tests of Electro-Impulse De-Ic1ng of an

Engine Inlet and High-Speed Wings,"AIAA 23rd Aerospace Sciences

Meeting, Reno, Nevada, January 14-17, 1985. AIAA Paper No. 85-0466.
V. References

(continued)

1-5. G.W. Zumwalt and A.A. Mueller. "Fl 1ghts and Wind Tunnel Tests of an

Electro-Impulse De-Icing System", AIAA/NASA General Aviation Technology

Conference, Hampton, VA, July 10-12, 1984. AIAA Paper No. 84-2234.

1-6. Mueller, A.A., D.R. Ellis and D.C. Bassett, "Flight Evaluation of an

Electro-Impulse De-1c1ng System on a Light General Aviation Airplane."

AIAA/AHS Aircraft Designs, Systems and Operations Meeting, San Diego,

CA., October 1984. AIAA Paper No. 84-2495.

1-7. "From the Sky to the Earth," Pravda, (unsigned), Dec. 10, 1978.

1-8. "Development of De-1c1ng System," Pravada, (unsigned), Feb. 11, 1980.

10
CHAPTER 2

ELECTRODYNAMIC STUDIES AND TESTS

I. Electro-Impulse Circuit

Characteristics

and Design Aspects

A. Theoretical Low Frequency Inductance and Resistance for Isolated

Pancake Shaped Coils

Procedures can be found in the literature for calculating the inductance

of flat pancake shaped coils wound spirally with rectangular wire. The

procedure given in Reference 2-1 was used to calculate a table of values

(Table 2-1) for coils having a fixed thickness of .188 inch and a fixed

inner diameter of 0.25 inch. Insulation thickness of .003 was assumed. The

table also lists theoretical d.c. coil resistance, assuming annealed copper,

and the ribbon wire thickness. It is particularly useful for design purposes,

once the outer diameter and inductance are selected.

One comparison was made between calculated inductance and a bridge

measurement. The coil had 30 turns and a 2 inch diameter. The bridge

measured 19.8 yH, compared with a calculated 18.9 yH. The difference may

have been due to about 18 inches of connecting leads.

11
TABLE 2-1

COIL DIAMETER, INCHES

TURNS

20

25

30

35

40

45

50

55

1.5

L

6.5

11.3

18.8

30.6

R

7.0

11.2

16.6

23.2

b

28.4

22.1

17.9

14.9

2.0

L

12.6

18.9

27.3

38.5

R

9.9

14.6

20.2

26.9

b

32.1

26.3

22.1

19.0

2.5

L

21.9

30.4

40.9

53.7

69.5

R

13.5

18.7

24.8

31.8

39.9

b

34.6

29.2

25.2

22.1

19.6

L = Inductance in yH

R = D.C. resistance in nfl

(Assumes annealed copper)

b = Ribbon wire thickness in Mils

Fixed Dimensions:

Ribbon width 0.188 inch

Insulation thickness 0.003 inch

Inner coil diameter 0.25 inch

3.0

L

35.0

46.3

59.7

75.1

93. 4'

R

17.7

23.4

30.0

37.5

46.0

b

36.4

31.4

27.6

24.6

22.0

12
B. Influence of Frequency on L and R QR1Q5MAL PA©1 SS

"DF POOR QUALITY

Bridge measurements were made at various frequencies on the 30-turn,

2 inch diameter coil. The inductance was found to be practically independent of frequency, decreasing only 5 percent as the frequency was scanned

from 500 Hz to 20 kHz.

The resistance exhibited a much greater frequency effect. Figure 2-1

shows a plot of the ratio of a.c. resistance to d.c. resistance vs.

frequency. The d.c. resistance value was measured by the ammeter-voltmeter

method, and exceeded the value given in Table 2-1 by 40 percent. This is

probably due to the fact that the wire was not annealed subsequent to a

rolling operating. The increase in R with frequency suggests the presence

of skin effect.

6-

5-

4-

3-

dc

2-

1-

0.5

FREQUENCY, KHz

Fig. 2-1 Coil Resistance vs Frequency

13
Assuming a metal plate of thickness "b" with symmetrical excitation

on both surfaces, one can derive the following formula for the resistance

ratio:

Ra „ b

RT^ = Re Kd.c.

where T = (1 + j)/6

and 6 is the skin depth given by

6 = /2/u>ya

In order to explain the experimental data, the "b" dimension had to be

taken as the width (not thickness) of the ribbon wire. With that

assumption, the calculated results agree well with the measurements

(see Figure 2-1) up to about 6 kHz. Beyond that frequency there

appears to be some added complication to the theoretical model.

It is concluded that the primary frequency affect on resistance

in the audio frequency range is due to energy diffusion into the ribbon

wire via the coil surfaces. Diffusion into the ribbon via its narrow

dimension should occur at a much higher frequency and may, in part, be

involved in the added complication referred to above.

C. Proximity Effects of Metal Plates

In designing a coil for electro-impulse de-icing, one should

recognize the fact that the apparent inductance and resistance of the

coil will be modified by the close proximity of the aluminum skin. A

suitable theoretical model would, presumably, account for this modification.

Limited bridge measurements were made in order to get a feel for the

magnitudes of the proximity effects. The coil used in the tests was the

30-turn, 2 inch diameter coil. Two "target" configurations were tried.

14
The first was a .032 inch 2024-T3 aluminum plate with a .05 inch gap

(approximately .078 inch between the plate and the copper surface of the

coil). For the second test a .021 inch copper "doubler" was added onto

the coil side of the .032 inch aluminum plate, still maintaining a .050

inch gap from coil-to-doubler. The experimental results are given in

Tables 2-2 and 2-3.

TABLE 2-2

INFLUENCE OF .032" Al PLATE

WITH .05" GAP

Frequency

1 kHz

2 kHz

4 kHz

Inductance

Factor

.85

.75

.65

Resistance

Factor

1.8

2.6

2.9

These are ratios of values with the plate present to

values without the plate (at the same frequency).

TABLE 2-3

Frequency

INFLUENCE OF .032" Al PLATE +

.021" Cu DOUBLER WITH .05" GAP

Inductance

Factor

Resistance

Factor

1 kHz

2 kHz

4 kHz

.68

.61

.58

1.8

1.9

1.8

Note that the addition of the doubler causes a greater

inductance reduction , but less resistance increase at

the higher frequencies.

15
D. Theoretical Waveforms for Current and Capacitor Voltage Assuming

an Ideal R-L-C Circuit

In spite of the fact that the coil parameters (especially resistance)

are functions of frequency, the experimental waveforms for current and

capacitor voltage resemble those for an ideal under-damped R-L-C circuit.

It would be useful to summarize the properties of the ideal waveforms,

sketched in Figure 2-2. These waveforms assume no clamping diode across

the capacitor.

•O

Fig. 2-2 Approximate Waveforms

for Coil Current and Capacitor

Voltage without a Diode Clamp

16
If we define a damping parameter by

1

m = ; (2-1)

4L _1

R

2C

then the period of damped" oscillation is

T = 2rr/LC"/m2+~l (2-2)

The time for the maximum current can be expressed in alternate forms

tm = /EC ,/m^TT tan ^(I)

(2-3)

If the Initial stored energy is W, then the maximum current is

- tan'1,!, (2.4)

Without the clamping diode, the capacitor will become reverse-charged

with the final voltage

-irm

vcf

 = -

Ve (2-5)

where V is the initial capacitor voltage. Finally, the maximum stored

inductor energy and the final re-stored capacitor energy are, respectively,

f. i* = W e'

£lluaM W (2-6)

and

¥

 vcf

 = W e "*"" (2-7)

17
A reliable technique for determining the effective circuit resistance

and inductance from the current and voltage waveforms is the measure T/2,

i'm, and vcf. (C and V are assumed to be known). Then the following sequence

of calculations can be made:

. - 1,. (-JL,

-eff

2 -2mtan"1(i)m ce

m

(2-8)

(2-9)

2irLm

Vf =

E. Circuit Parameters of the Energy System

Additional bridge measurements were made to determine the circuit

parameters contributed by the Simmonds Energy system. The following

Table lists the breakdown of parameter distributions for the discharge

circuit assuming the 30-turn, 2 inch diameter impulse coil (without a

target).

30-Turn

2 Inch Dia.

Coil

600uF

Capacitor

22 ft

Cable

Totals

Capacitance

(vF)

600.0

600.0

Inductance

(uH)

19.8

1.7

21.5

Resistance

(n)

.0273

(at 1 kHz)

.003

(at 1 kHz)

.043

.0733

18
Discharge waveform data for this same circuit were determined to be:

V =

vcf =

•v

T/2 =

500 volts

-260 volts

2025 amps

368 usec

Then, following the calculations outlined in Sec. D, one obtains,

m = 0.2083

Leff = 20.7 yh

Reff = 0.0736 n

These inductance and resistance values are in substantial agreement with

the total values deduced from bridge measurements.

F. Example of Coil Design

An elementary coil design prodcedure will be described assuming prior

determination (or estimation) of the following:

(1) Initial capacitor energy required

(estimate from de-icing test experience)

(2) Natural period of the surface mode

to be excited

(3) Initial capacitor voltage

(4) Cable resistance

(5) Cable inductance

PI.

(6) Capacitance, C = =£

r

(7) Coil Thickness

(8) Coil inner diameter

(9) Coil outer diameter

Sample Design Value

W = 250 Joules

1.2 msec

V = 1000 volts

.03 n

negligible

500 yF

0.188 Inch

0.25 Inch

2.5 Inches

19
Items (7) through (9) look like they ought to belong to the coil

"design" activity. We do not, to date, have a sufficiently extensive

design procedure to yield optimum dimensions. The coil thickness and

inner diameter values for the sample design are those that were assumed

for the calculation of Table 2-1. They also are the thickness and inner

diameter values for nearly all the coils that were used in the de-icing

tests. The selection-of 2.5 inches for the outer coil diameter is somewhat arbitrary. Ballistic pendulum tests have indicated that larger

diameter coils (other parameters remaining the same) result in increased

delivered force impulse. Diameter optimization most probably will be

based on weight-performance tradeoffs.

The remaining coil design steps, illustrated with the sample values,

proceed as follows:

(a) Electric Circuit Period

In order to minimize peak skin stresses, the force impulse should be

made as "soft" as possible, consistent with its ability to excite mechanical

motion. The rule of thumb for yielding the desired pulse duration will be

to make the (undamped) electrical period equal to half the natural period

of the surface mode to be excited. See sample design:

T = •£- (1.2 ms) = 0.6ms

(b) Coil Inductance

The required values of coil inductance is fixed by T and C, according to

L,« = - '* . I ">UAAU ] = 18pH Sample Design: eff ~ 500xlo-6 [ 2ir

This is the effective inductance for the mounted coil. When unmounted,

the inductance will be greater (see Tables 2-2 and 2-3). Taking the skin

"proximity factor" to be 0.6, then the required coil inductance for the

sample design is

L = 1M = 3QyH

The computer modeling activities should yield more definite bases for

selecting optimum dimensions.

20
(c) Number of Turns and Wire Thickness

These design values are read from Table 2-1. To achieve 30 yH with

a 2.5 inch diameter coil, then

N = 35 Turns

b = .029 Inch (Ribbon Wire Thickness)

(d) Effective Coil Resistance

The d.c. coil resistance for the sample design is also indicated in

Table 2-1 as

R , r = .0187 n d.c.

The effective a.c. resistance is higher. Since the electrical frequency is

f = Y = ;i.67 kHz

Figure 2-1 indicates an increase factor of about 1.25. Thus

R „ = 1.25(.0187) = .023 n

a.C.

Finally, applying a correction factor to account for skin proximity

(Tables 2-2 and 2-3)

R

a.c.

= 1.8(.023) = .0414

eff

(e) Doubler Design

A conservative criterion for the doubler design is to pick a doubler

thickness that equals one-half the electrical skin depth at the circuit

frequency. If this answer is comparable to the aluminum skin thickness,

then a somewhat thinner doubler might be used.

Sample design: Suppose 1145 Al is used as doubler material. It

has about 62 percent conductivity so the recommended thickness is

A

Doub

2Ti( 1667) (4nxlO-7)(. 62x5.8xl07)

= 1.028 x 10"3 m = .040 Inch

21
(f) Estimate of Peak Current

The peak current can be calculated from Eq. (2-4)

1 = Me -

m tan "^fe

m v L

where

1 m =

v

 R'C

For the sample design,

R ckt= R coil + Rcable = -0414 + '°3 = -0714

L = 18 yH , C = 500 yF

W = 250 J

Thus

4L = ili|] = 28 25

R^C (.0714)2(500)

So

m = =L== = .1916

x/27.25

and the current modification due to losses is

m = .767

Finally

18xlO-

(g) Design Iteration

Results from previous steps need to be inspected to see if they present

any particular system design problems, and adjustments made accordingly. For

example, one may be committed to use a thyristor for which 4.04 KA peak

current is dangerously high. An increased inductance is then called for. If

that results in too long a pulse, then C can be reduced and the operating

voltage increased. 22
II. Ballistic Pendulum Studies

A ballistic pendulum was constructed and utilized to study the effects

of parameter variations on impulse production, A schematic drawing of the

apparatus is shown in Fig. 2-3. The maximum pendulum swing (6 ) was

max

measured by an electrical output from the precision potentiometer, and the

delivered impulse was determined from the formula:

. . Q

Impulse =_J . sin-^*- Ib-sec (2-10)

where J = Moment of inertia of the

pendulum = .387 lb-ft-sec2 (calculated)

1 = Pendulum length, pivot to target

center = 3.727 ft.

T = Free oscillation period = 2.09 sec

(measured)

Numerically, 4*07(11) = 0.624 Ib-sec.

This section describes the various test conditions and the test results.

In most cases a "dummy coil" was included in series with the "active coil"

to simulate double coil operation. When dummy coils were used, they were

always identical to the active coils and were fitted with identical targets

(except where noted otherwise).

23
y

*

X

PRECISION

POTENTIOMETER

CAPACITOR

O.C.

POWER

SUPPLY

\

SCR

>£,

••••••

"T"

COIL

\

ni f

/ 9

/ ^

V\

\\

\\

\\

\\ PENDULUM

\\

\\

"

9

 \

\^

\

*\^>

S \ \

TARGET ^

/ -M U- GAP

Fig. 2-3 Ballistic Pendulum Apparatus

24
A. Varying Voltage

TEST 1

ELECTRICAL

Volts Var.

yF 600

Joules Var.

ACTIVE

COIL

Turns 30

Dia. 2.0"

Thickness.188"

DUMMY

COIL

Used yes

With Target

no

TARGET

Mat. Al 2024-T3

Thickness .032"

Gap .050"

.03

.02

o

o

o

O)

to

I

jQ

.01 ..

500 600

VOLTS

700 800 900 1000

100 200

ENERGY, JOULES

300

Fig. 2-4 Impulse Per Unit Energy

with Voltage Varied

25
TEST 2

ELECTRICAL

Volts Var.

yF 400

JoulesVar.

ACTIVE

COIL

Turns 30

Dia. 2.0"

Thickness 95

. DUMMY

COIL

Used yes

With Target

yes

TARGET

Mat. Al 1100

Thickness .050"

Gap .050"

.08

.06

o

o

O

<u

10

^.04

.02

VOLTS

600 700 800 900

100

ENERGY, JOULES

1000

200

Fig. 2-5 Impulse Per Unit Energy

with Voltage Varied

26
Coil Current Data For Test 2

Volts

600

800

1000

1m(KA)

1.40

1.85

2.31

KA/KV

2.33

2.31

2.31

t>s)

163

163

163

Comments: This test was intended to verify that the impulse is proportional

to the square of the capacitor voltage. The experimental data show some

small deviation from a proportionality, both for a thin poor-conductivity

target and a thicker good-conductivity target. This deviation is relatively inconsequential. Impulse values can still be extrapolated with fair

accuracy from one voltage to another by assuming proportionality to V2.

B. Varying Capacitance, Fixed Energy

TEST 3

ELECTRICAL

Volts Var.

UF Var.

Joules 150

ACTIVE

COIL

Turns 30

Dia. 2.0"

Thickness. 188"

DUMMY

COIL

Used yes

With Target

no

Fig. 2-6 Impulse vs Capacitance

with Fixed Energy

0)

«J

»

iu

.04

.03

.02

.0 1

TARGET

Mat.Al 2024-T3

Thickness .032"

Gap .050"

866 V

707 V

FIXED 150 J

200 400

CAPACITANCE, Jj F

600

27
TEST 4

ELECTRICAL

Volts Var.

yF Var.

Joules 120

ACTIVE

COIL

DUMMY

COIL

Turns 40 Used yes

Dia. 2.28" With Target

Thickness .195" yes

TARGET

Mat. Al 1100

Thickness .050"

Gap .050"

.10

.08

.06

O)

I-.04

.02

VOLTS

FIXED 120 J

200

CAPACITANCE, yF

Fig. 2-7 Impulse vs Capacitance

with Fixed Energy

400

28
Coil Current Data For Test 4

100

200

400

Volts

1549

1095

775

im(KA)

2.18

2.00

1.79

tm(v

80

115

163

Comments: This test explored the relative effectiveness of different

capacitance - voltage combinations with fixed energy. Both tests (3

and 4) showed small advantages of high voltage - low capacitance combinations.

C. Varying Coil-to-Target Gap

TEST 5

ELECTRICAL

Volts 800

yF 600

Joules 192

ACTIVE

COIL

Turns 30

Dia. 2.0"

DUMMY

COIL

Used yes

With Target

Thickness .188" no

TARGET

Mat. A1 2024-T3

Thickness .032"

Gap Var.

See Figure 2-8 on following page.

29
.06

,04

Ul

CO

.02

TEST 5

30 TURN, 2 INCH DIA. COIL,

TWO COILS IN SERIES

.02 .04 .06 .08 . 10

GAP, IN

Fig. 2-8 Impulse vs Coil-to Target Gap

Comments: As expected, the delivered impulse decreased with increased

coil-to-target gap. This reduction was about 20 percent for the gap

increase from zero to 0.05 inch.

30
D. Varying Target Material and Thickness

TEST 6

ELECTRICAL

Volts 800

pF 600

Joules 192

ACTIVE

COIL

•

Turns 30

Dia. 2.0"

Thickness .188"

DUMMY

COIL

Used yes

With Target

no

TARGET

Mat. Var.

Thickness Var.

Gap .050"

. 12 -

o

>u

O .06

•

«

•j

«

Uj

a. .

» .04

.03 .04 .OS .06 .07 .08

THICKNESS, IN

Fig. 2-9. Impulse vs Target Material

and Thickness

31
Comments: These data show large gains in impulse performance when high

electrical conductivity targets are used. The published conductivities

in percent of annealed copper for the three aluminum types are:

1145 - 62%

6061-T6 - 43%

2024-T3 - 30%

Increasing the target thickness also increases the impulse, but with

diminishing returns. A rule of thumb criterion is that the desired thickness from the standpoint of impulse production is about one electrical

skin depth calculated from the formula

<5 =

where a is the electrical conductivity of the target in mhos per meter,

= 47T xlCT7, and f is the resonant frequency (sec"1) of the electrical

circuit.

E. Varying Coil Diameter and Turns

TEST 7

ELECTRICAL ACTIVE DUMMY TARGET

COIL COIL

Volts 600 Turns Var. Used no Mat. Cu

yF 600 Dia. Var. With Target Thickness .021"

Joules 108 Thickness .188" n/a Gap .050"

32
o

o

V)

I

.11

.10

.09

.08

.07

\

2.5t DCo/I 3

30 40

TURNS

Fig. 2-10 Impulse Per Unit Energy

with Two Coil Diameters

and Variable Turns

50

Coil Current Data

For Test 7

Coil

Dia. (In)

2.09

2.09

2.56

2.56

2.56

Turns

30

38

32

38

48

\

(KA)

3.58

2.83

3.18

2.78

2.14

'm

(us)

117

137

137

146

174

33
Comments: This test measured the relative effectiveness of five coil designs

for which the number of turns and the outer diameters were varied. A greater

number of turns for a given coil diameter was achieved by rolling the ribbon

wire thinner.

No dummy coils were used in the impulse tests. Voltage and capacitance

were held fixed, so the energy was also fixed. In general, the data showed

that given diameter coils produced more impulse as turns were increased, and

that coils with a given number of turns produced more impulse as the diameter

was increased.

ELECTRICAL

Joules Var.

ACTIVE

COIL

TEST 8

DUMMY

COIL

TARGET

Turns 40 Used yes Mat. Al 1100

Dia. 2.28" With Target Thickness .050"

Thickness Var. yes Gap .050"

Current Data at V = 800 Volts

Coil Thickness

(In)

.195

.150

.100

UKA) m

1.85

1.79

1.66

t (y: m

163

158

146

Comments: For this test, six identical 40-turn coils were constructed.

Each had a thickness of 0.195" (the ribbon wire width after the rolling

operation). Then one pair was milled to a reduced thickness of 0.150",

and another pair to 0.100". Finally, the various pairs (one active, the

other dummy) were installed into the ballistic pendulum apparatus for

tests.

34
.07

o

o

o

I

JO

IS)

.06

O

COIL

A

o

0.150

0.100

•tt600 700 800

VOLTS

900 1000

Fig. 2-11 Impulse Per Unit Energy with

Three Coil Thicknesses

The results show no detectable degradation in performance for the 0.150"

thick coil, but an approximately 7 percent impulse reduction for the 0.100"

thick coil (relative to the 0.195" thick coil). The performance loss for

the thinnest coil was, presumably, due to its excessive resistance. Data

on resistance values were as follows:

0.195" thick coils

0.150" thick coils

0.100" thick coils

Supply cable

d.c. ohms

.035 each

.045 each

.068 each

.043

35
G. Series vs Parallel Coils

ELECTRICAL

TEST 9

ACTIVE

COIL

DUMMY

COIL

TARGET

Volts Var.

yF

Turns 40 Used yes Mat. A1 1100

Var. Dia. 2.28" With Target Thickness .050"

Joules 60/Coil Thickness .195" yes Gap .050"

.08

o

a)

t/0

Q.

S .07

.06

50 100 150

TIME TO CURRENT PEAK, ys

Fig. 2-12 Impulse vs Time to Current Peak,

with Series and Parallel Coil Operation

Compared

36
.(TEST 9 - continued)

Peak Currents For t = 80ys

KA

Series Coils (C.= 100 yF) 1.55

Single Coil (C = 200 yF) 1.49

Parallel Coils (C = 400 yF) 1.34/coil

Comments: This test utilized the two 0.195" thick coils from Test 8. The

object of the experiment was to compare impulse production with coils

(active and dummy) placed in series, to that with the coils in parallel, to

that for single operation. Various capacitance values were used, with

voltage adjusted for fixed 60 Joules/coil.

Relative to single coil operation, the series combination showed improved performance while the parallel combination showed substantially

worse performance. If the coil currents were identical (same peak values

and same time to peak) then the impulse values should be the same. The .

Table above, which lists peak currents with fixed risetimes,- shows the

greatest current for coils in series and least current for coils in parallel

Furthermore the impulse ratios agree approximately with the ratios of the

peak currents squared.

It appears that the explanation for the relative impulse performance

lies in the varying degrees of circuit damping. Without resistance in the

supply cable, then the damping factors should be the same in all cases.

However the (constant) supply cable resistance causes the damping to be

greatest with parallel coil operation.

H. Effect of the Diode Clamp

Some impulse measurements were made with and without the (reversecharge prevention) diode in place. The conditions and results were as

follows:

TEST 10

ELECTRICAL ACTIVE TARGET

COIL

Volts Var. Turns 30

yF Var. Dia. 2.0"

DUMMY

Used no

With Target

Mat. Cu

Thickness .021"

Joules Var. Thickness .188" Gap .050"

37
RESULTS

(TEST 10 - continued)

Impulse with Diode

C(yF) V (Volts) Impulse w/o Diode

600 500 .92

400 600 .92

200 800 1.00

Comment: The diode introduces a slight performance degradation, but it is

a small penalty to pay for increased capacitor life.

38
III. Magnetic Field Diagnostics

Experiment

This section describes a magnetic field measurement method for studying

the distribution of eddy currents within, and pressure distribution on, a

flat metallic plate subjected to an axially symmetric transient magnetic

field produced by an pulsed circular coil. The pulsing system, coil and

target ( including coil-to-target gap) closely simulated conditions that

existed during electro-impulse-de-icing tests conducted by Wichita State

University. For this experiment, the target was rigidly supported, and reduced voltage was used (energy about 15% of that required for de-icing).

A. Experimental Apparatus

Figure 2-13 shows the energy discharge system, omitting the capacitor

charging circuit and the thyristor firing circuit. Two identical pulsing

coils were operated in series, because that was the arrangement used in

most of the de-icing tests. However, only one of the two coils were

located in the coil-target assembly which is detailed in Fig. 2-14. The

effective gap between the coil (copper) surface and the near surface of

the target was .078 inch.

39
Fig. 2-13 Energy Discharge

Circuit

CdlL

COIL

.001

MOW/TOR

.OZB

POTTIK/&

POLYESTER

RESJN

vJOOPEW 5UPP6RT

\

'/////'///////////////s&//////////////////////////.'_jJi f</> it it,'//Hii/?(ii Hi/ii'in i J //////////

\

millilltlliiiiiNiililllll'illllllil!

COIL

MOU/VT

COIL DATA:

30 TURNS OF .024" x .

.188" RECTANGULAR COPPER

WIRE, SINGLE LAYER

SPIRALLY WOUND

*. \ \ \ \ \ \ \ \ \ \ \

SUPPORT

2-14 Details of the Coil-Target

Assembly

40
Each coil consisted of 30 turns of .024" x .188" rectangular copper

wire spirally wound in a single layer from an inner radius of .125" to

an outer radius of 1.00".

The initial capacitor voltage utilized for this experimental study

was 400 volts, and the resulting discharge current waveforms are shown

in Fig. 2-15. One trace is with the aluminum target removed, the other

is with the target in place.

1.6

1.2

oo

CL

LU

Ql

.4

WITHOUT TARGET

WITH TARGPT

A .6

TIME, MILLI-SEC

.8 1.0

Fig. 2-15 Discharge Current With and Without

the Aluminum Target Present

41
B. Magnetic Field Measurements

A field measuring plate was constructed in a manner illustrated in

Fig. 2-16. Shallow concentric grooves were cut into both sides of a

.050" thick phenolic disc with radius .increments of 0.2", starting at

r = 0.2" and ending at r = 2.0". Single-turn loops of .006" diameter

wire were then cemented into these grooves, and their twisted leads

brought out to solder tabs through radial channels. Figure 2-17 is a

photograph of the finished plate.

PHENOLIC

DISC

0.05 IN

THICK

TWO OF TEN

GROOVES

WITH WIRE

LOOPS

CHANNELS

WITH

TWISTED

LEADS

Fig. 2-16 Partial Illustration of the

Field Measuring Plate

42
ORIGINAL PAGE *!S'

OF POOR QUALITY

Fig. 2-17 Photograph of the Field

Measuring Plate

For measuring the fields on either side of the target, the measuring

plate simply substituted for the corresponding spacer plate in Fig. 2-14.

A measurement of the flux density component perpendicular to the plate

was derived from the induced voltage in any two neighboring loops connected

in series opposition. For the two loops illustrated in Fig. 2-16, for

example, ,• t

L v dt

BJL(t) = 1550 - Tesla s (2-11)

where r is in inches and v in volts. This value is the average perpendicular flux density over the area between the two induction loops. In

deriving further results, we will assume the answer to apply at a radius

midway between the two loops.

43
To measure the tangential (radial) component of flux density at any

radius, the front and back loops at that radius are connected in series

opposition, and calculations are made from

/c

v dt

Br(t) = 1550 -Teslas

27rrh

where r is the radius of the two induction loops and h is their separation,

both in inches.

The specific tests are identified in the following Test Table, and the

data appear on Figs. 2-18 through 2-24.

TEST TABLE

Test

1

2

3

4

Configuration

Target removed,

Field Measurement

Plate against

coil surface

Target removed,

.041" spacer

between coil and

Meas. Plate

Target in place,

Meas. Plate

between coil

and target

Target in place,

Meas. Plate

beyond target

Measurements

Made

Bx on both

sides of Meas. Plate

Br

Br

Bx adjacent

to target

Br

BA adjacent

to target

Br

Data

Peak Bj_ vs r (2-18)

plotted in Fig.

Peak Br vs r(2-19)

plotted in Fig.

(2-20)

Br vs time

in Fig.

Bx vs time (2-21)

in Fig.

Br vs time (2-22)

in Fig.

B± vs time (2-23)

in Fig.

Bf vs time (2-24)

in Fig.

44
Coil (Ce/.p«rJ S

. 08S" Frov*.

Co i'l (Copper ) Surf.

Vr i/o0 volti

. > Coils series

R, IMCHES

Fig. 2-18 Maximum BI vs Radius, no

Target Present

- 2

\;= MOO

C s 4oo

/' X

,05i

R I/JCHES

COIL

Fig. 2-19 Maximum B vs Radius, no

Target Present

45
0)

C/O

S_

(O

QJ

CU

e\j

CVJ

•

CT>

9V1S31

QJ

•(-> 4-> -a

c 3

QJ JO

O

4->

c

<u

>

O

CU QJ

CD O£

sQJ

CO 4->

+-

O

QJ

CJ1

S_

ro

<u

O

CVJ

I

CM

QJ

C

fO

CD Q.

SV1S3J.

46
C

O

-ooos-to0} COooCMICM-a•^oos-(O<DO) -PCoCOCMCMICM

cr>

SV1S31

47
set.tt

00

<

_l

00

LU

-100.0*

t.eoot •

Fig. 2-24 B on the Far Side

48
The purpose of Test 2 was to obtain Br data, due to the coil alone, at

the mid-plane position of the target to be added for Tests 3 and 4. '

These data will be used in the pressure calculation.

All Br data showed an anomolous behavior (irregularities) at r =.4

relative to r = .6. A separate check was made, in which the plate was

reversed (interchanging the two sides). This produced a reversal of the

irregularities, so the effect was probably due to an inaccuracy in the

construction of the plate.

C. Eddy Currents

Information on the magnitude and distribution of eddy currents in the

target was obtained from a finite increment version of Ampere's Law,

(2-13)

AIeddy = iT* B-dL

One of the circulation paths is illustrated by the dotted rectangle in

Fig. 2-25.

.20

©

HELD

MEASURING

WIRES

.5

CINCHES

Fig. 2-25 Illustration of a Circulation

Path For Evaluating Eddy Currents

49
It is imagined that the current in the drive coil (below the target)

is referenced into the view, and the eddy currents are referenced out of

the view. Thus, a counterclockwise circulation is taken, and the result

is divided by yQAr to obtain the current per inch of radius, assumed to

apply at r = .4". For the path illustrated in Fig. 2-25,

AI eddy

BlN(-5) + BlF(.5) Blp(.3)'+ BiN(.3)

h - h

|BrN(.4) - BrF(.4)JAr (2-14)

Numerically,

AI

- 4.14 [B1N(.5) + Blp(.5) - B1F(.3) - B±N(.3)J

+ 20.2 H N(.4) - B- F'(.4)] KA/In (2-15)

^eddy

where all B components are in Teslas.

Calculated results for Kedd vs time at various radii are graphed in

Fig. 2-26, and plots of Kedd vs radius are shown in Fig. 2-27. The

latter figure includes the drive coil current density for comparison. It

is observed that a reasonable (about 50%) current transfer efficiency

occurs very early in the process, whereas the eddy currents are only about

10% of the drive current at the time of 250 ys. Eventually, the eddy

currents reverse, due to the reversed EMF from the collapsing magnetic

field.

50
1*.000_

c

M

<

U)

ICAL

Q

a

LU

EDDY CURRENT

PER UNIT RADIUS

OF TARGET

-t.eooo _

SEC

Fig. 2-26 Calculated Eddy Current

Densities vs Time, at Various

Radii

51
t

oO) oD- 4J ' C(UJ-s-3-oUJCMCMoO)

NI/V>J

52
D. Pressure on the Target

Having evaluated the eddy currents, the pressure on the target can be

found by the "conductor force equation," utilizing data for B along the

center plane of the target. The Bp data used was that from Test 2,

corrected for change in coil current due to the presence of the target.

Thus the Br data from Test 2 were multiplied by the factor

Icoil(t) in Test 3

Icoil(t) in Test 2

The difference in the two currents is small, as seen from Fig. 2-15.

Figure 2-27 illustrates the factors involved in the pressure calculation. The normal force exerted on a typical ring from r to r+Ar is

AFi=BrAIeddy

COIL

Fig. 2-28 Illustration For Pressure

Calculations

53
yielding the following pressure formula:

p = 5'71 Br Keddy PSI

(2-16)

where Kedd is in KA/In, and Br in Teslas.

Calculated results for p vs t at various radii are graphed in Fig. 2-29,

and pressure "footprints" for several values of time are shown in Fig. 2-30.

35.000

10.000

NORMAL FORCE PER

UNIT AREA OF

TARGET SURFACE

.fc

FIG. n

I I I

t.O SEC

i r

701.00

Fig. 2-29 Pressure vs Time at

Various Radii

54
1.5

RADIUS, IN

Fig. 2-30 Pressure Distribution with

Radius

E. Total Normal Force and Impulse

The total normal force on the target was evaluated by the finite

interval version of

Force = / p.

Jr=Q

Zirrdr (2-17)

and is graphed in Fig. 2-31, and its time integral in Fig. 2-32. The latter

figure shows an impulse degradation of about 25%, due to the collapsing

magnetic field.

An independent measurement of the impulse was made with a ballistic

pendulum, and the pendulum value exceeded the .0112 Ib-sec value on

Fig. 2-32 by about 10%.

55
z:

Fig. 2-31 Total Normal Force vs Time

F. Radial Forces in the Target

If Bj^were used in the force calculation (Fig. 2-28), the result would

be the radially acting force tending to compress the target. This force

would turn to an expansive force when either

to the reference directions.

or K . . become opposite

56
1C.000 .

o

to

RCAL

UJ

0.

Fig. 2-32 Time Integral of the

Total Normal Force

57
The radial force acting on the incremental target area rA<j)Ar is

AFr = -( K

e ddy (

l - (2-18)

and the force per unit of target (surface) area is

The specific formula, for example, to evaluate p at r =.6" is

pr(.6) = -1.43 [B F(.5) + B^N( .5) + B±F(.7) + B N(.7l] Keddy(.6) PSI (2-20)

where the B components are in Teslas, and K dd in KA/In. B p data are

from Test 4 and B.., data from Test 3.

. Figure 2-33 shows graphs of p vs time, at various r locations, and

Fig. 2-34 shows p vs r at various times. At early stages, the radial

forces are compressive at small radii, and expansive at large radii. This

action is reversed in later stages.

58
c0•^

4-5Z3

.0

«r~S_

4_>

CO

•^Q

ro

•r~

T3

roC£

5^OJ Q. <UO5- Ou_

<L> Cr2Q.1CHH

roOJ SeX. QJ CDS- *DhI

OJ s-O) a.a)ao

racus- OJ CDsre

roCOi

evi

ISd

59
IV. Electrodynamics Modeling

Analysis work is in progress on the fundamental problem of a circular

impulse coil near an infinite sheet of flat stationary conducting plate

(target). Much of this work is being pursued by Mr. Robert Henderson for

a PhD dissertation.

A. The Field Problem

The electromagnetic field portion of the problem is being handled by

the Hankel transform technique described by El-Markabi, et al (2-2).

Figure 2-35 shows the coil-target geometry with nomenclature illustrated.

The Hankel transformed fields are defined by

0 COIL

i_

T^

R,

R

<r

d

— z

METAL

PLATE

Z.-o 2,

Fig. 2-35 Coil-Target Geometry

and Nomenclature

60
(2-21)

These are said to be the fields in "Hankel Space". El-Markabi shows that

when Maxwell's Equations for axially symmetrical fields are transformed into

Hankel Space, and displacement current ignored, then the resulting equations

are analogous to those characterizing one-dimensional transmission lines.

Thus, the field problem can be modeled by an equivalent transmission circuit

in Hankel Space. El-Markabi's solution assumes the coil current to be

sinusoidal, and the fields are then phasors. He also approximates the

volume-distributed coil current by several current sheets. A three-sheet

approximation is pictured in Fig. 2-35, and is assumed for the equivalent

transmission line model pictured in Fig. 2-36, on the following page.

In the T-line model, E (X,«) is analogous to voltage and H (X,w) is

analogous to current. The model consists of six line sections. The

extreme left section (z<z0) extends to z= - °°, so there is no reflected wave

on it. This section may then be replaced by its characteristic impedance

ZQ. The same reasoning applies to the section (z>z,J. All other sections

have been represented in Fig< 2-36 by their 4-terminal network parameters

(A,B,C,D). The three sheet current sources are assumed to be the same.

This assumption may be removed in some future work, to account for skin

effect. Then the middle source would have a phase delay and reduced

amplitude.

Formulas for the characteristic impedances,Z , and for the propagation

constants, y> are:

Air Sections

Metal Plate

V

vY = XZa

P

61
r\j

LU

1

3-

<

I

a. a A.

QQ O O

•eLU

in

M

QJ

•o

O

zn CD m 05

< CO O D o

«/)

to

-eLU

t

o u

< CD

O <J I

J Q ,

c

(O

OJ

O

Q.

00

QJ

^:

-a

O

SQT3

i—

QJ

O)

•e

LU

VO

CO

C\J

< CO O

o

•e

Lu — N

-Wvr
Formulas for the 4-terminal network parameters are:

AC = Dc = COSH (Ah/3)

B,. = I Sim (Ah/3) C a

Cc - Bc

 Za2

Ag = Dg = COSH (Ag)

Bg = Za SINH (Ag)

C9

 = B9

 Za2

Ap - Dp = COSH (Ypd)

Bp = Z

p

 SINH V)

CP ' BP V

The Hankel Space current sources are calculated from

ljAr)rdr (2-22)

where N is the number of coil turns, and I(co) is the Fourier transform of

the coil current.

Analysis of the network (Fig. 2-36) is done by a computer. If fields

are desired in real space, then the inverse Hankel transformation must be

calculated:

X(r,z)J_

(2-23)

and if they are desired as time functions, then an additional calculation

must be performed for the inverse Fourier transform.

63
B. The Circuit Problem

The electrical circuit can be represented as shown in Figs. 2-37 (a) and

(b). Both circuits have

/t)

U(t)

f?

•Mr

Fig. 2-37 (a)

Applies up to time t ,J\

when the diode clamp

begins to conduct.

Fig. 2-37 (b)

Applies for t>t

initial conditions represented as external sources. Direct time-domain

solutions for these circuits cannot be written because R and L are functions

of frequency.

Figures 2-38 (a) and (b) are the frequency domain circuits. They show

a breakdown of elements of the physical circuit. It is assumed that this

circuit contains a series-connected dummy coil in free space (no target).

DUMMY COIL

r

R, JuL, !

VV—'W r-^NA—^^

i

L

I, (co)

•ACTIVE:

COIL

Fig. 2-38 (a) Frequency Domain

Circuit (t<t )

64
AR(cu)

AA—\\^-

A

t) i,(tx)[TTSM -^

Fig. 2-38 (b) Frequency Domain

Circuit (t>t )X

The symbols in Fig. 2-38 represent:

R! = Equivalent resistance of the capacitor and the cable.

L! = Equivalent inductance of the capacitor and the cable.

Rri(u)

 = A.C. resistance of the dummy coil.

L.= Inductance of the dummy coil.

Rd(w) = Effective resistance of the active coil.

R'(w) = Portion of R,(w) that is the free-space A.C. resistance, a a

AR(to) = Portion of R,(u) that is field-induced. It is due to the a

presence of the metal plate.

L (to) = Inductance of the active coil in the presence of the metal

a

plate.

Data on RJ(U>) and R' (u>) can be estimated by procedures given in Section I. 0 a

The field-induced impedance of the active coil

Z(u>) = AR(u>) + jwLa(w) d

must also be found before the circuit problem can be solved. This impedance

is defined as the phasor voltage across the coil terminals, resulting from

E in Fig. 2-35, per ampere of coil current. Thus, a preliminary solution

4>

to the field problem must be carried out. The simplest procedure is to

set I(to) in Eq. 2-22 equal to unity, and then solve the transmission network

for Erf, , E* , and E<t, . T

0

 Yl 2

65
Basing the calculation on the average of these three fields, and following

El-Markabi's reasoning, the formula for calculating Z(w) becomes

dxNow the circuit problem can be solved.

Ii(u>) is first found from the circuit of Fig. 2-38 (a), and its inverse

transform, ii(t), calculated, along with the corresponding vc (t). The

value of t (called t in Figs. 2-37 (a) and (b))for the first zero crossing /\

of vc (t) is noted. The circuit in Fig. 2-38 (b) is then solved for i2(t),

using ij(t ) as the initial condition. The circuit current is finally

^

constructed from the two solution parts by

txj]+ i2(t-tx), (2-25)

and its Fourier transform I(o)) found. This is the current spectral information which enters into the Hankel Space current sources (Eq. 2-22).

C. Pressure and Force on the Target

Starting with the equation for the force per unit volume within the

target

' T = 7 x B" (2-26)

vol

where T is the local current density, and neglecting displacement current,

then the pressure becomes

P(r) =

(zs) B2 (z,) f' 8B -I

+ J Br j+- dzj . (2-27)

66
Also utilizing V-B = o, Eq. 2-27 can be further modified into

P(r) -I |Dr

- B*(z3)

F I <rBrB2)dz] a

The force contributed by the integral term in Eq. 2-28 is

dr

r=o

(rBrBz)drdz.

However,

r=o

|r(rBrBz)dr =

r=o

- (rBrBz) = o

(2-28)

at r=<» at" r=o

Thus, while the integral term in Eq. 2-28 may influence the distribution of

pressure, it does not contribute to the total force. This total force should

then be given by

F(t) = [

Br<Z3) - Bz(2a) " Br(z*) (2-29)

r=o

The calculation for pressure distribution appears to be difficult

(no formula has yet been discovered that utilizes only field information

at the two surfaces of the target). On the other hand, the force seems to

be more readily calculable.

67
D. Preliminary Results

All of the computational work to date has been for the problem conditions

that existed in the Magnet Field Diagnostics experiment, so that calculated

results might be compared with experimental results. A tabulation of the

conditions are:

Coils

«

(Active and Dummy Alike)

N = 30 turns, Ri = .125 inch,

R2 = 1.00 inch, h = .188 inch,

Rdc = .0235 fl, Ld = 22yH.

Target

- 2024-T3 Aluminum (a = 3.48 x 107 u/m).

d = .032 inch.

Gap

g = .1095 inch

Circuit

C = 600yF, Ri = .054 Q,

LI = 1 yH, vr(o) = 400 volts

C

Calculations have been completed for the coil impedance Z(w), the circuit

current i(t) and its spectrum I(u>), and the Hankel-space fields B (A,<jj) and

EL (A,u>). Time-domain fields on the target surfaces have not yet been cal- ri*

culated, nor has the target force.

Figures 2-39 through 2-43 show results pertaining to coil impedance.

Fig. 2-39 is a plot of the inductance of the active coil at three frequencies,

as calculated from Eq. 2-24, as the upper limit on A is varied. This plot

gives an indication of what A range provides the largest contribution.

Figure 2-40 is a similar plot for the real part of Z(u), which is the fieldinduced resistance.

68
lOh

50 Hz

5000 Hz

100 200

UPPER X LIMIT

300 400

Fig. 2-39 Calculated Coil Inductance,

vs Upper A Limit

69
100

.0

cc

<J

O.I

50 Hz

.01

I

100

UPPER X LIMIT

200

Fig. 2-40 Calculated Coil Resistance

Contributed by the Metal

Plate vs Upper X Limit

70
Figure 2-41 is a graph of coil inductance vs frequency, with math

model calculations, formula calculations, and experimental measurements

compared. These comparisons are fair, but could be better.

1

Fig. 2-41 Coil Inductance, Comparative

Results

ORJGHNSAL PAGE IS

POOR QUALITY

71
Figure 2-42 shows the calculated field-induced resistance as a function

of frequency, along with bridge measured data. The measured values are

the difference between the measurement without the target plate and a

re-measurement with the target brought into position.

25*0 5600

Fig. 2-42 Coil Resistance Due to the

Metal Plate vs Frequency

72
Figure 2-43 shows no additional results from the math model. Rather, it

puts the field-induced coil resistance in perspective, relative to the D.C.

and A.C. resistance values for the series combination of the active and

dummy coils. At 4000 Hz, for example, the resistance almost doubles as a

result of AR.

ORIGINAL PAGES

Of POOR QUALITY

Fig. 2-43 Comparison of Various

Contributions to Coil

Resistance

73
Figure 2-44 shows a comparison of the coil current calculated from

the math model Eq. 2-25 and the measured current.

1.0 -

I/O

Q-

, °-5 -

EXPERIMENTAL

/ MODEL

- /

/

/ , , .

°Nv

^*N*v.

Diode Clamp o —

^" Condoc-ts fti ^^^^_

t-«./*S. ^^^^^^

.

1 1 1 1 1 1

ZOO HOO 600 600

Fig. 2-44 Comparison of Coil Current

Calculated from the Math

Model with that Measured

74
V. References

2-1 Frederick W. Grover, "Inductance Calculations, Working Formulas and

Tables," Dover Publications, Inc., 1962

2-2 El-Markabi, M.H.S., et al, "Electromagnetic Properties of a Circular

Cylindrical Coil in a Set of Planar Ferromagnetic Regions", IEE

Proceedings (London), Vol. 129, Part A, No. 8, pp 582-589, Nov. 1982.

75
CHAPTER 3. STRUCTURAL DYNAMIC TESTS AND STUDIES.

I. Introduction

The structural dynamics associated with electro impulse de-icing has

proved to be a difficult and challenging problem. The EIDI project was begun

in May 1982 and the structural dynamic investigations conducted to date have

raised as many new questions as have been answered. The project has been a

modest activity from the standpoint of structures personnel and suitable

instrumentation. The majority of the resources have been utilized in the

proof of concept through subjective optimization of the numerous IRT test

demonstrations for the participants in the industrial consortium. Eight

IRT tests and two flight test demonstrations are summarized in Chapters 5

and 6. This Chapter is divided into the following sections:

II. Transient Strain Measurements near EIDI Coils.

III. Boeing (BCAC) 767 Leading Edge Slat Strain Survey.

IV. Semi-Cylinder Leading Edge Investigations:

A. Experimental Results.

B. Analytical Studies and Models.

V. Recent Developments Related to EIDI Force Measurements.

In Section II, the problem of strain measurement in both mild shock

environments and strong magnetic fields is discussed and the project solutions

to these problems are presented in Section III. The semi-cylindrical leading

edge structure was developed by Wichita State University personnel and

represents a portion of the graduate dissertation of Mr. Robert Friedberg.

The goal of this study is to gain some fundamental correlation

between experimental and analytical models for the simple geometry afforded

by a semi-cylindrical shell. For the past two years, Mr. Friedberg has

served as the principal designer for such items as coil winding and fabrication, coil mounts and the structural integrity of the coil supporting

structures, and he is largely responsible for the concept of skin-mounted

coils. In addition, he has also served as the University observer on the

two flight tests conducted to date. These demands have left little time

for any serious analytical investigations of the structural dynamics of the

EIDI process. Thus, during the first two years of the project, the

emphasis was on preparing the coils and their mountings and preparation of

the test article for the associated IRT test. During this same period the

project did manage to assemble certain resources that would serve to permit

some fundamental investigations of the structural dynamics associated with

EIDI.

PRECEDING PAGE BLANK NOT FILMED
The project began its activities with a single suitable acelerometer and

charge amplifier, and a HP-5423A Structural Dynamics Analyzer. This two

channel instrument has a 25 kilohertz bandwidth and a sample time

of approximately 10 microseconds, which proved to be adequate for acceleration measurements but inadequate for the proper resolution of strain

measurements in magnetic fields. This problem is discussed and demonstrated

in Section II. The project did acquire several transducers, amplifiers, and other miscellaneous instrumentation; and in May 1984, a

Nicolet 3091 Digital Oscilloscope was purchased. This two channel instrument has a digital sample time of 1 microsecond and channel subtraction

capability and was ideally suited for the proper resolution of the

strain measurement problem. Thus, during late 1984, the first reliable

strain measurements were available to the project and some of these

results are reported in Section IV. A. Unfortunately, the only output

was an analog xy-recorder, hence the signals had to be captured,

processed, and displayed immediately. During November 1984, the project

acquired a Zenith Z-150 micro-computer that was interfaced with the Nicolet

oscilloscope and the floppy disk units have provided a means of permanently

storing the various signals. A HP-7245B Digital Plotter is now interfaced

with the micro-computer and this plot capability is displayed in Section

III. The most recent improvement is an interface between the microcomputer and the University mainframe IBM-3081 computer, permitting the

transfer of signals to the 3081 for subsequent processing and display.

The second major resource was the installation of the MSC/NASTRAN code on

the IBM-3081, in October 1984. This finite element code has provided the

capability of performing structural dynamic investigations for various EIDI

applications. In September 1984, three graduate students majoring in

structural dynamics began to develop analytical models for the transient

response of pulsed leading edge structures. The majority of these studies

have been confined to the semi-cylinder and are reported in Section IV. B.

The simple concept of a semi-cylindrical leading edge simulation has also

attracted the attention of the McDonnell Douglas participants in the EIDI

program; Mr. David Blyther, Dr. Don Kerans, and Mr. Lou Barretto. The EIDI

project group has designed and fabricated a duplicate of the semi-cylinder

described in Section IV for McDonnell Douglas and they plan to fully

instrument the shell and conduct an IRT test in June or August, 1985. The

principal goal of this study is a time correlation of acceleration, strains,

and high speed photography in a serious attempt to answer the question of the

basic ice shedding mechanism in the EIDI pulse.

While all of these elements and resources are encouraging and appear to

be converging on a true analytical and experimental correlation of the EIDI

problem, there remains one major unsolved area, namely, an adequate

78
representation of the spatial and temporal behavior of the coil pulse. This

behavior was first shown by Sehrag, Ref. 3-1, for a single case of coil size,

gap, target material, and electrical parameters. The experimental result was

conducted with a special magnetic field measuring plate fabricated by WSU

personnel, described in Section 2. B of this report. This experimental result

was generalized into a separated solution of spatial and temporal behavior by

Bernhart, Ref. 3-2, and used successfully in a transient response study of a

flat plate subject to a EIDI pulse. While this limited study did show

favorable agreement with experimental results, it also clearly identified

the response sensitivity to both rise time and pulse duration. Mr. John

Walsh, who is a graduate student in Aeronautical Engineering, will be

repeating the experiment using the magnetic field measuring plate for a

wider range of parameters during the summer of 1985. These results will be

compared to ballistic pendulum results presented by Sehrag,Ref. 3-1, and

several British authors, Ref. 3-3. A very recent parallel experimental

development designed to measure the temporal behavior of the EIDI pulse is

reported in Section V. This experiment employs fundamental one dimensional

wave mechanics in a 48 inch long polycarbonate circular rod with a propagation time of approximately 850 microseconds. This time delay is adequate

to measure the plane wave behavior of the force pulse. Preliminary

results appear to be very encouraging, however a proper model for the

dispersion mechanism in the rod is currently under study.

The question of fatigue has been raised on several occasions by various

observers. It is important to recognize that a prolonged fatigue test of a

representative leading edge structure at this point in the EIDI program, would

essentially be a longevity test of the electrical components in the EIDI

pulsing system. The question does need to be addressed, but only after

certain question related to the potential usage of the system are answered;

e.g., number of hours of operation per year, and a suitable pulsing system

becomes available to the project. The EIDI project has developed several coil

attachment concepts that serve to reduce peak stress levels but raise other

questions related to bonding fatigue subject to shock environments. This

potential reduction is shown for the Boeing 767 slat in Section III.

The basic goal of these structural dynamics studies and tests is the

development of an analytic design procedure for the EIDI program. The studies

conducted to date are briefly summarized as follows:

A finite element model study of a semi-cylindrical leading edge has been

completed. The study has demonstrated that a proper finite element model is

adequate to duplicate, and hence predict, the peak accelerations and strains

that have been observed in various airfoil tests. The development of a finite

element model for a airfoil leading edge with more complex geometry and

boundary conditions is in progress. The EIDI project personnel are currently

79
conducting parallel studies to develop an analytical model for the EIDI pressure pulse with appropriate experimental verification. This model Hill become

an Integral part of the aforementioned structural dynamics computational code.

The pressure pulse format will In Include both the spatial and temporal

behavior of the Impulse for a variety of electrical parameters such as voltage

level* storage capacitance* and coll Inductance; as well as the target

material thickness and conductivity* and air gap.

When these basic elements are joined and supported by experimental verification; meaningful parametric studies may be conducted that will yield the

dynamic response of a proposed airfoil leading edge structure and coll

configuration. The computational code will have the flexibility of examining

the effects of one or more colls that can be positioned at arbitrary location

1n the leading edge cavity. This capability can then be easily extended to

simulate magnetic doublers* skin mounted colls* coll mountings* Internal ribs*

and various boundary conditions.

Dr. Don Kearns* McDonnell-Douglas Corp.* Is currently adding simulated

1ce accretions to his structural dynamic model of a semi-cylinder. The

McDonnel-Douglas tests* scheduled for the summer of 1985* will hopefully serve

to partially answer the question of the fundamental 1ce shedding mechanism In

the EIDI process. If this test Is successful* 1tw1ll lead to optimization

studies to maximize a specific performance criteria* e.g.* acceleration*

strain* strain rate* etc.* that 1s associated with optimum 1ce shedding.

80
II. Strain Measurements Near EIOI Colls.

Strain measurements 1n the magnetic field produced by a EIOI coll has

proved to be a difficult problem and the following demonstration problems are

Intended to portray some of the difficulties. The EIDI coll can generate

magnetic flux densities 1n excess of 10,000 gauss, and the field strength 1s

developed 1n 100 to 300 microseconds. Any strain gage 1n the proximity of

the coll 1s subject to transverse magnetic flux which results 1n Induced

voltage 1n the gage. If the resulting Induced voltage 1s not cancelled,

eliminated or minimized, 1t obscures the strain signal developed in the gage.

Thus, the problem 1s likened to a signal plus noise phenomena. The typical

strain signal 1s be on the order of 1 to 10 millivolts, whereas the Induced

or noise voltage may be of the same order and in some cases greater than the

strain signal.

Contacts with strain gage manufacturers revealed that the Measurement

Group, Inc., (Ref. 3-1) had solved a similar problem for the Princeton

University Plasma Physics Laboratory 1n connection with the Tokamak Fusion

Test Reactor. The application Involved magnetic flux densities as high as

50,000 gauss and a time duration of approximately one second. A special gage

was employed, Identified as a "dual element gage", which consists of two

Identical overlaid grids, laminated 1n precise alignment with one another, and

connected in series such that the current passes through them 1n opposite

directions. The gage 1s shown 1n conceptual form in Fig. 3-1. Since the

grids are very close together with respect to the field gradients, Induced

voltage in the grids tend to be equal and of opposite sign, rendering the

gages largely Immune to magnetic effects. The Measurements Group also

recommends a woven cable product commercially known as Inter-8 Weave to

minimize the magnetic noise 1n the cabling. Several of the "dual element

gages" (WK-13-125-WS-700) were obtained and Installed on a representative

leading edge structure. The lead attachments to the "dual element gages"

posed very special soldering problems and several gages were lost during

Installation. The recommended cabling also proved to be excessive in weight

for the relatively light weight structures common to aircraft leading edges.

The first time the "dual element gages" were tested on a EIDI pulsed

airfoil, the gages sheared along the bonding plane between the grids and the

resulting misalignment rendered them useless to the project. Thus, the

EIOI strain gage environment 1s not only harsh due to the magnetic flux, it

1s also subject to very high acceleration levels which can reach peak values

of 10,000 to 20,000 g's 1n 100 microseconds or less. This failure prompted

81
Or. R. L. Schrag to propose a modification to a standard strain gage which

1s herein Identified as a "compensated gage", and 1s also shown In Fig.

3-1. This modification Incorporates the Installation of a single loop

of 32 gauge enameled wire which serves to partially compensate or minimize the Induced magnetic effect. The compensation loop 1s also bonded

to the gage where appropriate. Standard Wheatstone bridges are used 1n

the EIDI project with a complete avoidance of any wire wound potentiometers or other electrical components. Gage leads consist of twisted

pairs of 28 or 32 gauge Insulated wire twisted with a drill motor. The

leads are both Inexpensive and light weight.

Strain measurement must be completed in two steps. After performing a

resistance balance of the bridge* the first step 1s to remove the bridge

excitation DC voltage source and pulse the EIDI coil to establish the induced

magnetic effect 1n the gage. In the second step, the bridge excitation 1s

restored and the EIDI coil 1s pulsed a second time to establish the Induced

magnetic effect plus the mechanical strain. Subtracting the two signals

yields the desired mechanical strain. This two step operation 1s of course

difficult to achieve in the IRT tests. Some data have been obtained by using

a pre-recorded magnetic effect pulse and subtracting it from the pulse

data obtained under icing conditions.

The two step process Is identified as a DC-bridge measurement by project personnel and has been used with varying levels of success. Figs. 3-2

and 3-3 depict some additional problems for the DC-bridge related to the

digital sample time of the transient capture Instrumentation. The data

contained in the two figures represent strain measurements for a 2.5 Inch

radius semi-cylinder* 66 Inches in length described in Section IV.

This simulated leading edge structure contains a single coll located at

the mid-length and compensated strain gages placed on the outer and Inner

surfaces of the 0.040 Inch thick skin. The data in Fig. 3-2 represents

the two signals for each location and the resulting subtraction when using

a HP-5423A Signal Analyzer* whose sample time Is approximately 10 microseconds. The desired strain signal 1s literally obscured during the

first 150 microseconds and the establishment of a peak strain is

difficult* Fig. 3-3 represents the same data acquired with a two channel

Nicolet 3091 Digital Oscilloscope using a 1 microsecond sample time. The

figure clearly indicates the desired levels of the two strain signals for a

well compensated gage. The resulting data can be reported with confidence as

has been done In Section IV. of this Chapter.

A recent Improvement in strain measurements has also been developed by

Dr. R. L. Schrag. The development is an AC-bridge* wherein the carrier

frequency 1s 100 kilohertz. which 1s considerably higher than any known

strain gage instrumentation. The principal advantage of the Instrument is the

82
single step measurement process. The AC-bridge does require both resistance

as well as capacitance balancing and 1s exceedingly sensitive to lead wire

movements during the balancing and pulsing operations. In spite of some of

these difficulties, the Instrument exhibits great promise and hopefully funds

can be obtained for future Improvements and the fabrication of multiple

channels. A schematic diagram of the AC-bridge 1s shown In F1g. 3-4.

F1gs. 3-5 through 3-7 are strain measurements for a single gage on the

Boeing 767 leading edge slat; details related to the gage location and the

slat structure may be found In Section III. Fig. 3-5 1s the data for

an noncompensated gage and standard strain gage leads using the DCbridge process; while Fig. 3-6 Is the same data with a compensating loop

added to the gage with twisted leads. The Improvement resulting from the

compensating loop Is obvious. F1g. 3-7 represents AC-bridge measurements for both the noncompensated and compensated gage. The compensating loop neither adds or detracts from the performance. A brief summary

of the peak strain level developed by the four measurements 1s

summarized below. The peak circumferential (chordwlse) strain taken

as the average of the four measurements 1s 2070 micro-strain.

Figure 3-5. 1940 micro-strain

Figure 3-6. 1850 micro-strain

Figure 3-7. 2220 micro-strain

Figure 3-7. 2270 micro-strain

83
III. Boeing (BCAC) 767 Leading Edge Slat Strain Survey.

A 76 Inch long section of the BCAC 767 leading edge slat was fitted with

EIOI colls and tested 1n the NASA Lewis IRT on August 13, 14; 1984. A sketch

of the leading edge cavity 1s shown 1n Fig. 3-8. Prior to the test» this

geometric definition was used 1n a simple two dimensional computer code to

approximate 'the fundamental mode shapes of the leading edge section. The mode

shape plots are displayed 1n F1g. 3-9 and are used 1n the preliminary design

of coll placement in both the upper and lower surfaces. A short 4 Inch long

section of the leading edge cavity was also dynamically tested to verify the

analytically predicted frequencies. The first two frequencies compared

favorably with the computer code.

The 76 Inch long section 1s divided Into two equal length bays of 38

Inches each with ribs on each end and at the center of the span. The selected

EIDI coll types and positions used 1n the IRT test are summarized below.

Coll Spanwlse Chordwlse Position

No. —Description— Position (measured from nose)

1 2 side coils-nose attached 6.3" from Lower Sfc Upper Sfc

lower Mb 1.6" 2.5"

2 2 side coils-spar mounted 20.0" from

lower rib

3 2 side coils-spar mounted 32.7" from

lower rib

4 2 side coils-spar mounted 32.7" from

upper rib

5 2 side coils-spar mounted 20.0" from

upper Mb

6 2 side co1ls-sk1n mounted 6.3" from

upper Mb

3.1"

3.1"

2.5"

2.7"

1.4"

2.5"

2.6"

2.5"

2.4"

1.4."

The colls used 1n the strain survey were colls 2, 3, and 6. Two strain

gages were Installed over the upper surface EIDI colls 1n the circumferential

(chordwise) and longitudinal (spanwise) directions. In addition, a single

chordwlse gage was utilized on the lower surface over coll 3. The test data

are presented 1n F1gs. 3-10 through 3-13.

These three colls were employed 1n five different IRT test runs. The

selected strain measurements correspond to two different sets of electrical

parameters for IRT test numbers (4) and (8). Each coll pair was pulsed twice

84
1n each test with the sequence; (2,2),

are summarized below.

(3,3), and (6*6). These two test runs

IRT

TEST

(4)

(8)

COIL (2)

400UF 900V

162 Joules

200UF 1250V

156 Joules

COIL (3)

400UF 900V

162 Joules

200UF 1250V

156 Joules

COIL (6)

400UF 800V

128 Joules

200UF 1120V

125 Joules

Subjective

Evaluation

Fair* some

residual

Clean!

The BCAC-767 slat 1s fabricated from 7075-T6 aluminium sheet 0.062 Inches

1n thickness. The published mechanical properties for this material are;

78,000 ps1 ultimate strength* 68,000 ps1 yield strength* and 10,400,000 ps1

for the modulus of elasticity. From an examination of the strain plots one

can conservatively state that the peak chordwlse and spanwlse strain have

nearly Identical occurrence times. Thus, the maximum peak stress 1s predicted

on this basis* using a value of 0.30 for Polsson's ratio. In each case* the

maximum stress will occur 1n the chordwlse direction. These data are

summarized below.

COIL LOCATION

IRT Test (4)

162 Joules

Upper Sfc.

IRT Test (8)

156 Joules

Upper Sfc.

COIL LOCATION

IRT Test (4)

162 Joules

Upper Sfc.

Lower Sfc.

IRT Test (8)

156 Joules

Upper Sfc.

Lower Sfc.

COIL LXATION

IRT Test (4)

128 Joules

NUMBER (2):

Chordwlse Peak:

Spanwlse Peak:

Maximum Stress:

Chordwlse Peak:

Spanwlse Peak:

Maximum Stress:

NUMBER (3):

Chordwlse Peak:

Spanwlse Peak:

Maximum Stress:

Chordwlse Peak:

Chordwlse Peak:

Spanwlse Peak:

Maximum Stress:

Chordwlse Peak:

NUMBER (6):

Chordwlse Peak:

Spanwlse Peak:

2*710 micro-strain

1,820 micro-strain

37,200 ps1

2,520 micro-strain

1,590 micro-strain

34*300 ps1

2,730

1,840

37,500

3,030

2,510

1*600

34,200

2*810

micro-strain

micro-strain

ps1

Micro-strain

micro-strain

micro-strain

ps1

micro-strain

2*050 micro-strain

1,010 micro-strain

85
Upper Sfc. Maximum Stress: 26,900 psi

IRT Test (8) Chordwlse Peak: 2,100 micro-strain

125 Joules Spanwlse Peak: 974 micro-strain

Upper Sfc. Maximum Stress: 27,300 ps1

Colls 2 and 3 produce maximum stress values of 37,000 ps1 at the desired

de-1c1ng levels on the upper surface; this 1s extrapolated to 40,000 ps1 for

the lower surface. While these values remain well below the yield strength

for the material, the question of long term fatigue remains unanswered at this

time. The skin mounted coll pair located at position 6, Is clearly superior

to the spar mounted configurations, with a maximum stress of 27,000 ps1. The

fatigue longevity of this configuration needs to be addressed.

86
IV. 2.5 Inch Radius Semi-Cylinder Studies.

A. Experimental Results.

During April 1984, a semi-cylindrical leading edge type structure was

designed and tested by Mr. Robert FMedberg as partial fulfillment of his

graduate study program at Wichita State University. The semi-cylinder was

selected as a representative candidate for delclng studies, an experimental

data base, and subsequent analytical modeling. The semi-cylinder 1s 66 Inches

1n length, an Inside radius of 2.5 Inches, and skin thickness of 0.040 Inches.

The material used 1n the fabrication was 2024-T3. The semi-cylinder 1s

attached to a reasonably rigid spar fabricated from a fiberglass covered 40

pound density 3/4 Inch thick foam core, 1t has no Internal ribs or bulkheads

except the closure ribs at each end. The semi-cylinder 1s attached to the

spar and closure ribs with No. 10 sheet metal screws spaced two Inches center

to center. The semi-cylinder was equipped with two coll configurations, a

single coll and a pair of series connected double colls for two different IRT

tests. The first IRT test was performed on May 19, 1984 and employed a single

coll located at the nose and mid-length of the semi-cylinder as shown 1n Fig.

3-16. The coll had a nominal radius of 1.25 Inches with 50 turns of

rectangular copper conducter and shaped to conform to the Internal radius with

provision for a 0.010 Inch air gap.

Acceleration data was established for the single coll configuration and

1s displayed 1n F1gs. 3-14 and 3-15 for the energy level of 36 Joules as

noted. The accelerometer used In the Investigation was an Endevco Model 22,

which has a total weight with the attached cable of 0.4 grams and a physical

size of 0.14 Inches diameter, 0.095 Inches In height and a response level of

approximately 12,000 g's. Mass loading of the light weight aluminium leading

structures 1s an Important consideration and this particular acceleronieter has

the largest g level per unit mass known to project personnel. Fig. 3-14

represents the spanwlse variations of acceleration measured 1n three Inch

Increments from mid-span. The data for S=+3" and S=-3W should of course be

Identical. The discrepancies shown are well within the variation of the shell

boundary conditions and un-symmetr1c positioning of the coll. The spanwlse

response delays, Indicated by the dashed line, 1s Interesting and probably

represents the longitudinal propagation of a circumferential bending wave.

The velocity 1s on the order of 65,000 Inches per second. The chordwlse

variations of acceleration, shown 1n F1g. 3-15, should again be equal for

positions either side of the nose, e.g. C=+2" and C=-2". It 1s further noted

that the 200 microfarad capacitance results produces peak accelerations whose

magnitude of 10 kilo-g's 1s approximately twice those corresponding to the 400

microfarad results. This 1s attributed to the Increased rise time for the

EIOI pulse for smaller values of capacitance, and dramatizes the structural

87
response sensitivity to changes In the EIDI electrical parameters.

Strain gages were Installed on both the Inner and outer surfaces of the

semi-cylinder at mid-span. The position and numerical designation of each

gage Is noted 1n F1g. 3-16. Strain data was established for the single coll

configuration for an energy level of 128 Joules; 400 microfarads and 800

volts. Compensated gages were Installed and the DC-bridge process was

employed for the measurements. The data 1s presented 1n F1gs. 3-17 through

3-20 for each of the four gage positions as analog records and Includes the

principal occurrence times for-strain peaks. Some data 1s further presented

as bending and membrane strains. Peak occurrence times for the chordwlse and

spanwlse peaks do not coincide* hence 1t 1s difficult to establish the maximum

stress values. The experimental strains Indicate maximum values of about 1100

micro-strain* which corresponds to maximum stresses on the order of 14,000 ps1

for the 128 Joule test condition. The energy level used In the 1RT test of

May 19, 1984 was 162 joules; 400 microfarads and 900 volts. The

corresponding stress level for this test condition 1s approximately 17,700

ps1.

An additional test was conducted on gage position (1) to establish the

spanwlse membrane strain for three different capacitance values and equal

energy levels. The data 1s displayed 1n F1g. 3-21 and clearly Indicates the

decreasing rise time of the strain response as the capacitance Is decreased.

The current delivered to the coll 1s also displayed with the membrane strain.

Case (3) of this test* evaluated for a capacitance of 100 microfarads* serves

as a measure of the maximum strain-rate response experienced for a EIOI pulse.

The peak of 1300 micro-strain occurlng at 141 microseconds represents an

approximate strain-rate of 10 per second* which could reach 40 to 50 per

second for Increased values of voltage. Most authorities* Ref. 3-5, consider

values below 100 per second as a Intermediate strain-rate regime where

quasistatic stress-strain properties are still applicable. Strain-rates 1n

excess of 1000 per second are normally associated with elastic-plastic shock

wave propagation and hence the EIDI pulse may be classified as a very mild

shock environment.

The semi-cylinder was configured with a pair of EIDI colls and a second

IRT test was conducted on August 17* 1984. The colls were connected 1n

series* placed at mid-span* and oriented at equal angles of 45 degrees

measured from the nose. The radius of each coll was 0.875 Inches and shaped

to conform to the cylindrical surface with an air gap. The established deIcing level was approximately 170 to 200 joules. The 400 microfarad and 1000

volt test condition was subjectively evaluated as optimum. The double coll

configuration failed to perform as well as the single coll and this 1s

attributed to the angular position of the colls from the nose. An angle of

less than 45 degrees or the ability to Independently pulse each coll would

88
probably have been more efficient. A limited amount of strain data was

acquired for the double coll configuration and this 1s presented 1n Fig. 3-22.

During the August 17 IRT test» some data was acquired for the double coil

configuration under 1c1ng conditions. The results confirmed earlier

observations related to acceleration in that the peak values are reduced by

approximately 20 to 30 per cent for the first pulse when ice is present on the

airfoil. Circumferential bending strains for the double coil test exhibited

reductions of approximately 15 to 25 per cent for the flrts coll pulse whereas

the memebrane strains remained basically unchanged.

89
8. Analytical Results.

The semi-cylindrical leading edge structure was chosen for analytical

study because of Its geometric simplicity and the vast literature which exists

for open cylindrical shells. Lelssa, Ref. 3-6, reports that there are no less

than 18,496 distinct problems for this shell corresponding to the numerous

combinations of boundary conditions that may be selected for the four edges.

An overwhelming majority of Investigations have been directed to but one of

these sets of boundary values and that 1s when all edges of the shell are

supported by shear diaphragms. The semi-cylinder 1s further classified as a

deep open shell requiring a consideration of the transverse shearing force

resultants and the Inclusion of tangential Inertia effects for dynamic

studies.

The boundary conditions for an open cylindrical shell supported by

shear diaphragms are exactly satisfied by choosing trigonometric displacement functions of the form:

u=A cos(\s) s1n(N0) longitudinal displacement*

v=B s1n(\s) cos(N@) transverse displacement,

w-C s1n(\s) sln(N-Q) normal displacement;

where, \s - MTTx/L R=shell radius L=shell length.

The parameters \- MTT/L and N are spatial wave frequencies with half wave

lengths defined as follows:

( L/M ) longitudinal direction,

( R777N ) circumferential direction.

A computer code was written based on these Ideal boundary conditions that

employed both the Donnell-MushtaH and Love-T1moshenko equations of motion.

The code was used In an attempt to duplicate some of the dynamic response

effects observed 1n the experimental results. The values of 'Mf and 'N' 1n

the spatial wave frequency expressions are Integers and combinations of each

will produce a corresponding discrete resonant frequency and vibratory mode

shape. The first analytic study was related to finding the maximum values of

these Integers for a selected bandwidth or cutoff frequency. The results of

this study are summarized 1n the table below for a sem1cyl1ndr1cal shell with

the following geometric shape and material properties.

L-66 Inches R=2.52 inches H=0.040 Inches (thickness).

E=10,500,000 ps1 U=0.30 Po1sson»s ratio

0.10 pounds per cubic Inch weight density.

90
M(max)

20

27

40

54

67

N(max)

7

8

9

11

12

Mode Count

95

148

265

434

623

Bandwidth Love-T1moshenko

Hertz

3,000

4,000

6,000

8,000

10,000

The calculated dynamic response of the spanwlse and chordwise acceleration

levels with a single coll placed at mid-span was performed using the

electrical parameters; 400 microfarads and 424 volts. The cutoff frequency

was selected as 10,000 Hertz and the results are displayed 1n F1g. 3-23. The

structural damping parameter used 1n this Investigation 1s noted as a 1 percent damping factor at 1,000 Hertz; thus a frequency of 4,000 Hertz would have

a corresponding damping factor of 4 per—cent. The calculated response has

several matching characteristics with the experimental response shown In Fig.

3-16, 1n spite of the fact that the boundary conditions do not match for the

two results. In addition, the temporal behavior of the 400 microfarad pulse

was synthesized from the 600 microfarad experimental result reported by

Schrag, Ref. 3-1. It 1s further observed that 155 of the potential 623 modes

were active or participated 1n the dynamic response, the remaining 468 modes

were Inactive due to the spatial symmetry of the pressure pulse. Moving the

coll a slight distance 1n both the spanwlse and chordwlse directions from the

symmetric nose and mid-length position will force a majority of the modes to

participate, however the peak response values will change only slightly as

shown In the upper portion of F1g. 3-24, even though 607 modes were active.

The peak accelerations and corresponding occurrence times from Fig. 3-24 are

noted below:

First Positive Peak: 3,619 g's at 112 mlcrosec.

First Negative Peak: -2,960 g's at 288 mlcrosec.

F1g. 3-25 1s a display of the contribution of each of these 607 active modes

to the two acceleration peaks noted above. The results are shown 1n an

ascending frequency order and the two peak response accelerations have

attained about 95 per-cent of their final value for a corresponding cutoff

frequency of approximately 4,000 Hertz. This result Is rather Important for

finite element shell studies, for 1t suggests that the approximate discrete

model need only match the modal frequencies for the reduced bandwidth to

produce a 92 percent confidence limit 1n peak response prediction. The lower

portion of Fig. 3-24 represents the acceleration response corresponding to a

91
4,000 Hertz cutoff frequency.

Two additional plots. Figs. 3-26 and 3-27 are Included to present the

analytical predicted mid-span circumferential strain response for the outer

and Inner surface and may be compared directly to the experimental results

displayed 1n F1gs. 3-17 through 3-20. The correlations agree 1n that

longitudinal strains are dominated by membrane effects, whereas bending

behavior 1s dominant In the circumferential strain response. Again, the

results only Indicate similarity to the experimental results and the peak

response occurrence times cannot be accurately matched until the boundary

conditions and forces pulse characteristics are properly represented.

An additional analytical study related to the semi-cylindrical leading

edge simulation 1s being conducted by two M.S. graduate students 1n

Aeronautical Engineering. Mr. Peter G1en 1s documenting the closed form

solution characteristics and Mr. Bryan Wilson 1s concerned with a matching

finite element representation using the MSC/NASTRAN code. Their semi-cylinder

1s Identical to the aforementioned study with the single exception of a

selected 30 Inch length. Their parallel Investigations will result 1n direct

comparisons of the acceleration and strain dynamic response using modal

expansion solutions.

The finite element model employs the popular 'QUAD41 element and the

study contains a correlation of the element size to the half wave lengths

described above. Letting 'a' and 'b1 be the longitudinal and chordwlse

dimensions of the 'QUAD4' element, the following requirements must be

satisfied.

a^ L/M and b£ R777N.

Thus, the required size of the 'QUAD4' elements 1s determined from the values

of 'M1 and 'N'» which 1n turn are dependent on the bandwidth or cutoff

frequency. These results are summarized below for a 30 Inch long semicylinder.

Bandwidth

Hertz

2,000

3,000

4,000

5,000

6,000

7,000

8,000

Love-T1moshenko

M(max)

6

9

12

15

18

21

24

N(max)

5

7

8

9

9

10

11

Mode Count

21

42

65

92

119

154

193

a=L/M

(Inches)

5.00

3.33

2.50

2.00

1.67

1.43

1.25

b=R/N

(Inches)

1.58

1.13

0.99

0.88

0.88

0.79

0.72

92
Based on these data* Mr. Wilson selected square "QUAD41 elements with a

dlrr.esnlon of 0.75 Inches. An elgensolutlon of this model exhibited a very

strong mode correlation with the closed form results which may be summarized

as follows:

Bandwidth Frequency Frequency

Hertz Mode Count Maximum Error Average Error

2,000 22 2.9 % 6 1948 Hz. 1.43 %

3,000 42 5.9 % % 2744 Hz. 2.14 %

4,000 65 11.6 % 6 3874 Hz. 3.44 %

5,000 92 21.3 % § 4556 Hz. 4.80 %

In general, the finite element model together with Its discrete boundary

conditions appears to present a slightly stlffer shell as 56 of the 65

frequencies are larger than the corresponding closed form solution values. It

1s also Interesting to note that the Love-T1meshenko equations of motion yield

the best correlation of frequencies with the finite element model and they are

clearly superior to the Donnell MushtaM shell equations. The anticipated

completion date for their final results and conclusions Is May 1985.

Mr. Dan ChHstmore, who 1s a Ph.D. candidate 1n Aeronautical Engineering

has presented a dlsseratlon proposal that would extend the work of Mr. Gien

and Mr. Wilson to the full 66 Inch length semi-cylinder with boundary

conditions selected to match the experimental model. This will hopefully

permit a more accurate correlation of experimental and analytical results.

His proposal also Includes a finite element model analysis of a representative

airfoil leading edge structure with a comparison to experimental results and

some parametric studies related to coll selection and placement for a desired

performance criteria.

93
V. Recent Developments Related To EIDI Force Measurements.

The Impulse strength of a EIDI pulse may be obtained from a basic

ballistic pendulum test. Schrag, Ref. 3-1. and others Ref. 3-3, have

published such results corresponding to various coll parameters* air gaps, and

metallic target properties. In addition* Schrag has conducted a single

experiment utilizing a special field measuring plate to establish both the

normal spatial pressure and temporal distributions. The spatial Integral of

the of the pressure distribution data yields the temporal force whose time

Integral agreed closely with the ballistic pendulum result. The results of

this experiment are summarized In F1g. 3-28.

Several different experimental procedures have been attempted to obtain a

direct measurement of the temporal behavior of a EIDI force pulse. Most of

these have given unsatisfactory results which are attributed to combinations

of magnetic effects* mass loading* and eccentricities 1n the measurement

transducer. A recent experimental Investigation has been Initiated that shows

considerable promise. The procedure Involves the longitudinal pulse

propagation characteristics of a small diameter polycarbonate non-magnetic

rod suspended as a pendulum. The basic experiment 1s summarized 1n Fig. 3-29.

The solution to the one dimensional wave equation* when applied to a elastic

rod* yields the following well known relationship.

f(0,t)=(pAc/2) V(L,t+L/c)

where: f(0,t)= force pulse at x=0,

t » time,

p = mass density of the rod*

A = cross sectional area of the rod*

L = length of the rod*

c = velocity of propagation; c = E/p *

V(L,t+l/c)= particle velocity at x=L,

L/C = time delay for the rod, (Tau).

This simple model 1s subject to the following limiting conditions:

(1) The force pulse must correspond to a uniform stress over

the cross section.

(2) Radial Inertia effects are not Included 1n the mathematical

model and thus dispersion Is neglected.

94
(3) The rod material 1s elastic and energy dissipation Is

negllgble; I.e.. no damping.

The force pulse Is approximately 1200 to 1400 microseconds 1n duration, thus

the rod should have a delay time of 600 micro-seconds or larger. The 48 Inch

length of polycarbonate rod used In the experiments has a delay time of

approximately 850 microseconds.

The experiment basically Involves the measurement of acceleration at the

free end of the rod, followed by an Integation to yield the particle velocity,

and then multiplying by the Impedance (pAc/2). A typical acceleration and

correpondlng particle velocity time response are presented 1n F1g. 3-30 for a

3 Inch diameter rod, 49.2 Inches 1n length. The velocity trace shows evidence

of dispersion; e.g., a broadlnlng of the pulse width for successive

reflections and some attenuation due to material damping. It 1s difficult to

establish the exact starting time for the force pulse due to the dispersive

effects. F1g. 3-31 1s a comparison of the force pulse developed by Schrag,

Ref. 3-1, and the measured force pulse for two different arrival times;

Tau=800 and 850 microseconds as noted.

F1g. 3-32 presents the predicted force pulse for a 2 Inch diameter polycarbonate rod 48 Inches 1n length. The upper portion of the figure represents

the measured response for a fixed value of capacitance, 600 microfarads, and

three different values of voltage; 400, 500, and 600 volts. Theoretically the

peak amplitudes should be proportional to the square of the voltage ratios.

The measured results are remarkably accurate. The lower portion of the figure

depicts the measured force pulse for three different values of capacitance

while maintaining a constant energy value. The decrease 1n both the rise time

and the first zero crossing time are entirely consistent with expectations.

Kolsky, Ref. 3-7, has presented a Fourier synthesis procedure that

employs experimental dispersion and attenuation versus frequency relations to

predict the correct form of a dispersed and attenuated pulse. The Inverse of

this technique 1s currently being Investigated to serve as a classical form of

pulse Identification problem. Preliminary dispersion data has been

established for the 2 Inch diameter polycarbonate rod by passing the

accelerometer signal through a narrow bandpass filter and noting the pulse

arrival time, Tau. The phase velocity (L/Tau) 1s thus established for the rod

corresponding to the center frequency of the filter. The phase velocity at

250 Hertz 1s approximately 33,000 Inches per second and increases to a

stationary value of approximately 55,000 Inches per second at 8,000 Hertz.

This agrees closely with a static test where the modulus of elasticity was

determined as 341,000 psi and a corresponding propagation velocity of 54,400

Inches per second. It is anticipated that this experimental Investigation

will be completed by August 1985.

95
VI. References

(3-1) R. L. Schrag and G. W. Zumwalt, "Electro-Impulse De-1c1ng:

Concept and Electrodynamlc Studies", AIM 22nd Aerospace

Sciences Meeting* Reno, NV, Jan. 9-12* 1984* Paper No.

84-0021.

(3-2) W. D. Bernhart and G. W. Zumwalt, "Electro-Impulse De-1c1ng:

Structural Dynamic Studies* Icing Tunnel Tests and

Applications"* AIAA 22nd Aerospace Sciences Meeting* Reno*

NV, Jan. 9-12, 1984, Paper No. 84-0022.

(3-3) R. M. Bowley, et. al. "Production of Short Mechanical

Impulses by Means of Eddy Currents"* IEEE Proc., V. 130,

Pi. B, No. 6, p.415-423, Nov. 1983.

(3-4) Measurements Group, Inc., Raylelgh, NC, "Strain Gages

Operate 1n 50,000 Gauss Magnetic Fields for Fusion

Research", Eps1lon1cs, Vol. II, Issue 3, Dec. 1982.

(3-5) J. A. Zukas, et. al., "Impact Dynamics"* John Wiley & Son*

New York, 1982.

(3-6) A. W. Lelssa, "Vibration of Shells", NASA SP-288, 1973.

(3-7) H. Kolsky, "The Propagation of Stress Pulses 1n V1scoelast1c

Solids", Phil. Mag., 8, 1956, pp. 693-710.

96 c -
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119

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ORIGINAL PAGE IS

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122

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123
80.000

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Impulse Strength.

124
0.050*

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POOR QUALITY

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125
0.0 SEC 5.000 0 m

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126
800.00 SEC

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127
SEC

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Figure 3-32 Variable Voltage and Capacitance

Effects (2" Rod).

128
CHAPTER 4. FABRICATION TECHNIQUES

Wichita State University supports EIDI research by fabricating coil

and coil mounts, by fabricating models for laboratory, icing tunnel, and

flight tests, and by installing instrumentation on these models.

Of these, WSU's first and most important function is the production

and mounting of high quality electro-magnetic coils in various test models.

Coil Mire Rolling

Coil wire starts with the purchase of annealed magnet strip wire

insulated with a polyimide/polyester alloy. Virgin insulated wire, as

received, measures approximately .033 x .180 inches. This is the smallest

standard cross-section wire available from our sources.* Rectangular wire

is used because it provides the largest copper cross-section for a given

coil cross-section.

However, the .033 inch thickness is too large for coil making. A

wire rolling operation reduces the wire to a thickness of .021 to .026

inches for coil winding. Rolling is done in one pass to minimize work

hardening of the wire and damage to the wire insulation. The virgin wire

is squeezed between a milling machine arbor and a ball bearing outer race

as shown in photo 4-1. Note that current practice has eliminated both the

steel feed pipe and the take-up reel. Both the pipe and the take-up reel

tended to scrape off the wire insulation. In addition, the take-up reel

work hardens the copper wire.

Photo 4.2 shows wire thickness being checked during wire rolling.

Typical finished dimensions are .024 x .195 inches (average). The word

average is used because the wire comes from the mill with non-uniform

dimensions, a condition which is worsened by our secondary rolling operation.

Coil Making

Photo 4-3 shows the first step in coil winding, which is attaching an

electrical connector to the wire. In this case, that connector is a gold

plated brass pin. In addition to serving as a connector, the pin also serves

to anchor the wire in the coil winding tool.

A custom-wire supplier has been found who may greatly simplify the process

described here.

129
ORIGINAL PAGE iS

OF POOR QUALITY

Photo 4-1

Photo 4-2

130
OWGHVAL

POOR

Photo 4-3

Photo 4-4 shows the actual winding operation. Winding is done by hand.

Photo 4-4

131
Many automated schemes of winding were tried, but hand winding has

proven to produce better quality coils in about the same amount of time.

Coils are either wound flat or curved depending on the coil installation.

Photo 4-5 shows the second electrical connector being soldered into

place. Here again, the pin serves a secondary function which is holding

the coil together during manufacture. In this case, the coil was wound

flat.

The next step is adding a layer of 9.8 oz fiberglass, as shown in

photos 4-6 and 4-7. The fiberglass serves three functions: 1) it provides

insulation; 2) it gives the coil mechanical strength and integrity; and

3) it provides impact and scuff protection for the coil.

Photo 4-5

132
ORIGINAL PAGE IS

OF POOR QUALITY

Photo 4-6

Photo 4-7

133
Photo 4-8 shows a heat lamp being used to speed epoxy/fiberglass cure

time. Care must be taken at this time, because too much heat will degrade

the insulating properties of the epoxy by making it porous from too much

"out-gassing."

Photo 4-8

As mentioned earlier, coils must often be curved to match the contours

of a leading edge. Photo 4-9 shows form blocks, which are one of three

ways currently used to produce curvature. These particular form blocks are

3

made of 40 Ib/ft polyurethane foam. Photo 4-9 shows the form blocks and

the newly finished coil. Photo 4-10 shows the form blocks pressing a coil

into shape.

OWGWAL PAGE IS'

OF POOR QUALITY

134
Photo 4-9

ORIGINAL PAGE sS

OF POOR QUALITY

Photo 4-10

The second method is to wind the coil in a curved shape, a job which is

more easily said than done.

The third method is to actually machine the coil to shape. Photo 4-11

shows this operation on a rotary die mill . This operation is only performed

for high precision laboratory coils because it is a time consuming and

difficult job. The very soft copper in the coils is nearly impossible to

machine. Many different methods have been tried but nothing seems to work

very well. After machining, the coils must be etched in concentrated nitric

acid to remove electrical shorts induced by this machining.

135
After acid etching, manufacture is completed by repeating the steps

shown in Photos 4-6 through 4-8.

Coil Mounts

Fabricating coil mounts is the next step. Photos 4-12 through 4-15 illustrate

the basic method used for making almost all coil mounts. While details may

vary from coil mount to coil mount, the basic design philosophy is constant.

Photo 4-11

136
ORIGINAL PAGE IS

OF. POOR QUALITY

Photo 4-12

Photo 4-13

137
ORiQtNAL PA9E I

OF POOR QUALIT

Photo 4-14

Photo 4-15

138
Coil mounts are made with epoxy/fiberglass and polyurethane rigid

foam . Composites are used because they do not conduct electricity and

because composite structures can be made both rigid and light. The first

question is, "Why use fiberglass and not Kevlar or graphite?" Graphite

is not used because it conducts electricity and Kevlar is not used because

it's resistance to abrasion is much inferior to fiberglass.

Fiberglass seems to have the best mix of desired properties. These

desired properties are high abrasion and impact resistance, structural

stiffness, and good electrical insulating properties.

The second question is, "Why use polyurethane rigid foam as a composite

sandwich material?" Polyurethane foam is used because it is compatible with

epoxy/fiberglass construction and because polyurethane foam can be formulated

to have good bearing strength and good impact resistance.

At this point, comments on the quality of available materials are appropriate. Regardless of manufacturer's guaranteed specifications, components

must be chosen carefully. Not all epoxies work equally well, nor are all

urethane foams equal. Specifications seem to mean very little. For epoxy,

we use Gougeon "Brothers" West System. It's specifications are unimpressive

when compared to competitive products, but it's performance is outstanding.

Our two part pour-in-place urethane foam comes from General Plastics Inc. who

also supplies our pre-cast rigid urethane blocks.

Now we will return to actual construction techniques. Photo 4-12 shows

the first stages of mount construction. This photo shows the interior surface of a LearJet Model 55-Century III leading edge. The interior surface

is treated with an epoxy release agent and then a two ply fiberglass spacer

is laid up using the leading edge as a mold. This spacer is about .03 inches

thick and appears as the long rectangular part in Photo 4-12. Spacers can

be made of sheet wax or just about anything else. The spacer's sole purpose

is to provide a .05 inch stand off distance between the interior surface of

the leading edge and the opposing coil face and an .03 inch air gap between

the coil mount and the interior surface of the leading edge. The spacer is

discarded upon final installation.

The spacer is then trimmed to the desired size and treated with an

epoxy release agent. A single ply of 9.8 oz fiberglass is laid up on the

spacer and the leading edge. This ply is the outer layer of the coil mount

and is the square part shown in Photo 4-12. The coil or coil pair is then

139
bonded to this outer layer. This layer provides abrasion protection, impact

protection and electrical insulation for the coil, and is most important.

Two more plies of fiberglass are added behind the outer layer completely

capturing and encapsulating the coil. Total ply count now stands at 3 layers.

For lighter gauge leading edges of .03 inch or less only 2 layers of

9.8 oz fiberglass are required. The LearJet leading edge requires three

plies because it is .063 inch thick and is extremely stiff. We now have

a three ply arch with the exact shape of the leading edge less the thickness

of the standoff spacer.

A three ply back plate arch is now added for extra strength. Photo 4-13

shows the spacer and the 3-ply back plate arch. The encapsulated coil appears

as a dark shadow behind the semi-transparent back plate. Note that for

lighter gauge aluminum leading edges 2 plies have the necessary strength.

The six ply double arch is allowed to cure until it is almost hard. At this

time the double arch is separated from the mold which is the leading edge and

spacer. Next, the double arch is sawed to the desired size. Photo 4-14 shows

the mount after removal and trimming to size.

Wire leads are then soldered to the coil, and the space between the two

arches is filled with 2 part foam-in-place rigid urethane foam, Photo 4-15.

3 3 In this case, 12 Ib/ft foam was used, however foam as light as 2 Ib/ft will

work for light gauge leading edges.

Electrical Leads

Care must be taken or the leads to the coil will fatigue with repeated

coil firings. Photo 4-15 shows the best method found to date of connecting

up coils; note the fiberglass strain relief which protects the two lead wires.

Placing screw terminals on coil mounts can lead to broken mounts or

broken power cable terminals. The illustrated design arrangement eliminates

exposed terminals and wires and provides double and triple insulation for all

electrical components.

Weight Estimates

Single coils have an outside diameter of about 2.5 inches and 40 turns.

The wire measure about .025 x .195 inches, and completed coils weigh about

5 ounces.

Typical coil pairs are composed of two 30 turns coils having an outside

diameter of 1.8 inches, and weighing 5.5 to 6.0 ounces.

140
Reducing coil thickness will reduce weight. A coil pair made from

.025 x .100 inch wire should weigh 2 to 10 ounces depending on size and

strength requirements. The mount in Photo 4-15 weighs about 3 ounces

after subtracting the weight of the coil and lead wires.

An absolute minumum design weight using aluminum coils would appear

to be in the 3 to 4 ounce range. A more practical design weight would be

5.0 to 7 ounces, a figure which has already been achieved.

Coil Attachment

Coil mounts are attached to models in two ways. The first way is

screw attachment to the front wing spar or to a false spar positioned

ahead of the front wing spar.

The second method is direct bonding to the leading edge interior

surface. The mount pictured in Photo 4-15 is a direct bonding type mount.

The bonding surfaces are the flat, slightly extended strips adjacent to

the table in the photograph. Bonding to the interior surface of the leading

edge skin is a difficult task because the bond lines are subject to both

peel and impact. To this point, only two adhesives have proven capable

of performing this job, Loctite 324 and Gougeon Epoxy with natural fiber

reinforcement. Of the two, Loctite 324 is easier to use. Aluminum 1100

series electro-magnetic doublers are also bonded to thin skins using

Loctite 324 or Gougeon Epoxy.

Icing Tunnel Models

Model preparation is another major task performed at WSU. WSU uses

sheet metal and composite fabrication techniques which are common to all

aircraft experimental shops. Unlike some wind tunnel test models, most

EIDI models are actual parts or pieces taken from real aircraft. Photo 4-16

shows a LearJet Model 55 wing being prepared for Icing Tunnel tests. The

wing has been sawed off to proper length and threaded steel angles have been

added to accept a wind tunnel mounting plate. The LearJet wing has quite

thick skin (.125 in) and so the angle brackets are attached to the skin.

More commonly, angle are attached to wing spars. Photo 4-17 shows the

model mounting plate. The plate bolts to the steel angles in the reinforced

LearJet wing and in turn the plate/wing assembly bolts to the wind tunnel

turntable. The plate measures 1/2 inch thick and is made of 6061 T-6 aluminum,

141
ORIGINAL PAGETffif

OF POOR QUALITY

Photo 4-16

Photo 4-17

142
Test Model Instrumentation

The third major WSU shop function is instrumenting test models.

This primarily involves the installing of strain gages, but with a sinister

twist. EIDI pulsers do terrible things to strain gages. Firstly, EIDI

coils physically knock strain gages off the surface of the leading edges.

Secondly, The coils electrically energize the gages, turning every strain

gage into a miniature high voltage transformer.

Photos 4-18 through 4-20 show partial solutions to both problems.

The first problem, physical de-bonding, is eased by installing a thin

sacrificial layer of fiberglass above the strain gage.

The second problem, electrical interference, is limited by special

wiring techniques developed by Dr. Robert Schrag.

Strain gage installation starts with metal preparation. Metal preparation steps are:

1. De-grease the metal

2. Sand the metal with 60 to 100 grit sand paper

3. De-grease the metal

4. Mark the desired strain gage location with 6H drafting pencil

5. De-grease the metal

6. Acid etch the metal

7. Neutralize the metal with ammonium hydroxide solution

8. De-grease the metal twice

9. Bond strain gages to metal surfaces using M-Bond 200 or PermaBond 200 and adhesives. Step 9 is shown in the center of Photo 4-18.

10. Solder a small gauge insulated wire to one of the solder tabs (shown

in bottom of Photo 4-18).

11. Solder a second small gauge insulated wire to the second solder tab

and double back the wire across the middle of the strain gage,

Photo 4-19. This doubling back causes the gage to self-cancel much

of the induced error signal.

12. The two leads are now twisted together to again reduce induced

signals; see Photo 4-19.

13. Finally a 4 mil layer of fiberglass cut on a 45° line is bonded over

the gage to greatly slow gage debonding, Photo 4-20. The fiberglass

is oriented to cause little or no strengthening to the metal while

at the same time retaining the gage.

143
Photo 4-18

ORIGINAL PAGE "IS*

OF POOR QUALITY

Photo 4-19

144
ORIGINAL

OF POOR QUALITY

Photo 4-20

Strain gages being used have a resistance of 350 ohms, are temperature

compensated and have a foil, grid which measures 1/8 x 1/8 inch.

145
Chapter 5. ICING TUNNEL TESTS

An essential part of developing the EIDI system was testing In the NASA

Icing Research Tunnel (IRT) at the Lewis Research Center 1n Cleveland, Ohio.

This 1s a unique facility. The test section of 2 by 3 meters permits fullscale testing of many aircraft components such as wing sections, engine Inlets

and tall sections. A1r speeds over 250 MPH and temperatures below 20°F can be

obtained. The spray system provides a range of subcooled water droplets with

median volume diameters from 10 to 20 microns and a fair range of liquid water

content values.

The first test 1n Oct.-Nov. 1982 was a feasibility demonstration. This

resulted 1n the decision to proceed with a full development project for the

EIDI system. The eight test periods 1n the IRT are briefly summarized below.

I. Oct. 25 - Nov. 5, 1982 Tests

A. Two Models Tested.

1. Beech Bonanza wing; aluminum leading edge, skin thickness

0.032 Inches; tapered; very small distance from spar to

high-light, giving a stiff leading edge.

2. Cessna 206 wing; aluminum leading edge, skin thickness

0.025 Inches; no taper or twist; over 10 Inches from spar

to high-light gave a soft, flexible leading edge. R1bs

were spaced from 9 to 15 Inches apart.

B. Test Descriptions:

1.Beech Bonanza W1ng

* Colls 1n 3 positions, 14 Inches apart, racetrack shaped

colls, two side colls 1n series at each position.

* 31 runs

* Tunnel conditions:

110 MPH

110

160

160

230

230

110

160

160

230

110

25°F

25

25

25

25

25

5

5

5

5

20

l^g/m3

2.4

0.83

1.7

10.6

1.2

1.2

0.83

1.7

0.6

2.4

15 microns

20

14

20

12

20

15

14

20

12

20

PAGE BIANK NOT RLMED PRECEDING PAGE BLANK NOT FILMED
2. Cessna 206 W1ng

* Colls 1n 3 positions between ribs, supported from beambetween-rlbs. Colls were round; side pairs of colls were

series wired. After 21st run, the center coll was

replaced by a "racetrack" shaped coll at the nose.

* 36 runs

* Tunnel conditions:

110

110

160

160

160

160

110

110

110

160

160

110

110

160

160

160

MPH 25°F

25

25

25

5

5

5

5

-15

-15

-15

29

29

29

29

15

l^g/m3

2.4

0.83

1.7

0.83

1.7

1.2

2.4

1.2

0.83

1.7

1.2

2.4

0.83

1.7

1.7

15 microns

20

14

20

14

20

15

20

15

14

20

15

20

14

20

20

In addition to the above variables, angle of attack and

capacitance were also varied.

C. Major Results and Conclusions

1. The EIDI system can de-1ce two General Aviation wings* one

stiff and one flexible 1n leading edge properties* over a

wide range of atmospheric and 1c1ng conditions.

2. The energy required for a de-1c1ng cycle was about 800

joules per foot of span.

3. High speed movies gave Insight Into the de-1c1ngphenomenon.

D. Participants

These tests were performed under W.S.U. direction with participation at the IRT by persons from Beech and Cessna Aircraft

Companies and S1mmonds-Prec1s1on. Beech and Cessna provided

wing sections for the test models* while Slmmonds designed and

fabricated the power-and-sequenclng box. This SlmmondsPredslon power box has been used 1n all IRT tests to date

with some modifications. It can supply up to 1800 volts at

capacitances from 100 to 750 micro-farads.

148
II. April 18-22, 1983 Tests

A. Two Models Tested;

1. Cessna 206 wing with 0.025 aluminum leading edge skin.

2. Cessna 206 wing with 0.040 aluminum leading edge skin.

Both were on an unswept, untapered, untwisted wing.

B. Test Descriptions (See diagrams on the next two pages).

1. Cessna 206 wings; 0.025 Inch aluminum skin.

* Colls 1n 5 bays between ribs; 5 different coil-mount

combinations.

* 31 runs

* Tunnel conditions were kept constant: 160 MPH, 15°F, 1.7

g/m3, 20 microns, 2° angle of attack, 600 uFd.

* Primarily a comparison between different coll or mount

designs.

* Variations on a few runs were:

(a) reduce capacitance to 400 uFd.

(b) "full-cycle" voltage by removing the diode clamp.

(c) Insertion of false spars to stiffen skins In 3 bays.

(d) reduced voltage to find number of Impulses vs voltage

to de-1ce.

2. Cessna 206 wing; 0.040 Inch aluminum skin.

* Colls 1n 5 bays between ribs; 5 different coil-mount

designs.

* 28 runs

* One bay had skin replaced by a graphite composite

leading edge with a copper doublet-.

* Primarily a comparison between different coll/mount

designs, two of which were different from those used for the

previous, th1n-sk1n, leading edge.

* Variations on a few runs were:

(a) reduce capacitance to 200 uFd.

(b) "full-cycle" voltage cycle.

(c) Insertion of false spars.

(d) application of de-1c1ng fluid, ICEX, on the composite

leading edge.

(e) gap between coll and skin filled with plastic material

to see structural damping due to "no gap" condition.

(f) one run at air temperature 29°F.

149
ELECTRO-IMPULSE DE-ICER PLACEMENT IN WING FOB ICING

TUNNEL TESTS, APRIL 18-22, 1983

WIND

1

72

Wfe

t

6

i

i

• ~T

13.5"

|

^T"

*

0" fAT

TT

i

7~

Vt

12

_L

J

j

ToTs

Bay

BayBay

Bay

Bay

^- TUNNEL

' CEILING

t\/Biblets

*J

C

CESSNA 206

D WING

SECTION

E

^-TUNNEL

/ FLOOR

$ .

Thin-Skin Leading Edge (0.025 in.)

Bay A. Two Side Coils, With Doublers

(Spar)*(Zero Gap)*

Bay B. Two Side Coils

(Spar) (Zero Gap)

Bay C. One Nose Coil, With Doubler

Bay D. One Nose Coil

(Zero Gap)

Bay E. Push-Pull Design

(Spar) (Zero Gap)

Thick-Skin Leading Edge (O.OfrO in.)

Bay A. Two Side Coils, With Doublers

(Spar)

Bay B. Two Side Coils

(Spar)

Bay C. Two Side Coils Off-set gpanwise

Bay D. One Nose Coil

Bay E. Composite Leading Edge Inserted

One Nose Coil of Inrerse Design

•Indicates changes which can be made without reaoring the

wing from the tunnel. A false spar will be inserted or a

semi-soft plastic will be placed between the coil and skin

to simulate the effect on skin morement of placing the coil

directly against the skin.

150
COIL DESIGNS TESTED APRIL 18-22, 1983

SIDE COILS

NOSE COIL

PUSH-PULL

TYPE

INVERSE TYPE

(COIL MOUNTED

ON THE SKIN)

INSERTABLE SPAR

NON-METALLIC COIL SUPPORT

SUPPORT BEAM

COPPER"DOUBLER"

ON SKIN

METAL SUPPORT

METAL"TARGET"

COMPOSITE SKIN

151
C. Major Results and Conclusions

1. Using 0.050 Inch aluminum doubler with the 0.025 Inch

skin decreases the required energy by 50* both for nose

and side colls. For the 0.040 Inch thick skin with side

colls, the doubler advantage was (surprisingly) over

50%.

2. With doublers, de-1c1ng energy was under 600 Joules per

foot of span.

3. One bay had two side colls off-set spanwise from each

other, at about the 1/3 span positions,1n an attempt to

twist the leading edge. This was Inferior to aligned

colls.

4. Ant1-1c1ng attempts were disappointing. Almost the same

energy 1s required to expel 1/8 Inch and 2 Inches of Ice

thicknesses.

5. The optimum number of Impulses at each coll 1s two or

three. The general pattern 1s shown below.

Voltage

\

2 3 4- S

Number of Impulses Required

D. Participants

In addition to W.S.U. personnel, Alan Mueller of Cessna

Aircraft Company* conducted the test. Cessna provided

the leading edges; the wing model from the 1982 test was

used again.

152
III. August 15-18, 1983 Tests

A. One Model Tested;

A 50 Inch span "glove" or "cuff" placed on a six-foot section

of DHC-6 Twin Otter wing. The glove was Identical 1n shape

and material to the base wing, but extended three Inches

forward of the leading edge. Skin thickness was 0.025 Inches

with leading edge radius of 2.5 Inches.

B. Test Descriptions:

* four nose colls, one each 1n four 12.5 Inch bays;

separated by ribs. Colls were skin-mounted by rivets about 3

Inches behind the stagnation line.

* 35 1c1ng runs.

* Tunnel conditions:

100 MPH

100

124

124

170

170

170

170

124

100

124

100

170

170

124

124

124

124

28°F

28

28

28

28

28

15

15

15

15

15

15

15

15

15

15

15

30

1.49/m3

2.2

1.1

2.0

0.85

1.65

0.85

1.65

2.0

2.2

2.0

2.2

1.65

0.85

1.1

0.85

1.1

2.0

= 10°

10

7

7

5

5

5

5

5

10

7

14

5

5

7

4 to 9 (varying)

7

7

* Capacitance was fixed at 400 uFd. Doublers of 0.050 Inch

thick No. 1145 aluminum were riveted opposite the colls.

* Colls were Impulsed both singly, simultaneously by wiring

adjacent-bay colls 1n series, and by series wiring of every

other bay. Series connecting was done for 2, 3 and all 4

bays.

153
C. Results and Major Conclusions.

1. When Impulsed separately* energy required for de-1c1ng

was about 600 Joules per foot of span (3 Impulses of 200

Joules each). This was reduced for series-connected

colls as follows:

Voltage Energy

Coll Connections Required (Joules/Ft.)

Single Bay 1000 600

Two Bays 1n Series 1100 363

Two Alternate Bays 1200 280 (center bay not

completely clean)

Three Bays 1n Series 1200 280

Four Bays In Series 1400 292

2. Some runs were made with continuous 1c1ng for up to 21

minutes with Impulses 3 or 6 minutes apart. De-1dng

Improved after the first Impulse sequence 1n all cases.

3. When angle of attack was varied during the run* 1ce

width was greater 1n the chord-wise direction* but 1ce

expulsion was not affected.

D. Participants

W.S.U. was joined by Robert Goehner of SlmmondsPredslon for these tests.

154
ORIGINAL ,

POOR QUALITY

LJJ

uj

O K

^ CO

<

LJJ

155
JRIGWAL PAGE IS

OF POOR QUALITY

Coil Mounts Attached to the Skin

of the Wing Glove for the DHC-6

Twin Otter

156
IV. November 7-9, 1983 Tests

A. Two Models Tested:

1. The 50 Inch Twin Otter glove (same as previous test).

2. A 34 Inch span section from a LearFan wing made of Kevlarepoxy composite. Colls were placed at the quarter and

three-quarter positions* giving 17 Inch spacing between

colls. The leading edge had no ribs. A new type of coll

mount was tested for use with composite wings. The coll

was separated from the leading edge by a fiberglasswrapped foam piece mounted on a thin aluminum plate. The

plate was opposite the coll and provided the force to the

foam wedge which transferred the Impulse to the leading

edge nose. Half of the composite surface was painted.

B. Test Descriptions

1. Twin Otter glove.

* Colls In the 4 bays were series-connected 1n two pairs

of alternating bays (l-and-3; 2-and-4) for the first

glove test series. Capacitance was kept at 400 micro

Farads.

* Only one coll was Impulse for a second series while the

electrical pulse time was varied by changing capacitance

and adding external "dummy" colls 1n series. The purpose

was to determine the optimum capacitance value.

* 8 runs for first series; 11 runs for second series.

* Tunnel Conditions:

100 MPH 28°F l^g/m3

124 28 1.1

170 28 0.85

2. LearFan Composite Wing.

* The two colls were Impulsed separately and simultaneously.

* 10 runs

* Test Conditions:

160 MPH 25°F O.SSg/m3

* Capacitance of 200* 400 and 600 mlcroFarads.

C. Major Results and Conclusions.

1. For the Twin Otter (DHC-6) glove, the use of odd-and-even

coll pairs reduced the needed Impulses from 3 to 2 for

good de-1c1ng. The energy was thus reduced to about 200

Joules/ft.

157
2. For the second series of glove tests, the effect of

capacitance was clear. The minimum energy for de-ic1ng

was obtained for 400 mlcroFarads for this wing.

3. The Kevlar composite LearFan leading edge de-1ced, but

not well. Side coll Impulses are probably needed. About

500 Joules/ft, was required to de-1ce.

4. Painted and unpalnted Kevlar surfaces de-Iced alike.

D. Participants

Mr. James Chase of LearFan took part 1n the test along

with W.S.U. personnel.

NOTE: At this point the energy required to de-1ce a general

aviation wing had been reduced dramatically, as shown

1n the table and plot on the following two pages.

158
SUMMARY OF RESULTS OF

ICING RESEARCH TUNNEL TESTS

OF

ELECTRO-IMPULSE DE-ICING

DATE WING METHOD

n = NO.IMPULSES

PER POSITION

E = ENERGY ENERGY TO

(Joules/Ft) DE-ICE = nE

Oct.1982 Beech Two Side Coils;

Bonanza Single Bay Hits.

Nov.1982 Cessna Two Side Coils;

206 Single Bay Hits.

Apr.1983 Cessna

206

Aug.1983 DHC-6

Twin

Otter

Nov.1983 DHC-6

Twin

Otter

Two Side Coils; 2

Single Bay Hits.

One Nose Coil; 2

Single Bay Hits.

Two Side Coils With 2

Skin Doublers;

Single Bay Hits.

One Nose Coil With 2

Skin Doubler;

Single Bay Hits.

One Nose Coil With

Skin Doubler;

Single Bay Hits. 3

Two-Bay Hits. 3

Two-Bay With Bay 3

Skipped Between

Three-Bay Hits 3

Four-Bay Hits 3

One Nose Coil With 2

Skin Doubler;

Two-Bay Hits, With

"Odds-then-Evens"

Sequence. (Bays 1 & 3

hit twice simultaneously,

Then Bays 2 s. 4 hit twice

simultaneously.

383

440

450

414

289

272

766

880

900

828

578

544

200

121

98

96

98

90

600

363

294

288

294

180

Note: For this method,

four impulses

clean four bays.

These are plotted on the following page, with the "Total Energy

Per Position Required To De-Ice" plotted against time when the

tests were performed.

159
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160
V. May 14-23, 1984 Tests

A. Four Models Tested;

1. Learjet win g with the standard (but modified) leading

edge; skin thickness 0.063 Inches; sweepback 17°.

2. Learjet wing with a composite material replacement leading

edge which was 50% thicker and 100X stlffer, but had the

same shape as the metal one.

3. A semi-cylinder of 5.0 Inch diameter mounted with no

sweepback; thickness 0.040 Inches.

4. A Cessna 206 horizontal tall section; skin thickness 0.025

Inches. This was tested to support the planned fullaircraft prototype program. Of the four models 1n this

test* only this one had ribs 1n the leading edge.

B. Test Descriptions:

1. Learjet; metal leading edge.

* Colls 1n 6 positions; 4 coil-mount combinations.

* 15 runs with doublers added and 9 runs without doublers.

* Tunnel conditions:

144 MPH 28°F l.Og/m3

144 15 1.0

225 15 1.4

225 -10 0.6

2. Learjet; composite leading edge.

* Colls 1n 6 positions; 5 types of coll/mountings.

* 12 runs

* Tunnel Conditions:

144 MPH 15°F l.Og/m3

144 15 1.4

144 -10 0.6

3. Semi-cylinder

* One nose coll at mid-span* spar mounted.

* 13 runs

* Tunnel conditions:

100 MPH 24°F l^g/m3

100 24 2.2

100 10 2.2

160 15 1.7

161
Cessna horizontal stabilizer

* Colls 1n 4 positions; one without doublers.

* 11 runs with the production-type rubber abrasion shield;

tunnel conditions for this were:

100 MPH 28°F 1.49/m3

124 28 1.1

160 28 0.83

160 15 0.83

* 14 runs with the rubber abrasion shield removed;

tunnel conditions for this were:

100 MPH 28°F 1.49/m3

100

100

124

124

124

160

160

15

15

28

15

15

28

15

1.4

2.2

1.1

1.1

2.0

0.83

0.83

C. Major Results and Conclusions.

1. Learjet leading edge needs re-designed coll/mount system

due to Its peculiar* unsymmetrlcal shape.

2. Doublers are still helpful for 0.050" thick skins.

3. The composite material posed no special problem. It was

only slightly more difficult to de-1ce than Its metal

counterpart.

4. Painting the composite skin made no difference 1n deicing.

5. The one coll In the semi-cylinder expelled 1ce from over

five feet of length. Good high-speed movies were obtained

for this.

6. The Cessna tall section was easily de-1ced.

D. Participants

1. Gates-Learjet supplied the wing model and participated 1n

the IRT test.

2. Cessna supplied an empennage for making the tall model and

participated 1n the test.

3. Slmmonds-Predslon sent a participant for the IRT test.

162
PAGE IS"

OF POOR QUALITY

20 micron droplet size

1.4 g/m3 LWC

5 minutes spray

700 volts/coil station

12 micron droplet size

0.6 g/m3 LWC

10 minutes spray

800 volts/coil station

Test Conditions For Both:

225 Miles/Hour Airspeed

15°F Air Temperature

4° Angle of Attack

400 MicroFarads

LEARJET ALUMINUM LEADING EDGE AFTER

DE-ICING: ILLUSTRATION OF GREATER

DIFFICULTY DE-BONDING SMALL-DROPLET

ICE

163
VI. August 13-17, 1984

A. Three Models Tested;

1. A slat from the wing of a Boeing 767; sweepback 34°; skin

thickness, 0.062 Inches. This portion of the slat was

from a wing station about 65% of span from the fuselage.

2. The Learjet wing with Its standard metal leading edge was

returned for a second IRT test.

3. The 5 Inch diameter semi-cylinder was returned with a

s1de-co1l pair at mid-span.

B. Test Descriptions;

1. Boeing 767

* Colls 1n 6 positions; 3 coll/mount types.

* 21 runs.

* Tunnel conditions

170 MPH 28°F O.SSg/m3

170 28 1.65

240 27 0.85

240 15 0.6

240 15 1.1

240 -10 0.54

240 -10 1.1

2. Learjet

* Colls 1n 4 positions; 4 coll/mount types.

* 22 runs

* Tunnel conditions:

144 MPH 28°F l.Og/m3

144 15 1.0

225 15 1.0

225 15 1.4

3. Semi-cylinder

* Two side-colls were spar-mounted 1n one mid-span

position.

* 14 runs

* Recordings were made for signals from strain gages and,

sub-miniature accelerometers mounted near the colls.

* Tunnel conditions:

100 MPH 24°F 2.2g/m3

100 15 1.4

164
C. Major Results and Conclusions.

1. The Learjet win g still defied de-1c1ng without

unacceptable residual 1ce particles.

2. The Boeing 767 slat de-1ced quite well using spar-mounted

side-pair colls.

3. Best de-1c1ng was achieved using very short electrical

pulse times for the 767 slat; this gave higher skin

stresses than desired. An optimized coll design and retest 1s needed.

4. For the sem1-cyl1nder» a nose coll 1s clearly superior to

a pair of side colls. Good high-speed movies were

obtained for the cylinder.

5. The skin-mounted coil-pair gave the best performance and

lower stresses.

D. Participants

W.S.U. was Joined by three men from McDonnell-Douglas, one

each from Boeing and Gates Learjet and two observers from

the Lockheed-Georgia Co. Boeing supplied the B-767 slat.

165
VII. Sept. 17-21, 1984

A. Two Models Tested:

1. The Learjet wing was returned for a third set of tests.

Three new coll/mount types were Installed.

2. A Falcon Fanjet engine Inlet (supplied by Rohr Industries)

was equipped with EIDI colls.

B. Test Descriptions;

1. Learjet

* 6 types of coll-and-mount configurations were tried 1n 6

positions.

* 10 runs

* Tunnel conditions:

144 MPH 28°F l.Og/m3

144 15 1.0

225 15 0.6

225 15 1.4

2. Falcon Fanjet Engine Inlet

* colls were placed 1n 8 equally-spaced positions; 6

coll/mount types.

* 29 runs

* Tunnel conditions:

110 MPH 27°F l^g/m3 = 0°

110 27 1.2 10

110 15 1.2 0

110 15 1.2 10

110 15 2.4 0

170 15 0.85 0 (with 66% blockage)

170 15 1.65 0

170 26 0.85 0

225 15 0.6 0

225 27 0.6 0

C. Major Results and Conclusions

1. A new transverse skin-mounted coll proved to be the most

effective de-1cer for both the Learjet wing and the engine

Inlet. It has the added advantage of very light weight.

The only question concerning Its usefulness 1s 1n Its

resistance to de-bonding during long-time use.

2. The Learjet wing was finally de-1ced well by either of two

co1ls» both using a single coll well off-set from the

nose.

166
The engine Inlet's curvature-Induced stiffness was no

problem for electro-Impulse de-1c1ng.

For the 36 Inch diameter engine nacelle 1nlet» six coll

positions are easily adequate.

Ice fragments were caught 1n a net and photographed for

the engine Inlet tests. These seem to be small enough to

be safely Ingested by a turbofan engine.

D. Participants

W.S.U. personnel were assisted by one Industry participant

from S1mmonds-Prec1s1on» Cessna, Boeing, and Gates

Learjet. Three participants were present from Rohr

Industries.

167
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COBRA HELICOPTER BLADE ICED

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POOR QUALITY

170
COBRA HELICOPTER BLADE AFTER DE-ICING

ORIGINAL PAGE IS

OF POOR QUALITY

171
VIII. Nov. 26 - Dec. 4 1984 Tests

A. Six Models Tested

1. Cessna 206 horizontal stabilizer modified for retrofitting

EIDI colls. A cuff extending the leading edge 3 Inches

was attached by rivets at the front spar.

2. Cessna 206 wing strut modified for retrofitting EIDI

colls. A sheet metal (aluminum) airfoil shape was wrapped

around the original elliptical-shaped strut .

3. Cessna 206 wing strut modified for new production to be

more compatible with EIDI Installation. Aluminum sheet

airfoil was supported by an Interior* load bearing* Ibeam.

4. Cessna 206 wing section* outboard * skin thickness of

0.025 Inches.

5. Cessna 206 wing section* Inboard* modified for optimum

EIDI rib spacing (18 Inches) with 0.040 Inch skin

thickness.

6. Helicopter blade, mid-span section from AH-1 Cobra; this

1s an all-composite blade with 31 Inch chord and slight

twist.

B. Test Descriptions

1. C-206 Horizontal Stabilizer; 0.032 Inch aluminum skin.

* Two coll positions between 22 and 19 Inch Mb spadngs;

One position had side-pair colls mounted on an arch

supported from the skin 3 Inches behind the nose. The

second had a composite beam version of the transverse

skin-mount (band-aid) single coll on the pressure side of

the leading edge. Doublers were used with all colls.

* 48 runs

* Tunnel conditions:

110 MPH 27°F 2.4g/m3 = 4°

110 27 1.2 4

160 27 1.7 0

160 27 0.83 0

110 10 2.4 -4

110 10 1.2 -4

160 10 0.83 0

160 10 1.70 0

160 -10 0.83 0

All of these were run with both metallized and

electrolytic capacitors except the last two, which were

done only with electrolytic.

172
* To represent two or three colls 1n series, each coll was

connected to one to two Identical colls, outside the

tunnel* which had "dummy11 aluminum skins.

2. and 3. C-206 W1ng Struts (Both types).

* Two coll positions 24 Inches apart; one position had

side-pair colls supported on a skin-mounted arch; the

second was a single "band-aid" mounted coll, one one side

only Immediately behind the nose.

* 42 runs

* Tunnel conditions

110 MPH 27°F 2.49/m3

110 27 1.2

160 27 1.7

160 27 0.83

110 10 2.4

110 10 1.2

160 10 1.7

160 10 0.83

160 -10 0.83

* Electrolytic capacitors were used (500 mlcroFarads).

4. C-206 W1ng, 0.025" skin thickness.

* A current-production leading edge was used which had 3

bays of differing span. In a 13 Inch bay* a nose coll was

mounted on an aluminum beam mounted to the ribs; In a 15

Inch bay* a nose con was mounted on a stiff* 6 Inch long*

composite bar attached (spanwise) to the skin at Its ends;

In a 17 Inch bay* a band-aid type coll mount was placed on

both upper and lower surfaces behind the nose; (these were

wired separately). Doublers were used with all colls.

* 19 runs

* Tunnel Conditions

110 MPH 27°F 2.4g/m3

110 27 1.2

160 27 1.7

160 27 0.83

* Electrolytic capacitors; 375 mlcroFarads.

5. C-206 W1ng; 0.040" skin thickness.

* Three 18 Inch bays each had Identical nose colls* but

were mounted differently. One was skin-mounted on a

composite arch* one was a band-aid type* and one was

supported by a composite rib mounted to the ribs.

Doublers were used with all colls.

* 32 runs

* Electrolytic capacitors; 375 and 500 mlcroFarads. Two

173
runs were made with metal1zed capacitors for comparison.

* Tunnel conditions were the same as listed above for the

horizontal stabilizer.

Helicopter blade

* The outer 32 Inches had an aluminum leading edge of

0.032 Inch thick aluminum; the composite underneath was

cut back to the first spar and two band-aid type coll

mounts were placed 16 Inches apart on the lower (pressure)

surface. The Inner 38 Inches of the blade had an

aluminum* leading edge stretched against the solid

composite leading edge, but bonded to 1t only aft of the

nose region. Small "racetrack" shaped colls were recessed

Into the composite material and aluminum doublers bonded

to the skin opposite them. Coll dimensions were 0.75 x

1.3 Inches, and 0.125 Inches thick. These were recessed

Into positions 8 Inches part at three lower surface

positions and 12 Inches apart at two upper surface

positions.

* 15 runs

* Tunnel conditions

230

230

230

230

230

230

230

MPH 27°F

27

27

10

10

-10

-10

l^g/m3

0.6

0.6

1.2

0.6

1.2

0.6

* 10°

10

0

5

5

5

5

* Metallized Capacitors, 400 mlcroFarads.

C. Major Results and Conclusions

1. For the horizontal stabilizer, rib spadngs of 22 Inches

are are acceptable. Colls can be connected for three

side-pairs 1n series and Impulsed three times to de-1ce

with the 375 mlcroFarads electrolytic capacitors. This

gives a total energy of 246 Joules per foot of span. The

band-aid type, similarly connected, would require 215

J/ft.

2. The wing struts de-1ced well with three Impulses using

about 200 J/ft. The band-aid mounts were somewhat

superior to side-pair types.

3. For the wing leading edges, the beam-between-rlbs and

simple band-aid coll mounts de-Iced well. The attempts to

devise a rigid band-aid mount were not successful. Energy

required was about 400 J/ft. with three Impulses.

174
4. For the helicopter blade, the outer section was used as a

comparison with similar sized wing tested previously.

This de-1ced easily, but 1s not typical of a helicopter

leading edge structure. The Inner blade section was

believed to be a viable helicopter EIDI design. Best deicing was achieved using the three lower surface minicolls 1n series with 1300 volts and three Impulses. This

gave 400 Joules per foot of span as energy requirement.

Ice shedding was good for all cases. A continuous Icing

run was made with de-1c1ng every 2 minutes for 10 minutes;

Results were very good. Attempts to use only two colls 16

Inches apart were unsuccessful. Upper surface colls used

alone or In series with lower colls were Inferior to lower

colls Impulsed alone.

D. Participants

W.S.U. test conductors were joined by Industrial

participant representatives from Cessna* SlmnondsPrec1s1on» and Kaman Aerospace Co.

IX. References

5-1. G.W. Zumwalt and A.A. Mueller, "Flight and Wind Tunnel

Tests of an Electro-Impulse De-Ic1ng System," AIAA/NASA

General Aviation Technology Conference, Hampton, VA, July

10-12, 1984. AIAA Paper No. 84-2234.

5-2. G.W. Zumwalt, "Icing Tunnel Tests of Electro-Impulse DeIdng of an Engine Inlet and High-Speed Wings," AIAA 23rd

Aerospace Sciences Meeting, Reno, Nevada, January 14-17,

1985. AIAA Paper No. 85-0466.

175
Chapter 6. FLIGHT TESTS

At the date of writing, two flight test series have been completed and

a third 1s beginning. The two completed flights were reported 1n Ref. 6-1.

I. NASA Icing Research Aircraft

During January.1994, the EIDI system was tested 1n twenty-one flights

from NASA Lewis Research Center over the Lake Erie area 1n the wellInstrumented NASA/Lewis Icing Research Aircraft, a DeHavllland DHC-6 Twin

Otter. Robert Frledberg of H.S.U. operated the EIDI Equipment on the flight.

NASA pilots were Rich Ranaudo and Robert McKnlght. The purposes were:

1. To make direct comparisons between the EIDI performance 1n the

1c1ng tunnel and 1n natural Icing conditions.

2. To explore possible electro-magnetic Interference (EMI) problems.

3. To obtain flight operation experience safely 1n natural 1c1ng with

an aircraft and crew having extensive prior experience flying 1n

1c1ng conditions.

Twenty-one flight were made from NASA/Lewis 1n Cleveland, Ohio, over the

Lake Erie area 1n January, 1984.

A. Equipment Used

A power supply-and-sequendng box was designed and built by SlmmondsPrec1s1on. While considerably lighter than that used 1n the IRT tests, this

was still a research tool rather than a production model. It provided

capacitance of 400 uFd and voltage from 800 to 1200, with automatic

sequencing of either 2 or 3 Impulses per station as fast as the capacitors

recharged, an Interval of 3 to 4 seconds.

Since the aircraft was already fitted with delclng boots, a cuff

section was constructed for the tests which would not Interfere with boot

operation. The cuff (or "glove") was a 50 Inch long section constructed

from a cannibalized DHC-6 wing, and Its leading edge extended 3 Inches

forward of the wing's leading edge. Thin fences were placed at each end to

minimize end effects on the flow. Like the main wing, the cuff section was

unswept with ribs at 12.5 Inch spanwlse spacing, and had a leading edge made

of 2024 T-3 aluminum, 0.025 Inches thick, with radius of approximately 2.5

Inches. The cuff was placed at about two-thirds semi-span position from the

fuselage on the right wing as shown 1n Figure 6-1. The cuff assembly was

fabricated and colls Installed by HSU, and tested first 1n the Icing

Research Tunnel, as described 1n 5. IV.

r RECEDING PAGE BLANK NOT FILMED

177
PNEUMATIC BOOT DE-ICER

I—MING GLOVE: 3 INCH EXTENSION

H EID1 COILS BETWEEN RIBS

Fig. 6-1 DHC-6 Twin Otter Aircraft

for E.I.D.I.

A single, contoured coll was placed at the center of each bay, mounted

on a composite plate which was riveted to the skin about 4 Inches aft of the

leading edge. The colls had 40 turns of copper ribbon wire* 0.024 x 0.190

Inches, giving a 2.5 Inch coll diameter. An aluminum doubler, 0.050 thick,

was bonded to the Inside of the skin opposite the coll, with a gap of 0.050

Inches between doubler and coll. The colls were connected 1n alternating

pairs and were Impulsed two times each (bays l-and-3 twice, then 2-and-4

twice) with 1000 volts at 400 uFd. This gave an energy of 96 joules per

Impulse per foot of span per Impulse.

Cameras are Installed 1n right side of the forward fuselage, Including

a steroscoplc system to permit determination of 1ce shapes.

B. Test Procedures

Due to the effect of Lake Erie on the atmosphere above 1t, a wide

variety of 1c1ng conditions are available during the winter months. The

pilots sought a range of temperatures and 1ce types. R1me 1ce was

encountered on 3 flights, clear glaze Ice on 8 flights, and mixed on the

remaining 10.

Indicated airspeeds were 85 to 140 knots at altitudes 3000 to 6000

feet. Ice thickness ranged from 0.1 to nearly 1.0 Inches. A1r temperatures

were -8 to -3°C (17.5 to 27°F). Generally, 1ce was collected at almost

constant speed, but on four tests the speed was varied (three decreasing and

one Increasing), with resulting angle of attack change, to obtain a wider

1ce collection on the wing.

178
ORIGINAL PAGE IS

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179
After the desired 1ce was accreted, the altitude was Increased to reach

a cloud-free atmosphere for obtaining good photographs.

C. Results

Figure 6-2 shows a sequence during de-1dng, with pictures taken at

about 4 second Intervals. For this test, conditions were: TAS 139 knots,

altitude 3075 ft, outside air temperature -5°C, LWC 038 g/«3, 1ce thickness

about 0.7 Inches.

Ice removal was very good, being (by pilots' testimony) always better

than the adjacent boots. Two de-1c1ng attempts were judged to be falr-topoor. If a very thin layer of 1ce (about 0.1 Inch) was pulsed, then only

half of the 1ce was shed. A second sequence of Impulses removed half of the

residual. Changing speed during 1ce accretion had no noticeable effect.

When 1ce was accumulated beyond about 2% chord on top or bottom, some

residual 1ce tended to be left after de-1c1ng. This suggests that for

leading edges of this 2.5 Inch radius or greater, two colls (upper and

lower) may be needed rather than the single nose coll.

No EMI problems could be found, even though all flight and data

acquisition Instruments were turned on and monitored. The sound made by the

electrical discharge could not be heard Inside the cabin 1n flight.

D. Conclusions

The system worked well and de-1ced the aircraft well and reliably.

The natural 1c1ng was easier to remove than that encountered 1n the 1c1ng

tunnel.

II. Cessna TV 206 Flights (1984)

In February-March 1984, Cessna Aircraft company Bade fifteen flight

over the Wichita, Kansas area 1n a Cessna 206 partially protected by an EIDI

system. Alan A. Mueller of Cessna's Pawnee Division was Project Engineer

and operated the EIDI system during these flights. The 206 pilot was Doug

Bassett and tanker pilot was Tom Wai 11s. Prior to this time, the C-206 wing

had two test series 1n the NASA Icing Research Tunnel with EIDI colls

Installed. A grant from the State of Kansas Advance Technology Commission

enabled W.S.U. to give more specific support than would otherwise have been

possible. The purposes were:

1. To expand experience with the EIDI system/206 vlng combination to

Include flight 1n tanker 1c1ng spray and natural Icing conditions

1n addition to the previous wind tunnel experience.

2. To bring experience In-house for both the operational and design

aspects of the system.

180
3. To Increase the visibility and credibility of the program within

the company.

Fifteen flights were made over the Wichita, Kansas, area during FebruaryMarch, 1984.

A. Equipment Used

The test vehicle was the engineering prototype Turbo-206, a single

engine, propeller-driven, high wing, six-place aircraft. Wichita State

University personnel fabricated and Installed seven, 2.25 Inch diameter

electro-Impulse colls 1n a production leading edge assembly supplied by

Cessna. The EIDI colls were supported by composite beams suspended between

ribs. Doublers and gaps were similar to those 1n the Twin Otter. The

modified leading edge was then Installed on the TU-206 1n the Cessna

experimental shop. The portion of the right wing that was to be de-1ced was

painted black to make the 1ce more visible and the rib locations were

highlighted with a yellow stripe to mark the bays; F1g. 6-3. Coll locations

were marked, and the bays Identified, with three Inch numbers. Seven bays

were equipped, having rib spadngs varying from 7 to 18 Inches, for a total

span of 8 feet. Metal tabs with 1/2 Inch wide stripes were mounted on the

wing leading edge as an aid 1n gauging the 1ce thickness. In addition to

the wing Installation, a pair of colls were Installed 1n the wing strut.

Two, 1.5 Inch diameter colls, wired 1n series, were embedded 1n foam Inside

the strut. (Note that the number 8 on the strut 1s not over the colls but

1s about 8 Inches above the colls).

The colls were connected 1n alternating pairs (1-3, 2-4, 5-7, and 6-8)

and were pulsed two or three times each. The EIDI control box and power

supply were the same as those used 1n the DHC-6 tests, above.

The tanker used was a Cessna 404 equipped with two 100 plus gallon

water tanks, a pump system, and spray bars mounted on top of the vertical

tall. About forty minutes of spray time was available on each flight, with

a spray plume about 4 feet wide. The 206 1s shown 1n formation with the

tanker during 1ce accretion 1n Figure 6-4.

Documentation was provided by camera(s) mounted 1n the blister on top

of the 206 cowl and cameras 1n the tanker and/or a separate photo chase

plane.

B. Test Procedures

After locating an altitude where the desired temperatures were

available, the tanker pilot would turn on the spray pump. Once started the

pump could not be stopped without freezing the spray system. Unable to see

the test vehicle, the tanker pilot was required to hold altitude, heading

and airspeed so that the pilot of the 206 test vehicle could maneuver behind

him and position the spray over the portion of the alrframe to be Iced.

181
Fig. 6-3 Cessna TU 206 Test Aircraft

ORIGINAL PAGE IS"

OE POOR QUALITY

Fig. 6-4 Cessna 206 with Icing Tanker

182
As the ice accumulated the airspeed decreased due to drag* and the

pitch attitude Increased. A typical flight would start at 115 to 120 KIAS

and would decrease to about 95 KIAS by the time 1/2 Inch of 1ce had built

up. This required close coordination between the two aircraft on airspeed

control.

For most of the test program, a photographer rode 1n the tanker which

required the 206 to move out of the water spray and reform on the left wing

of the tanker. Starting with flight 12* a separate photo chase plane was

use which allowed the 206 to remain In the water spray during the de-1ce

cycle. This was found to have a significant effect on the operation of the

system as discussed below.

The primary form of data acquisition was photographic. On the early

flight a video camera was mounted 1n the cowl camera housing, and a 35mm

still camera was carried 1n the tanker. The video camera allowed the flight

test engineer to monitor the Image while 1t was being recorded. Since no

processing was required* It provided Instant review of the flight on the

ground after the test. On later flights* the video camera was replaced with

a 500 frames per second movie camera and a 35mm still frame camera with a

motor drive. The video camera was used from the tanker as a back-up 1n case

the other cameras should fall to operate. On the last flights the video

camera 1n the tanker was replaced by a 16mm movie camera running at normal

speed and a 35mm still camera 1n a photo chase plane. Each media added Its

own unique contribution to the understanding of the de-1c1ng mechanism on

the electro-Impulse system.

C. Tanker Ice

Twelve flight were made with tanker 1c1ng. The photo sequence shown 1n

F1gure6-5 was taken on flight 5. This sequence Illustrates the clearing

pattern adopted for the bulk of the tests. Three pulses were sent to

alternating pairs of bays wired 1n series. Bay 6 was wired 1n series with

the coll pair 1n the strut. The two stage clearing action of fracture and

expel are clearly shown. In this test* two pulses were adequate to clear

the bays. However* on other tests three pulses were required. In all

cases* each Impulse had energy of less than 140 joules per bay.

Test conditions were varied on subsequent flights to cover a range of

temperatures and Ice thicknesses. The system was successful 1n clearing 1ce

ranging 1n thickness from l/10th of an Inch to 1n excess of one Inch. The

temperature was varied from 5° to 30°F. The most difficult 1ce to clear was

the extremely thin* warm 1ce. In this case* some residue was left on the

rib locations with approximately 90S of the area between the ribs clean.

On the flights using a separate photo chase plane the system was

activated while the airplane remained 1n the water spray. This caused a

significant degradation 1n system effectiveness. Up to three cycles of the

183
\ ,

- \ -

Fig. 6-5 De-icing Sequence of Cessna 206 Wing

with Tanker Spray Ice

JRiGINAL PAG£ iS

OF POOR QUALITY

184
system were required to clear 1ce that had cleared easily 1n one cycle on

previous tests. There were three significant variables different on these

tests. The presence of liquid water on the surface of the 1ce caused the

fractures 1n the 1ce to fill with water. This water apparently damped the

action of subsequent pulses and allowed the Ice to refreeze between cycles.

It was also estimated that the 206 was flown closer to the tanker than on

previous flights and that the water 1n the tanker was warmer than on earlier

flights. This tended to create a greater amount of runback with clear 1ce

extending 12 to 14 Inches along the lower surface of the wing. Examination

of the high speed movies revealed that Ice was being fractured and pushed

free of the wing, only to be pushed back Into Its original location by the

air pressure and held by the 1ce ridge behind 1t.

D. Natural Ice

Natural 1ce flights were limited due to safety consideration. Other

than the de-1c1ng provided by the EIDI system and propeller de-1ce, there

were no other de-1ce provisions on the airplane. We were therefore limited

to 1c1ng conditions with warm air below and clear skies above. The two

natural 1c1ng encounters were 1n thin stratus layers with a low liquid water

content (not measured). One quarter Inch of 1ce took approximately 20 to 25

minutes to accumulate and was a soft agglomerate of 1ce particles. This

resulted 1n a considerable residue after the first clearing cycle. This

residue provided seed points from which the next accretion grew and seemed

to promote a more rapid and solid 1ce growth resulting 1n more complete

clearing of the second 1ce bu1ld-up. Figure 6-6 shows the clearing cycle 1n

light natural 1c1ng.

E. EMI/RFI

There was no evidence of any electo-magnetlc or radio frequency Interference problems on any of the flights. The equipment Installed on all

flights Included two digital display NAY/COMs, ADF, RNAV, Autopilot, and

Weather radar. In addition there was a LORAN-C Installed for part of the

flights with no detrimental effect on Its operation.

F. Noise

The noise level when operated on the ground 1s quite startling but 1n

the air 1s noticeable but not disconcerting. It nay even be comforting to

know thatthe system 1s working without requiring constant monitoring. A

production aircraft would be quieter with a full Interior Installed 1n the

cabin and some soundproofing 1n the wing root.

185
ORIGINAL

IS

OF POOR QUALITY

Sequence

186
G. Conclusions

It 1s Important that the coll mounts be stiff. The original mounts

were too soft and absorbed some of the de-1c1ng energy.

System components must be properly matched. Colls wound a little small

had the compounding effect of reducing efficiency and limiting the voltage

of the system due to approaching the current capacity of the SCR's.

Ice thickness ranging from l/10th Inch to over one Inch can be shed

successfully. As 1n the tunnel* warm soft 1ce 1s the most difficult to

clear.

The presence of liquid water on the surface of the 1ce can effect the

system operation. There are probably two mechanisms at work here: (1) The

water coming out of the tanker was quite warm and did not cool to near

freezing before Impinging on the aircraft. This kept the 1ce temperature

near the freezing point. (2) Liquid water on the surface quickly fills the

cracks after the first pulse* damping subsequent pulses, and then refreezes.

The quantity and temperature of this water probably does not represent

natural 1ce. It does* however* point out the need to look at the effect of

flight In natural 1ce near the freezing point.

In general* tanker 1ce 1s more difficult to shed than natural 1ce.

Since the system was able to shed the tanker 1ce on all tests* natural 1ce

should provide no significant surprises. Reference 6-2 gives further

details of this test.

As a result of this test* Cessna Aircraft Company decided to Install a

complete prototype EIDI system 1n the C-206 for a thorough evaluation 1n

natural 1c1ng. Anticipating both new production Installations and retrofit*

several design changes were made from the wing and strut Installation tested

above. In addition* empennage Installations were designed. Changes were

made 1n wing rib spadngs and skin thickness to be more congenial to the

EIOI components* and to the strut and tall to facilitate EIDI Installation

under both field and factory conditions. All these changes were tested 1n

the IRT on Nov.-Dec. 1984 and are currently being Incorporated 1n the C-206

for flight tests to address certification Issues.

III. References

6-1 Zumwalt* G.W. and A.A. Mueller* "Flight and Wind Tunnel Tests of an

Electro-Impulse De-Icing System** AIAA/NASA General Aviation Technology

Conference* July 10-12* 1984, Hampton, VA. AIAA Paper No. 84-2234.

6-2 Mueller, A.A.* O.R. Ellis and O.C. Bassett* "Flight Evaluation of an

Electro-Impulse De-Icing System on a Light General Aviation Airplane."

AIAA/AHS Aircraft Designs* Systems and Operations Meeting* San Diego* CA.,

October 1984. AIAA Paper No. 84-2495.

187
CHAPTER 7. PLANS AND PROSPECTS

At this Juncture, the EIDI group at Wichita State University can design a

system for general aviation aircraft with considerable confidence. This

ability rests on the empirical data and experience reported here. For larger*

transport category aircraft, more testing of similar type needs to be done to

define system parameters and design optimums* but no Intrinsic difficulties

are seen. Helicopters pose special problems* which are discussed briefly 1n a

Section of the Appendix; more testing Is planned 1n the Icing Research Tunnel.

More engine Inlet de-1c1ng testing 1s scheduled.

From the outset 1t was proposed to go beyond this empirical base for the

design of Electro-Impulse De-Icing applications. The ultimate aim Is to be

able to mathematically model (1) the structural dynamics; (2) the electrodynamics and (3) the 1ce removal effects so that a system can be designed

without tunnel testing. From the start* an approximate* two-dimensional

analysis of the natural vibratory modes of the leading edge was computed as

the first step 1n the design. Then circuit computations gave an estimate of

current pulse times to match the structural mode frequencies. The computer

modeling being performed as two Ph.D. dissertations should lead to a design

method which m1m1m1zes the reliance on Insight and Intuition based on the type

of experience now residing only 1n the W.S.U. group.

After three years of rather heavy funding* the NASA support 1s expected

to diminish. The technology has been developed and demonstrated to the point

where U.S. Industries or other government agencies can continue much of the

development and provide the support for further research.

.RECEDING PAGE BLANK NOT FILMED

189
CHAPTER 8

APPENDIX

W.S.U. Participants on E.I.D.I. Project

(to March 1985)

Glen W. Zumwalt, Ph.D., Distinguished Professor

of Aeronautical Engineering* Project Director, May 1982 to March 1985.

Robert L. Schrag, Ph.D., Professor of Electrical Engineering,

Leader of Electro-Dynamic Research, May 1982 to March 1985.

Walter D. Bernhart, Ph.D., Professor of Aeronautical Engineering,

Leader of Structural Dynamic Research, May 1982 to March 1985.

Thangavel Panamas1vam, Ph.D., Ass't. Professor of Aeronautical Engineering,

Research on Composite Materials Dynamics, May 1982 to March 1983.

Robert Frledberg, M.S.A.E., Research Associate 1n Aeronautical Engineering,

Leader of Manufacturing Methods, May 1982 to March 1985.

Robert E. Henderson, M.S.E.E., Instructor of Electrical Engineering, January

1984 to March 1985.

Curtis Lug1nb1ll, Chief Technician, Dec. 1982 to Dec. 1983.

Fred Cook, Chief Technician, Jan. 1984 to Apr. 1984.

Vernon L. Hagnauer, Chief Technician, May 1984 to March 1985.

B. Keith Wilson, B.S.A.E., Research Assistant, Jan. 1984 to March 1985.

Peter G1en, B.S.A.E., Research Assistant, May 1984 to March 1985.

Ellas Bounajm, M.S.A.E., Research Assistant, Sept. 84 to March 1985.

Dan ChrlSMore, M.S.A.E., Doctoral Fellow, Aug. 1984 to March 1985.

Michael Schmltt, Student Assistant, Feb. 1984 to March 1985.

John Schwartz, Student Assistant, Feb. 1984 to March 1985.

Mondhes Yah1aou1, Student Assistant, Feb. 1985 to March 1985.

191 ['RECEDING PAGE BLANK NOT FILMED
INDUSTRIAL PARTICIPANTS

Beech Aircraft Company, 9707 E. Central, Wichita KS 67201

Harold Riesen 316/681-7944

Boeing Commercial Aircraft Co., P.O. Box 3707, Seattle, WA 98124

Douglas Cozby, M.S. 47-03 206/237-6003

Cessna Aircraft Company, P.O. Box 7704, Wichita KS 67277

Alan Mueller 316/946-6956

Electro-Delta, Inc., P.O. Box 8, White Oak TX 75693

George Massey 214/759-3942

Gates Learjet Corp., P.O. Box 7707, Wichita KS 67277

Allyn Heinrich 316/946-2881

Lear Fan Corp.(U.S.), Box 60000, Reno NV 89506

Ian Gilquist 702/972-2647

Douglas Aircraft Div. of McDonnell-Doug!as Co., 3855 Lakewood Blvd,

Long Beach CA 90846

David Blyther 213/593-8126

Rohr Industries, Inc., P.O. Box 878, Chula Vista CA 92010

Herman A. Rosenthal, M.Z. 19T, 619/691-3420

Don Nelepovitz, M.Z. 19T, 619/691-3004

Simmonds-Precision Co., Engine Systems Div., P.O. Box 310, Norwich NY 13815

Jack Chearmonte 607/335-5113

Sikorsky Aircraft Co., North Main St., Stratford CT 06601

Robert Flemming 203/386-5789

192
Prospects and Problems for Applying

EIDI to Helicopter Blades

The need for de-icing is greater for helicopters than for any other aircraft.

Only recently has a helicopter been certified by F.A.A. for flight into known

icing conditions; one "limited conditions" certification has been given Sikorsky

by the British BCAR. The small blades collect ice proportionately better than

fixed wings and can tolerate only small accumulations since power requirements

rise quickly with icing. Pneumatic boots are only beginning to be available for

rotors and electro-thermal power requirements impose a severe penalty. Thus,

electro-impulse must be examined for meeting this need.

Several design decision can be made immediately. The power-and sequencing

box would be placed in the core section of the rotating mechanism. Power for

this would be supplied through commutator rings at low voltage and nearly

continuous flow to recharge the capacitors. The small space available in the

blades will probably force the SCRs to be in the center column also. Thus,

the coils will be connected by a rather large number of wires to the central

power box, and will be pulsed in opposing blades simultaneously to maintain

symmetry.

There are several problems, and these will require some re-designing of

rotor blades is EIDI is to be used. Many rotor blades have a main spar which

is a hollow, flattened-oval, tube. Spars have been metallic but are now

beginning to be made of composite materials. It seems likely that electroimpulse can be used only with the composite spars. Forward of the spar is a

plastic (e.g., phenolic) material forming the leading edge. This is often

protected by a metal abrasion strip, a thin sheet having good abrasion resistance (usually titanium) wrapped around and bonded to the leading edge for

about 15% of the chord. Alternately, a rubber-like abrasion shield may be

needed.

Electro-impulse coils could be recessed into the plastic leading edge

material with a gap between the coil and the metal shield. A doubler would

probably be needed, bonded to the inner surface of the abrasion shield immediately over the coil to provide adequate electrical conductivity.

193
The abrasion shield must be free to move normal to its surface to expel ice,

so bonding of the shield would have to be restricted to its upper and lower

sides well downstream of the nose. Some gaps will be necessary between the

shield and its supporting composite under-structure, since the metal must be

able to move at -least 0.2mm in the nose region. Realizing that this metal

shield frequently impacts objects at high speeds (birds, hail, sand, bugs,),

maintaining adequate free movement will be a challenge. . . but it may be

possible.

Coils may be made in the familiar flat spiral shape with two opposing

coils in the upper and lower nose at the same span station, or an elongated

oval winding might be used, with only one coil at a given span location.

There seems to be adequate space for a coil to be recessed into the plastic

nose material.

A major problem is the lack of space forward of the spar for running the

needed number of insulated wires to the coils. This problem is compounded

by the space required for coils and by the practice of imbedding lead weights

in the nose at frequent span locations.

Where can space be found for the power cables to the EIDI coils? The

most obvious available space is inside the hollow spar, but this suggestion

is rejected by some rotor designers since it implies drilling holes in the

spar at each coil station to attach the power wires. Holes are undesirable

due to stress concentrations which compromise structural integrity. In

addition, it is a common practice to apply a constant small air pressure

(or vacuum) inside the hollow spar to detect cracks. Some sort of feedthrough electrical connector may be needed which will be acceptable to the

structural designers of the rotor blade.

Of course, the distributed weights of the EIDI components must be included

in the aeroelastic analysis and design of the blade. This calls for the EIDI

system to be a part of a redesigned blade; EIDI is not an add-on system for

helicopter blades.

194
1. Report No.

NASA CR-174919

2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle

Analyses and Tests for Design of an Electro-Impulse

De-icing System

5. Report Date

May 1985

6. Performing Organization Code

7. Author(s)

G.W. Zumwalt, R.L. Schrag, W.D. Bernhart,

and R.A. Friedberg

8. Performing Organization Report No.

AR 85-1

10. Work Unit No.

9. Performing Organization Name and Address

Wichita State University

College of Engineering

Wichita, Kansas

11. Contract or Grant No.

NAG 3-284

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546

13. Type of Report and Period Covered

Contractor Report

14. Sponsoring Agency Code

505-45-54

15. Supplementary Notes

Interim report. Project Manager, John J. Reinmann, Propulsion Systems Division,

NASA Lewis Research Center, Cleveland, Ohio 44135.

16. Abstract

De-icing of aircraft by using the electro-magetic impulse phenomenon is an idea

which has been proposed and demonstrated in several European countries. However,

it has not been available as a developed system due to lack of research on the

basic physical mechanisms and necessary design parameters. The de-icing is accomplished by rapidly discharging high voltage capacitors into a wire coil rigidly

supported just inside the aircraft skin. Induced eddy currents in the skin create

a repulsive force resulting in a hammer-like force which cracks, de-bonds and

expels ice on the skin surface. The promised advantages were very low energy, high

reliability of de-icing and low maintenance. In 1982, under NASA Lewis Research

Center sponsorship, Wichita State University organized a consortium of aerospace

companies to work with the University to develop the method, which was termed

Electo-Impulse De-Icing (EIDI). This interim report summarizes nearly three years

of effort and includes analytical studies and the results of testing done in the

laboratory, in the NASA Icing Research Tunnel and in flight. It is shown that the

design must be done with clear understanding of both the electro-dynamic and structural dynamic phenomena, and the interaction between them. If properly designed,

EIDI has been demonstrated to be an effective and practical ice protection system

for small aircraft, turbojet engine inlets, elements of transport aircraft, and

shows promise for use on helicopter rotor blades. Included in the report are practical techniques of fabrication of impulse coils and their mountings. The use of

EIDI with nonmetallic surface materials is also described.

17 Key Words (Suggested by Authors))

Electro-impulse de-icing

Aircraft

Ice Protection

18. Distribution Statement

19. Security Classif. (of this report)

Unclassified

20. Security Classif. (of this page)

Unclassified

21. No. of pages

196

22. Price'
National Aeronautics and

Space Administration

Lewis Research Center

Cleveland, Ohio 44135

Official Business

Penalty for Private Use $300

SECOND CLASS MAIL

ADDRESS CORRECTION REQUESTED

Postage and Fees Paid

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Space Administration

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