NASA CR-175061

IXIAQ • W 1^1A 1

National Aeronautics and

Space Administration

.tun

EVALUATION OF CAPILLARY

REINFORCED COMPOSITES

^WASA-CR-175061)EVALDftTIC N O F CflPILLARYN86-2476 0

REINFORCED COMfCSITES Contractor Report,

Sep. 1984 - Sep. 1985 (General Electric Co.)

141 p HC A07/KJF AC1 CSCX 11D Unclas

G3/24 43350

s, ,

by

Samuel W. Ciardullo and Lewis J. Stoffer

General Electric Company

Aircraft Engine Business Group

Cincinnati, Ohio 45215

Prepared for

National Aeronautics and Space Administration

NASA Lewis Research Center

Contract NAS3-24386
1. Report No.

NASA-CR-175061

Government Accession No

4. Title and Subtitle

Evaluation of Capillary Reinforced Composites

7. Author(i)

J.E. Cahlll, J.F. Halase, W.K. South, L.J. Stoffer, R.D. Zerkle

9. Performing Organization N«me *nd Address

General Electric Company

Aircraft Engine Business Group

Cincinnati, Ohio 45215

12. Sponsoring Agency Name and Addreu

National Aeronautics and Space Administration

Lewis Research Center

Road

15. Supplementery Notes

3 Recipient's Catalog No.

5. Report Date

September, 1985

6. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

11. Contract or Grant No.

NAS 3-24386

13. Type of Report and Period Covered

Contractor Report

Sept. 1984 - Sept. 1985

14. Sponsoring Agency Code

Project Manager, K. Bowles

NASA-Lewis Research Center

Cleveland, Ohio 44135

16. Abstract

Anti-icing of the inlet of jet engines is generally performed with high pressure

heated air that is directed forward from the compressor through a series of pipes

to various manifolds located near the structures to be anti-iced. From these

manifolds, the air is directed 'to all flowpath surfaces that may be susceptible

to ice formation. There the anti-icing function may be performed by either heat

conduction or film heating. Unfortunately, the prospect of utilizing lightweight, high

strength composites for Inlet structures of Jet engines has been frustrated by the

low transverse thermal conductivity of such materials.

It was the objective of this program to develop an advanced materials and design

concept for anti-icing composite structures. The concept that was evaluated used

capillary glass tubes embedded on the surface of a composite structure with heated

air ducted through the tubes. An analytical computer program was developed to predict

the anti-icing performance of such tubes and a test program was conducted to demonstrate

actual performance of this system.

Test data and analytical code results were in excellent agreement. Both

indicate feasibility of using capillary tubes for surface heating as a means for

composite engine structures to combat ice accumulation.

17. Key Words (Suggested by Author(sl)

cnpi]lar y

anti-ice

heat transfer

reinforce composites

18 Distribution Statement

Unclassified - Unlimited

19. Security Oassif. (of this report)

Unclassified

20. Security CJassif. (of this page)

Unclassified

21. No. of Pages ! 22 Price'

131 |

* For sale by the National Technical Information Service. Springfield. Vuginia 22161

NASA-C-168 (Rev 10-75)
PREFACE

The work described herein was conducted by the GENERAL ELECTRIC

AIRCRAFT ENGINE BUSINESS GROUP (GE AEBG) under Contract NAS3-24386

for the NASA LEWIS RESEARCH CENTER, Cleveland, Ohio. The NASA

Project Manager was Mr. Kenneth Bowles.

For General Electric AEBG, the overall program was performed by the

Advanced Composites Technology Development group under Mr. Stephen

C. Mitchell. Mr. Samuel W. Ciardullo was Program Manager and Mr.

Lewis J. Stoffer was the Technical Project Manager. Other key

General Electric contributors to this program were:

James E. Cahill - Specialist, Heat Transfer Lab/Windtunnel

Pauline M. Eichelberger, Specialist, Thermographic Imaging

John F. Halase - Manager, Infra-Red Inspection

William K. South - Manager, Development Heat Transfer Unit

Dr. Ronald D. Zerkle - Manager Heat Transfer Design Methods
TABLE OF CONTENTS

1.0 Summary 1

2.0 Introduction 2

3.0 Program Guidelines 4

4.0 Technical Approach 5

4.1 Thermal Model 5

4.2 Steady State One Dimensional Thermal Model 5

4.3 Steady State Two Dimensional Thermal Model 11

4.4 Experimental Test Panels 24

4.5 Physical Properties of Test Panel 31

5.0 Wind Tunnel Test Rig 45

5.1 Wind Tunnel Testing 45

5.2 Test Data Analysis 53

5.3 Heat Transfer Analysis 53

5.4 Conclusional From the SIMCAP Analysis 92

5.5 Boundary Layer Results 93

5.6 Conclusions From the Boundary Layer 94

6.0 Infrared Measurement of Surface Temperature 95

6.1 Infared Temperature Measurement Sensor 95

6.2 Measurement System Calibration . . 95

6.3 Preliminary Emmisivity Test 96

6.4 Data Acquisition 96

6.5 Data Analysis 100

6.6 Results 100

7.0 Combined 2D Wall Heat Conduction and ID Compressible

Tube Air Mathematical Model 104

7.1 Thermal Analysis of Capillary Reinforced Composites . 104

7.2 Two-Dimensional Heat Conduction Analysis 104

7.2.1 Analysis of a Simple Test Case: A Plane Wall. 106

7.2.2 Analysis of the Reference Case 108

7.2.3 Effect of Heat Transfer at the Hot Wall

Surface 110

7.2.4 Effect of a Second Layer of Tubes 110

7.2.5 Effect of Tube Spacing on Heat Conduction . . Ill

7.3 Axial Variation of TWALL 115

7.4 Comparison of Test Results with Computer Code ... . 117

8.0 Summary and Conclusions 126

9.0 Projections for Future Captube Effort 127

10.0 Proposed Follow-on Effort 131
LIST OF TABLES

Page

Table I Flow Check Data Sheet - No. I 55

Table II Flow Check Data Sheet - No. II 62

Table III Capillary Tubes W/Ragged Ends Flow Check 63
LIST OF ILLUSTRATIONS

Figure ' Page

1 Convection Heated Configuration 6

2 Hot-Air Passages for Anti Icing 7

3 SS1D Heat Transfer Model Effect of YWALL and TKWALL 9

4 SS1D Heat Transfer Model Effect of HTCO and HTCI 10

5 SS1D Heat Transfer Model Effect of Tube Diameter 12

6 Capillary Tube 13

7 SS2D Heat Transfer Model Effect of Captube Mass

Flowrate on TCAP and TWALL 15

8 SS2D Captube Heat Transfer Model Effect on Captube

Diameter and Length on Mass Flowrate 19

9 SS2D Captube Heat Transfer Model Effect of Tube

Diameter on Mach Number Distribution 20

10 SS2D Captube Heat Transfer Model Effect of Tube

Diameter on Pressure Distribution 21

11 SS2D Captube Heat Transfer Model Effect of Tube

Diameter on TUiTT and TCAp 22

12 SS2D Captube Heat Transfer Model Effect of Tube

Diameter and Tgource

 on HTCI 23

13 SS2D Captube Heat Transfer Model Effect of PSource

on MC Within Larger Tube 25

14 SS2D Captube Heat Transfer Model Effect of YWall

on MC Within Larger Tube 26

15 SS2D Captube Heat Transfer Model Effect of HTCO on

MCap Within Larger Tube 27

16 SS2D Captube Heat Transfer Model Effect of PSource

on TWall and TC Within Larger Tube 28

17 SS2D Captube Heat Transfer Model Effect of HTCO on

T

Wall

 and TCap within

 Larger Tube ' 29

18 SS2D Captube Heat Transfer Model Effect of YWall

on TWall and TC Within Larger Tube. 30

19 Final Experimental Designs 32
LIST OF ILLUSTRATIONS

/

20 .Wind Tunnel Test Panels, 90° 33

21 Wind Tunnel Test Panels, 0° 34

22 Tube Panels for Property Testing 35

23 Capillary Panel for Physical Test 36

24 Capillary Tubing Panel A Flexural Strength Test

Report 37

25 Capillary Tubing Panel B Flexural Strength Test

Report 38

26 Capillary Tubing Panels A & B Compressive Strength

Test Report ' 39

27 Capillary Tubing Panels A & B Tensile Strength

Test Report 40

28 Capillary Tubing Panels A & B Short Beam Shear

Test Report 41

29 Microphotograph of Capillary Tube Panels Showing

0.012 inch ID Tubes Above and 0.020 inch Tubes

Below 43

3"0 Cross Ply Capillary Tube Test Panel 44

31 Wind Tunnel Test Features 46

32 Wind Tunnel Test Section with Captubes at Crossflow. 47

33 Wind Tunnel Test Rig Assembly 48

34 Wind Tunnel Test Rig Horn 49

35 . Wind Tunnel Test Rig Plenum Chamber 50

36 Wind Tunnel Test Rig 51

37 Test Panel Plenum, Bottom View 52

38 Test Panel Plenum, Side View 52

39 Wind Tunnel Test Set-Up for Check Out 53
LIST OF ILLUSTRATIONS

40 Plot of Predicted and Measured Adiabatic Flow

Check Thru 204 Tubes 56

41 Exit Mach Number Vs. Source Pressure of

Adiabatic Flow Check 57

42 Exit Reynolds Number Vs. Source Pressure of

Adiabatic Flow Check . 58

43 204 Tubes Arranged in Parallel for Manifold

Adiabatic Flow Check 59

44 Flow Check Data Sheet 61

45 .F/F Vs. Press Ratio 64

46 Flow Rate Vs. Source Pressure - Small Tubes 65

47 Exit Mach Numbers Vs. Sources Pressure - Small Tubes 66

48 Tube Exit Reynolds Number Vs. Source Pressure -

Small Tubes 67

49 • Mass Flowrate Vs-. Source Pressure 68

50 Prediction of Exit Mach Number Vs. Source Pressure 69

51 Prediction of Reynolds Number at Tube Exit Vs.

Source Pressure 70

52 Prediction for Heat Transfer Test at T = 95F/S=70F. 71

53 Prediction for Heat Transfer Test at T - 120F/S = 70F 72

54 Prediction for Heat Transfer Test at T - 170F/S = 70F 73

55 Prediction for Heat -Transfer Test at T = 250F/S <=• 70F. 74

56 Prediction for Heat Transfer Test Effect of HTCO

at T = 120F, S = 70F, P = 50 psia. 75

57 Prediction for Heat Transfer Test Effect of P Vs.

T - 120F, S = 70F. 76

58 Prediction for Heat Transfer Test of Larger Tubes

with T - 120F, S = 70F for Mass Flowrate 77

59 Prediction for Heat Transfer Test of Larger Tubes

with T = 120F, S = 70F for Mach Number at Tube Exit 78
LIST OF ILLUSTRATIONS

60 Prediction for Heat Transfer Test of Larger Tubes

With T - 120F,. S = 70F for Reynolds Number at

Tube Exit 79

61 Prediction for Heat Transfer Test of Smaller Tubes

Effect of P vs. T - 120F, S = 70F. - Short Panel 80

62 Prediction for Heat Transfer Test of Smaller Tube

Effect of P Vs. T - 120F, S = 70F - Long Panel 81

63 Prediction for Heat Transfer of Small Tube

Effect of P Vs. T - 120F, S - 70F on Mass Flowrate 82

64 Prediction for Heat Transfer Test of Smaller Tube

With T - 120F, S - 70F on Mach Number at Tube Exit 83

65 Prediction for Heat Transfer Test of Smaller Tubes

With T - 120F, S - 70F on Reynolds Number at Tube

Exit 84

66 KEPS Boundary Layer Analysis for Various Mach Number

Flows in Test Duct 85

67 KEPS Boundary Layer Analysis for Heat Transfer

Coeficient Along the Duct. 86

68 KEPS Boundary Layer Analysis for Variations of Heat

Flux_and Velocity BL Thickness Along Duct Walls 87

69 KEPS Boundary layer Analysis of Heat Flux and

Velocity BL Thickness Along Duct Wall. 88

70 KEPS Boundary Layer Analysis of Heat Transfer Tests

on 2 Inch Pre-Heat Section/Counterflow Air Streams 89

71 KEPS Boundary Layer Analysis of Heat Transfer.Tests

on 2 Inch Pre-Heat Section - HTC Effect. 90

72 KEPS Boundary Layer Analysis of Heat Transfer Tests

on 2 Inch Pre-Heat Section - Q^ALL

 Effect

-

 91

73 Infrared Measurement Test Apparatus 97

74 Infrared Camera Set-Up to View Test Panel 98

75 Cross Flow Test Panel in Wind Tunnel Throat 99

76 Computer Color Enhanced Images for Test

Point 120 - showing successively narrower

temperature windows 101
LIST OF ILLUSTRATIONS

77 Colorized Images with Corresponding Temperature

Maps, Describing Two Test Points 102

78 Basic Conduction Problem 105

79 Simple Plan Wall Test Case 107

80 Reference Case: One Layer of Tubes 109

81 Control Area with Two Layers of Tubes 112

82 Effect of Tube Spacing on Dimensionless Wall

Surface Temperature 113

83 Effect of Tube Spacing on Dimensionless Wall

Surface Heat Transfer 114

84 Effect of Tube Spacing on Heat Conduction

Shape Factor 116

85 SS2D Captube Heat Transfer Model Effect of

Captube Spacing on TWALL and TCAP Distributions 117

86 Actual -Vs. Predicted Heat Transfer Tests •- Mass

Flowrate Vs. Source Pressure 119

87 Actual Vs. Predicted Heat Transfer Tests - Wall

Temperature Vs. Panel Length - Test 101A 120

88 Actual Vs. Predicted Heat Transfer Tests - Wall

Temperature Vs. Panel Length - Test 103A 121

89 Actual Vs. Predicted Heat Transfer Tests -

Wall Temperature Vs. Panel Length - Test 112 122

90 Actual Vs. Predicted Heat Transfer Tests - Wall

Temperature Vs. Panel Length - Test 114 . 123

91 Actual Vs. Predicted Heat Transfer Tests - Wall

Temperature Vs. Panel Length - Test 117 124

92 Actual Vs. Predicted heat Transfer Tests -

Wall Temperature Vs. Panel Length - Test 120. 125

93 Graphite/PMR15 Composite Anti-Iced IGV Assembly 128

94 Graphite/PMRlS Composite Anti-Iced Front Frame

Assembly 129

95 Concept for Anti-Icing IGV Flap 130
1.0 SUMMARY

The objective of this program was to advance the technology

needed for designing advanced anti-iced composite structures which

inherently have low transverse thermal conductivity. Anti-icing of

the metal components in the inlet of jet engines is usually achieved

with high pressure heated air. The air is directed forward from the

compressor through ducts to manifolds that direct the air to the

internal cavities of areas that are susceptible to ice formation.

The heat is then conducted rapidly through metal to the air flow

surface to prevent ice from forming. This program was conducted to

investigate the prospect of conducting heated air through small

capillary tubes (CAPTUBES) that are embedded in the surface of

composite inlet components. The heated air is then delivered

directly to the surface of such components rather than relying on

heat being conducted through thick composite walls that are

inherently poor conductors of heat.

An analytical computer program was developed to predict the

anti-icing performance of heated pressurized air flowing through

small glass capillary tubes embedded in a composite wall. A physical

wind tunnel test conducted on a flat test panel with capillary tubes

clustered on its surface in tight parallel array was found to display

excellent agreement with the analytical code. The heat transfer ' test

results and the analytical code results both indicate feasibility of

using capillary tubes for surface heating of composites.

Issues remaining to be evaluated for this captube anti-icing

system are:

o Long term influence of stress corrosion and contamination

of capillary tubes.

o Impact damage and possible methods of repair.

o Mechanical design, engine systems and further heat transfer

studies.

o Proof-of-concept test on a composite frame sector,

o Maintenance methods for unplugging clogged tubes.

o Engine performance benefits.
2.0 INTRODUCTION

The use of composite materials is essential in meeting the high

thrust " to-weight and low-cost goals of the new General Electric

family -of commercial and military engines. Consequently, the

application of advanced composite structures in aircraft gas turbine

engines is expanding as weight and cost reductions become more

difficult to achieve with metal designs. Over the past 15 years, the

basic feasibility of fabricating major engine structures utilizing

advanced composite materials has been demonstrated through the

fabrication and test of a number of components. Most of these

components, however, have been in the cooler portions of the engine

because of the temperature limitation of graphite/epoxy, which has

received the most attention due to its avilabability and advanced

stage of development. The use of advanced composites in the higher

(300 F to 600 F) temperature regions has been paced by the slower

emergence of the polyimide-type matrix systems.

The first series of polyimides to be investigated were of the

condensation cure type which were difficult to process.

Subsequently, reaction type systems were developed that were some

improvement, but still presented difficult processing problems. One

of the first polyimide-type matrix systems to offer both relative

ease of processing with good, consistent part quality at a

competitive cost was the NASA-Lewis Research Center-developed PMR

system. Although somewhat more limited in temperature capability

than a conventional condensation-type polyimide system, PMR15 has the

highest potential for opening up the 300 F to 550 F temperature

range to the economic application of advanced composites.

An application of high temperature composites that would exist on

many military (T700, F404, F110) and commercial (CT700, GE36, GE38)

engines is in the front frame structure. A typical front frame is a

primary structural component that supports the front end of the

engine, and controls the airflow to the fan. Consequently, the front

frame may contain the main engine mounts and/or support the forward

thrust bearing. Since the frame is the first component, it must be

capable of being anti-iced and the struts must be able to withstand

various types of foreign object damage. Unfortunately, the design

requirement for anti-icing places significant cost, weight, and

performance penalties on both the frame and engine. Increased

attention, therefore, is being focused on new anti-icing techniques

that will allow for the application of composites to anti-iced engine

structures.
Initially both electrical and air anti-icing concepts were

investigated. Unfortunately, the initial work showed large power

requirements for electrical anti-icing and the subsequent lack of

on-board power terminated the use of electrical anti-icing.

Consequently, all efforts have been focused on hot air anti-icing

concepts.

Anti-icing in jet engines is generally performed with high

pressure, heated air that is directed from the compressor through a

series of pipes to manifolds located on the exterior of the anti-iced

structures. From the manifold the air is directed to all flowpath

surfaces that may be susceptible to ice formation. Depending on the

size and shape of the surface to be heated, the anti-icing function

may be performed by either heat conduction or by film heating.

It was the objective of this program to develop an advanced

materials and design concept needed for designing advanced, anti-iced

composite structures. The concept that was evaluated uses flexible

micro-capillaries embedded in a composite structure as a hot air

distribution system for anti-icing. This method has the potential of

providing an efficient, anti-icing technique that could be better

than present approaches. The success of this program has helped

remove a major barrier in the application of lightweight, low cost,

composite designs to anti-iced structures.
3.0 PROGRAM GUIDELINES

The program to accomplish the objectives of this capillary

heating evaluation was divided into three tasks as shown in the

following Flow Chart.

FLOW CHART

Evaluation of Capillary Reinforced Composites

Task I - Analysis and Preliminary Testing

A. Numerical Analysis of Heat

Transfer Characteristics

*.

B. Preliminary Test Program

Task II - Fabrication and Testing

A. Experimental Design

1. Recommendation of composite

configuration

2. Thermal Analysis

3. Stress Analysis

4. Final recommendation of

specimens and tests*

B. Fabrication and Testing

C. Fabrication of Deliverables

I

Task III - Reporting

*NASA Project Manager Approval
4.0 TECHNICAL APPROACH

4.1 Thermal Model

The concept for this analysis is illustrated in Figure 1. Air at

elevated pressure and temperature would enter and flow through very

small tubes located near the surface of a composite wall exposed to

very cold moist air. Heat would flow from the hot air to the cold

air, causing the wall outside surface temperature to be warmer than

if the tubes were not present. The tubes would then exhaust to low

pressure surroundings.

It was of interest to determine if available capillary-size

tubing (0.0126 and 0.0209 inch inside diameter) could provide

sufficient energy flow at typical pressures and temperatures to

achieve effective anti-icing protection. Therefore, an objective of

this analysis was to derive mathematical models which relate

performance characteristics to design parameters. The models could

then be used to examine the feasibility of this anti-icing concept

and to guide an experimental validation program.

4.2 Steady-State One-Dimensional Thermal Model

Consider the cross-section of a tube assembly shown in Figure

2(a). and the cross-section of a single capillary tube' shown in

Figure 2(b). For the initial study, it was assumed that no . spacing

exists between the tubes, and all of the heat flows directly from a

tube to the nearest wall outside surface.

The local stagnation temperature of the hot air in the tube is

TCAP, and the temperature of the cold air is TSINK. Heat flows from

the hot gas to the tube inside surface having a heat transfer

coefficient HTCI. It is then conducted through the tube wall having

thermal conductivity TKTUBE. A protective layer of material with

thickness YWALL and thermal conductivity TKWALL is located between

the captube and the wall outside surface having a heat transfer

coefficient HTCO.

A circuit representation for the SS1D model is shown in

2(c). The heat flow per unit tube length, q', is related to

individual thermal resistances in the circuit as follows:

q' - (TCAP-TTI)/RI (1)

(TTI-TTO)/RTUBE (2)

(TTO-TWALD/RWALL (3)

(TWALL-TSINK)/RO (4)
Coating

Capillary

Tube

Sturctural

Laminates

Hot Air Plenum

Figure 1. Convection Heated Configuration.
HTCO Wall

rn

 A

I I Diameter

^-L- I

^Wall

a

/

Adiabatic

(b)

•RO

RI

Sink

T

^all

TTO

T

Cap

CCap

(c)

Figure 2. Hot-Air Passages for Anti-Icing.
where

RI - 2/(HTCI*PI*DI) (5)

RTUBE = ALOG(DO/DI)/PI*TKTUBE) (6)

RWALL - [YWALL+DO(l-PI/4)/2]/(DO*TKWALL) (7)

RO - 1/(HTCO*DO) (8)

RTOT - RI + RTUBE + RWALL + RO ' (9)

Equations (1-4) can be solved simultaneously to eliminate the

intermediate temperatures TTI, TTO and TWALL. The resulting

expression for the heat loss from the captube per unit length as a

function of the model parameters is

q' = (TCAP-TSINK)/RTOT (10)

Equation (10) can be combined with Equation (4) and rearranged to

obtain the following relation for the wall outside surface

temperature as a function of TCAP and the model parameters:

CTWALL-TSINK) RO

(TCAP-TSINK - RTOT (11)

The model parameters are TSINK, TKWALL, WALL, HTCI, HTCO, DI,

DO, AND TKTUBE. TKTUBE is assumed to have a value of 1

(Btu/hr-ft-F), corresponding to glass. Two tube sizes were available

for testing. The smaller tube had an insider diameter of 0.0126

inch, and the larger tube had an inside diameter of 0.0209 inch.

Both tubes had a wall thickness of 0.004 inch. HTCI was expected to

have a value in the range from 100 to 500 (Btu/hr-sq. ft-F), and HTCO

from 50 to 150 (Btu/hr-sq. ft-F). TKWALL was expected to have a

value in the range from 0.2 to 0.5 (Btu/hr-ft-F).

The effect of YWALL and TKWALL on TWALL is shown in Figure 3.

Values of model parameters listed at the top of Figure 3 were used

with Equation (11) to obtain these results.

An illustration of the use of Figure 3 is as follows. Typical

temperatures for the T700 engine at a "worst case" are TSINK = 4F and

TCAP - 152F. If the desired value of TWALL is 40°F in order to

provide anti-icing protection, then

TWALL-TSINK 40-(-4) 44

= - --- 0.282

TCAP-TSINK 152-(-4) 156

From Figure 3 at the above dimensionless temperature level, if

TKWALL = 0.2 (BTU/hr-ft-F), then YWALL = 0.050 inch. That is, if

YWALL were less than 0.050 inch, then TWALL would be warmer than

40°F. Also, if YWALL were fixed at 0.050 inch, then TWALL would be

warmer than 40°F if TKWALL were greater than 0.2 (Btu/hr-ft-F).
•rl

co

ex

CB

O

•H

H

I

0.9

0.8

0.1

HTCI = 300. (Btu/hour-f t2

T° t

HTCO = 100.

ID « 0.0209 inch\ ,

OD = 0.0289 inch}

 Large Tube

1. (Btu/hour-ft-° F)

TKWall = 10*° (Btu/hour

~

f

t

I I !

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14

Wall'

Figure 3. SS1D Heat Transfer Model Effect of Y .... and TK^ ..
•H

OT

n.

td

10

E-i

=0.2 (Btu/hour-ft-° F)

0.050 (inch)

Wall

ID = 0.0209 (inch)

1.0 (Btu/hour-ft-° F)

HTCI = 100.—f

(Btu/hour-f t2-

0

 F)

0.1

50 75 100 125 150 175

HTCO, Btu/hour-ft 2-

0

 F

Figure 4. SS1D Heat Transfer Model Effect of HTCO and HTCI.

10
The effects of heat transfer coefficients, HTCO and HTCI on

dimensionless wall outside surface temperature, are shown in Figure

4. It can be seen that TWALL is very sensitive to HTCO, and

relatively insensitive to HTCI. This is because the variation of

HTCO occurs at the location of TWALL, whereas the wall and tube

resistances exist between the locations of HTCI and TWALL, and tend

to damp the influence of HTCI on TWALL.

The effect of tube diameter on dimensionless wall outside surface

temperature is plotted in Figure 5. The SS1D model indicates that

the dimensionless wall temperature is insensitive to tube diameter.

However, this result can be misleading, because tube diameter will

have an important effect upon the mass flowrate of hot air through

the captube, and the temperature distribution of the hot air in the

captube, TCAP, will decrease significantly with decreased tube

diameter. This effect is demonstrated by the two-dimensional model

described in the following section.

4.3 Steady-State Two-Dimensional Thermal Model

The temperature of the hot air will decrease as it flows through

the capillary tube and loses heat through the wall material to the

cold air. Equation (11) indicates that the wall outside surface

temperature, TWALL, is proportional to the local hot air temperature,

TCAP. Therefore, the wall surface temperature will also become

colder in the direction of the captube flow, and. can lose its

anti-icing ability.

The energy carrying capacity of the hot air is proportional to

its mass flowrate through the captube. Thus, mass flowrate has a

strong influence upon temperature decrease along the captube, and the

relationship between these quantities can be derived as follows.

Consider the side view of the capillary tube imbedded in a

composite wall as represented in FiEure 6. Hot air at pressure

PSOURCE and temperature TSOURCE enters the captube and flows at mass

flowrate MCAP in the axial direction z with origin at the tube

entrance. The local

stagnation temperature of the hot air in the tube is TCAP. Heat at

the local rate q' is transferred from the captube through the wall

layer to the cold air at pressure PSINK and temperature TSINK. It is

assumed here that q' is described by Equation (10), and the local

wall surface temperature TWALL is related to TCAP and the model

parameters by Equation (11). If conservation of energy is applied to

a differential control volume of length dz within the captube, then

the heat transfer rate from the control volume is equal to a decrease

of stagnation enthalpy. That is,

q' dz = -MCAP*CP*d(TCAP) (12)

11
1.0,

0.9

0.8

•n _ -

en 0.6

o.

CO

u

0.5

w 0.4

0.3

0.2

0.1

HTCI

HTCO

300. (Btu/hour-f t 2-

0

 F)

100.

=0. 2 (Btu/hour-ft-° F)

= 1.0

Larger Tube, ID = 0.0209 (inch)

Smaller

Tube,

ID = 0.0126 (inch)

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14

YWall » inch

Figure 5. SS1D Heat Transfer Model Effect of Tube Diameter.

12
T

Sink'

 PSink

HTCO

Wall

Diameter

Figure 6. Capillary Tube.

13
where CP is the specific heat of air at constant pressure.

Substituting Equation (10) for q' into Equation (12) and separating

variables gives

TCAP ,z

d(TCAP) - - / dz (13)

(TCAP-TSINK) / (MCAP*CP*RTOT)

TSOURCE -/ o

If MCAP, CP and RTOT are constant along the captube length, then

Equation (13) can be integrated to give

(TCAP-TSINIO = EXP [-Z/(MCAP*CP*RTOT)] . (14)

(TSOURCE-TSINK)

and Equation (11) can be rearranged to the following form:

(TWALL-TSINK1) = (TCAP-TSINK") * (RO/RTOT) (15)

(TSOURCE-TSINK) (TSOURCE-TSINK)

Equations (14) and (15) relate TCAP and TWALL to z, MCAP and the

model parameters including TSOURCE, • TSINK, YWALL, TKWALL, HTCI, HTCO,

DI, DO, TKTUBE, and CP.

The effect of MCAP on TCAP and TWALL distributions along the

captube are shown in Figure 7. Values of model parameters listed at

the bottom of Figure 7 were used with Equations (14) and (15) to

obtain these results. Figure 7 shows that the decrease of TWALL

along the captube directions can be large when the captube mass

flowrate is less than one Ibm/hr, and the decrease of TWALL is

relatively small when MCAP is greater than one Ibm/hr.

Equations (14) and (15) can be useful in evaluating the

anti-icing thermal performance of capillary tubes. However, they are

not complete because MCAP is also a function of design parameters

including the source and sink pressures and the captube length.

Therfore it is necessary to develop a mathematical model which

relates MCAP to design parameters.

-In general, air flow in a capillary tube must be treated as a

compressible flow. The length to diameter ratio of a one foot long

tube with an inside diameter of 0.0209 inch is 574. Thus, viscous

friction has an important effect upon the flow. That is, an inlet to

exhaust pressure ratio of five or more is required in order to

achieve significant mass flowrate through the tube. Since mass

density for air is proportional to pressure then density can also

14
1

CO

0)

o

o

co

•H

CO

•S

co

0)

U

l-i

O

CO

•H

CO

1.0

0.9

0.8

0.7

0.6

0.5

0.4

^^js<r<r<=*S.V"f- •*• lcl

p

i.

v ^^ "^"••*«. P"TM*TM*""

X x I xx

\_i !^

N x

.\ y

\

\

\

\

V

CP = 0.24 (Btu/lbm-0 F)

YWall = °'050 (inch)

TKWall =0. 2 (Btu/hour-ft-° F)

HTCI = 300. (Btu/hour-ft 2-

0

 F)

HTCO = 100.

=1. 0 (Btu/hour-ft-° F)

5.0 (Ibm/hour)

3.0

MCap =

5.0 (Ibm/hour)

3.0

1.0

0.5

0.3

Figure 7. SS2D Heat Transfer Model Effect of Captube Mass Flowrate

°

n TCap and TWall.

15
decrease by a factor of five or more. Since local air speed in the

tube is inversely proportional to mass density, the air speed can

also accelerate by a factor of five or more along the tube length,

and can approach the local acoustic speed of the air. In fact,

compressible flow theory indicates that the air speed can accelerate

along a constant diameter tube only until it reaches the limiting

value of its acoustic speed at the tube exit; that is, until the Mach

number is one at the exit. This compressible flow phenomenon is

called "choking".

If the tube inlet pressure at a choked flow condition is

increased, then the mass flowrate will also increase, however, the

Mach number at the exit will remain equal to one, and the pressure in

the exit plane will rise above the exhaust pressure. Thus, the

choked mass flowrate in a tube becomes independent of its exhaust

pressure.

Heat transfer can also affect compressible flow in a tube. For

example, if the air is heated, then its temperature will increase,

its density will decrease, and the local air speed will tend to

increase. Thus, compressible flow in a tube is very complex and, in

general, it is necessary to numerically integrate a coupled set of

differential equations satisfying conservation of mass, momentum and

energy as well as the equation of state for the gas.

For flow in a constant diameter tube with heat transfer and

viscous friction, the general equations reduce to the following forms

convenient for calculation. The equation describing the variation of

Mach number is

dM - M(1+0.5(GAM-1')MSQ') f Q+GAM*MSQ) dTO

dz 2(1-MSQ) TO dz

+ GAM*MSQ*4*FF/DI] (16)

where MSQ is M squared, GAM is the specific heat ratio for the air,

TO is the stagnation temperature of the air, and FF is the friction

factor for flow in a tube. Therefore, the last term on the RHS of

Equation (16) describes the effect of viscous friction, and the next

to last term includes the effect of heat transfer. That is, the

variation of stagnation temperature is related to the local heat

transfer rate by

dTO - - a'

dz MCAP*CP (17)

16
The third and last differential equation needed for numerical

solution is

dPO - - PO*GAM*MSO fl dTO + 4*FF/DI]

dz 2 TO dz (18)

Where PO is the local stagnation pressure of the air in the captube.

After M, TO, and PO have been determined, then all remaining

properties of the flow can be determined from algebraic equations .

Several additional relations are required in order to integrate

Equations (16-18). Correlations for friction factor, FF, and tube

inside heat transfer coefficient, HTCI, were taken from the

literature; and a computer subroutine containing thermodynamic and

thermophysical properties of ait as a function of temperature was

constructed. Equations (5-10) were used to evaluate the local heat

transfer rate, q', appearing in Equation (17); however, the

temperature of the captube air was taken as its recovery (adiabatic

wall) temperature which is related to stagnation temperature and Mach

number as follows :

TCAP = TO*

[l+(GAM-l)MSQ/2] (19)

.Note that TCAP - 687. 17R for TO = 700R, GAM -1.4, and M - 1. Thus,

the recovery effect becomes significant at high Mach number,

especially in devices requiring small variation of wall surface

temperature .

A computer program called SIMCAP was developed to simulate the

complete steady-state two-dimensional captube model. Equations

(16-18) are integrated using the Euler predictor-corrector method

beginning at the tube inlet where PO = PSOURCE and TO = TSOURCE.

However, the inlet Mach number is unknown at the beginning of

calculation. Therefore, SIMCAP uses the bisection method to iterate

for the inlet Mach number which satisfies both the inlet and outlet

conditions . That is , the lower and upper bounds for inlet Mach

number are known to be zero and one. The inlet Mach number for each

iteration is assumed to be the average of the upper and lower bounds

(e.g., 0.5 for the first iteration), and integration proceeds from

the tube inlet. If the local Mach number reaches a value of one

upstream of the tube exit, or if the local pressure falls below

PSINK, then the assumed inlet Mach number is too large and it becomes

the upper bound for the next iteration. Alternately, if the

integration reaches the tube exit with a calculated exit Mach number

less than one, but the calculated exit pressure is greater than

PSINK, then the assumed inlet Mach number is too small and it becomes

the lower bound for the next iteration.

17
The iteration procedure in SIMCAP is converged when the exit Mach

number is acceptably close to a value of one or, if less than one,

the calculated exit pressure must be acceptably close to the value of

PSINK. The process typically requires less than twenty iterations to

converge to an acceptable solution. After convergence all quantities

of interest, including MCAP and TWALL, are written to output files

for printing or plotting.

The steady-state two-dimensional captube model contained in the

SIMCAP program includes the following input parameters: PSOURCE,

TSOURCE, PSINK, TSINK, HTCO, TKWALL, WALL, DI, DO, and tube length.

The program has been used to investigate the influence of several of

the-se parameters upon captube performance.

For example, the effects of captube diameter and length upon mass

flowrate are shown in Figure 8. The parametric values listed at the

top of Figure 8 were used for the calculations. The results shown

that values of MCAP are generally less than one Ibm/hr for the

smaller tube, and greater than one Ibm/hr for the larger tube over a

wide range of tube lengths. This means that the larger tube will

have relatively small variations in wall surface temperature as

illustrated at the higher mass flowrates shown in Figure 7.

SIMCAP results showing the effect of diameter on a captube having

a length of one foot are plotted in Figures 9-12. Mach number

distributions in the smaller and larger tubes are shown in Figure 9.

Note that the larger tube has a larger inlet Mach number. Also, the

exit Mach numbers are equal to one (both tubes are choked), and most

of the Mach number increase occurs in the last 20% of the tube

length.

The corresponding pressure distributions are plotted in Figure

10. Note that the exit pressure for these choked flows are 18.1 and

25.5 psia, both greater than the specified exhaust pressure of PSINK

- 15 psia.

The corresponding captube air and wall surface temperature

distributions are plotted in Figure 11. Note that the decrease in

TWALL from tube inlet to exit is twice as great for the smaller tube

as for the larger tube. Also.note that the temperature distributions

in Figure 11 curve downward near the tube exit, whereas the

temperature distributions in Figure 7 have no such curvature. This

is because the SIMCAP results plotted in Figure 11 include the

effects of recovery temperature, variable air properties, and

variable HTCI.

The effects of tube diameter and captube inlet temperature on

HTCI are illustrated in Figure 12. Note that the larger tube has a

value of HTCI about 20% greater than for the smaller tube, because of

its greater mass flowrate. However, Figure 4 shows that this level

of HTCI variation has a small influence upon wall temperature.

18
3

O

f

e

XI

CO

u

S

(U

*J

td

M

00

V

•s4J

s- u

4.5

3.5

0.5

Source

T » Sink

p

' Source

P

Sink

 =

HTCO

Wall

- 300. (° F)

0. (° F)

= 150. (psia)

15. (psia)

100. (Btu/hooir-ft 2-

0

 F)

=0. 2 (Btu/hour-ft-° F)

0.050 (inch)

Smaller Tube , ID = 0.0126 (inch)

0.6 0.8

Tube Length, ft

Figure 8. SS2D Captube Heat Transfer Model Effect of Captube Diameter and

Length on Mass Flowrate.

19
1.0

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

T0 = 300.

Source

Sink

>

'Source

o F)

0. (° F)

= 150. (psia)

15. (psia)

Sink

HTCO = 100. (Btu/hour-ft"-0 F)

'"Stell

 = °'2

 (Btu/

hour

-

ft

-°

 F>

Yt, ni = 0.050 (inch)

Wall

ID = 0.0209 (inch)

OD = 0.0289 (inch)

Smaller Tube

ID = 0.0126 (inch)

OD = 0.0206 (inch)

0.2 0.4 0.6 0.8

Z, ft

1.0 1.2 1.4

Figure 9. SS2D Captube Heat Transfer Model Effect of Tube Diameter on

Mach Number Distribution.

20
W

o.

150

135

120

105

300. ° F)

0. (° F)

150 (psia)

15. (psia)

Sink

HTCO = 100. (Btu/hour-ft~-

= 0.2 (Btu/hour-ft-° F)

0.050 (inch)

Larger Tube

Smaller Tube

Figure 10. SS2D Captube Heat Transfer Model Effect of Tube Diameter

on Pressure Distribution.

21
CO

o

H

•a

500

450

400

350

300

250

200

150

100

50

"^-^

X

^5=

^^

— *. N XX]

- -

-

^^

^^

\

•

-- — —.

T

Soui

T

Sink

p

Sour

P

Sink

HTCO

"Stal

YWall

T

Cap

^^ «_,...

^««»

>.

^,

TWall

•- —

•

.ce = 300. (• F)

= 0 (°F)

ce = 150 (psia)

= 15. (psia)

= 100. (Btu/hour-ft2-°F)

l

 = 0.2 (Btu/hour-f t-° F) '

= 0.050 (inch)

"*1*^^

">

"--,

" >

^

Larger

Smaller

Larger

Smaller

Tube

Tube

lube

Tube

0.2 0.4 0.6 0.8

Z, ft

1.2 1.4

Figure 11. SS2D Captube Heat Transfer Model Effect of Tube Diameter

on TTT n n and T.,

Wall Cap

22
600

400

Larger Tube

Smaller Tube

400. (°F)

= 0. <" F)

150. (psia)

15. (psia)

Sink

HTCO = 100. (Btu/hour-ft"-° F)

0.2 (Btu/hour-ft-° F)

0 0.2 0.4 0.6 0.8 1.2 1.4

Figure 12. SS2D Captube Heat Transfer Model Effect of

Tube Diameter and Tc on HTCI.

Source

23
SIMCAP results showing the effects of PSOURCE, YWALL and HTCO on

MCAP within the larger tube for a range of tube lengths are plotted

in Figures 13. 14 and 15. These results show that PSOURCE has a big

effect upon MCAP, as expected; however, YWALL and HTCO have a

negligible effect upon MCAP.

Results showing the effects of PSOURCE, YWALL, and HTCO upon

TWALL and TCAP distributions along the larger tube having a length of

one foot are plotted in Figures 16. 17. and 18. These results show

that YWALL and HTCO have a big effect upon the value of TWALL near

the tube inlet, and a small effect upon the slope of the TWALL

distribution. Conversely, PSOURCE has a negligible effect upon the

value of TWALL at the tube inlet, and a significant influence upon

the slope of the TWALL distribution.

The effects of PSOURCE, YWALL, and HTCO can be understood by

considering Equations (14) and (15). That is, the primary effect of

PSOURCE is to change MCAP which appears only in the argument of the

exponential function in Equation (14). However, the primary effects

of HTCO and YWALL are to change the thermal resistance terms (RO and

RTOT) which appear on the RHS of Equation (15) and influence the z =•

o intercept of the TWALL distribution.

This parametric study has demonstrated the capability and

usefulness of the SIMCAP computer program. It has also demonstrated

that Equations (14) and (15) can assist in interpreting the influence

of design parameters upon captube performance.

The SIMCAP computer program developed here includes the following

ten design parameters which are defined above in the text of the

derivation: PSOURCE, TSOURCE, PSINK, TSINK, HTCO, YWALL, TKWALL,

TKTUBE, tube diameter, and tube length. Negligible spacing between

adjacent captubes is assumed; however, the model can be easily

extended to include tube spacing. It is concluded here that the

anti-icing ability of the larger tube is feasible, and its

performance is much better than the smaller tube.

4.4 Experimental Test Panels

Five 2" x 2" x .05" graphite/epoxy panels with parallel clusters

of capillary glass fibers bonded to one full surface were prepared

with the following adhesive systems in order to select the most

accomodating system for further evaluations.

1. HT-424

2. • HT-424 Premeir

3. Metalbond 328

4. FM-123-2

5. 3109 (3M)

24
M

3

O

JS

p.

(0

CJ

S

01

4J

ffl

0}

CO

T0 - 300.

Source

0. (° F)

15. <psia)

Sink

HTCO = 100. (Btu/hour-ft"-° F)

=0.2 (Btu/hour-ft-° F)

(inch)

'0.0209 (inch)

Source

200. (psia)

1.6

0.5

0 0.2 0.4 0.6 0.8 1

Tube Length, ft

Figure 13. SS2D Captube Heat Transfer Model Effect of P

Within Larger Tube.

_ on MSource Cap

25
3

O

a.

td

I

f*

CO

CO

4.5

3.5

300. (° F)

0. (° F)

= 150. -(psia)

Sink

HTCO = 100. (Btu/hour-ft -° F)

0.2 (Btu/hour-ft-0 F)

0.0209 (inch)

Wall

0.0

0.025

0.050 (inch)

0.4 0.6 0.8

Tube Length, ft

Figure 14. SS2D Captube Heat Transfer Model Effect of Y on M

Within Larger Tube.

 Wail Cap

26
u

J3

O

Cu

0

S2

CO

CO

5

4.5

4

3.5

' 3

2.5

2

1.5

1

0.5

0

\

\

1

\

X

-

x^

1 i i

TT s inn f° v\

Source

 300' ( F)

TSink

Sour

PSink

TTf

^al

~ Wall

ID = (

^

= 0. (° F)

ce - 150. (psia)

^ = 0.2 (Btu/hour-ft-° F)

= 0.050 (inch)

3.0209 (ii

^ .

nch)

HTCO =

/ 150. (Bt

1 ^o'

 f

.u/hourt

2-8 F)

0.2 0.4 0.6 0.8

Tube Length, ft

1.2 1.4

Figure 15. SS2D Captube Heat Transfer Model Effect of HTCO on M Within

Larger Tube.

 ap

27
500

450

400

a

CO

u

H

•a

CO

350

300

250

200

150

100

T

Source =

 30°- <° F>

T

Sink

 S °« <° F)

P

Sink

 e 15' (Psia>

HTCO = 100. (Btu/hour-ft -° F)

Jail

 = °'2

 (Btu/hour-ft-0 F)

Wall

0.050 (inch)

ID = 0.0209 (inch)

T

c-,.

 Cap

<•* '^^•C""1' —^. p ^•^. ^^.^

 c*>—*.^. Source =

^^••^ ^""^-*.^% 200. (psia)

*

Source

(psia

150.

100.

200. (psia)

150.

100.

0.2 0.4 0.6 0.8

Z, ft

1.2 1.4

Figure 16. SS2D Captube Heat Transfer Model Effect of P. on T ni and

_ »,... „ m i_ Source Wall

T Within Larger Tube.

28
ra

u

500

450

400

350

300

250

E-T 200

150

100

50

Source

300. (° F)

TSink = °- (°'F)

PSource ' 15°' ("sia)

PSink = 15' (

"

sia)

0.2 (Btu/hour-Ft-° F)

0.050 (inch)

Wall

ID = 0.0209 (inch)

Cap

HTCO

»^. 50. (Btu/hour-ft -" F)

*"" *•*!> i nn

Wall

^100.

150.

50. (Btu/hour-ft -° F)

1.2 1.4

Figure 17. SS2D Captube Heat Transfer Model Effect of HTCO on T .... and

TCap Within

 Larger Tube.

29
500

450

400

350

300

Q.

CO

•o

S

250

200

150

100

50

Cap

TSource ' 300'

F)

'

 (psia)

Sink

Source

Sink '

HTCO = 100. (Btu/hour-f t - F)

= 0.2 (Btu/hour-f t-° F)

0.025

0.050 (inch)

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4

Z, ft

Figure 18. SS2D Captube Heat Transfer Model Effect of YT: .. on TrT ,. and

_ „.•%.. T tn i_ Wall wall T Within Larger Tube.

30
Cosmetically, all five panels looked good, however there was some

concern that the aluminum in the first three panels might react

electrochemically with the graphite over long duration. The fourth

adhesive is limited to about 250 F service so with the prospect of

300°F compressor bleed air, it was dropped as a candidate. The

3109 adhesive from 3M will display 1250 psi sheer at 300°F and has

no galvanic concern, so it was selected for use on the eventual wind

tunnel test panels.

Thermocouple instrumentation was then designed to fit within the

physical geometry of one capillary tube for future panels. Work was

initiated to provide a series of such thermocouples for integration

in discreet areas of the array of cluster surface capillaries on the

test panels. By keeping the thermocouples within the profile of a

tube, they would not interfere with the laminar flow of test air.

Figure 19 illustrates the final • experimental wind tunnel panel

designs.

A total of twenty-four special thermocouples were prepared (with

5 spares) for incorporation in the wind tunnel capillary tube test

panels. The thermocouples were made by precision butt welding the

wires end-to-end then benching away any weld flash at the joint. A

very light touch of contrasting paint was applied directly at the

weld joint so it would be easy to observe exactly where the

thermocouple joint was being located during test panel assembly.

Half of the thermocouples were made with thin wire and half with

thick wire to match the respective 0.017 inch and 0.027 inch outside

diameter of the capillary tubes being used for this program. The

thermocouples were delivered to the fabricator where they were bonded

in columnar array on graphite/PMPIS test panels using 3109 adhesive

as illustrated in Figure 20 and 21. Four wind tunnel test panels

were prepared. Two incorporated the 0.017 O.D. inch tubes and two

had the 0.027 inch O.D. in the two respected 0° and 90°

directions. Tubes initially extended about three inches beyond the

edges of the panels to facilitate manifold connection.

4.5 Physical Properties of Test Panels

Two more 12-inch by 3-inch test panels of identical construction

to the wind tunnel panels were prepared for physical evaluation as

shown in Figure 22. Tensile, flexural, beam shear and compression

tests were performed on these two panels which incorporate the

0.017-inch and 0.027-inch outside diameter tubing respectively.

Figure 23 depicts the method of cutting the test speciman from these

two panels. Cincinnati Testing Laboratory evaluated the physical

properties of these two panels which are listed on their test data

sheets in Figures 24 thru 28.

31
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36
TEST REPORT

CINCINNATI TESTING LABORATORIES, INC.

REPORT NO. TO 6041

FLEXURAL STRENGTH

CUSTOMER: GENERAL ELECTRIC COMPANY

Material:

Specification:

Preconditioning; 40 hrs ./2 3 °C/50 %R. H.

Test Condition: 23°C/50%R.K.

Span (L) 2.9 L/d Ratio: 32/1

Date:April 23, 1985

graphite PMR-15/glass capillary tubing (f ^"o^XSSS) PaneJ, A

ASTM D790-81 (Method I) Nose radius: 1/4"

Test

Speed: . 20 in./min.

Specimen length: 4V

Rexural Strength (S) = 2bd"i

Modulus of Elasticity (EB)

S * flexural strength in psi

Eg- Modulus of elasticity in psi x 10*

P = Break load in Ibs.

b = specimen width in inches

d = depth of beam in inches

L = span in inches

m = initial slope of load-deflection curve

in Ibs./in.

ibes

>wn

>wn

Specimen

(No.)

Al

A2

A3

Avg.

S

(psi)

68,200

61,190

42,470

57,290

d

(in.)

.089

.086

.098

.091

b

(in.)

.496

.495

.497

P

dbs,.)

61.6

51.5

46.6

m

(Ibs./in.)

217

207

237

(psix 10* )

3.78

4.01

3.09

3 . fi 3

Respectfully Submitted

CINCINNATI TESTING LABORATORIES, INC.

417 NORTHLAND ROAD

CINCINNATI, OHIO 45240

Figure 21*

37

TP<;T Tprhnirian _^J''"s^f"*' ~" Approved:

lyars D. Browning SBS
TEST REPORT

CINCINNATI TESTING LABORATORIES, (NC.

REPORT NO. TO 6041

FLEXURAL STRENGTH

Date. April 23, 1985

graphite PMR-15/glass capillary tubing (sk-31985) Panel B

Support radius: 1 / 8 "

Nose radius: 1/4"

Test

Speed: . 20 in./min.

Specimen length: 4V

CUSTOMER: GENERAL ELECTRIC COMPANY

Material:

Specification: ASTM D790-81 (Method I)

Preconditioning: 40 hrs . /23°C/50%R. H.

Test Condition: 23°C/50%R.H.

Span (L) 3.6 L/d Ratio: 32/1

Flexural Strength (S)

Modulus of Elasticity (Eg)

3m

S * flexural strength in psi

Eg= Modulus of elasticity in psi x 10*

P = Break load in Ibs.

b = specimen width in inches

d = depth of beam in.inches

L = span in inches

m = initial slope of load-deflection curve

in Ibs./in.

Specimen

(No.)

Bl

B2

B3

Avg.

S

(psi)

41,340

62,820

40,440

48.200

d

(in.)

.111

.108

.118

.112

b

(in.)

.494

.493

.492

P

(Ibs,.)

46.6

66.9

51.3

m

(Ibs./m.)

194

186

190

EB

(psix 10* )

3.35

3.49

2.74

3.1Q

Tubes

Down

Up

Down

Respectfully Submitted

CINCINNATI TESTING LABORATORIES, INC.

417 NORTHLAND ROAD

CINCINNATI, OHIO 45240

Figure 25

Test Technician:

38

Approved
TEST REPORT

CINCINNATI TESTING LABORATORIES, INC.

REPORT NO. TO 6041

COMPRESSIVE STRENGTH

GENERAL ELECTRIC COMPANY Date: April 23,1985

graphite PMR-15/glass capillary tubing (sk-31985) Panels A & B

CUSTOMER:

Material:

Specification: ASTM D695-80

Pre-Conditioning: 40 hrs ./23°C/50%R. H.

Test Condition: 23°C/50%R.H.

Specimen Type: dogbone (fig. 5)

p

Compression Strength (S) = j

Modulus of Elasticity (Ec) =

Testing Speed:.

.05 Jn./min.

S = Compression strength in psi

Ec= Modulus of elasticity in psi x 106

P = break load in Ibs.

b = specimen width in inches

d = specimen thickness in inches

Y = strain in in./in.

Specimen

(No.)

A1

A2

AVG

B3

B4

Avg.

S

(psi)

37 , 090

39,270

38,180

43,650

46,160

44,910

P

(Ibs.)

1592

1702

2319

2466

d

(in.)

.085

.086

.105

.106

b

(in.)

.505

.504

.506

.504

EC

(psix 10«)

4.69

4.51

4.60

4.0 4

4.17

4.11

Respectfully Submitted

CINCINNATI TESTING LABORATORIES, INC.

417 NORTHLAND ROAD

CINCINNATI, OHIO 45240

Figure 26

Test Technician 39 Approved:.

D. Browning
TEST REPORT

CINCINNATI TESTING LABORATORIES, INC.

REPORT NO. TO 6041

TENSILE STRENGTH

CUSTOMER: GENERAL ELECTRIC COMPANY Date April 23,1985

Material: graphite PMR-15/glass capillary tubing (sk-31985) Panels A & I

Specification: ASTM D638-82a

Preconditioning: 40 hrs/23 °C/50%R. H. Testing Speed: 0.20 in./m.n

Test Condition: 23°C/50%R.H.

Specimen Type I

S - Ultimate Tensile Strength in PSI

Sy - yield strength in PSI

Et - modulus of elasticity in PSI x 10*

Tensile Strength (SI - -^ P = break load in Ibs.

b - specimen width in inches

A

Modulus of Elasticity (Et) - d - specimen thickness in inches

Y = strain in in./in.

@ break

Specimen

(no )

Al

A2

AVG

B3

B4

Avg.

S

(PSII

35,430

35,950

35,690

34,400

33,150

33,780

Sy

(PSI)

P

(Ibs 1

1747

1758

1962

1944

d

(in.)

.099

.098

.115

.118

b

(in.)

.498

.499

.496

.497

Et

(psix 10*)

3.85

3.82

3.84

3. 91

3.67

3.79

Elongation

<%)

0.94

1.00

0.97

* 0.89

0.96

0.93

Respectfully Submitted *Failed outside 2" gage

CINCINNATI TESTING LABORATORIES, INC.

417 NORTHLAND ROAD

CINCINNATI, OHIO 45240

Figure 27

Test Technician

40

Approved:.

D. Browning
TEST REPORT

CINCINNATI TESTING LABORATORIES, INC.

Report No. TO 6041

SHORT BEAM SHEAR

CUSTOMER: GENERAL ELECTRIC COMPANY Date: April 23, 1985

Material: graphite PMR-15/glass capillary tubing (sk-31985) Panels A & B

Specification: ASTM D2344-76

Precondition: 40 hrs./23°C/50%R.H. TestSpeed: -05 in./Min.

L/d ratio: 5/1

M Load Points Radius:

 1/8" nose

1/16" supports

Test Temp: 23 °C Ret. Humid. 50% Specimen Size: L...

Specimen

(No.)

A

1

A 2

A

3

A

4

Thickness-d

(in.)

Width-b

(in.)

.097 ! .259

.096

.098

.097

.263

.259

.266

Break Load

(Ibs.) - P

185

150

193

150

3P

Shear Strength : S= 77-7 PSI Span: 485"

5520

4460

5700

4360 Code: panei A

Avg. S =5010

 PSI

>es

in

m

Specimen

(No.)

B 1

B 2

B 3

B 4

Thickness-d

(in.)

.114

.112

.115

.113

Width-b

(in.)

.252

.258

.251

.259

Break Load

(Ibs.) -P

240

156

225

140

Shear Strenc

6270

4050

5850

3590

th- C 3P PQI <?nan-

~ 4bd .570"

Code:

 Panel B

Avg. S= 4940 PSI

>es

m.

Respectfully Submitted,

CINCINNATI TESTING LABORATORIES, INC.

417 NORTHLAND ROAD

CINCINNATI, OHIO 45240

Test Technician:

Figure 28

41

T. Burns

Approved:

D. Browning
Panel A as referred to in Figures 24 through 28. was composited

with the smaller 0.012 inch I.D. tubes and Panel B had the larger

0.020 inch I.D. tubes. The test identified "down" or "up" refers to

which side the bonded tubes were placed during testing. Figure 29

shows a microphotograph of sections taken from both panels. In each

case, test failure was typical with no unexpected results. However,

conclusions expressed at the end of this report caution against

relying on any physical influence from the bonded tubes due to long

term stress corrosion and contamination corrosion that may occur

within the tubes.

In addition to the fabrication of composite panels with capillary

tubes on one surface, an all-capillary tube panel was also fabricated

under this program as a possible means for displaying isolated

material properties of captubes alone.

An order was placed to fabricate a 3ply 0 /90 all tube panel

9-inches by 7-inches for delivery to the NASA Project Manager as

required by Task II-C. This panel, illustrated in Figure 30 together

with the tested panels described above constitute an excess of the

one square foot (active surface) of capillary reinforced composites

to be submitted to the NASA Project Manager at the completion of this

program.

42
ORIGINAL PAGE IS

OF POOR QUALFTY

Microphotograph of Capillary Tube Panels Showing

0.012 inch ID Tubes Above and 0.020 inch

Tubes Below

Figure 29

43
Figure 30 Cross Ply Capillary Tube Test Panel.

44
5.0 Wind Tunnel Test Rig

In order to simulate an isolated engine inlet arrangement, a wind

tunnel test rig was designed with a six inch test section as

illustrated in Figure 31. A "Sodium Chloride Window" would allow

undiminished monitoring of infra-red detail during testing. The

capillary tubes would emerge into the two manifold sections as shown

with rubber sealing their respective connection to either manifold.

The 40/60% screen blockage in the air supply plenum would act to

smooth out airflow turbulence before it enters the test section.

A satisfactory method for manifolding the ends of the capillary

tubes for the wind tunnel tests was worked out as illustrated in

Figure 32. The tubes were cut off flush with the edge of the base

graphite/PMR15 test panel, then a clamp bracket was capped over the

top surface of the clustered tubes to create a flow path into each

manifold. Some plastic filler caulking material was applied to

smooth the transition areas as indicated.

The wind tunnel test rig was fabricated as illustrated in Figures

33. 34 and 35. Figure 35 shows the interior of the feed air plenum

with porous interior wall intended to even-out the feed air

turbulence which enters equally from four walls. Figure 36 shows the

wind tunnel test rig set up. Figure 37 and 38 show the capillary

tube panel manifold box which fits to the opening in the wind

tunnel. The original plan to simply bend the tubes into slots

feeding air to and from the manifold plenum chambers did not work.

The tubes broke where they joined the rigid panel. Earlier tests

showed that the tubes could be bent to the required radius, but it

was found that after the tubes were composited into panels, the

excess resin that flowed out along the tubes changed their bending

characteristics such that they broke when forced into the required

radius necessary to penetrate the plenum slots. Accordingly, another

approach was investigated that allowed the tubes to be cut off

abruptly at the edged of the test panel exactly where the tubes

broke.

5.1- Wind Tunnel Testing

Wind tunnel testing was initiated and several points were taken

before an air leak was detected in the vicinity of the test panel.

In order to repair the leak, it was necessary to remove the test

section manifold/panel system from the wind tunnel. It was then

discovered that the test panel was warped severely in two

directions. The side with the 3 inch glass tubes bonded in the

parallel cluster simply expanded in both directions. In the 6 inch

direction, the center was lifted about 3/16 inch and in the 3 inch

direction the center was lifted about 1/16 inch relative to the bowed

45
C

O01

!-i 3

4-1

CO0) H01ccEHTD300

46
oT3to

C/J

co•H

4-> 5 O O CO i-l CA u-01

4-1 CO

en o

cu )-i H CJ (U tO C « 3 0) H XI 13 4-1 C O

.

CMCO0)MM•H

47
Figure 33. Wind Tunnel Test Rig Assembly

48
ORIGINAL PAGE IB

OF POOR QUAUIt

Figure 34. Wind Tunnel Test Rig Horn.

49
Figure 35. Wind Tunnel Test Rig Plenum Chamber.

50
ORIGINAL PAGE IS

OF POOR QUALfTY

VD

oc

M

•H

Pi

01

c

-a

c

51
Figure 37 Test Panel Plenum,

Bottom View.

Figure ^g . Test Panel Plenum,

Side View.

52
edge. During testing, air temperature in the tubes got as high as

185°F with pressure at about 80 psia. Since 185 F was not high

enough to cause permanent thermal degradation of any of the

materials in the capillary tube test panel, it is assumed that the 80

psia must have induced some stretching or micro-cracking to cover the

observed residual warp which remained after the panel was released

from its clamped assembly at room temperature. Since the panel had

been rigidly restrained on all four sides, any expansion from

temperature or pressure would manifest in an upward bulge at the

center. However, if temperature was the main contributor to

expansion, the bulge should have retracted at cool down, but it

remained. Accordingly, it is suspected that pressure induced

stretching .and/or micro-cracks may be a contributing factor. At any

rate, the panel was reassembled to the test section by bonding it

clamped flat in place using 630 RTV adhesive leaving clearance on all

four sides for expansion. Testing was resumed and no further bulging

occurred at the original 185 F and 80 psia as before. Figure 39

shows the wind tunnel test rig arranged for this program.

5.2 Test Data Analysis

The SIMCAP computer program was used to simulate the conditions

existing in the composite wall/capillary tube flow check No. 2 dated

4-23-85, listed in Table I. .Predicted results for total mass

flowrate, Mach number and Reynolds number at tube exit are plotted

versus source pressure in Figures 40 through 42. Test data for mass

flowrate are also plotted in Figure 40. Conclusions based on these

results are:

1. Predicted and measured mass flowrates agree within 1%. This

is a remarkable agreement considering the number of

parameters and the large ranges of Mach and Reynolds numbers

involved in this test.

2 Figure 41 shows that the flow was choked for source pressure

greater than about 51 psia. Therefore, the three data

points at the highest source pressure were at choked flow

conditions.

3 Figure 42 shows that the Reynolds numbers for all of the

test runs were in the turbulent range ( 2300).

5.3 Heat Transfer Analysis

An air flow check or measurement was conducted on a total of 94

parallel capillary tubes with inside diameter of 0.0126 inch and

length of 10 inches using air at 75 F. The 94 tubes were potted

air tight through a slotted panel that formed on one side of a

manifold as shown in Figure 43. Air fed into the manifold was

53
ORIGINAL PAGE IS

OF POOR QUALTTY

54
Table I.

«•«•*«• F~I_OUJ CHECK D#*T#* SMEET **•«••«-

!Oil2i47 O4-23-1985

PART DESCRIPTION ISt.O2DIA.ID CAPILLARY PANEL W/T TEST

PURPOSE OF FLOW CHECK ISi DETERMINE FLOW CHARACTERISTICS

TECHNICIAN!JIM CAHILL EXT. 3-93O2

BAROMETRIC PRESSURE IS 29.37 IN HG. / 14.42067 PSIA

>*»***********»******«*****«***************«»«**«*********«*********«*****

I DENT STEP PLEN * W * PRESS FLOW PLEN AIR ROTO ROTO

NO. NO. PRESS (ACTUAL) RATIO FUNC PRESS TEMP PRESS - SCALE

IN.HG LB/SEC PSIA DEG. IN.HG RDG.

2.O* l.O* 8.81* O.O1O239* 1.3OO* 0.01263* 18.8* 75.O* 101.0* 0.004680

!= = = = = •» = = = = = » = = = = =: = * = = = = = = = = = # = = == = = # = = = = = =r = = *= = = = = *= = = = s* = = = = = *»t = = ««:«ss = =

2.O* 2.O* 14.66* O.01379O* 1.499* 0.01476* 21.6* 76.0* 101.O* O.O063OO

= = = = = =* = = = = = « = = = = = = * = = = = = = = = = » = = = «===» = = = = = = = = » = = = = = « = «=» = * = = = «=«»a=m»«»=== =

2.O* 3.O* 2O.56* O.O17182* 1.70O* 0.01622* 24.5* 76.O* 101.O* O.O07850

2.O* 4.O* 26.43* O.O2O451* 1.900* 0.01727* 27.4* 76.O* 1O1.0* O.OO9350

2.O* 5.O* 29.37* O.O21868* 2.0OO* 0.01754* 28.9* 76.O* 1O1.0* O.01OOOO

2.O* 6.O* 32.30* 0.023265* 2.1OO* 0.01778* 30.3* 76.O* 101..0* 0.01O650

!•••••*»••••*••••••*•••«»•••*••••••«••••••••*•••••*•••••*••••••*••••*•««•

2.'O* 7.O* 38.18* 0.026165* 2.3OO* 0.01825* 33.2* 76.O* 101.O* 0.0120OO

1» = = = = * = = = = = * = = = C = = »= = = = = = = = I» = =«=«=*»« ===««Z»C = ««M * = « = = = » = = = = = «*= = = = = = « = =

2.O* 8.O* 44.05* O.O29145* 2.5OO* 0.01B72* 36.1* 77.O* 1O1.0* O.O134OO

:=»= = = »== = « = »=« = = «M »« = ======*» ====== »«======»»»»»»«»=zc»E* ====»«*«»««=== = =

2.O* 9.O* 49.93* O.O31659* 2.700* 0.01883* 39.0* 77.O* 100.5* 0.0146OO

2.O* 1O.O* 55.80* O.O34497* 2.900* 0.01911* 41.8* 77.O* 100.5* 0.016OOO

B»««==»==s::=«#=«« = s=» = ====s = K = »35 = oa=«»=sj===ss»»«=s«»»ss = « = = # = = = = = 3 * = = — = = = = = =

2.O* 11.O* 58.74* O.O35653* 3.000* 0.0191O* 43.3* 78.0* 100.5* O.0166OO

= = = = = = » = = = = = » = = = = = =•» = = = = = = = = = » = = ===a» = K=r = = = = =» = = = 3«K» = = = « = » = = = = «:= * = = *: = = = = = =

2.0* 12.0* 61.67* O.O37O38* 3.1OO* O.O1921* 44.7* 78.0* 1O0.5* O.O173OO

^Z = Z = =* = = = = S# = = = » = S* = = = = = = =: = = .* = = = = ^=<= = 3=s= = =;zr=: + = = = =::r#= = =;=::=* = ==: = = =* = = = = = = Z = :=

2.O* 13.O* 67.55* O.039733* 3.3OO* 0.01936* 47.6* 78.O* 1OO.O* O.O1B7OO

= _—._- = _* _____•!»• ______*__ _ s = = =: :r = *:= = ===::=*:= = := = = = — = <(• — — — :sz»s»ssa<|S = ««K»*»s«S!SE = .-s

.2.0* 14.0* 73.43* O.O42696* 3.5OO* 0.01961* 5O.5* 78.O* 1OO.O* 0.0202OO

tmmmmm+mmmmm+mmmmmm+mmmmmmmmm+mmmmmm+mmmmmmmm+mmmmm+mmmmm+mmmmmm»mmMmmmmmm

2.0* 15.O* 79.30* O.045500* 3.7OO* O.O1977* 53.4* 78.O* 1OO.O* O.O213OO

= = = = = = # = = = = = *= = = = = =*= = = = = = = = = * = = ==:* = * = = = = = = = = * = = = = = »=: = = ZK*» = K*s«#K««««:3B = = =

2.O* 16.O* 85.17* O.O47864* 3.9OO* O.01975* 56.3* 79.0* 1OO.O* 0.0225OO

r = = = = = # = = = = = » = = =: = = = »=:= = = ==: = =: = * = = — = = =* = = = = = =:=:= * = = = = K # = = = = =:#=:== = = = = «-== = =====:== = =: =

2.0* 17.O* 88.11* O.O49268* 4.0OO* O.O1982* 57.7* 79.0* 1OO.O* O.O232OO

55
ADIABflTIC FLOUI CHECK «2

204 TUBES ! 10=0.0209 INCH L=3 INCHES : TSOURCE=77.F

u

UJ

en

x

z:

CD

o:

o

_j

LL

CO

CO

<I a

I—

o

SIMCAP Prediction

10 30 40

SOURCE PRESSURE (PSIfl)

Figure 40.

56
ADIABATIC FLOW CHECK »2

10=0.0209 INCH : L=3 INCHES : TSOURCE=77.F

U

CD

m

a

a

n

30

SOURCE PRESSURF (PSIft)

Figure 41.

57
ADIABflTIC FLOU CHECK «2

o 10 = 0.0209 INCH ; L = 3 INCHES TSOURCE=77.F

x

UJ

Ul m

a -

LJ tfl

Q ~

C/5

o

10 30 4D

SOURCE PRESSURE (PSIA)

Figure 42.

58
ORIGINAL PAGE IS

<* POOR QUALfTY

Figure 43,

59
measured for flow rate and pressure. The only escape for the air was

through the 94 tubes.. Results of this flow check are tabulated in

Figures 44. Table 2 & 3 and Figure 45. The SIMCAP computer program

was then used to simulate the conditions established in Table 2.

Predicted results for mass flowrate, Reynolds number at tube exit,

and Mach number at tube exit are plotted versus source pressure in

Figures 46 through 48. Test data for mass flowrate are also plotted

in Figure 46.

Conclusions based on these results are as follows:

1. Transition from laminar to turbulent flow begins to occur at

a source pressure of about 30 psia, there the Reynolds

number at the tube exit is about 2300.

2. There is a very good agreement between predicted and

measured mass flowrates, even within the transition range.

3. Figure 47 shows that the largest Mach number achieved in

these tests was about 0.45 at a source pressure of about 50

psia. In order to achieve choked flow conditions, the

source pressure would have to be increased above 105 psia.

4. If the relatively small source pressure level used in the

flow check is also used in the heat transfer tests, then

small flowrates and very large temperature drops will be

produced.

In the text that follows there are twenty-four figures containing

predictions for heat transfer tests projected for four different

assemblies of captubes. Two samples have an external heat transfer

surface which is six inches wide by three inches long, and two

samples are three inches wide by six inches long. One sample of each

size contains the larger diameter tubes (0.0209-inch inside diameter)

and one sample of each size contains the smaller diameter tubes

(0.0126-inch diameter).

The captube simulation code (SIMCAP) was used to calculate

results plotted in Figures 49 through 65. In addition, the KEPS

boundary layer clode was used to simulate the fluid flow in the cold

air duct to which the samples would be exposed. Boundary layer

results for the duct wall and the sample external surface are plotted

in Figures 66 through 72

60
****** FLOW CHECK. DATA SHEET ******

FROM THE LABORATORIES OF THE DEVELOPMENT HEAT TRANSFER UNIT

W. K. SOUTH MANAGER

ITEM DESCRIPTION

FLOW METER:

PRESSURES:

TEMPERATURE:

BAROMETRIC PRESSURE:

DATA:

COX SERIES 12 ROTOMETER

CALIBRATED AT 5O PSIB.

HEISE GAUGES , IN.HG

'MENSOR'DIGITAL PRESS. GA. IN.HG \ H20

DORIC TYPE 400A

HEISE GAUGE , ABSOLUTE PRESSURE

MANUALLY ENTERED INTO I.B.M. P-C

ALL HARDWARE AND PROCEDURES CONFORM TO 'AEBG 'SPECIFICATION

** E5OTF218 **

UNLESS OTHERWISE NOTED.

DEFINITIONS;

IDENT NO.:

STEP NO.:

PLEN PRESS:

IN.HG

* W *

(ACTUAL)

PRESS RATIO:

FLOW FUNC:

PLEN PRESS:

PS IA

AIR TEMP:

ROTO PRESS:

ROTO SCALE:

RDG.

A NUMBER ON THE PART THAT IS UNIQUE

THE ORDER OF PROGRESSION OF A PREDETERMINED PROCEDURE

THE WALL STATIC PRESSURE IN THE SUPPLY

PLENUM

THE CORRECTED TRUE MASS FLOW

IN LBM / SECOND

THE RATIO OF PLEN PRESS TO P. BARO

PR =( P. BARO + P. PLEN) / P. BARO

A CALCULATED VALUE,DOESN'T INCLUDE AREA

W ACTUAL* (AIR TEMP+460) '". 5/P. PLEN PS IA

A CALCULATED VALUE,

(P.PLEN + P.BARO)*(14.699/29.92)

THE TEMPERATURE OF THE AIR SUPPLY

THE INTERNAL PRESSURE OF THE ROTOMETER

THE NUMERICAL VALUE READ FROM THE

SCALES ON THE ROTOMETER TUBES

Figure 44.

61
•«-••*-••«- F="L_OUJ CMEC±><: D^Tft SHEET •*••*+•*-

15:35:12 03-11-1985

PART DESCRIPTION IS: (94)CAPILLARY. TUBES W/RAGGED ENDS

PURPOSE OF FLOW CHECK IS:DETERMINE FLOW CHARACTERISTICS

TECHNICIAN;JIM CAHILL EX. 3-9302

BAROMETRIC PRESSURE IS 29.O45 IN HG./ 14.2611 PSIA

h**««***«**»***************«**» «»** ****«*********«*****«>««-***4l>*»***«*«

I DENT STEP PLEN * W * PRESS FLOW PLEN AIR ROTO ROTO

NO. NO. PRESS (ACTUAL) RATIO FUNC PRESS TEMP PRESS SCALE

IN.HG LB/SEC PSIA DEG. IN.HG RDG.

l.O* l.O* 8.70* O.OOO424* 1.300*. 0.00053* 18.5* 75.0* 103.0* 0.000187

l.O* 2.O* 14.50* O.OOO738* 1.499* O.OOO8O* 21.4* 75.O* 103.0* O.O00345

1.0* 3.0* 20.30* O.O011OO* 1.699* O.OO105* 24.2* 74.0* 103.0* O.O00525

l.O* 4.O* 26.10* O.O01446* 1.899* 0.00123* 27.1* 74.0* 103.0* 0.000650

l.O* 5.0* 29.OO* O.O01611* 1.998* 0.00131* 28.5* 74.0* 103.0* 0.000730

l.O* 6.0* 32.OO* O.O01797* 2.1O2* O.00138* 30.0* 74.O* 103.0* 0.000820

= = = # = = = = = * = = = = = = # = = = = = = = = = # = = = = =r = # = = = = = sr= = » = = s:=!=*== = = ss»=: ==si:s = #rr = = = = s: = 

1.0* 7.O* 37.70* O.O0215O* 2.298* 0.00152* 32.8* 75.O* 103.0* O.O00990

1.0* B.O* 43.50* O.O02339* 2.498* O.00152* 35.6* 75.0* 103.0* O.O010BO

===*=====#======#=========•*======*========#=====•».=====#======*=========

l.O* 9.O* 49.40* O.002444* 2.701* 0.00147* 38.5* 75.O* 103.0* 0.001130

===»=====•»======#=========*======»========»=====*=====#======*=========

l.O* 10.O* 55.20* O.O0257O* 2.900* 0.00144* 41.4* 75.O* 103.0* O.001190

l.O* 11.0* 58.00* O.O02633* 2.997* O.00142* 42.8* 75.0* 103.0* 0.001220

l.O* 12.O* 63.OO* O.O028OO* 3.169* O.00143* 45.2* 75.O* 103.0* O.001300

= = =*==== = * = = = = = = # === = == = = = #= = = ===» = = = = = = = = * = = = = = # = = = ==# = = = = = = # = = = 

l.O* .13.0* 68.00* O.O02948* 3.341* 0.00143* 47.7* 75.0* 103.0* 0.001370

= = = * = = = = = #= = = = = = # = = = = = = = = = # = = = = = = » = = == = = = = * = = ===* = = = = = #== === = * = = 

1.0* 14.0* 73.OO* O.O03065* 3.513* 0.00141* 50.1* 75.O* 103.0* 0.001425

Table II

62
CAPILLARY TUBES W/RAGGED ENDS FLOW CHECK

PLEN PRES

8.7

14.5

20. 3

26. 1

29

32

37.7

43.5

49.4

e- c- ^

*J*J« --

58

63

68

73

W ACTUAL

O.O00424

0.000738

O.0011

0.001446

0.001611

0.001797

O.O0215

0.002339

O.002444

0.00257

0.002633

0.0028

O.OO294B

0.003065*

PRESSURE RATIO

1.3

1.499

,699

,899

,998

2. 102

2.298

2.498

2.701

2.9

2.997

3. 169

3.341

3.513

FLOW FUNC

0.00053

0.0008

O.00105

0.00123

0.00131

0.00138

O.O0152

0.00152

O.00147

O.O0144

0.00142

O.O0143

O.00143

0.00141

Table III

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ORIGINAL PAGE IS

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ORIGINAL PAGE IS

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70
ORIGINAL PAGE 13

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PRETEST PREDICTION FUR b-IN. WIDE C*PiUbE htflT TftMNiFEK TESi

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72
ORIGINAL PAGE IS

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PRETEST PREDICTION FOR b-iN. HIDE CflPTUBE HEftT TR»

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78
ORIGINAL PAGE 13

OF POOR QUALITY

PRETEST PREDICTION FOR HEAT TRANSFER TESTS OF LARGER C&PT'JBE

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PRETEST PREDICTION FOR HEAT TRftNSFER TESTS OF SrtftLLER TUBE

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Z (FT)

Figure 61

80
ORIGINAL PAGE IS

OF POOR QUALITY

PRETEST PREDICTION FOR HEAT TRANSFER TESTS OF SnALLER TUBE

EFFECT OF PSOURCE : TSOURCE=120.F A TSINK=70.F

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81
PRETEST PREDICTION FOR HEAT TRflN?F£f? TESTS OF SMALLER TUB:

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ORIGINAL PAGE IS

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84
KEFS BL ANALVSIS FOR CAPTU8E HEAT TRANSFER TESTS

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85
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86
KEPS BL SNfiLVSIS FOR CAPTUBE HEftT TRANSFER TESTS

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87
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88
KEPS Bl_ ANALVSIS FOR CAPTuBE HEAT TRANSFER TESTS

8-INCH PREHEAT SECTION ; COUNTERFLOU AIR STREAMS

Mach =0, 5

= 2%

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Preheat

Section

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Length

X (IN.)

Figure 70

89
KEPS BL flNALVSJS FOR CAPTUBE HE«T TRANSFER TES^S

2-INCH PREHEAT SECTION : COUNTERFLOw AIR STREAMS

Preheat

Section

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Length

X (IN. )

Figure 71

90
ORIGINAL PAGE IS

OF POOR QUALrTY

KEPS B'_ ANflLVSIS FOR CAPTUBE HEAT TRANSFER

 TESTS

e a-INCH PREHEAT SECTION : COUNTERFLOUI AIR STREAHS

at

a

i 1

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Figure 72

91
SIMCAP Results

Results for the six inch wide by three inch long sample

containing captubes with the larger diameter are plotted in Figures

49 through 65. The effect of hot-air source pressure and temperature

on total mass flowrate of air through the 204 parallel captubes, and

on Mach number and Reynolds number at the tube exit are shown in

Figures 48 through 51. The effect of PSOURCE and TSOURCE on the air

recovery temperature in the captubes (TCAP)) and on the outside wall

surface temperature along the sample (TWALL) are plotted in Figures

52 through 55. The effect of heat transfer coefficient at the

outside surface of the sample (HTCO) on TCAP and TWALL are shown in

Figure 56.

Similar results for the other three samples with a hot-air source

temperature of 120°F are plotted in Figures 57 through 65.

5.4 Conclusions from the SIMCAP Analysis

1 It is not necessary to have extreme values of PSOURCE and

TSOURCE in order to verify the captube simulation model.

For example, a source pressure of 50 psia and a source

temperature of 120 F give typical results.

2 There are very large step jumps in the sample surface

temperature (TWALL) at the ends of the captubes where the

sample is exposed to cold air at the sink temperature

(TSINK). However, the temperature drop across the heat

exchange length of the captubes is relatively small. (about

3.5°F in Figure 52. up to 24°F in Figure 55).

3 Figure 56 shows that a uniform heat transfer coefficient at

the outside surface of the sample (HTCO) has a big influence

on the mean temperature level of TWALL, but a negligible

effect on the temperature drop along the heat exchange

section of the surface.

4 The temperature distribution on the sample surface will be

very sensitive to variations in the value of HTCO over the

sample surface, and to end/edge effects. Such variations

would make it difficult (if not impossible) to verify the

captube thermal model, and variations in tube depth should

therefore be reduced as much as possible.

92
5.5 Boundary Layer Results

The KEPS BL code was used to simulate the boundary layer on the

duct wall preceding the sample surface, and on the sample surface.

The length of the upstream duct wall was assumed to be eight inches

followed by a sample length of six inches.

A thermal boundary condition input to the BL program is plotted

in Figure 66. The duct wall temperature was specified as 70 F

(529.7 R), and the sample temperature was assumed to decrease

linearly from 97°F (556.7°R) at x - 8 inches to 85°F (544.7) at

x - 14 inches. Note that a step jump in the wall temperature of 27

degrees is specified between the cold duct wall and the hot sample

surface at x - 8 inches.

BL simulations were made for three different mainstream Mach

number flows in the duct of 0.1, 0.3 and 0.5. The turbulent

intensity in the duct mainstream was specified at 2%, with a total

pressure of 20 psia and a total temperature of 70 F. The recovery

temperature for the mainstream duct air at the three Mach number

levels is also plotted in Figure 66.

The predicted variation of heat transfer coefficient along the

duct and sample surface is plotted in Figure 67 for each Mach number

about 0.5, 1.0 and 3.5 inches from the beginning of the duct wall for

Mach number flows of 0.5, 0.3 and 0.1. Note also that a thermal

spike occurs on the first inch of sample length following the step

jump in wall temperature.

The predicted variations of heat flux and velocity BL thickness

along the wall are plotted in Figure 68 and 69. A thermal spike

again appears in the heat flux distribution at the beginning of the

sample; however, the BL thickness is unaffected by the thermal BC,

and it is not significantly affected by the mainstream Mach number.

The velocity BL thickness is about 0.25 inches at the sample trailing

edge, which is small compared to the duct gap of 2.5 inches.

The thermal spike at the leading edge of the sample followed by

the large variation of HTC over the sample surface, as shown in

Figure 67 is unacceptable. However, the situation can be improved by

preheating the duct wall upstream of the sample to the same

temperature as the sample leading edge, and by orienting the sample

such that the duct air and the captube air streams are in

counterflow. Thus, the thermal spike would occur upstream of the

sample, and the rise of sample wall temperature in the direction of

duct air flow would tend to diminish the variation of HTC on the

sample surface.

93
To illustrate the effect of a preheat section followed by a

linearly increasing sample wall temperature, the wall temperature

distribution plotted in Figure 70 was input to the KEPS BL program.

The mainstream Mach number was set at 0.5, and all other parameters

were set at the same values used in the previous simulation.

The new results for HTC and wall heat flux are plotted in Figures

71 and 72. The thermal spike now occurs on the preheat section of

the duct wall, and HTC varies from about 80 to 90 Btu/hr-sq. ft. -F

on the sample wall. This variation of HTC on the sample could be

reduced even more by providing an even longer preheat section;

however, the average value of HTC on the sample surface would also

decrease.

5.6 Conclusions from the Boundary Layer Analysis

1. An undesirable variation of Heat Transfer Coefficient (HTC)

can develop on the outside surface of the test samples.

2. The variation of HTC on the sample surface can be reduced by

heating the duct wall preceding the sample to the same

temperature that occurs at the sample leading edge. Any

step change of surface temperature will cause a thermal

spike on the sample surface.

3. The variation of HTC can be further reduced by orienting the

sample such that the duct air and the captube air streams

are in counterflow.

94
6.0 Infrared Measurement of Surface Temperature

6.1 Infrared Temperature Measurement Sensor

Infrared measurement of composite material surface temperature

was performed using a scanning radiometer with two spectral

sensitivity bands of 3 to 5 microns and 8 to 12 microns.

Measurements were taken using a radiometer setting which provided

minimum temperature resolution of 1 F over a range of approximately

100 F. The spatial resolution of the thermal image was dependent

on the surface area contained in the radiometer's field of view; each

infrared image frame consists of 200 horizontal lines containing 250

resolvable elements per line. Data was available in standard

television video format which was viewed in real time, and also

stored on videotape for transfer to computer disk via a video image

digitizer.

6.2 Measurement System Calibration

In order to accurately determine the composite material surface

temperatures, the infrared temperature measurement system was

calibrated based on the actual operating conditions and physical

characteristics of the equipment utilized in. the experimentation.

Three primary items influence the accuracy of the infrared

temperature measurement: (1) " the radiometer measurement

characteristics, (2) the transmission properties of the viewing

optics, and (3) the emissivity of the thermal surfaces being

observed. Experiments were conducted to determine the effect of

these items on the wind tunnel measurements. In each case, the

radiometer output test results were recorded on videotape for later

processing by computer.

The calibration of the radiometer equipment requires a source of •

radiation with known intensities and spectral distribution. The

"blackbody" source is ideal for this purpose because it provides

radiation having well-known theoretical relationships between

temperature and radiated energy. A reference source traceable to the

National Bureau of Standards was used to generate radiometer image

data over a temperature range which spans the composite surface

experimental test temperature.

Sodium chloride (NaCl) was used as the optic material to provide

both a visual and infrared viewing window for observing the composite

surface under test. To determine the transmission properties of the

NaCl, thermal radiation from a source of known emissivity was

measured as viewed through the NaCl with the radiometer. During this

testing, the NaCl material was heated to temperatures representative

of actual test conditions. This information was used to compensate

the infrared image data for changes in transmission due to

temperature or local material variations in the NaCl.

95
Knowledge of the composite surface emissivity at the experimental

test temperature is essential to accurate infrared temperature

measurement. To determine surface emissivity, a sample of the

composite material was heated, and radiometer images were recorded

for a series of test temperatures spanning expected composite surface

test conditions.

6.3 Preliminary Emissivitv Test

The composite sample had a thin electric resistance heater

mounted to the back surface, and a piece of insulating material next

to the heater to minimize heat loss to the environment.

Three thermocouples were surface mounted, equidistant along the

axial centerline, and four thermocouples were embedded near the

sample edges. Temperature readings from these thermocouples were

used as reference points to ensure stablized and uniform heating

during the test. Radiometer video images of the sample were recorded

at 100, 125, 150, 175 and 200°F. The area surrounding the center

surface thermocouple was selected as representative of the sample

temperature.

Background data was obtained by maintaining all infrared camera

settings constant and recording the video image of a front silvered

mirror. A computer program was then used to adjust, the digitized

grey level data for the digitized background data and to reference

the data with the known emittance from the blackbody. For each test

temperature using this method, the emissivity was found to be

constant over the temperature range within the accuracy of the

experiment.

6.4 Data Acquisition

Steady-state temperature data were acquired using a combination

of thermocouple sensors and the scanning radiometer. Figures 73 and

74 illustrate the infrared measurement test apparatus used to acquire

composite surface thermal images of the test panel viewed in Figure

75. Thermocouples provided mainstream and capillary air

temperatures. The video output of scanning radiometer was recorded

on videotape to capture infrared image data of the composite surface

thermal patterns at each experimental test point. All test data

including calibration results were maintained in laboratory

notebooks. Infrared image data was referenced to written data using

laboratory notebook entries of tape count and electronically

generated text headers which were inserted on the videotape

immediately prior to recording the infrared image data. In addition

to recording the infrared data on videotape for later analysis in the

GE-AEBG Inspection Automation image processing laboratory, each test

was observed in real time with a television display to provide

monitoring of the thermal flow patterns during testing.

96
ORIGINAL PAGE 15

OF POOR QUALITY

97
ORIGINAL PAGE IS

OF POOR QUALITY

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3RIGINAL PAGE IS

OF POOR QUALITY

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6.5 Data Analysis

The recorded infrared images were digitized and transferred to

computer disk for access by a variety of image processing and data

analysis software routines. Digitized video image data were adjusted

for intensity level using a known black and white reference generated

electronically in the video image. Using image data files for NaCl

transmission properties and composite surface emissivity in

conjunction with calibration data for the radiometer obtained from

blackbody measurements, the thermal image data was converted to

temperature images. The resulting surface temperature computer files

served as input to computer image display and analysis software.

Data analysis of the digitized images and the subsequent

conversion from grey levels to temperature values was accomplished

via several independent computer programs. In addition, a high

resolution graphics terminal was utilized to generate statistical

histograms displaying the composite material eraissivity data and

detailed temperature numerical printouts for each test point.

To assist in the visualization of the surface thermal pattern,

the infrared temperature data were color enhanced for display on a

high resolution video display unit. Plots of temperature variation

were generated along lines or within regions as defined on the image

of the composite surface using interactive graphics techniques.

6.6 Results '

Color enhanced images for test point 120 are shown in Figure 76.

The four images display the composite surface temperatures about a

midpoint temperature of 151 F. In each image, the total range of

displayed temperature data was divided into seven equal segments.

These segments then were assigned a color which generated surface

thermal patterns of equal temperature bandwidth. For example, Figure

76a illustrates color "isotherms" having a temperature bandwidth of

T^F. Likewise, Figures 76b. 76c. and 76d have 6°F, 5°F and

4 F per color band, respectively.. The heated capillary tube flow

direction was left to right as shown in Figure 76. The resulting

color patterns show a decrease in temperature across the composite

surface from left to right resulting from the cooling effect of the

mainstream flow.

Color enhanced images for test points 112 and 114 are shown in

Figures 77a and 77b with their accompanying temperature printouts

below each in Figure 77c and 77d. Both points were taken with

primary flow at 2.02 pps, primary temperature at 60.5 F. Capillary

pressure for point 112 was 10" Hg, and for point 114 was 30" Hg.

These examples of image and numerical data are representative of

the large amount of both qualitative and quantitative information

obtained from the infrared thermal image measurements. The

visualization of the thermal surface patterns obtained in real-time

during the experimentation and in computer enhanced digital images

100
ORIGINAL PAGE

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102
provided information useful for qualitatively evaluating the effect

of different capillary tube and mainstream flows. Comparison of

infrared-based and thermocouple data indicated that a good

correlation could be obtained between the two measurement

techniques. The infrared method, however, provided more detailed

temperature maps of the composite material thermal behavior than was

available from conventional techniques.

103
7.0 COMBINED 2D WALL HEAT CONDUCTION

AND ID COMPRESSIBLE TUBE

AIR MATHEMATICAL MODEL

7.1 Thermal Analysis of Capillary Reinforced Composites

A computer program named SIMCAP was previously developed to

simulate the steady-state thermal performance of capillary tube

embedded within a composite wall. Negligible spacing between

adjacent capillary tubes was assumed.

The SIMCAP code includes a one-dimensional heat conduction model

at a typical cross section of a tube combined with a one-dimensional

compressible flow model along the tube.

It is of interest to determine the effect of tube spacing upon

thermal performance, and to determine if multiple layers of tubes can

significantly influence the surface temperature of a composite wall.

Therefore, a two-dimensional heat conduction analysis was made at a

typical cross section of the wall, and results are described in this

report.

A correlation for two-dimensional conduction shape factor for a

single layer o'f spaced tubes was also derived and inserted into a

modified version of SIMCAP. Computed results for composite wall

surface temperature as a function of tube spacing are also presented.

7.2 Two-Dimensional Heat Conduction Analysis

Consider the cross section of a tube/wall assembly represented in

Figure 78. One wall surface is exposed to hot air at temperature

TSOURCE, and the other wall surface is exposed to cold air at

temperature TSINK. Hot air at TSOURCE enters one end of the tubes

and flows through the tube length.

A layer of tubes with outside diameter D-2R and uniform spacing,

SPACE, is located at depth YWALL from the cold wall surface. An

optional second layer of tubes centered in the spaces between the

upper layer of tubes, and at a greater distance from the cold wall

surface, is also represented in Figure 78.

The conduction analysis will focus on a basic area bounded by

lines of symmetry for a single tube. The width of such an area is

designated in Figure 78 by "CONTROL WIDTH"; this width is equal to

(R+SPACE/2).

A computer program called BIE4 is available at General Electric*

for analyzing two-dimensional heat conduction problems. It is

particularly convenient and economical for analyzing the basic

conduction problem represented in Figure 78 and will, therefore, be

used in this analysis.

104
<

SPACE

R +

SPACE

« *i

CONTROL

WIDTH

\

Figure 7 Basic Conduction Problem,

105
7.2.1 Analysis of a Simple Test Case: A Plane Wall

In order to test the BIE4 program, a simple plane wall was first

considered so the BIE4 results could be compared to exact analytical

results.

The control area for the plane wall test case is shown in Figure

79. The heat transfer coefficient is HTCO at the cold wall surface,

and HTCL at the hot wall surface. The wall temperature is TWO at • the

cold surface, and TWL at the hot surface. The thermal conductivity

of the wall material is k.

The width (3R) and thickness (17R) of the control area are

representative of a composite wall. The vertical sides of the

control area are specified as adiabatic.

An equation for the rate of heat transfer (per unit depth) from

TSOURCE to TSINK through the wall is:

FTSOURCE - TSKINK1 (20)

1 + 17R + 1

3R(HTCO) 3Rk 3R(HTCL)

Equation (20) can be rearranged to the following dimensionless

form:

(q' /kAT) - 3 (21)

[ 1/BO + 17 + 1/BL ]

Where the following parameters are defined:

£T - TSOURCE - TSINK (22)

BO - R(HTCO)/k (23)

BL - R(HTCL)/K (24)

The dimensionless temperature at the cold wall surface is given by

the following equation:

TWO* - (q'A^-T) (1/3BO) (25)

Where TWO* is defined as

TWO* - F TWO - TSINK 1 (26)

[ TSOURCE - TSINK ]

Gibbs, R.W. and Kilduff, K.A., "2-D Heat Transfer Computer

Program Using the Boundary Integral Equation Method: BIE4",

TIS Report 82-GTD-12, April 1981.

106
/

/

x /

HTCO, TWO

\

/

/

/

/

/

• k

3R

\

17R

HTCL, TWL

f TSOURCE }

Figure 79 Simple Plane Wall Test Case,

107
If the following representative parametric values are specified,

HTCO - 100 (Btu/hr-ft**2-° F) (27)

HTCO - 10 (Btu/hr-ft**2-° F) (28)

k - 0.2 (Btu/hr-ft-° F) (29)

R - 0.01445 inch (30)

then the dimensionless surface heat transfer coefficients become:

BO - 0.60208 (31)

BL - 0.060208 (32)

After these values are substituted into Equations (21) and (25), the

following values of dimensionless heat transfer and cold wall surface

temperature are obtained:

(q'/kAT) _ 0.08506 (33)

TWO* - 0.04709 (34)

Note that the cold wall surface temperature is only 4.709% of M

greater than TSINK. That is, if TSINK-0°F and TSOURCE - 200°F,

then TWO - 9.418°F. Thus, the cold wall surface temperature would

be colder than the freezing temperature of water, unless capillary

tube heating is added to the wall.

The simple plane wall problem was input to the BIE4 program, and

the computed values for dimensionless heat transfer and wall surface

temperature agreed with the above values to within five decimal

places.

7.2.2 Analysis of the Reference Case

The effect of adding a single layer of capillary tubes was

considered next. The control area for the reference case is shown in

Figure 80. The tubes were specified at a depth of 3R, and the

dimensionless heat transfer coefficient at the tube surface was

specified as

Bl - 1.64206 (35)

The reference problem was input to the BIE4 program, and the

following values of dimensionless heat transfer at the cold wall,

tube, and hot wall surfaces were computed:

(q'A'T)n - 0.48938 (36)

(q'A'TK - 0.47962 (37)

- 0.00976 (38)

108
TWMAX

/

X

/

• k

3R

HTCO

TWMIN

17R

HTCL

Figure 80 Reference Case: One Layer of Tubes.

109
The dimensionless cold wall surface temperatures closest to the tube

(TWMAX) and farthest from the tube (TWMIN) were also computed as:

TWMAX* - 0.27414 (39)

TWMIN* - 0.26795 (40)

Note that when TSINK - 0°F and TSOURCE - 200°F then TWMIN -

53.39 F, which is well above the freezing point of water. Also,

the difference between TWMAX and TWMIN would be only 1.238°F.

7.2.3 Effect of Heat Transfer at the Hot Wall Surface

Equations (36-38) show that 98% of the heat flow through the top

surface comes from the tube, and only 2% comes from the bottom

surface. A simple but more conservative approach would be to neglect

the heat contribution from the bottom surface. Therefore, the

reference case was repeated with an adiabatic bottom surface: that

is, set HTCL - BL - 0.

The BIE4 program computed the following values for dimensionless

heat transfer and cold wall surface temperature:

(q'A'T)- - 0.48845 (41)

. (q'/kAT)-j - 0.48845 (42)

TWMAX* - 0.27365 (43)

TWMIN* - 0.26738 (44)

Comparing these results with Equations (36-40) for the reference

case, it can be seen that the bottom surface has negligible influence

upon the top surface. Therefore, the bottom surface was assumed to

be adiabatic in all later calculations.

7.2.4 Effect of a Second Layer of Tubes

The effect of adding a second layer of tubes was considered

next. The control area for this case is shown in Figure 81. The

second layer of tubes was assumed to be located at a depth 3R below

the first layer of Tubes, and the dimensionless surface heat transfer

coefficient for both tubes were the same as the reference case; that

is,

B2 - Bl - 1.64206 (45)

The BIE4 program computed the following values for dimensionless

heat transfer and wall surface temperature:

110
- 0.49338 (46)

- 0.44331 . (47)

(q'/kAT);: _ 0.05007 (48)

TWMAX* - 0.27620 (49)

TWMIN* - 0.27029 (50)

Equations (46-50) show that 90% of the heat flow through the top

surface comes from the upper tube, and only 1.0% comes from the lower

tube. Also, comparing these results to Equations (41 - 44) for a

single layer of tubes shows that adding the second layer of tubes

increases the heat transfer and wall surface temperature at the top

surface less than 1%. That is, if TSINK - 0°F and TSOURCE -

200 F, then adding the second layer of tubes would increase TWMAX

by only 0.5°F.

Therefore, adding the second layer of tubes has a negligible effect

on the cold wall surface temperature for the specified geometry.

This is because the upper layer of tubes "masks the view" of the

lower layer of tubes to the top wall surface; that is, the upper tube

causes a temperature in the region between the upper and lower tubes

that is very nearly equal to the temperature of the lower tube. Thus

the lower tube has only a small temperature gradient to cause heat

conduction from its surface.

There are, of course, cases where the second layer of tubes would

have a more significant influence on the cold wall surface. For

example, if the air temperature in first layer of tubes falls well

below TSOURCE, or if the spacing between the upper tubes is very

large, then the second layer of tubes would have a greater influence.

7.2.5 Effect of Tube Spacing on Heat Conduction

Calculations were carried out for the reference case illustrated

in Figure 80 with a single layer of tubes and an adiabatic bottom

surface. The width of the control area was varied from R to 6R; that

is, the spacing between adjacent tubes was varied from zero to 10R

(five tube diameters).

The results computed by the BIE4 program are plotted in Figures

82 and 83. Results for dimensionless wall surface temperature are

plotted in Figure 82. Note that TWMAX decreases more than TWMIN as

SPACE increases; that is, the wall surface temperature not only

decreases with spacing, but it also becomes nonuniform.

Results for dimensionless heat transfer from a tube to the upper

wall surface are plotted versus tube spacing in Figure 83. Note that

heat transfer from a tube increases as spacing increases (tubes are

eliminated); however, the heat transfer increase does not compensate

for the loss of tubes.

Ill
TWMAX'

TSOURCE

HTC1-

HTCO

/

/

/

/

t

t

1

1

1

'• V1

sc - 3:

2R

1

3R

2R

3R

"'2

- W

- ^^_ £_ /

17R

^ HTC2

= HTC1

3R

Figure 31 Control Area with Two Layers of Tubes,

112
0)

J-l

3

0)

O.

§

EH

en

en

<D

rH

C

'O

•H

CO

0)

) 1 2 3 4

[SPACE/D]

Figure 82 Effect of Tube Spacing on

Dimensionless Wall Surface

Temperature.

113
Figure 8.5 Effect of Tube Spacing on Dimensionless

Wall Surface Heat Transfer.

114
It is of interest to determine a heat conduction shape factor for

the control area. An equation for the rate of heat transfer (per

unit depth) from a captube to the top wall surface is:

F TSOURCE - TSINK (51)

(HTC1) Sk (R+SPACE/2)HTCO

Where the rate of heat conduction is proportional

factor, S. Equation (51) can be rearranged to give

dimensionless form:

to the shape

the following

(52)

L- r Bl (1+SPACE/D)BO

Equation (52) can be solved for conduction shape factor to give:

-1

S -

_ (q'A> Bl (1+SPACE/D)BO

(53)

Shape factor can be determined by substituting values of BO and • Bl

into Equation (53) along with computed values of (q'kAT) plotted in

Figure 83. The resulting shape factor values are plotted in Figure

84 versus tube spacing.

A least-squares regression fit was made to the curve in Figure 84

to give the following equation for conduction shape factor:

S - 0.28985 + 0.28278 * (SPACE/D)

- 0.022247* (SPACE/D) **2

(54)

Note that Equation (54) is valid only for a single layer of tubes at

a depth of 1.5D from the cold outer wall surface. The rate of heat

transfer can then be determined from Equations (52) and (54), and the

average wall surface temperature at an axial cross section can be

determined from:

TWALL* - (q'A'T)

(1+SPACE/D)BO

7.3 Axi, L Variation of TWAL]

(55)

The effect of tube spacing on the surface temperature of a

composite wall of arbitrary length can be determined by inserting

Equations (52), (54), and (55) into the SIMCAP computer program.

This was done and the modified code was named SIMCAP2.

115
1 3 4

[SPACE/D]

Figure 84 Affect of Tube Spacing on Heat

Conduction Shape Factor.

116
SIMCAP2 predictions showing the effect of tube spacing on a

captube having a length of one foot are shown in Figure 85. The

axial variations of air temperature in the captube(TCAP) and average

cross-sectional wall surface temperature (TWALL) are plotted versus

distance from the tube inlet for tube spacing from zero to five tube

diameters. Note that a tube spacing of only one diameter will

significantly reduce TWALL.

7.4 Comparison of Test Results with Computer Code

Analysis of test data was conducted with emphasis on the

sensitivity of difference between predicted assumptions and actual

test hardware with related test results. For example, a difference

of only a few thousandths of an inch between the assumed depth of a

capillary tube below the air flow surface versus its actual measured

depth could display a significant difference between predicted and

measured test results. Test data was carefully reviewed with this

sensitivity in mind.

The following Figures 86 thru 92 illustrate how closely actual test

results compare with the computer SIMCAP predictions.

117
320

280

TSOURCE = 300.(" F)

TSINK = 0.( ° F)

PSOURCE = ISO.(psia)

_PSINK = IS.(psia)

HTCO = 100.(B/HR-FT2-° F)

TKWALL =0. 2 (B/HR-FT -° F)

YWALL = 1.5 DO

DI = 0.0209 (in.) .

DO = 0.0289 (in.)

Figure 85 SS2D Captube Heat Transfer Model Effect of Captube Spacing on

TWALL and TCAP Distributions.

118
HEAT TRANSFER TESTS 1 0 1 A, 1 03A, 112, 114, 117, 120

^204 TUBES ! 10=0.0209 INCH ; L=3 INCHES

0

if)

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T

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Ul

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TESTS

117, 120

15 30 35 40 45 50

SOURCE PRESSURE (PSIA )

Figure 86
SinCAP RESULTS FOR 6-IN. UIIDE^ CAPTUBE HEAT TRANSFER TESTS

TEST 101A ; EFFECT OF VUIALL

T/C OPEN

IR SOLID
SiriCAP RESULTS FOR fe-IN. UIDE CAPTUBE HEAT TRANSFER TESTS

TEST 103A ; EFFECT OF VUALL

T/C OPEN SYMBOLS

IR SOLID shfMBOLS

HTCO

35 (BTU/HR-FT

35

Figure
SiriCAP RESULTS FOR S-IN. WIDE CAPTUBE HEAT TRANSFER TESTS

TEST 112 ; EFFECT OF VU1ALL

T/C OPEN SYMBOLS

IR SOLID SYMBOLS
!' U C

SIMCAP RESULTS FOR 6-IN. UIDE CAPTUBE HEAT TRANSFER TESTS

TEST 114 ; EFFECT OF VUALL

T/C OPEN

IR SOLID

0.05 0.1

Figure 90

0.3

(FT)

0.35

123
' SINCAP RESULTS FOR 6-IN. UIDE CAPTUBE HEAT TRANSFER TESTS

TEST 117 : EFFECT 6F VUALL

g

8

i

V

1

I

!

T/C OPEN SYMBOLS

IR SOLID SYMBOLS

'0 O.OS

Figure 91

o. i 0.15 0.8

Z (FT)

0.2S 0.3 0.35

124
SIMCAP RESULTS FOR 6-IN. WIDE CftPTUBE H'-.AT

 TPAN3.-ER TESTS

TEST 120 : EFFECT OF VUA*LL

<r

rz

Figure 92

Z (FT)
8.0 SUMMARY AND CONCLUSIONS

A heat transfer analysis was made to determine the thermal

performance of capillary tubes embedded within a composite wall. The

primary objectives were to determine the effects of spacing between

adjacent tubes, and of multiple layers of tubes within the wall.

A two-dimensional heat conduction analysis using the BIE4

computer program was made at a typical axial cross section of the

wall. Results show that a second layer of tubes had a small effect

on heat transfer and temperature at the cold outer surface of the

composite wall for the assumed geometry. Also, the cold wall

temperature decreased significantly as spacing increased between

adj acent tubes.

A correlation for two-dimensional heat conduction shape factor

for a single layer of tubes was also derived, and inserted into a

modified version of the SIMCAP captube simulation code. Computed

results show that a tube spacing of only one diameter can

significantly reduce wall surface temperature.

The BIE4 2D heat conduction program could be combined or coupled

with the SIMCAP simulation code to obtain a versatile tool for

analyzing complete tube/wall assemblies. However, such an effort is

beyond the work scope of this investigation.

The correlation between infra-red monitoring, thermocouple

recording and calulated analysis was exceptionally close for the

entire scope of this program.

126
9.0 PROJECTIONS FOR FUTURE CAPTUBE EFFORT

The prospective need for larger diameter capillary tubes for

anti-icing during lower engine performance conditions such as at

loiter, descent or taxiing when compressor bleed air-pressure and

temperature is relatively low, is being reviewed. No fixed

assessment has yet been made, however, as all engines are different

in this regard. Several tube vendors have been approached with

request for tube avaiability up to .050" I.D./.070" O.D. During the

course of discussing tube features with several vendors, it became

apparent that tube stress corrosion and contamination corrosion are

factors that must be addressed. Several glass tube suppliers

cautioned that structural stress could induce stress corrosion cracks

in their capillary tubes in a relatively short time. The suppliers

advised that their tubes should be composited as straight and

unstressed as possible. They also advised that contamination such as

acid rain would accelerate failure of a stressed tube. However, even

if tubes should crack, if they are composited sufficiently deep

enough into the surface, they could still function as an anti-icing

duct of hot air without catastrophic failure. Perhaps such tubes can

be coated internally with a barrier coating sufficient to inhibit

corrosive attack. The Hewlett-Packard Company that supplied the

tubes used in this contract had the tubes jacketed externally with a

very thin coat of polyimide as a means of inhibiting external

scratching and corrosive attack. It is obvious that larger diameter

tubing should be easier to coat internally, but evaluation of this

prospective solution was beyond the scope of this contract.

Several conceptual studies were made to project future

utilization of capillary tubes in anti-icing structures. Figures 93

and 94 show two versions of front frames for jet engines where hot

air can be ducted through surface capillary tubes clustered in

tangent assembly around radial struts. Figure 95 shows how one

vulnerable area on a metal or composite inlet flap might be anti-iced

with capillary tubes bonded in one particular area.

127
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130
10.0 PROPOSED FOLLOW-ON EFFORT

1. A detailed design of a captube anti-iced jet engine front

frame should be made. Detailed design will require input

from performance, systems, heat transfer and mechanical

design engineers.

2. -An evaluation should be made on impact damage of captube

structures and possible methods of repair.

3. A proof-of-concept test of an anti-iced captube frame should

be a culminating effort for a future program.

131