LL1 21

N ii 637O - MR No. A4L11

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

It'l-IRTlillis' RISPORT ORIGINALLY ISSUED

December 1944 as Memorandum Report A4L11

TESTS OF A THERMAL ICE-PREVENTION SYSTEM FOR A WING

LEADING -EDGE LANDING -LIGHT INSTALLATION

By Wesley H. Hillendahi

Ames Aeronautical Laboratory Moffett Field, California

FILE COPY Ja be returnd to

the files of the National Advisory Cornrntee

for Aeronautics Washington, 01 C

A

/

WASHINGTON

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre-viously held under a security status but are now unclassified. Some of these reports were not tech-nically edited. All have been reproduced without change in order to expedite general distribution.

A-3

https://ntrs.nasa.gov/search.jsp?R=19930093543 2018-09-01T05:53:42+00:00Z

MR No. A11.L11

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

MM0TA.NDT M REPORT

for the

Air Technical Service Command, U. S. Army Air Forces

TESTS OF A THERMAL ICE—PREVENTION SYSTEM FOR A

WING LEADING—EDGE LANDING—LIGHT INSTALLATION

By Wesley H. Hillendahi

SUMMARY

The thermal ice—prevention system of a bomber--type airplane has been modified. in an attempt to provide protection against ice and fog formations on the trans parent fairing over, the landing light in the wing leading edge. A comparison of the design performance with the 4ctual performance measured on the ground and in flight in dry air indicated that the prediction of the outer—surface heat—transfer coefficient was satisfactory, but that the inner--surface heat—transfer coefficient was approximately four times as large as expected. This difference is attributed to the impinging action of the heated air on the transparent fairing, a factor which could not be evaluated in the idealized, design analysis. The failure of the transparent plastic fairing due to overheating, coupled with the return of the airplane to service, precluded modification of the system and further testing.

INTRODUCTION

As an extension to the application of thermal ice—prevention systems, which up to the present have been concerned only with the prevention of ice on the -metallic surfaces of the wings and empennage, an investigation of a method of ice prevention on wing—leading—edge landing lights has been conducted. Ice prevention of such installa-tions is necessary because the light beam is interrupted by ice forma-tions on the outer surface of the leading—edge transparent fairing and by fog on the ianer sur:ace. In addition, the ice destroys the aerodynamic efficiency in the region of the wing surrounding the light.

The analytical part of the investigation was reported in reference 1 wherein it was indicated that the wing thermal ice—prevention system could be utilized to afford protection for the light installation. The current investigation was conducted on a bomber—type airplane and includes ground and flight tests in dry

2

MR No. A1.L11

air to provide experimental verification of the analysis..

DESCRIPTION MY THE II\ISTALIATION

• Photographs of the standard leading-edge landing—light insta l la-tion in the test a:trplane and of the test installation, modified in accordance with :ol:'erence 1, are shown in figures 1(a) and 1(b), respectively. A Dian view of the test installation showing its relationship to the outer—wing —panel thermal ice—prevention system is presented in figure 2.

The analysis of reference 1 indicated that a reduction in the space between the sealed—beam light and the leading edge would be necessary to prcven the formation of ice on the transparency, and that a reduction in the surface area of the transparent fairing would be desirable from strength considerations. Accordingly, the landing light was moved forward from a location at approximately 6 percent chord to about 3 percent chord and the surface area of the transparent fairing reduced in accordance with reference 2, which governs the area of the light beam.

The lending light was incorporated into the existing wing do—icing systrn by' extending' the spanwise plenum through the wing splice into the light well, as shown in figure 2. The heated—air supply duct which originally bypassed the light was directed to the inboard end of the extended plenum at station 18. A double skin was extended to the region around the transparent fairing, as shown in figure 3.

Instrumentation included a venturi meter and shielded thermo-couple, both located in the duct supplying heated air from the heat exchger, as shown in figure 2, and five small--gage wire thermocouples mounted on each surface of the plastic, as shown in figure 3.

The transparent fairing was fabricated from 1/8—inch--thick CR-39 plastic, since that plastic retains its strength at higher temperatures than any other plastic known to be available.

TESTS AND RESULTS

Flight and ground tests were conducted with the test installa-tion to obtain dry—air performance data. The flight—test data were taken during (1) normal— and rated—powerc1imbs to determine if the plastic became overheated, (2) cruise and high power in level flight at 7,000, 10,000, and 18,000 feet pressure altitudes to check the design analysis, and () descent to determine if the plastic received sufficient heat under this low—power condition. The value of thermal conductivity of the CR-39 plastic fairing was determined

MR No. A1L1l

experimentally at the Ames Aeronautical Laboratory to be approximately 1.75 Btu per hour, square foot, F per inch.

Table I contains a summary of the results of tests at the conditions tested. Average temperatures on the surfaces of the plastic were obtained by means of the thermocouples shown in figure 3. Heat-transfer rates through the plastic were obtained from the average temperature gradients, thickness, and thermal conductivity Of the CR-39 plastic fairing. The inner-- and outer-surface heat-transfer coefficients were calculated from the heat-tranfer rate, and from the difference between the average surface temperatures and adjacent air tempera' Wires-

A comparison of the 'oorLe. tal results at two heated-air-flow rates with the analytical results of refOrence 1 is presented in table II and in figure 1 where urface-tomporaturo profiles are plotted.

A comparison is made it table III o the values measured during rated-power cl:brbs at two heated-air flow rates. The plastic fairing failed under the condition of run D, as shown in figure 5.

iiSCtJS ION

A comparison of the analytical and experimental results of figure 4 shows the surface temperatures of the plastic to exceed the predictod values even though test A of table II shows the flow rate of heated air to be about 60 percent of the design flow rate. Although the outer-surface heat-transfer coefficient Is of the same order of magnitude as the predicted value, the inner-surface coefficient is about four times as large for eompaehle flow rates as its predicted value. This difference may be attributed to the impinging action of the heated air on the plastic, a factor inherent in the test installation, since higher heat-transfer coefficients are known to result when air impinges on a aurfaco rather than when it flows parallel to the surface. In the analysis, which was based upon the flow in straight pipes, no account was taken of this factor.

The heat-transfer rates and. temperatures shown in table I, being of the so-no order of magnitude as those in systems which have been successfully tested in ice, are considered adequate for ice prevention under most conditions.

The upper temperature limit of a properly mounted plastic fairing lies botwoen the conditions shown in tests C and D of table III since failure occurred under the conditions of test D. Accordingly, the maximum temperature of the inner surface of the CR-39 plastic should not exceed. 2200 F when transferring a9uantity of heat corresponding to an average temperature drop of 100° F

MR No. A4Lll

through the plastic.

A previous failure which occurred during a ground runup is attributed to faulty mounting. Insufficient clearance had been allowed around the bolt holes to allow for ex pansion of the plastic.

Since the wing ice-prevention system requires a larger amount of heated air than was obtained in the present tests, the landing-light installation must be modified so as to prevent overheating of the plastic at the required flow rate. Such modifications include (1) a redirection of the supply duct to allow the heated air to flow parallel to the surfaco of the plastic fairing, (2) an enlarge-ment of the dross-sectional area of the plenum in the region of the plastic, and (3) the installation of a. bypass duct to allow a portion of he heated air to flow around the landing light. The first two modifications reduce the heat-transfer coofficiont on the inner surface of the plastic, while the third allows the larger flow rate to be supplied to the wing ico-prevention system without modifying the lending-light system.

CONCLtJS IONS

The following conclusions are drawn and recommendations are made:

1. Satisfactory agreement was obtained between analytical and ex3erimental values of the outer-surface heat-transfer coefficient; however, the inner--surface heat--transfer coefficient was four times as large as the predicted value, the difference being attributed, to the impinging action of the heated air on the surface of the trans--parent fairing.

2. The rate of heat transfer and the outer-surface temperatures are considered to be adequate for the prevention of ice on the landing-light fairing.

3. The maximum inner-surface temperature of the CR-39 plastic fairing should not exceed 220° F when the average tempera- ture gradient through the plastic is 1000 F.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Moffett Field, Calif., December 11, 19.

MR No. AL11

REFERENCES

1. HIllendahi, Wesley H.: Analysis of a Thermal Ice—Prevention System for Wing Leading—E1ge Landing—Light Installations. NACA ARR No. )4A11, 194,

2. Anon.: General Specifications for Installation of Lighting in Aircraft. U.S. Army Specification No, 9-32265--B. Aug. 10, 1942.

I: o. A' ' T 11

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MR No. AtL11

TABLE II.— COMPARISON OF ANALYTICAL AND TEST RESULTS

Analytical Test A e.et B

Fl I ght condition Level Level Levolflight flight flight

Pressure altitude, ft 18,000 18,000 18,000

Indicated airaoaed, mph —__155 153-

155

Freeair temperature, OF 30 27 32 -

315 Heabe—airtemeratureF 20 305

3280 Total flow rate from heat 270 3490

Flow ra;e of heated air to lenaing iigr,_lb/nr 2730 1550 184o

Average temperature of inner surface of plastic fairing, 1 165 210 235 OF

Maximum temperature of inner 185 230 270

Average temperature of outer 90 110 110 surface,

Maximum temperature of outer surface, 0^ 120 115 127

Average temperature gradient through _plastic fairing, OF 70 1 100 120

Heat—transfer rate through 1 plastic fairing, Btu/hr, 980 1100 1680 Sd ft

Inner---surface heat—transfer cce±'ficient, Btu / r, sq ft, 6.5 15 21

Outer—surface heat—transfer coefficient, Btu/hr, sq ft, 16 17 21

It?. 1: 7, A4Lil

TABLE III.- COMPARISON OF RESULTS AT MAXflvIUM POWER

TestC Test ____ _____

F1iht Ratsd-powor Rated-power -

Pressure t

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climb

Indicated airDeed, mph 135 1 135 - Free-air temperature, OF 55 25

Heated-air temperature, O F 320 350

Total flow rate from heat 1 -_changer, lb,ir

120 1710 Flow rate of heated air to landing light, lb/hr_______

Average temperature of inner surface of plastic, 225 270 oF

----Maximum temDerat,rre of () 235 25 imer surface, F

Average tem perature of 115 135 OF outer surfa ce,

Maximum temperature of 1

155 °F outer surface,

Average temperature gradient through plastic fairing,'

110 130

Average heat-transfer rate through plastic 1570 1800 fairing, .Btujhr, sq ft

MR NO. A4LI I

(a) Stendard. Installation.

(b) Installation fitted with ice -prevention equipment. Figure 1.— Landing—light installation on test airplane.

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- DOUBLE SKUJ (DIMPLED)

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-- CR-39 PLASTI / - BAFFLE AND LIGHT RETAINER FAIRING

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- A --LANDIPIG LIGHT

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TRUE SURFACE AREA 64 OF PLA5TIC FAIRING .0.715 FT

•- NATIONAL ADVISORY

a 6 C COMMITTEE FOR AERONAUTICS

PtOJeCTEO SURFACE OF PLASTIC FAIRING

FIGURE 3.- PLENUM AREA AND THERMOCOUPLE LOCATIONS ON PLASTIC FAIRING. -

•1

MR No. A41-11

320 .L HOT MR. TEMP. (ANALYTICAL)

0 MEASURED INNER SURFACE TEMP. OF PLASTIC FAIRING

.G. MEASURED OUTER 5U9,FACE TEMP. OF PLASTIC PAIRING 280

CALCULATED INNER SURFACE TEMP. PROFILES AT STA1 ION S

--CALCULATED OUTER SURFACE TEMP. PROflLES J .

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WING CHORD, PERCENT

.FIGURE 4. -. COMPARISON OF ANALYTICAL AND 'TEST VALUES OF SURFACE TEMPERATURES OF PLASTIC FAIRING.

MR No. A4LH

Figure .- Failure of Plastic fairing resulting frQu overheating in flight.