10 IRE TRANSACTIONS ON AERONAUTICAL AND NAVIGATIONAL ELECTRONICS March

Problems Arising in High-Speed Aircraft

Due to Cooling Requirements

of Electronic Equipment*NATHAN A. CARHARTt

Summary-High-speed aircraft require the successful operationof large quantities of electronic equipment. The problem of providinga suitable environment for such equipment is of interest to both air-plane and electronic designers. A proposal is offered by which coolingprovisions may be standardized for airplanes of widely varying per-formance. Basic design parameters for the development of such asystem are outlined, as are the economic factors involved.

IGH performance aircraft require large quan-tities of electronic equipment. Since this equip-ment must be designed to dissipate the inter-

nally generated heat, provisions must be made to main-tain an environment in the aircraft that not only pro-vides proper dissipation of the iinternal heat generatedbut also protects the equipment from overheat due toexposure to an environment that acts as a heat sourcerather than a heat sink. A simple method for establish-ing a standardized method of temperature control andan approximation of the costs involved are offered inthis discussion.Heat transfer may be effected by radiationi, conduc-

tion, convection, or some combination of these. For ourconsideration, radiation is not an attractive means be-cause the limited maximum temperature of electronicequipment precludes the practical application of coldwalls to the degree that would be sufficient for effectivecooling. In addition, some conduction and/or convec-tion would be necessary to cool such hot spots as mightbe radiantly shielded due to the architectural arrange-ment.

Conduction is frequently used for local heat transferwithin equipment packages; however, it offers a poorsolution to the transfer of heat to the aircraft structureor to some other sink because of shock mounting re-quirements and the weight involved in providing suit-able heat paths. Further, the immediate sink chosenmay itself require cooling by some supplemental means.

Convection shows promise as a readily useable coolingmethod offering a maximum of flexibility and goodefficiency. Evaporation is considered a means of con-vective cooling, for which water can be used quite ef-ficiently, as it dissipates 300-watt hours per pound.Problems of packaging and freezing, however, makethis, for the present, a special type of solution.

For universal application to air-breathing vehicles ofsubstantial flight endurance it is proposed that convec-tion be considered as the cooling means and that air be

* Manuscript received by the PGANE, November 11, 1957.t Douglas Aircraft Co., Inc., El Segundo, Calif.

used as the immediate heat transfer medium. The de-sired optimum characteristics of air, as defined by theelectronics designer, differ from those desired by the air-plane designer, although it is possible to so define thesequalities that both designers can achieve a practicalinstallation with a minimum expenditure of time andeffort.Common denominators for use of air for equipment

cooling are weight flow and differential temperature.Aircraft environment is a variable. Static temperaturesmay vary from above 100°F to below -100tF, and op-erating altitudes may range from sea level to above50,000 feet. Fig. 1 defines approximately the speed-alti-tude range of existing combat aircraft. Both speeds andaltitudes are well below existing record values.

60000,

40000 >ALTITUDEIN FEET

20000

n

0 200 400 600 800 1000

TRUE SPEED - KNOTS

Fig. 1 Combat aircraft limits.

Through its effect on absolute pressure, differentialpressure due to velocity, and temperature, aircraft per-formance affects the basic air supply. Since speed andaltitude are not quantities of importance to electronicequipment, the replot in Fig. 2 shows the temperaturevs pressure envelope of an aircraft, making allowancefor intake losses. Temperatures per AN-A-421 are usedto establish limits. Superimposed is a plot of the tem-perature requirements of Mil-E-5400. The extent towhich we are beyond the limits of standard equipmentshould be noted.

In the not too distant future we may expect aircraftto operate at Mach 3 and above. Fig. 3 depicts what atypical performance envelope may be for such speeds.Here the velocity scale is plotted as Mach number withthe upper limit of the previous example included forcomparison. The predicted environment of a vehicle atsuch speeds is plotted in Fig. 4. Note that very signifi-can-t increases in both temperature and pressure are

U. .a 9 a a a

Carhart: Cooling Requirements in High-Speed Aircraft

-100 - 50 0 50 i00STAGNATION TEMPERATURE -C

Fig. 2 Existing environment.

Fig. 5-Module standardization.

EXHAUSTED

EXHAUSTEDRAM AIR

. 4,.

0 1 2

MACH NUMBER

Fig. 3-Predicted envelope for Mach 3.

3-HERMETICALLYSEALED MODULE

RAM AIR FROMPLENUM CHAMBER

--OPEN MODULE30

20PRESSUREP. S. I.

AT MODULEINLET

10

RAM AIR FROMPLENUM CHAMBER

-100 0 100 200 300 40

STAGNATION TEMPERATURE - C

Fig. 4 Predicted environment.

evident. Later discussion will develop how these in-

creases may be accommodated.An arrangement such as that depicted by the module

shown in Fig. 5 is well suited to the needs of the air-plane designer and the electronic designer. It offers a

standard package whose mechanical characteristics can

be defined to the satisfaction of both.' The mountingsurface is shown with a seal and suitable air inlets, and

I Structural design, including details of mechanical anid electricalconnection, are not a part of this discussion.

Fig. 6 Typical module treatment.

through a plenum chamber this surface may be suppliedwith air of a requisite quality. The internal configura-tion of the module may be such that adequate heattransfer is attained.

Fig. 6 suggests two possible designs chosen from avariety of heat exchanger techniques. If pressurizationis not required, the air flow can be directed over the hotcomponents. With the exception of high voltage cir-cuits, absolute pressures as low as one psi present noproblem in electrical operation. Where high voltagesdemand a high pressure environment, this can be fur-nished within a hermetically sealed module in which thecooling air is run through heat exchanger tubes that ac-cept heat from the hot components either by radiationor conduction or both. Choice of an actual heat transferconfiguration should be made by the electronic equip-ment designer. The heat-rejection requirements andtheir solution are subject to engineering techniqueswhich are often analogous to those in use in electricaldesign. An optimum solution will be obtained onlywhen good electrical, mechanical, and heat transfer en-gineering are combined at a single location.

1958 11

12 IRE TRANSACTIONS ON AERONAUTICAL AND NAVIGATIONAL ELECTRONICS March

The temperature of air is readily defined and meas-ured. Weight flow across a given resistance is a func-tion of density and pressure drop available. Since amultiplicity of modules would be required in a typicalinstallation, it is desirable to establish a uniform re-sistance to airflow because the units would be in paralleland the flow of air should be proportional to the heatgenerated within the package. A value for this resistancewill be chosen to illustrate design limits for density andtem perature.The simplest driving force for an airborne forced-air

system is the pressure head developed by the vehicle'smotion through the atmosphere. Fig. 7 defines the en-velope of static pressure vs pressure drop available inthe aircraft on a hot day as described by Fig. 1. Alsoshown are lines of equal Mach number and equal tem-perature. The minimum power available as pressuredrop is about 0.05 psi. It is relatively independent of al-titude since minimum flying speeds are a function ofindicated airspeed and not ground speed. The maximumpressure drop is about 2.5 psi, which is far more thannecessary for cooling purposes and which may presentstructural problems. The line drawn at 0.38 psi (10inches of water) represents a suggested maximum valueto be made available to the module inlet. It providesadequate driving force, a reasonably low structuralloading, and a workable value for regulation. The maxi-mum temperatures as determined by aircraft velocityremain unchanged.

Suggested design values of pressure drop for a modulesystem are 1.5-10 inches of water. Efficient coolingsystems can be achieved with pressure drops as low as0.1 inch of water at static pressures of 15-20 psi; this lowa value is not recommended because it is proposed toallow static pressures to drop as lowv as one psi.

For universal application the internal module re-sistance must be standardized. A value of 5 pounds of airper minute per kw results in a heat rise across themodule of 47.1 °F (26 °C), as has been suggested.2 Thecurve of Fig. 8 is constructed about this point and de-picts the variation in air flow required with change intemperature rise. Much presently available electronicequipment will operate at temperatures far in excess ofthose specified in Mil-E-5400. Let us assume an allow-able air exhaust temperature of 110 'C. Pressure drop,weight flow, temperature, and pressure may be relatedby\

KW2T

p

where

Ap =pressure drop,W=weight flow,T = temperature related to absolute zero,

P=pressure related to absolute zero.

2 SAE Aeronautical Information Rep. No. 62; December 15,1956.

1.0 1.5 2.CPRESSURE DROP ACROSS MODULE - PSIG

Fig. 7-Equipment inlet conditions.

TEMP. TO GIVE EXAUST TEMP. OFRISE °C TEMP.283 1OC 383eK AT 60,000 FT. (PT. A)

80 -303

60 -323 ECS LV

EXCESS HEAT40t ~EXCESHEA aS FLOW CURVE FOR CONSTANT

20 -363 1LINE OF CONSTANT. EXHAUST TEMP.20363 ~ AP =1.5" H20P 15 psi

1 3 4 5 6 7 8 9 10 11 12FLOW */MIN./ KW.

Fig. 8-Airflow required per kw of heat dissipation.

Experience indicates that cooling with ram air tendsto be critical at altitude. If a point on the curve of Fig.8 is chosen for 60,000 feet and an inlet temperature of-10 °C, the value of K is 0.007. Using this value and theperformance envelope of Fig. 2, the weight flow vs ATis plotted on the figure with pressure drop regulatedto a maximum p= 10 inches of water. An excess ofcooling air is available for all conditions except of fullthrottle at sea level. The temperature deficiency of 5 °Cis conisidered acceptable because such operation is ex-tremely rare since it must be coupled with high ambienttemperatures, and short time overheating of equipmentby this magnitude should be tolerable. Also shown inFig. 8 is the weight flow attainable in a pressurizedsystem at 15 psi if the Ap is limited to 1.5 inches ofwater. A temperature rise of 50 °C per kw results underthose conditions.The type of module proposed earlier can be used in

a variety of systems and as previously noted, the sim-plest relies on the forward motion of the aircraft. Fig. 9illustrates the elements of such a scheme. External air is

'AP =

Carhart: Cooling Requirements in High-Speed Aircraft

Fig. 9-Ram cooling. Fig. 12-General arrangemiieiit of experimental pressurized module.

Fig. 10-Refrigeration system.

\[ze > > > >\ > > 7, aE~~~~XHAUST AIR

MODULES

-PLENUM CHAMBER

Fig. 11-Isolation cooling.

brought into the aircraft through an intake scoop. Pres-sure across the modules is controlled through a pressure

regulating valve located in the duct ahead of a plenumchamber that supplies the modules. This valve senses

plenum chamber pressure as compared to compartmentstatic pressure and regulates this to a predeterminedmaximum value.

Should the air temperature as measured in the plenumchamber be so high that cooling is inadequate (becauseof a mismatch in aircraft performance and heat toler-ance of the electronic equipment) air cycle refrigerationcan be provided from jet enginie bleed, as shown in Fig.10. Except for the immediate source of air to the pres-

sure regulator, the system is identical to the ram airsystem. Refrigeration units, utilizing cooling due to ex-

pansion from a high pressure semicooled source, are

proven and practical. As compared to the ram system

Fig. 13-Internal detail of experimonital pressurized module.

though, there are penalties both in weight and com-plexity.

Special flight conditions or the limitations of certainelectronic componenits may require isolating the equip-ment from the primary cooling air. In such a casemodules may be mounted inside a pressurized box in-corporating a secondary heat exchanger and means forinternial circulation of air. It should be noted that sucha system as shown in Fig. 11 adds additional heat in theform of power to drive a blower and requires a doubleheat transfer. Consequently, this is less efficient thanithe other two systems and is generally more difficult tomaintain since access to individual components may beachieved only by opening the pressurized box.

In order to verify the practicability of the secondaryheat exchanger, two autopilot circuits were built andtested in accordance with the packaging techniquesproposed. Because of contract cancellationi, the testwork was terminated before it could be considered com-plete; however, results were sufficient to demonstratethat adequate control of pressure drop is practicalthrough sizilng of the intake and exit orifices.One of the modules, as shown in Figs. 12 and 13, was

designed to allow pressurization of the componenits andtheir isolation from the cooling air. This air was directedthrough a tube heat exchanger made a part of the heatconductive structural mountinig plate. Acceptable in-ternal heat control of the hot elements was obtained.Cooling runs were made within a box whose walls weremaintained at a temperature of 200 °F in order to elim-

1958 13

14 IRE TRANSACTIONS ON AERONAUTICAL AND NAVIGATIONAL ELECTRONICS March

6

,/WATT

4

2

Fig. 14 Airplane growth factor.

FLIGHT TIME

Fig. 16 Cost of electrical poNTer.

- 20 0 20 40 60 80 100 120

INLET TEMPERATURE C

Fig. 15-Power generation.

mnate any appreciable heat loss from the module due toradiation.A penalty in aircraft performance is inievitable with

any power requirement not associated with producingthrust. Weight is of primary importance to the aircraft

designer, and the size of an airplane is directly propor-tional to the amount of payload that it is designed to

carry. For purposes of this discussion, payload con-

sists of those elements of an airplane exclusive of the

structure, flight controls, engine, and fuel. Present day,high performance, combat aircraft accommodate a

design payload of approximately 10 per cent. For equalperformance, airplane size must be modified by a

growth factor (Fig. 14) applied to the weight of those

items conistituting the payload. Substantial over-all sav-

ings may be made through reductions in electronic

equipment weight. Aircraft costs on a gross weightbasis are of the order of $40 per pound. Utilizing the

10-to-1 ratio it can be shown that dollar values ap-proaching $400 per pound may be assigned to weightsaved in the electronic equipment or in its use of

power.Electrical power supplied through an ac alternator is

normally available in an aircraft. Fig. 15 plots weightper kva against cooling air temperature at two different

altitudes for the installation of a Class C 20-kva blast-

FLIGHT TIME

Fig. 17 Cost of cooling.

cooled alterniator including its conistanit speed drive.Installed weight is a funiction of inlet temperature at

constant altitude. Points marked oni the curves showapproximate operating limits for existing airplanes. Thereductioni in air temperature with altitude is sufficientto provide the same output per pound in spite of a lowerair densitv.Power requirements for operating the alternator in-

clude both the energy to drive it and the drag associatedwith its supply of cooling air. Representative values forthese lead to a fuel expenditure of 4.25 pounds per hourper kva developed. The cost of supplying power vs

flight times may be plotted as showni in Fig. 16.Cooling costs will vary with the choice of packaging

techniques and the amount of refrigeration required.Ram air systems may be equated directly to the dragassociated with the routing of the air through the air-craft. At representative cruise conditions one-halfpound of fuel per hour will supply one pounid of air per

minute. Costs may be equated against time, resultingin a slope of $1 per watt per hour, as shown in Fig. 17.The installed cost (the point where the curve crosses

zero time) is based on a ducting system weight of 10pounds per kw of equipment rating.

Refrigeration systems, as noted, are complex and re-

quire power. Including ducting it is estimated that

INSTALLEDWEIGHTPOUNDSPER KVA S/WA1

198Welsh: Temperature Ratings and Cooling Procedures

FLIGHT TIME

Fig. 18-Total cost.

to provide a 50 °C drop in effective temperature, such a

system, as compared with a ram system, would weighabout 15 pounds per kw, or cost $6 per watt. Fuel re-

quirement of 8.5 pounds per hour per kw result in acost of $3.40 per watt per hour as shown.

Total energy requirements must include both thecost of generation and the cost of dissipation. Fig. 18plots total costs of the ram and the refrigeration sys-tems. It is readily seen that the costs of electronic equip-ment do not stop at the time of installation. Appreci-able effort can be justified to reduce the power require-ments of equipment and to make it amenable to opera-tion in high temperature environments.

CONCLUSION1) Convection cooling by air is suitable for the ma-

jority of electronic equipment cooling requirements inaircraft.

2) It is practical to standardize the characteristics ofair supplied to and its use in electronic packages.

3) The cost of energizing and cooling electronic equip-ment is appreciable. Extensive efforts can be justifiedto reduce wattage input since the least expensive wattto dissipate is that which is designed out of a piece ofequipment.

Temperature Limits, Ratings, and NaturalCooling Procedures for Avionic

Equipment and Parts*JAMES P. WELSHt

Summary-The results of adequate cooling of electronic partsare gains in part life and reliability. An engineering compromise be-tween ideal electronic part temperature and the thermal point ofdiminishing return must be evaluated not only with respect to desiredlife, but also in terms of the electronic circuit and cooling efficiencies.This paper outlines the flow of heat within, through, and from heatproducing electronic parts in terms of internal thermal limitations,part surface and environmental ratings, and cooling indices. Naturalheat flow design data pertinent to conduction cooling of heat sources,tube shields, the placement and mounting of parts, and "sink con-

nectors" are presented.

INTRODUCTION

SIGNIFICANT gains in electronic part reliabilityand life expectancy can be achieved through properthermal design since the life of most parts is an

inverse nonlinear function of the operating tempera-tures of their elements. The exact relationships are notcompletely known because of the complicated electrical,chemical, and mechanical phenomenona which occur.

* Manuscript received by the PGANE, December 16, 1957.f Cornell Aeronautical Lab., Inc., Buffalo, N. Y.

However, it has been established that life decreasesrapidly with increasing temperature in excess of anoptimum level. This is illustrated by Fig. 1, which pre-sents percentage of normal electron tube life vs per-centage of cathode temperature, and by Fig. 2, whichdisplays percentage of normal tube life vs percentage ofnormal plate temperature. The data for both figures arerough approximations obtained empirically with con-ventional receiving tubes.The benefits of adequate cooling have also been dem-

onstrated. In one instance, an over-all reduction of 20°Cresulted in a five-fold improvement in reliability. In an-other instance, thermal improvements increased themean-time-to-failure of an equipment from 35 to 1200hours. It cannot be overemphasized that reliability isprobably more closely related to the adequacy of heatremoval than any other single factor.One of the primary problems in the thermal design of

electronic equipment is determining the maximum tem-peratures which electronic parts can withstand. Maxi-

151958