NASA/TM-- 1998-208826

Impact of Radiation Hardness and Operating

Temperatures of Silicon Carbide Electronics

on Space Power System Mass

Albert J. Juhasz, Roy C. Tew, and Gene E. Schwarze

Lewis Research Center, Cleveland, Ohio

December 1998

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NASA/TM--1998-208826

Impact of Radiation Hardness and Operating

Temperatures of Silicon Carbide Electronics

on Space Power System Mass

Albert J. Juhasz, Roy C. Tew, and Gene E. Schwarze

Lewis Research Center, Cleveland, Ohio

Prepared for the

Space Technology and Applications International Forum

cosponsored by the Boeing Company, Lockheed Martin, National Aeronautics

and Space Administration, U.S. Air Force, and the U.S. Department of Energy

Albuquerque, New Mexico, January 31 - February 4, 1999

National Aeronautics and

Space Administration

Lewis Research Center

December 1998
Acknowledgments

Larry Matus and Philip Neudeck of the NASA-Lewis SiC High Temperature Integrated Electronics

and Sensors Team for providing us with information about the development of SiC electronics

at NASA-Lewis and elsewhere and encouraging us to do this study

This report is a preprint of a paper intended for presentation at a conference. Because

of changes that may be made before formal publication, this preprint is made

available with the understanding that it will not be cited or reproduced without the

permission of the author.

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Price Code: A03
Paper Prepared for

Space Technology & Applications International Forum

Albuquerque, New Mexico, January 31-February 4, 1999

Impact of Radiation Hardness and Operating Temperatures

of Silicon Carbide Electronics on Space Power System Mass

Albert J. Juhasz 1, Roy C. Tew 2, and Gene E. Schwarze 3

NASA Lewis Research Center, 21000 Brookpark Road, Cleveland, OH 44135-3191

1Mail Stop 301-3, Phone." 216-433-6134, e-mail: Albert.Juhasz@lerc.nasa.gov, FAX: 216-433-6133

2Mail Stop 301-2, Phone." 216-433-8471, e-mail: Roy.Tew@lerc.nasa.gov, FAX." 216-433-6133

3Mail Stop 301-5, Phone." 216-433-6117, e-mail: Gene.Schwarze@lerc.nasa.gov, FAX." 216-433-8311

Abstract. The effect of silicon carbide (SIC) electronics operating temperatures on Power Management and

Distribution (PMAD), oi Power Conditioning (PC), subsystem radiator size and mass requirements was evaluated

for titree power output levels (100 kWh, 1 MW_, and 10 MW_) for neat term teclmology (i.e., 1500 K turbine inlet

temperature) Closed Cycle Gas Turbine (CCGT) power systems witit a High Temperature Gas Reactor (HTGR)

heat source. The study was conducted for assumed PC radiator temperatures ranging from 370 to 845 K and for three

scenarios of electrical energy to heat conversion levels wlficit needed to be rejected to space by means of rite PC

radiator. It] addition, during part of rite study rite radiation hardness of rite PC electronics was varied at a fixed

separation distance to estimate its effect on rite mass of rite instrument rated reactor shadow slfield. Witit botit rite

PC radiator and rite conical shadow slfield representing major components of rite overall power system rite influence

of rite above on total power system mass was also determined. As expected, results show titat rite greatest actual

mass savings aclfieved by rite use of SiC electronics occur witit Itigit capacity power systems. Moreover, raising rite

PC radiator temperature above 600 K yields only small additional system mass savings. The effect of increased

radiation hardness on total system mass is to reduce system mass by virtue of lowering rite slfield mass.

INTRODUCTION

During the past few years considerable R&D eftort has been devoted to raising the operating temperature

limits as well as the radiation hardness (tolerance to neutron fluence and gamma radiation dose) of electronic

components (Matus et al., 1997). Significant advances in the use of silicon carbide (SIC) materials for solid

state semiconductor components have been especially promising. An expected payoff is that resulting

electronic subsystems utilizing such components will be able to operate longer and more reliably, even in a

more severe operating environment. For space power systems, high temperature electronics offers additional

benefits, namely the reduction of PC radiator size and mass. Moreover, for power systems with nuclear heat

sources, increased radiation hardness enables relaxation of shielding requirements for the on-board electronics.

The present study is directed at providing an estimate of just how much system mass savings can be expected

from the use of electronic subsystems capable of operation at elevated temperatures and in higher radiation

environments. The hypothetical mission for the power system, especially for the 10 MW c power plant, was

assumed to be a precursor cargo spacecraft to Mars with a high specific impulse (Isp) electric propulsion

system. Since this power plant was also to provide surface power during the approximately one year Martian

stay time, the total mission time of 3 years was assumed for reactor sizing. To perform the study a "Closed

Brayton Cycle" (CBC) power system code developed by one of the authors (Juhasz) at the NASA Lewis

Research Center was utilized along with a reactor and shield code that was essentially the same as employed

by (Dobranich et al. 1987).

To reduce the mass of the high temperature part of the system it was desired to circumvent the need for an

intermediate heat exchanger between the reactor and the CCGT. Hence the helium working fluid was assumed

to be heated directly in a high temperature epi-thermal pellet bed gas reactor (HTGR), as described in several

NASA/TM--1998-208826 1
publications at the 10_1 SSNPS, e.g., (El Genk et al. 1993). Although the above references are to reactors

capable of heating the working fluid to near 2000 K, the turbine inlet temperature assumed for this study was a

relatively modest 1500 K, thus obviating the need for shield cooling. In agreement with prior system studies for

megawatt level power systems (Juhasz et al. 1993) required for inter-planetary nuclear electric propulsion

(NEP) applications, an initial system optimization showed that for the power system considered here a nonregenerated power plant would actually result in a lower overall system mass. Hence all system calculations

were done for a non-regenerated cycle with the temperature ratio kept constant at a value of 3.1, representing

the minimum system mass baseline case.

Regarding the low temperature side of the cycle, the need for a gas-to-liquid heat exchanger was eliminated

by ducting the He working fluid from the turbine exhaust section to a heat pipe (HP) main radiator featuring

direct forced convection heat transfer from the working fluid to the HP evaporator sections. The cycle reject

heat would then be conducted from the HP condensers to carbon-carbon (C-C) fins and radiator faceplates for

eventual heat transfer by radiation to the space environment. The C-C composite radiator design was based on

development work conducted during the SP-100 program (Rovang et al. 1994) and highlighted in an

anniversary publication (Juhasz and Peterson, 1994). This radiator conceptual design has also been proposed

for a lunar outpost power system based on the CBC, as well as for the multimegawatt NEP power system

referred to above (Juhasz et al. 1993). The PC radiator was also based on the same technology, albeit simpler

since it was required to operate at a constant temperature. For this study there was assumed to be no

interaction between the PC and main radiators. The sink temperature for the radiators was assumed to be

230 K.

To determine the heat load on the PC radiator three scenarios of electrical energy to heat conversion levels

were postulated. These included assumed combined PMAD and payload efficiencies of 0.9, 0.5, and 0.01, with

the highest value representing a power system for which 90% of the power generated was exported as work

energy, while the lowest value implied that practically all the power generated was consumed on-board, being

eventually converted into heat. The middle value represented an in between case. It should be noted that heat

pump subsystems were not included in this study. Study results for the baseline radiation hardness case for the

10 MWc system show the effect of electronic operating temperatures on the total system, main radiator, PC

radiator and the balance of platform (BOP) mass. The BOP included the reactor and shield, the CCGT turboalternator components, the PC sub-system and the connecting structure. Effects on total system mass, only, are

shown for the larger combined PMAD and payload efficiencies. The effect of radiation hardness, in terms of

neutron fluence and gamma dose, on shield weight were also studied and results are reported for the baseline

PC temperature case.

RESULTS AND DISCUSSION

First, a study of the effect of electronics temperature on system and PC radiator masses is discussed; for this

study the limiting values of neutron fluence (1 x 1013 neutrons/cm 2) and gamma radiation (5 x l0 s rad) used

are assumed to correspond to silicon electronics. Then a study of radiation hardness on shield mass is

discussed; for this study the PC electronics temperature was held constant at 400 K (127 C).

Effect of PC Electronics Temperature on System and PC Radiator Masses

The temperature of 400 K (127 C) is approximately equal to the military specification of 125 C which is a

current limit that is used for most silicon based electronics applications. It is believed that silicon carbide

electronics will be able to eventually operate reliably at 600 C or approximately 875 K, the upper limit used in

this study. The information below shows the effect of electronics temperature on system and PC radiator

masses. The average PC radiating temperature was assumed to be 30 K less than the electronics temperature;

this is believed to be characteristic of a heat pipe radiator. Results are shown for 10, 1 and 0.1 MWc

(megawatt electric) systems.

99% of Power Rejected through PC Radiator

The first results shown assume that 99% of the electrical power generated must be rejected through the PC

radiator. This would likely be the case if an insignificant amount of power is rejected via radar, other

electromagnetic signals, thermal radiation from the surfaces of the spacecraft, or via direct conversion to work.

NASA/TM--1998-208826 2
Figure1 showshowtotal system mass varies with PMAD and payload electronics temperature for the 10, 1

and 0.1 MW c power levels; it is

100

_90

09

co80

o70

e60

50

09

09

_4o

E30

_20

°910

0

2OO

m

m

m m

't'......................... _.......-............

400 600 800 1000

Temperature (K)

FIGURE 1. System mass as a function of electronics

temperature for 99% of power radiated via PC radiator

6O

09

09

50

E

40

09

O9

._..E 30

E

O

o 20 -3

"0

n-

_ 10

+10 1 MWe MWe /

_0.1 MWe_

0 --

200 400 600 800 1000

Temperature (K)

FIGURE 2. Per cent reduction in system mass, relative to

that at 400 K for 99% of power radiated via PC radiator.

90 T +Main Radiator

09_0980 _r \ _PC Radiator

09_ 1 \ _BOP

- 7ot O9_

_v

>,c_

e_40

50

30

x

0

20

10

0

200 400 600 800 1000

Temperature (K)

FIGURE 3. 10 MW, system masses as functions of

electronics temperature for 99% of power radiated via

the PC radiator.

16

14

09

0912

glO

E_-

o8

__

o9 •

N 4

Main Radiator

)__ PC Radiator

BOP

_System

,I, J2 *

.

0

200 400 600 800 1000

Temperature (K)

FIGURE 4. 1 MW, system masses as functions of elec

tronics temperature for 99% of power radiated via the

PC radiator.

seen that the larger power systems are more significantly affected by changes in the electronics temperature.

Figure 2, based on the same computed data, shows percent reduction of system mass relative to the system

mass at 400 K electronics temperature and provides better resolution for the smaller power systems. The plots

also show that most of the potential reduced mass benefit can be achieved by increasing the electronics

operating temperature to 600 K (or 327 C).

Figures 3, 4, and 5 show system, radiator, and Balance of Platform (BOP) masses for each of the power

systems, respectively, as functions of electronics temperature. The BOP mass is the total system mass minus

that of the main and PC radiators; it includes the reactor, shield, power conversion system, electronics, and

the structural mass.

NASA/TM--1998-208826 3
Forthe10MWsystemmassesshowninFigure3,it isseenthatat400K,thePCradiatordominatestheother

system components; PC radiator mass becomes less significant if the electronics temperature can be

increased. In

ffl

ffl 4

ffl

E__3

O

_, "r-

_N2

_v

O

0

2OO

---0-- Main Radiator

PC Radiator

BOP

_System

400 600 800

Temperature (K)

1000

FIGURE 5. 0.1 MW_ system masses as functions of

electronics temperature for 99% of power radiated via

PC radiator.

70-- 5 +10 MWe

60 _ _1 MWe

O

_- __,__MWe

._ 50

'6

_ 40

_9

,_ 30

E 20

>, {o10 - --

_- _- _- ,,

0 .......................... I I........................... 4

200 400 600 800 1000

Temperature (K)

FIGURE 6. System mass as a function of electronics temp

erature for 50% of power radiated via the PC radiator

comparison, Figures 4 and 5 illustrate that as the power level of the system becomes smaller, and less waste

electronics heat needs to be dissipated--the mass of the PC radiator becomes less significant.

50% of Power Rejected through PC Radiator

If 50% of the electrical power is emitted "directly" from the spacecraft via radar, other electromagnetic

radiation or by direct conversion to work, then only the remaining 50% of the electronics waste heat must be

rejected via the PC radiator. For each system power level, reducing PC radiator rejected heat from 99% to

50% of the electrical power has the effect, primarily, of reducing the mass of the PC radiator by approximately

50% and also reduces the structural mass. The effect of increasing the operating temperature of the

electronics on the system masses for the 50% case is shown in Figure 6. Figure 6 shows that for a 10 MW

system, increasing the electronics temperature from 400 to 875 K, produces only about half the reduction in

system mass (- 27,000 kg) that is shown for the 99% case of Figure 1 (which shows a reduction of

54,000 kg).

10% of Power Rejected through PC Radiator

If 90% of the electrical power is emitted directly from the spacecraft, leaving only 10% of the electronics

waste heat to be rejected via the PC radiator, the impact on the spacecraft mass of increasing the electronics

operating temperature is greatly reduced. This is shown in Figure 7. Here, for a 10 MW system, increasing

the electronics temperature from 400 to 875 K produces a reduction in system mass of less than 5500 kg. This

is, of course, because the PC radiator is only a little more than 5100 kg at the lowest, 400 K, electronics

temperature.

The above results show only the effect of increasing the electronics operating temperature on reducing the P C

radiator and system masses. Another benefit to be derived by replacing silicon with silicon carbide electronics

is increased radiation hardness. Increasing the radiation hardness of the electronics would reduce the mass of

the instrument rated shield for a nuclear reactor power system; it would also reduce any requirement to protect

spacecraft electronics from other radiation that might be encountered in space. The results shown below

address only the impact of electronics radiation hardness on the mass of instrument rated shielding.

NASA/TM--1998-208826 4
Effect of PC Electronics Radiation Hardness on Shield Mass

The results discussed here are based on a 46 MWt (megawatt thermal) reactor heat source for a 10 MWc

CCGT power system designed for three years of operation with a shield having a fixed cone half angle of 15

degrees and 25 m separation between the reactor and the payload electronics.

0_

g

I----

O

.mx._

v

0_

0_

t_

E

09

50 8

45

40

35

30

25

20

15

10

5

0

200 400

"-.._.___.

---0--- 10 MWe

1 MWe

+0.1 MWe

l ii ii m m

A A A A

600 800 1000

Temperature (K)

FIGURE 7. System mass as a function of electronics

temperature for 10% of power radiated via PC radiator

a

t-..-

O

.m

x._

0_

0_

t_

-05

(-

09

4

1.00E+12 1.00E+13 1.00E+14 1.00E+15

Neutron Fluence (neutrons/cm 2)

FIGURE 8. Shield Mass as a function of neutron fluence

dose limit for a 10 MW, (46 MW t ) 3 year system with

25 m reactor to PMAD separation distance.

A neutron fluence of 1013 neutrons/cm 2 is assumed to be the upper radiation limit of silicon based electronics.

A neutron fluence about an order of magnitude larger, i.e., 10 TM neutrons/cm 2, is assumed to be a viable goal for

silicon carbide based electronics.

For this study, the neutron fluence was allowed to vary from 5 x 1012 neutrons/cm 2 (about one-half the assumed

limit for silicon electronics) to 5 x 1014 neutrons/cm 2 (about 5 times the assumed limit for silicon carbide

electronics). The variation in shield mass (composed of lithium hydride, or LiH, and tungsten, or W, layers for

neutron and gamma attenuation, respectively) over this range of limits is shown in Figure 8 as a function of

neutron fluence; the gamma radiation dose was held constant at 5 x l0 s rad. Increasing the neutron fluence

from 1013 neutrons/cm 2 (upper limit of silicon electronics) to 10 TM neutrons/cm 2 (viable goal for silicon carbide

electronics) was found to allow a reduction in shield mass of 435 kg.

The gamma radiation dose was held constant at the silicon electronics limit of 5 x 10 s rad for this study

because (1) the gamma dose is limited by the gate oxide such as that used in power MOSFETs, and (2) the

same oxide is used for both silicon and silicon carbide electronics. If (1) the SiC electronics are limited to

semiconductor devices that don't use a gate oxide (such as bipolar junction devices) or (2) a more gamma

resistant gate oxide is developed, then it might be possible to increase gamma dose by an order of magnitude,

also, from 5 x l0 s to 5 x 106 rad. Such an increase, for constant neutron fluence of 1013 neutrons/cm 2, would

reduce the mass of the shield used in this study by 1630 kg. Total shield mass is much more sensitive to

gamma dose than to neutron fluence because of the higher mass density of the gamma-tungsten shield as

compared to the neutron-LiH shield.

CONCLUDING REMARKS

Substitution of silicon carbide for silicon electronics would produce substantial reductions in PC radiator and

system masses for space power systems of 0.1 MW (100 kW) and larger--when most of the electronics waste

heat must be rejected through the PC radiator. When 99% of the electronics waste heat must be rejected

through the PC radiator, substitution of 875 K (-600 C) silicon carbide electronics for 400 K (-125 C) silicon

electronics, was found to produce reductions in system mass of 56%, 39% and 11% for 10 MW, 1 MW and

NASA/TM--1998-208826 5
0.1 MW power systems, respectively. These percentage reductions in system mass correspond to absolute

system mass reductions of approximately 54000, 5400 and 540 kg, respectively. These system mass reductions

are almost entirely due to reductions in the size of the PC radiator (there is some reduction in structural mass).

When these percentage reductions in system mass were plotted as a function of system power level on a semilog plot and extrapolated to 0.01 MW (10 kW) and 0.001 MW (1 kW), it seemed clear that the substitution of

silicon carbide for silicon electronics would have negligible impact on the PC radiator and system masses for

10 kW and smaller power systems.

If the percentage of electrical power rejected through the PC radiator is reduced from 99 to 50%, then the

percentage reduction in system masses fall to 40%, 25% and 6% for 10, 1, and 0.1 MW systems, respectively.

A further reduction in the percentage of power lost via the PC radiator to 10% means that the substitution of

875 K capable silicon carbide electronics for 400 K capable silicon electronics would produce reductions in

system mass of only 12%, 6%, and 1% for 10, 1, and 0.1 MW systems, respectively.

Suppose that silicon carbide electronics has developed to the point of being capable of operation at "only"

600 K (-325 C) instead of 875 K (-600 C). When 99% of the electrical power is rejected via the PC radiator,

substitution of 600 K capable silicon carbide electronics for 400 K capable silicon electronics would produce

reductions in system mass of 49%, 34% and 10% for 10, 1, and 0.1 MW systems, respectively. This is not

much less than the 56%, 39% and 11% reductions that could be achieved by substitution of 875 K capable

silicon carbide for silicon electronics. A similar study reported by (Dobranich, 1987) assumed substitution of

600 K capable for 400 K capable electronics and a system designed for one year of operation, instead of the

three years assumed for this study. Dobranich's study showed reductions of 40%, 17% and 7% in system mass

for 10, 1 and 0.1 MW systems. Only the 17% reduction reported for the 1 MW system seems somewhat

inconsistent with the results found in the current study.

The radiation hardness study showed that substitution of SiC for silicon electronics, with an increase in neutron

fluence from 1013 to 10 TM neutrons/cm 2, would reduce instrument rated shield mass by 435 kg for a 10 MW

system (for constant gamma dose of 5 x l0 s rad). If the gamma dose could also be increased an order of

magnitude, this would reduce shield mass by 1630 kg (for constant neutron fluence of 1013 neutrons/cm2).

These benefits can be compared with system mass reductions that were found by substituting 600 K capable

SiC electronics for 400 K silicon electronics. Such a substitution produced a 47100 kg reduction in system

mass when 99% of the power is rejected through the PC radiator; the reduction in system mass was about half

as large when only 50% of the power was rejected through the PC radiator.

REFERENCES

Dobranich, D. "The Effect of Maximum-Allowable Payload Temperature on the Mass of a Multimegawatt

Space-Based Platform," Sandia National Laboratory Report # SAND87-1449, 1987.

E1-Genk, M. S.; Morley, N. J.; and Juhasz, A. J. "Pellet Bed Reactor Concept for Nuclear Electric Propulsion,"

10.11Symposium on Space Nuclear Power and Propulsion, Albuquerque, New Mexico, January 10-14,

1993.

Juhasz, A. J.; E1-Genk, M. S.; and Harper, W. B. "Closed Brayton Cycle Power System with a High

Temperature Pellet Bed Reactor Heat Source for Nuclear-Electric Propulsion (NEP) Applications,"

NASA TM-105933, Oct. 1992. (10 .11Symposium on Space Nuclear Power and Propulsion, Albuquerque,

New Mexico, January 10-14, 1993).

Juhasz, A. J. and Bloomfield, H. S. "Development of Lightweight radiators for Lunar Based Power Systems,"

NASA TM-106604. (24 .11 International Conference on Environmental Systems), Friedrichshafen,

Germany, June 20-23, 1994.

Juhasz, A. J. and Peterson, G.P. "Review of Advanced Radiator Technologies for Spacecraft Power systems

and Space Thermal Control," NASA TM-4555, June 1994.

Juhasz, A. J. and Rovang, R.D. "Composite Heat Pipe Development Status: Development of Lightweight

Prototype Carbon-Carbon Heat Pipe with Integral Fins and Metal Foil Liner," NASA TM-106909, May,

1995.

Matus, L.G., Beheim, G.M., Kuczmarski, M.A., Larkin, D.J., Neudeck, P.G. and Powell, J.A.; "Silicon Carbide

High Temperature Integrated Electronics," Paper 52 HITEMP Review 1997, April 30, 1997.

Rovang, R.D., Hunt, M. E., and Juhasz, A. J. "Testing of a Liquid Metal Carbon - Carbon Heat Pipe," 11_1

Symposium on Space Nuclear Power and Propulsion, Albuquerque, New Mexico, January 10-14, 1994.

NASA/TM--1998-208826 6
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1998 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Impact of Radiation Hardness and Operating Temperatures

of Silicon Carbide Electronics on Space Power System Mass

6. AUTHOR(S)

Albert J. Juhasz, Roy C. Tew, and Gene E. Schwarze

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135- 3191

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Washington, DC 20546- 0001

WU-632-1A-1G-00

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E-11453

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NASA TM--1998-208826

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Prepared for the Space Technology and Applications International Forum cosponsored by the Boeing Company,

Lockheed Martin, National Aeronautics and Space Administration, U.S. Air Force, and the U.S. Department of Energy,

Albuquerque, New Mexico, January 31 - February 4, 1999. Responsible person, Albert J. Juhasz, organization code 5440,

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12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

The effect of silicon carbide (SIC) electronics operating temperatures on Power Management and Distribution

(PMAD), or Power Conditioning (PC), subsystem radiator size and mass requirements was evaluated for three

power output levels (100 kWe, 1 MWe, and 10 MWe) for near term technology ( i.e. 1500 K turbine inlet temperature) Closed Cycle Gas Turbine (CCGT) power systems with a High Temperature Gas Reactor (HTGR) heat

source. The study was conducted for assumed PC radiator temperatures ranging from 370 to 845 K and for three

scenarios of electrical energy to heat conversion levels which needed to be rejected to space by means of the PC

radiator. In addition, during part of the study the radiation hardness of the PC electronics was varied at a fixed

separation distance to estimate its effect on the mass of the instrument rated reactor shadow shield. With both the

PC radiator and the conical shadow shield representing major components of the overall power system the influence of the above on total power system mass was also determined. As expected, results show that the greatest

actual mass savings achieved by the use of SiC electronics occur with high capacity power systems. Moreover,

raising the PC radiator temperature above 600 K yields only small additional system mass savings. The effect of

increased radiation hardness on total system mass is to reduce system mass by virtue of lowering the shield mass.

14. SUBJECT TERMS

Silicon carbide; Space power systems

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Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-1B

298-102