High Performance Thermal Switch for Lunar and Planetary

ICES-2020-145

Copyright © 2020 Jet Propulsion Laboratory, California Institute of Technology

High Performance Thermal Switch for Lunar and Planetary

Surface Extreme Environments

David C. Bugby1 and Jose G. Rivera2

Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA 91109

This paper describes a high performance thermal switch for lunar/planetary extreme

environments. This device has been given the name Reverse-Operation DTE Thermal Switch

(or ROD-TSW). Two different prototypes were designed, built, and tested. Their basis of

operation is the mating/de-mating of parallel flat metal surfaces driven by the differential

thermal expansion (DTE) between mid-to-high CTE metal/polymer or metal-only end-pieces

and a low CTE, low thermal conductivity (k) metal/polymer support rod. The requirements

were to be fully ON above 300 K and fully OFF below 260 K. A series of four tests were carried

out to verify/qualify the prototypes. In the first (non-vacuum) test, the thermal switches were

cooled with freeze-spray and the non-conductive rod in the OFF thermal path allowed

electrical resistance to confirm the 273 K as-designed ON/OFF actuation temperature. In the

second (vacuum) test, a calibrated Q-meter was used, which indicated 5 W/K ON, 0.002 W/K

OFF conductance (2500:1 ON/OFF ratio). To increase readiness to TRL6, a vibration test with

pre/post thermal cycling followed by a 36-day test in a relevant environment were also carried

out. This paper describes the development program and on-going related work at JPL.

Nomenclature

CTE = coefficient of thermal expansion (K-1)

DTE = differential thermal expansion

* = effective emissivity

G = conductance (W/K)

GEVS = Goddard environmental verification standard

k = thermal conductivity (Wm-1K-1)

MER = Mars Exploration Rover

QM = Q-meter

RHU = radioisotope heater unit

ROD-TSW = reverse-operation DTE thermal switch

TS = thermal strap

TSW = thermal switch

I. Introduction

VER the next several years, NASA and the commercial sector are reportedly planning lunar/planetary exploration

missions that will require a new class of surface science instruments that can make sustainable high-value

measurements over an extended number of day/night temperature/power cycles. Instruments that can provide this

capability without radioisotopes, by combining advanced thermal management capabilities and solar/battery power,

will have a higher likelihood of NASA selection and mission implementation due to their: (1) lower cost, easier/faster

producibility, and reduced public scrutiny, as they will have avoided radioisotope heat sources, which are costly,

relatively scarce, and (from a lay perspective) environmentally risky; and (2) extended operability/survivability, which

will enable them to generate science data over a long period of time that will likely extend well beyond the operational

lifetimes of early commercial lunar/planetary landers (or rovers) upon which they will reside.

To support this new class of instruments, including magnetometers, seismometers, IR spectrometers, various other

instruments, and instrument suites, JPL is developing several improved thermal management “toolbox” elements

1 Technologist, Special Programs Thermal Engineering, T1708/107, 4800 Oak Grove Dr, Pasadena, CA 91109. 2 Group Supervisor, Instrument/Payload Thermal Engineering, B125/220B, 4800 Oak Grove Dr. Pasadena, CA 91109.

O

International Conference on Environmental Systems

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including: (a) high performance thermal switches; (b) thermal switch capable enclosures; (c) easily fabricated

parabolic reflector radiators (PRRs); (d) low effective emissivity (*) multi-layer insulation (MLI); and (e) low

conductance (G) thermal isolators. With the aforementioned list of thermal toolbox elements as a complementary

backdrop, this paper focuses on a patent-pending JPL invention known as the reverse-operation DTE thermal switch

(ROD-TSW). This thermal switch has demonstrated an ON/OFF ratio of 2500:1 (5 W/K ON, 0.002 W/K OFF), which

outperforms the NASA state-of-the-art paraffin thermal switch (flown on MER) by about 25 times.

The paper is organized into the following six remaining sections. Section II (Background) describes the

development history of the reverse-operation DTE thermal switch. Section III (Concept) describes the thermal switch

requirements and operating principles. Section IV (Design) describes the thermal switch design features and pre-build

predictions of their thermal performance. Section V (Fabrication) describes the key manufacturing methods developed

for thermal switch fabrication. Section VI (Testing) provides an overview of a four-step sequence of tests that were

utilized to take the ROD-TSW from idea to TRL6 in about one year. Finally, Section VII (Conclusion/Future Work)

summarizes the key findings of this research, provides the major conclusions, and describes on-going related work at

JPL to further enhance the performance of the ROD-TSW. One of those enhancements is an extended stroke ROD-

TSW that utilizes multiple stages and a negative CTE material known as Allvar to increase DTE stroke by ten times

for operation in non-vacuum extreme environments.

II. Background

The reverse-operation DTE thermal switch (ROD-TSW) has a novel albeit lengthy development history. Figure 1

illustrates that history, which involves several thermal switch forerunners developed over the last 23 years as follows.

In 1997, JPL contacted Swales Aerospace (now Northrop Grumman) to repair a thermal short in a conical copper (Cu)

gas-gap thermal switch. In 1998, Swales obviated the need to repair the JPL thermal switch by replacing it with a flat

gas-gap version with beryllium (Be) end-pieces, a titanium (Ti) support rod, and a stainless steel (SS) bellows. In

1999, a young (at that time) Swales engineer (Brian Marland) came up with a new thermal switching idea that totally

eliminated the gas and hydride pump from the gas-gap thermal switch. That novel idea was the first generation normal-

operation DTE thermal switch, which had Be end-pieces, a SS rod, and an exceedingly small 1-2 mil gap. In 2001,

Swales came out with the second generation normal-operation DTE thermal switch, which used 99.999% (5-9s) pure

aluminum (Al) end-pieces, an Ultem 1000 rod, and a slightly larger 4 mil gap. In 2003, Swales developed two second

generation normal-operation DTE thermal switches for the James Webb Space Telescope (JWST) using Al and Ultem

1000 (neither of those thermal switches made it onto JWST). Finally, in 2012, a slightly different normal-operation

DTE thermal switch with Al and Al/Invar end-pieces and a SS rod was flown as part of the SAM instrument on the

Mars Curiosity Rover. How this development history led to the ROD-TSW is discussed in the next section.

Figure 1. ROD-TSW development history – thermal switch forerunners and their flight heritage.

1997Conical Gas-Gap: Cu/SS

1999Normal DTE: Be/SS

2003 … JWSTNormal DTE: Al/Ultem

1998Flat Gas-Gap: Be/Ti/SS

2012 … SAM FlightNormal DTE: Al/SS/Invar

2001Normal DTE: Al/Ultem


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III. Concept

The normal-operation DTE thermal switches described in the preceding section are not able to thermally isolate

lunar (or planetary) instruments from the cold nighttime environment because that type of thermal switch starts out

OFF at room temperature and turns ON as the environment cools. A normal-operation DTE thermal switch is typically

utilized to thermally couple two cryocoolers to a single cryogenic component where it is desired to: (a) operate just

one cryocooler at a time; and (b) minimize the parasitic heat leak to the cryogenic component from the OFF cryocooler.

In a normal-operation DTE thermal switch, the support rod has a higher CTE than the end-pieces. What was needed

for lunar (or planetary) extreme environments was a reverse-operation DTE thermal switch, where the support rod has

a lower CTE than the end-pieces.

In 2014, the lead author of this paper left Swales Aerospace (then ATK), joined JPL, and over the next three years

submitted several unsuccessful internal JPL and external NASA proposals to build and qualify a ROD-TSW. Finally,

in 2017, JPL provided a small amount of R&TD (Research and Technology Development) funding to give the idea a

try. That idea was to use DTE as the thermal switching mechanism, but reverse the actuation direction so that the

thermal switch starts out ON at room temperature and turns OFF as the environment cools. To reverse the actuation

direction of a normal-operation DTE thermal switch, the high CTE, low k support rod needed to be replaced by a low

CTE, low k support rod, as illustrated in Figure 2. The design requirements for the ROD-TSW were to be fully ON

(closed gap with a pre-load of roughly 1 kN) at 300 K and fully OFF (open gap) at 260 K.

IV. Design

Two ROD-TSW prototypes were designed and built as illustrated in Figure 3. DESIGN-1 has a 6063-Al large end-

piece, a 6061-Al small end-piece, an Invar 36/Ultem 2300 rod, and a compact mounting flange. DESIGN-2 has an

Ultem 1000/6063-Al large end-piece, a 6061-Al small end-piece, an Invar36/Ultem 2300 rod, and a heritage mounting

flange. The compact mounting flange had a few performance abnormalities, thus future ROD-TSWs will all use the

heritage mounting flange or variations thereof. The key dimensions of both prototypes and their predicted ON/OFF

conductance values are included in Figure 3. Actual ON conductance will be seen later to exceed these initial

predictions. Figure 4 illustrates how the two ROD-TSW prototypes are designed for box-like instrument enclosures.

Figure 2. Conceptual underpinnings of normal-operation and reverse-operation DTE thermal switches.

Normal-Operation Reverse-Operation

Figure 3. ROD-TSW prototypes, envelope dimensions, and (pre-build) predicted thermal performance.

DESIGN-1: Compact Mounting Flange DESIGN-2: Heritage Mounting Flange

Predicted Performance

GON = 2.8 W/K

GOFF = 1.8E-3 W/K

G-Ratio = 1560

TON/OFF = 273 K

DESIGN-1: Compact Mounting Flange

Envelope: Length = 126 mm, Max. Diameter = 35 mmMass: 137 grams

Envelope: Length = 86 mm, Max. Diameter = 55 mmMass: 142 grams

DESIGN-2: Heritage Mounting Flange


GON = 2.8 W/K

GOFF = 1.8E-3 W/K

G-Ratio = 1560

TON/OFF = 273 K


GON = 1.8 W/K

GOFF = 1.1E-3 W/K

G-Ratio = 1640

TON/OFF = 283 KEnvelope: L = 86 mm, D (max) = 55 mmMass: 142 grams

Envelope: L = 126 mm, D (max) = 35 mmMass: 137 grams


GON = 1.8 W/K

GOFF = 1.1E-3 W/K

G-Ratio = 1640

TON/OFF = 283 K


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V. Fabrication

Conventional machining methods were used to fabricate ROD-TSW prototype parts. In fact, the entire fabrication

process is completely standard except for mating surface preparation. To ensure superior ON/OFF performance, the

mating surfaces must be exceedingly smooth and highly parallel. Figure 5 illustrates the key elements of that process.

To achieve a 1 kN contact force at room temperature, the Invar 36/Ultem 2300 rod needs to be sized shorter than the

adjacent space it occupies by a small offset 1. The magnitude of 1 varies as a function of the desired ON force at

room temperature and the desired actuation temperature. If 1 is increased, the ON force at 300 K increases and the

actuation temperature decreases. But the actuation temperature can only be varied over a relatively narrow range. The

target actuation temperature for lunar/planetary instruments is 273 K. The magnitude of 1 is about 0.05 mm (2 mils).

VI. Testing

The testing performed to verify ROD-TSW performance and raise its readiness to TRL6 involved a sequential

series of four test segments as follows: (a) benchtop tests to ascertain ON/OFF actuation temperature; (b) thermal

vacuum tests with a calibrated Q-meter to determine ON/OFF conductance; (c) vibration test to GEVS protoflight

levels with pre/post-test thermal vacuum cycling; and (d) long duration (36-day) thermal vacuum test in a relevant

lunar environment. Those four test segments are described briefly in the next four subsections.

Figure 4. ROD-TSWs are designed to attach to box-like instrument enclosures (DESIGN-2 shown here).

Figure 5. ROD-TSW fabrication process is conventional but requires special steps for mating surface prep.

• Mating surfaces hand-lapped and polished to near-mirror finish

• Bosses in Invar 36 and Ultem 2300 keep parts in stable position

• Machine, profile scan of rod while on end-piece for parallelism, offset (1)

• Bellville washers (used on some previous DTE switches) not used

DE

SIG

N-1

DE

SIG

N-2

6061-Al

6061-Al

6063-Al

Ultem 1000

6063A

l

Invar 36

Invar 36

Ultem

2300

Ultem

2300

Parallel to

<<< 1 mil

Parallel

but out of

plane

(lower) by

design

offset (1)


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A. Benchtop Actuation Test

To ascertain ROD-TSW actuation temperature, a few simple tests were performed. As indicated in Figure 6, each

prototype was sprayed with a hand-held aerosol freeze-spray while monitoring temperature and electrical resistance

across the mating surfaces. At room temperature, the electrical resistance is very low (closed circuit). When cooled

enough so that the mating surfaces are no longer in contact, the electrical resistance becomes exceedingly high (open

circuit). Both prototypes indicated ON-to-OFF and OFF-to-ON actuation temperatures within a few degrees of their

design values. An additional actuation test, as indicated in Figure 6, was performed where both prototypes were put

into a freezer for one hour and then removed. This test ascertains just the OFF-to-ON actuation temperature, and again

electrical resistance/temperature monitoring indicated actuation temperatures of the two prototypes in agreement with

design values. It should be noted that the presence of the Ultem 2300 plug in the support rod allows these simple tests

to work even if the end-pieces are all aluminum (as is the case with DESIGN-1).

B. Thermal Vacuum Tests with Calibrated Q-Meter

Following the successful benchtop actuation test, which demonstrated that the two prototype thermal switches

turned ON and OFF as designed, the next testing step was to measure their conductance and cycle them ON/OFF in a

thermal vacuum chamber. This second test segment would incorporate an intra-chamber cryocooler cold head for

cooling and a calibrated Q-meter to measure heat flow. Figure 7 illustrates the cryocooler expander used and the five

distinct test setups that were needed to perform this test.

Setups 1-3 were used for conductance and temperature cycling while Setups 4-5 were used for Q-meter calibration.

The conductance values measured for both prototypes were 5 W/K (ON) and 0.002 W/K (OFF), respectively.

Temperature cycling data from Setup 4 indicated that DESIGN-2 is sensitive to cooling/heating rate (low k body)

while DESIGN-1 is not (high k body). Figures 8-9 illustrate DESIGN-1 and DESIGN-2 temperature cycling results,

respectively. As can be inferred from Figure 9, if the cooling rate is lower than about 0.5 K/min, DESIGN-2 is

insensitive to cooling rate. To assess (rate-sensitive) DESIGN-2 viability, lunar surface cooling rate was postulated as

the controlling parameter. Diviner1 data in Figure 10 indicates a max cooling rate of about 0.06 K/min. Thus, both

ROD-TSWs appeared viable at this point. But two key issues associated with Figures 8-10 needed to be looked into.

Figure 6. ROD-TSW actuation tests with temp./elec. resistance monitoring (photo of DESIGN-2 OFF gap).

DESIGN-1 DESIGN-26061-Al 6061-Al

6063-A

l

Ult

em

1000

6063 Al

Inva

r 3

6

Inva

r 3

6Ultem

2300

Ultem2300

Perspective View Top View Side ViewDesign-2 Gap After Cooling

During Part 1 TestingFreeze-Spray TestON-to-OFF-to-ON

Freezer TestOFF-to-ON

DESIGN-2 OFF Gap

Figure 7. ROD-TSW thermal vacuum test setups for conductance measurements and temperature cycling

Setup-1: ON Testing

CH QM TS

TSW

ENCL

HTR

g

CH QM TS

TSW

ENCL

HTR

g

Setup-2: OFF Testing

TS

W

CHQM

BRKT

g

CH QM TS

TSW

ENCL

HTR

g

Setup-3: Cycling

TS

W

BRKTQM

MINI-

ENCL

CH

HTR

g

CH QM TS

TSW

ENCL

HTR

g

TS

W

CHQM

BRKT

g

Setup-4: QM Cal w/ Heater

TS

W

CHQM

BRKT

g CH QM TS

TSW

ENCL

HTR

g

TU

TL

Setup-5: QM Cal Adj.

TS

W

CHQM

BRKT

g

CH

QM

TS

TSW

ENCL

HTR

g

CH QM TS

TSW

ENCL

HTR

g

AR

S, Inc. Exp

and

er DE-1

02

F

GM Cold HeadARS DE-102F

LEGEND*BRKT = bracketCH = cold headENCL = enclosureg = gravityGM = Gifford-McMahonHTR = heaterTL = lower temperatureTU = upper temperature

*terms not defined in Nomenclature


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Figure 8. ROD-TSW DESIGN-1 temperature cycling results indicate no sensitivity to cooling rate.

0

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100

150

200

250

300

350

0 500 1000 1500 2000 2500 3000 3500 4000

Tem

per

atu

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Time (minutes)

Design-1/Setup-3 Test Data

T_sw_cold_avg_SW1 T_sw_hot_avg_SW1

Actuation OFF

270 K … Cold

Head at 163 K

(3rd cycle)

Actuation OFF

270 K … Cold

Head at 163 K

(1st cycle)

Heater ON

(16.5 W, 315 K)

Heater ON

(6.6 W, 315 K)

Actuation OFF

273 K … Cold

Head at 250 K

(4th cycle)

Actuation OFF

275 K … Cold

Head at 265 K

(5th cycle)

Heater OFF

(300 K Control,

Cold Head OFF)

Heater ON

(16.5 W, 315 K)

Heater ON

(16.5 W, 315 K)

Actuation OFF

270 K … Cold

Head at 163 K

(2nd cycle)

Actuation OFF

270 K … Cold

Head at 33 K …

Rapid Cooldown

(6th cycle)

Figure 9. ROD-TSW DESIGN-2 temperature cycling results indicate sensitivity to cooling rate > 0.5 K/min.

0

50

100

150

200

250

300

350

0 500 1000 1500 2000 2500 3000 3500 4000

Tem

per

atu

re (

K)

Time (minutes)

Design-2/Setup-3 Test Data

TSWC_AVG TSWH_AVG

Heater ON

(16.5 W, 315 K)

Actuation OFF

195 K … Cold

Head at 163 K

(1st cycle)

Heater ON

(16.5 W, 315 K)

Heater ON

(8 W, 315 K)

Actuation OFF

260 K … Cold

Head at 33 K …

0.5 K/min Rate

(6th cycle)Actuation OFF

187 K … Cold

Head at 163 K

(2nd cycle)

Actuation OFF

188 K … Cold

Head at 163 K

(3rd cycle)

Actuation OFF

260 K … Cold

Head at 250 K

(4th cycle)

Heater ON

(5.5 W, 315 K)

Actuation OFF

275 K … Cold

Head at 270 K

(5th cycle)

Heater OFF

(300 K Control,

Cold Head OFF)

Figure 10. Diviner lunar surface temperature measurements indicate a max cooling rate of ~ 0.06 K/min.

Slope = 6.28”/1.22”(from PowerPoint Size button)5.95” = 350 K8.00” = 24 lunar hrsdT = 350*(6.28/5.95) Kdt = 30*24*1.22/8 Earth hr(30 Earth hr = 1 Lunar hr)dT/dt = 369.4/109.8dT/dt = 3.36 K/(Earth) hr

Lunar Surface Temperature Diviner Data

Max Cooling Rate = 0.056 K/min … so DESIGN-2 is Viable

Straight line overlaid onto Diviner lunar surface temp

data to get max cooling rate

Scaling of line + curves yields dT/dt = 3.36 K/(Earth) hour

max cooling rate (0.06 K/min)

Paper (by J.P. Williams)where Diviner lunar surface temp data extracted from


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The first technical issue involves both ROD-TSWs during the 5th cycle, as shown in Figures 8-9, where the cold

head is at 265 K. Due to the small hot-to-cold side Ts that are apparent, it is likely that neither ROD-TSW actuated

fully OFF during this cold (OFF) half-cycle. DESIGN-1 appears to be making grazing contact while DESIGN-2

appears to be making a little firmer contact. Since the cold head is at 265 K and the two ROD-TSWs actuate OFF at

273 K, this result is not surprising. Full OFF actuation will undoubtedly occur upon further cooling, but at these (near

273 K) temperatures, the ROD-TSWs provide a more gradual (variable conductance-like) actuation behavior.

The second technical issue, which is associated with Figures 9-10, involves the following interrelated concerns:

(a) whether the rate of change of lunar surface temperature is the sole time parameter against which ROD-TSW

viability should be judged; and (b) whether the inference that DESIGN-2 is insensitive to heating/cooling rates less

than 0.5 K/min is valid. Relative to concern (a), the lunar surface response is just one part of the overall problem as

other key parameters are instrument power (Q), mass (m), and heat capacity (Cp). In the limit of no cooling, or in the

limit of max cooling when power is turned off, instruments heat up or cool down at a max rate of Q/(mCp). If the heat

capacity is 900 Jkg-1K-1 (value for Al), a heating/cooling rate less than 0.5 K/min requires Q/m < 7.5 W/kg, which will

likely be true for virtually all lunar instruments. Relative to concern (b), once it was observed that DESIGN-2 actuated

OFF at a temperature near 273 K (and the stated value in Figure 9 for cycle 6 is 260 K), further testing at lower cooling

rates was unnecessary. This statement is based on four things: (i) ON/OFF actuation is driven by average temperature;

(ii) with high heating/cooling rates, average temperature changes more slowly due to Ultem 1000 low k, resulting in

a large temperature gradient; (iii) transition near 273 K implies the Ultem 1000 temperature gradient is small; and (iv)

once rates are low enough, and internal gradients stay small, further rate reductions won’t affect actuation temperature.

Thus, from a thermal performance standpoint, both ROD-TSW prototypes are viable for lunar applications.

C. Vibration Test to GEVS Protoflight Level with Pre/Post-Vibe Thermal Cycling

The third test performed to raise ROD-TSW technical readiness was a vibration test to GEVS protoflight level

with pre/post-vibe thermal cycling. The thermal cycling test results are illustrated in Figure 11, which includes

pre/post-vibe temperature response curves for the hot and cold sides of both prototypes. DESIGN-1 was cycled three

times between 300 K (27 °C) and 173 K (-100 °C) and there was no change in its thermal response as the pre-vibe and

post-vibe curves are identical. Using the same heater power that the DESIGN-1 test used for heating half-cycles,

DESIGN-2 very briefly reached a hot side temperature of 373 K (100 °C). This very brief high temperature exposure,

which DESIGN-2 materials can withstand indefinitely, is due to the previously discussed cooling/heating rate

sensitivity. In addition, there was a slight difference in the DESIGN-2 hot side thermal response during cold half-

cycles. While the reason for this slight difference is not clear, it had no material impact on DESIGN-2 thermal

performance and the vibration test was deemed a success. However, before proceeding to the next test (the long

duration test in a relevant environment), a brief discussion of DESIGN-2 pre/post-vibe thermal results is included.

Figure 11. ROD-TSW pre/post-vibe thermal cycling indicates no material change* in performance (*reason

for small differences in DESIGN-2 profile unclear, but there has been no impact on its thermal performance).

No change in DESIGN-1 Thermal Response(3 cold-hot cycles before vibe, 3 cold-hot cycles after vibe)

No material change* in DESIGN-2 Thermal Response(3 cold-hot cycles before vibe, 3 cold-hot cycles after vibe)

-140.0

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)

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Alum-TSWH

Pre-Vibe Post-Vibe

DESIGN-1 HOT SIDE

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DESIGN-2 HOT SIDE

THERMAL TESTCONFIGURATION

COLD PLATE

HTR

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)

COLD(OFF)

HOT(ON)


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With respect to the DESIGN-2 hot side in Figure 11, cold (OFF) half-cycle temperatures traces, although modestly

different in temporal profile, appear as if they might have ended up very close had the cycles been longer, while hot

(ON) half-cycle (post-peak) temperature traces are very close. With respect to the DESIGN-2 cold side in Figure 11,

cold (OFF) half-cycle temperatures are nearly identical, while hot (ON) half-cycle pre-vibe temperatures are a few K

lower. These observations suggest a slightly higher DESIGN-2 post-vibe ON conductance (highly unlikely), but

unchanged OFF conductance. Subsequent DESIGN-2 testing indicated no change in ON or OFF performance.

D. Long Duration Test in Relevant Lunar Environment

The final ROD-TSW qualification test was a long duration thermal vacuum test in a relevant (lunar surface)

environment. The key features of this test are as follows. Each ROD-TSW prototype was mounted to a cube-shaped

enclosure denoted as an Instrument Cube (IC) identical to that shown in Figure 4. Each IC, made from 6061-Al and

left bare (uncoated) inside/outside, was a 0.127 m x 0.127 m x 0.127 m enclosure with 0.004 m thick walls. A small

radiator plate of dimensions 0.178 m x 0.356 m x 0.001 m was screwed onto the small end-piece of each ROD-TSW.

The 6061-Al radiator plates were black anodized on their radiating side (surface opposing the IC) and left bare on

their non-radiating side (surface facing the IC). The ICs were covered with a twenty-layer conventional multilayer

insulation (MLI) blanket with a circular opening to accommodate the ROD-TSW flange. The ICs were suspended by

Kevlar cables from the inside of a black anodized (inside/outside) lunar shroud (LS) that was a small environmental

enclosure sized large enough (0.381 m x 0.381 m x 0.254 m x 0.003 m thick walls) to accommodate two ICs and their

radiators. The LS was bolted to an LN2-cooled cold plate inside a thermal vacuum (TVAC) chamber. The test was

conducted over 36 days with six hot half-cycles (280 K cold plate, 8W to each IC internal heater) and six cold half-

cycles (88 K cold plate, 0-1.5 W to each IC internal heater) that were three days each. This temperature cycling

approach represents an approximately 5X acceleration of the lunar day. The selection of 280 K as the hot case

temperature represents the worst case hot sink temperature of a parabolic reflector radiator (PRR) on the 380-400 K

lunar surface. Photos of the test setup and a thermocouple placement diagram are provided in Figure 12.

The long duration test results are illustrated in Figure 13. This graph shows the hot side temperature of the two

ROD-TSW prototypes. For DESIGN-2, this temperature is TC4 from Figure 12, recorded as TC104. For DESIGN-1,

this temperature is TC10 from Figure 12, recorded as TC110. Overall, the two prototypes exhibited highly repeatable,

predictable (identical) ON/OFF performance except for the following issue. During the 4th cold cycle, DESIGN-1 hot

side got about 15 K cooler than the DESIGN-2 hot side. Then, during the 6th cold cycle, the cooling effect had

vanished. Upon removal of the setup from the chamber, two DESIGN-1 Kevlar cables had broken (leaving it totally

unsupported by the Kevlar) and one DESIGN-2 Kevlar cable had broken (leaving it partially supported by the Kevlar).

As illustrated in Figure 14, the DESIGN-1 radiator ended up resting benignly (totally level, still thermally isolated

from the LS) on the eye-bolts from which the Kevlar cables were strung, while the DESIGN-2 unit was canted slightly

because it was still being held up partly by its Kevlar cables. The theory as to what happened is provided below.

It is postulated that during (or prior to) the 4th cold cycle, one DESIGN-1 cable broke creating a side load on the

thermal switch support rod which caused its thermal switch mating surfaces to grazingly touch, which slightly

increased the OFF conductance and caused it to cool a little more than it should have. From thermal model results,

that grazing contact increased OFF conductance by about 25%. Then, during (or prior to) the 6 th cold cycle, a second

DESIGN-1 cable broke and the unit came to rest on the eye-bolts in a perfectly unstressed manner. The break of just

one DESIGN-2 Kevlar cable caused it be slightly canted, but because of its shorter rod and higher resistance to side

Figure 12. Setup, Thermocouple Diagram for ROD-TSW Long Duration Test in Relevant Environment

TC13

TC16

TC15

Switch 2 Switch 1

LN2-Cooled Cold Plate

Rad 2 Rad 1

Cube 2 Cube 1

Shroud

Photos of Test SetupD1,2 in Lunar ShroudD2 Instr. Cube (IC) Kevlar Suspension Design-2 IC MLI Design-1 IC MLI

LS Front/Top AddedD1,2 in Lunar Shroud (LS) w/ Rads LS in TVAC (SLI added)LS in TVAC (before SLI)

Test Setup, Instrumentation Diagram


9

loads, it was unaffected by a single broken Kevlar cable. Overall, the impact of the broken cables was negligible. The

reason the Kevlar cables broke is excessive surface roughness and small cable bend radii at the IC attach points.

In Figure 13, only hot side temperatures were presented because the two ROD-TSWs have nearly identical ON

and OFF conductance. Thus, just one side is needed to discern a performance change. In addition, the only cycle where

the two ROD-TSWs deviated in response was Cycle 4. Since this deviation vanished in Cycle 6, it was concluded

Cycle 4 was the anomaly. Thus, we hypothesized that DESIGN-1 suffered one broken cable prior to Cycle 4, causing

a side load, and a second broken cable prior to Cycle 6, relieiving the side load. While not conclusive, this hypothetical

explanation appears to be entirely plausible.

Figure 13. Long duration test results show repeatable, predictable behavior (cable failures had no effect).

-150

-125

-100

-75

-50

-25

0

25

50

75

0 10000 20000 30000 40000 50000 60000

Tem

per

atu

re (

deg

C)

Time (min)

TC104 (Switch 2-Hot), TC110 (Switch 1-Hot) Data

Data Switch 1

Data Switch 2

0 W, 88 K

8 W280 K

0.5 W88 K

1.0 W88 K

1.5 W88 K

8 W280 K

8 W280 K

8 W280 K

8 W280 K

8 W280 K

1.0 W88 K

1.5 W88 K

IC Heater PowerLN2 Plate Temp.

Figure 14. DESIGN-1, 2 after removal from chamber showing benign impact of broken Kevlar cables.

DESIGN-1

DESIGN-2

DESIGN-1

DESIGN-2


10

VII. Conclusion/Future Work

The conclusion of this effort is that the reverse-operation DTE thermal switch (ROD-TSW) is now at TRL6 and

is therefore ready for use on an actual mission. In fact, each ROD-TSW prototype – DESIGN-1, with its Al end-pieces,

and DESIGN-2, with its Al and Al/polymer end-pieces – has demonstrated repeatable, reliable, totally passive thermal

switching with ON/OFF actuation at 273 K, ON conductance of 5 W/K, and OFF conductance of 0.002 W/K. This

2500:1 ON/OFF ratio outperforms the NASA state-of-the-art MER paraffin thermal switch by 25 times.

With respect to the likelihood of future flight implementation, two DESIGN-1 style ROD-TSWs, under NASA

Lunar CATALYST program funding provided by NASA Marshall Space Flight Center (MSFC) to JPL, have been

recently delivered by JPL to Astrobotic as part of a thermal switching system that will fly on the initial Astrobotic

Peregrine lunar lander. This thermal switching system, which includes an Al bracket, two DESIGN-1 ROD-TSWs,

and two Pyrovo graphite thermal straps built by Thermotive, is intended to prevent a high power transponder (similar

to the transponder made by Space Micro for the LADEE mission) from overtemperature during operation by providing

passive thermal switching to an adjacent radiator panel. A future paper will describe this system.

Lastly, the two authors of this paper are currently working on three related efforts at JPL that include the following:

(1) a NASA Center Innovation Fund (CIF) project in conjunction with NASA/MSFC to design, build, and test an

extended stroke ROD-TSW with multiple stages and a negative CTE material known as Allvar to provide a larger

OFF gap for use in non-vacuum or vacuum environments; (2) an internal JPL Strategic Research and Technology

Development (SR&TD) program to develop advanced thermally-switched enclosures for lunar instrument suites; and

(3) a recently awarded NASA STMD project, run by the NASA Game Changing Development (GCD) program at

NASA Langley Research Center, called Planetary and Lunar Environment Thermal Toolbox Elements (PALETTE),

which aims to develop a new class of better performing thermal management tools (thermal switched enclosures,

PRRs, MLI, thermal isolators, and other elements) for future instruments operating in extreme environments. Funding

for the three-year PALETTE project had just been provided to JPL by NASA as this paper was being written.

Appendix

One topic not covered in the main body of this paper is thermal modeling. One of the final steps taken near the

end of this program was the development of a correlated thermal model of the long duration test that was able to

accurately reproduce the measured data with just three correlating parameters: (1) instrument cube (IC) MLI effective

emissivity (*); (2) lunar shroud (LS) top-to-bottom radiative coupling (GR); and (3) LS top-to-bottom conduction

coupling (GC). The reasons for selecting correlation parameters 2 and 3 are as follows: (a) each IC radiator looked at

the LS top; (b) the LS bottom was bolted to the LN2-cooled cold plate; (c) the LS did not have a very high top-to-

bottom conductance; and (d) there was a modest LS top-to-LS bottom temperature difference. As a result of these

factors, the two conductance values were relatively uncertain, hence they needed to be adjusted (from initial estimates)

so the model would acceptably match the test data. The MLI e* that was selected was 0.04. Eventually, after adjusting

parameters 2 and 3 in an iterative fashion, the thermal model was able to reproduce nearly all the long duration test

data with surprising accuracy. The GR and GC values that were ultimately chosen were 0.13 m2 and 0.15 W/K,

respectively, values that are not far from initial estimates. This modeling work may be presented in a future paper.

Acknowledgments

The authors would like to acknowledge several individuals for their contributions to this effort. First, thanks to

Tim O’Donnell, Garry Burdick, Dave Eisenman, Sabrina Feldman, and Ying Lin for their great help in securing

internal JPL funding for this development work. Next, thanks to Virgil Mireles, Chuck Phillips, and Gani Ganapathi

for their line management support of this work at JPL. Lastly, thanks to the JPL scientists, engineers, and/or managers

who contributed to the success of this program including Pamela Clark, Doug Hofmann, Bob Kovac, Bart Patel, Mike

Dombrower, Tri Huynh, Mason Mok, Chris Hummel, Eric Sunada, Rob Staehle, and John Elliott. This research was

carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National

Aeronautics and Space Administration (80NM0018D0004).

References 1Williams, J. P., Paige, D. A., Greenhagen, B. T., and Sefton-Nash, E., “The global surface temperature of the Moon as

measured by the Diviner Lunar Radiometer Experiment,” Icarus, Vol. 283, February, 2017, pp. 300-325.

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High Performance Thermal Switch for Lunar and Planetary