Technical Memorandum 104822

Advanced Active Thermal Control Systems

Architecture Study

A. J. Hanford, Ph.D., and M. K. Ewert

October 1996

Technical Memorandum 104822

Advanced Active Thermal Control Systems Architecture Study

A. J. Hanford, Ph.D.

Lockheed Martin Engineering and Science Services

Houston, Texas 77058

and

M. K. Ewert

Lyndon B. Johnson Space Center

National Aeronautics and Space Administration

Houston, Texas 77058

October 1996
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

ACKNOWLEDGMENTS

The authors would like to express their thanks for the extensive review that this and earlier

versions of this document received from the following people:

From the National Aeronautics and Space Administration, Lyndon B. Johnson Space Center

(NASA/JSC), Houston, Texas: A.F. Behrend, J.D. Cornwell, C.D. Cross, K.M. Hurlbert,

Dr. C.H. Lin, P. Petete, J.G. Rankin, and Dr. E.K. Ungar.

From Lockheed Martin Engineering and Science Services (LMES), Houston, Texas:

B.C. Conger, J.F. Keener, and Dr. J.R. Keller.

From the National Aeronautics and Space Administration, Goddard Space Flight Center

(NASA/GSFC), Greenbelt, Maryland: T.D. Swanson.

From the National Aeronautics and Space Administration, Langley Research Center

(NASA/LaRC), Hampton, Virginia: W.A. Sasamoto.

From the National Aeronautics and Space Administration, Lewis Research Center (NASA/LeRC),

Cleveland, Ohio: M. Hill.

From the National Aeronautics and Space Administration, Marshall Space Flight Center

(NASA/MSFC), Huntsville, Alabama: J. Holiday.

The authors would also like to recognize those individuals who added their technical input or

provided literature sources in support of various topics throughout this report: Dr. K.K. Andish

(LMES), B.H. Blades (LMES/Dual), J. Bryan (Texas A&M University, College Station, Texas),

R. Cataldo (NASA/LeRC), J.D. Cornwell (NASA/JSC), C.D. Cross (NASAJJSC), D.L. Farner

(LMES), R.A. Henson (LMES), A.J. Juhasz (NASA/LeRC), J.F. Keener (LMES), Dr. J.R. Keller

" (LMES), Dr. J. Ku (NASA/GSFC), G.W. Lucas (LMES), J. Oren (Loral Vought Systems, Dallas,

Texas), T.H. Paul (LMES), W.P. Ruemmele (LMES), and Dr. E.K. Ungar (NASAJJSC).

This publication is available from the NASA Center for AeroSpace Information,

800 Elkridge Landing Road, Linthicum Heights, MD 21090-2934, (301) 621-0390.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

TABLE OF CONTENTS

Section Page

1.0

2.0

INTRODUCTION ............................................................................................... 1

ADVANCED THERMAL CONTROL TECHNOLOGIES .................................. 5

2.1 Two-Phase Thermal Control Systems ....................................................... 5

2.1.1 Two-Phase Thermal Control System

With Mechanical Pump/Separator ................................................. 5

2.1.2 Low-Power Two-Phase Thermal Control System .......................... 7

2.1.3 Two-Phase Thermal Control System

With Electrohydrodynamic Pumping ............................................. 8

2.1.4 Capillary Pumped Loops ............................................................... 9

2.2 Heat Pumps ............................................................................................ 12

2.2.1 Vapor Compression Heat Pump .................................................. 13

2.2.2 Solar Vapor Compression Heat Pump ......................................... 14

2.2.3 Complex Compound Heat Pump ................................................. 16

2.2.4 Zeolite Heat Pump ...................................................................... 16

2.3 Heat Pipe Radiators ................................................................................ 18

2.3.1 Arterial Heat Pipe Radiators ....................................................... 19

2.3.2 Arterial Heat Pipe Radiators

With Electrohydrodynamic Pumping ........................................... 23

2.3.3 Axial-Groove Heat Pipe Radiators .............................................. 24

2.4 Lightweight Radiators ............................................................................ 25

2.4.1 Composite Flow-Through Radiators ........................................... 27

2.4.2 Composite Reflux Boiler Tube Radiators .................................... 28

2.4.3 Composite Heat Pipe Radiators .................................................. 30

2.4.4 Unfurlable Radiators ................................................................... 32

2.5 Other Heat Rejection Technologies ........................................................ 33

2.5.1 Phase-Change Thermal Storage .................................................. 33

2.5.2 Parabolic Radiator Shade ............................................................ 36

2.6 Additional Technologies ......................................................................... 37

2.6.1 Rotary Fluid Coupler .................................................................. 37

2.6.2 Plant Chamber Cooling Improvements ........................................ 38

2.6.3 Carbon Brush Heat Exchanger .................................................... 40

2.7 Summary ................................................................................................ 44

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Section Page

3.0 ORBITAL MISSIONS ...................................................................................... 47

3.1 INTERNATIONAL SPACE STATION EVOLUTION .......................... 47

3.1.1 Reference Mission ...................................................................... 47

3.1.2 Baseline Case .............................................................................. 47

3.1.3 Implementing the Reference Mission ........................................... 54

3.1.4 Parametric Study Using the Baseline Case ................................... 56

3.1.5 Advanced ATCS Architecture for

International Space Station Evolution ......................................... 63

3.1.5.1 Two-Phase Thermal Control System

With Mechanical Pump/Separator ................................. 63

3.1.5.2 Low-Power Two-Phase Thermal Control System ......... 64

3.1.5.3 Capillary Pumped Loops .............................................. 64

3.1.5.4 Vapor Compression Heat Pump ................................... 65

3.1.5.5 Solar Vapor Compression Heat Pump .......................... 68

3.1.5.6 Arterial Heat Pipe Radiators ......................................... 69

3.1.5.7 Axial-Groove Heat Pipe Radiators ................................ 74

3.1.5.8 Arterial Heat Pipe Radiators

With Electrohydrodynamic Pumping ............................. 75

3.1.5.9 Lightweight Radiators .................................................. 77

3.1.5.10 Rotary Fluid Coupler .................................................... 80

3.1.5.11 Carbon Brush Heat Exchanger ..................................... 80

3.1.6 Summary .................................................................................... 81

3.2 SPACE TRANSPORTATION SYSTEM UPGRADE ............................ 84

3.2.1 Reference Mission ...................................................................... 84

3.2.2 Baseline Case .............................................................................. 84

3.2.3 Parametric Study Using the Baseline Case ................................... 90

3.2.4 Advanced ATCS Architecture for

Space Transportation System Upgrade ........................................ 93

3.2.4.1 Low-Power Two-Phase Thermal Control System ......... 93

3.2.4.2 Lightweight Radiators .................................................. 94

3.2.4.3 Phase-Change Thermal Storage .................................... 96

3.2.5 Summary .................................................................................... 99

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Section Page

4.0 PLANETARY MISSIONS .............................................................................. 101

4.1 FIRST LUNAR OUTPOST LANDER ................................................. 101

4.1.1 Reference Mission .................................................................... 101

4.1.2 Baseline Case ............................................................................ 101

4.1.3 Parametric Study Using the Baseline Case ................................. 104

4.1.4 Advanced ATCS Architecture for

First Lunar Outpost Lander ....................................................... 107

4.1.4.1 Low-Power Two-Phase Thermal Control System ....... 109

4.1.4.2 Two-Phase Thermal Control System

With Electrohydrodynamic Pumping ........................... 110

4.1.4.3 Capillary Pumped Loops ............................................ 110

4.1.4.4 Solar Vapor Compression Heat Pump ........................ 111

4.1.4.5 Lightweight Radiators ................................................ 115

4.1.4.6 Parabolic Radiator Shade ........................................... 117

4.1.5 Summary .................................................................................. 121

4.2 PERMANENT LUNAR BASE ............................................................ 123

4.2.1 Reference Mission .................................................................... 123

4.2.2 Baseline Case ............................................................................ 123

4.2.3 Parametric Study Using the Baseline Case ................................. 128

4.2.4 Advanced ATCS Architecture for Permanent Lunar Base ......... 130

4.2.4.1 Two-Phase Thermal Control System

With Mechanical Pump/Separator ............................... 131

4.2.4.2 Low-Power Two-Phase Thermal Control System ....... 132

4.2.4.3 Two-Phase Thermal Control System

With Electrohydrodynamic Pumping ........................... 133

4.2.4.4 Capillary Pumped Loops ............................................ 134

4.2.4.5 Vapor Compression Heat Pump ................................. 134

4.2.4.6 Solar Vapor Compression Heat Pump ........................ 142

4.2.4.7 Complex Compound Heat Pump ................................ 143

4.2.4.8 Zeolite Heat Pump ..................................................... 145

4.2.4.9 Lightweight Radiators ................................................ 147

4.2.4.10 Parabolic Radiator Shade ........................................... 149

4.2.4.11 Plant Chamber Cooling Improvements ........................ 152

4.2.4.12 Carbon Brush Heat Exchanger ................................... 153

4.2.5 Summary .................................................................................. 154
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Section Page

5.0

6.0

7.0

4.3 MARS LANDER (ML) ........................................................................ 157

4.3.1 Reference Mission .................................................................... 157

4.3.2 Baseline Case ............................................................................ 157

4.3.3 Parametric Study Using the Baseline Case ................................. 160

4.3.4 Advanced ATCS Architecture for Mars Lander ........................ 167

4.3.4.1 Low-Power Two-Phase Thermal Control System ....... 167

4.3.4.2 Two-Phase Thermal Control System

With Electrohydrodynamic Pumping ........................... 168

4.3.4.3 Capillary Pumped Loops ............................................ 169

4.3.4.4 Vapor Compression Heat Pump ................................. 169

4.3.4.5 Solar Vapor Compression Heat Pump ........................ 172

4.3.4.6 Lightweight Radiators ................................................ 177

4.3.5 Summary .................................................................................. 179

CONCLUSIONS ............................................................................................. 181

5.1 General ................................................................................................ 182

5.2 Two-Phase Thermal Control Systems ................................................... 183

5.3 Heat Pumps .......................................................................................... 183

5.4 Heat Pipe Radiators .............................................................................. 183

5.5 Lightweight Radiators .......................................................................... 184

5.6 Other Heat Rejection Technologies ...................................................... 185

5.7 Additional Technologies ....................................................................... 186

RECOMMENDATIONS ................................................................................. 187

6.1 Additional Mission Scenarios ............................................................... 187

6.2 Additional Thermal Control Technology Studies .................................. 188

6.3 Additional Mission Specific Work for Current Study ............................ 189

REFERENCES ................................................................................................ 191

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

NOMENCLATURE

item

ABS

ATCS

COP

DCSU

DDCU

DDCU-E

ETCS

FCA

FES

FLO

GSE

GSE HX

H20

ISS

ITCS

low alpha

LP

LTL

LVS

MBSU

ML

MP/S

MTL

N2

NASA

NH3

02

ORU

P0

PCM

PLB

Definition

ammonia boiler subsystem

active thermal control system (ITCS plus ETCS)

coefficient of performance

dc switching unit

dc-to-dc converter unit

external DDCU

external thermal control system

flow control assembly

flash evaporator subsystem

First Lunar Outpost

ground support equipment

ground support equipment heat exchanger

water

International Space Station

internal thermal control system

low solar absorptivity

low-power

low temperature loop

Loral Vought Systems

main bus switching unit

Mars Lander

mechanical pump/separator

moderate temperature loop

nitrogen (elemental)

National Aeronautics and Space Administration

ammonia (R717)

oxygen (elemental)

orbital replacement unit

a proposed segment to be added to ISS as part of the baseline ISS

evolution mission to support additional radiator ORUs

phase-change material

Permanent Lunar Base

vii
Advanced Active Thermal

Item

PV-TCS

radiating area

SO

S1, P1

S4, S6, P4, P6

Shuttle

SINDA/FLUINT

STS

STS-41

TCS

TRL

TRRJ

TSS

VPGC

Z-93

Control Systems Architecture Study NASA TM 104822

Definition

thermal control system for photovoltaic power arrays

the total surface area of a radiator through which heat is rejected,

equal to the plan or projected area for a single-sided radiator (for

example, a horizontal radiator mounted on a planetary surface), and

equal to twice the plan or projected area for a two-sided radiator

(for example, the ATCS radiator panels on ISS's radiator ORUs or

a vertical radiator)

base truss segment for U.S.-led portion of ISS

Truss segments outboard to either side of Segment SO on ISS, to

which the TRRJs and the structure supporting the radiator ORUs

attach directly. "S" designates a starboard segment while "P"

designates a port segment.

segments outboard of the solar alpha rotary joints (Segments $3

and P3) on ISS which support the solar photovoltaic power arrays

an altemate name for Space Transportation System

Systems Improved Numerical Differencing Analyzer and Fluid

Integrator, a thermal and fluid analysis package which uses a

network analysis approach. This program was prepared under a

contract from NASA Lyndon B. Johnson Space Center. Cullimore

and Ring Technologies, Inc., continues to distribute and support

this product.

Space Transportation System (which is often called "Shuttle")

Space Transportation System Mission 41 (Discovery, October 6 to

10, 1990)

thermal control system

technology readiness level

thermal radiator rotary joint

Thermal Synthesizer System, a thermal analysis package for space

systems. This program was developed by Lockheed Missiles and

Space Company, Inc. of Sunnyvale, California under a contract

from NASA Lyndon B. Johnson Space Center. Currently TSS is

maintained and supported by Lockheed Martin Engineering and

Science Services of Houston, Texas.

Variable Pressure Growth Chamber, also known as the Johnson

Space Center 10-foot Regenerative Life Support Systems Test

Chamber

a radiator surface coating

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

ADVANCED ACTIVE THERMAL

CONTROL SYSTEMS ARCHITECTURE STUDY

SUMMARY

STUDY OBJECTIVE

The study which follows quantifies potential mass savings offered by various

technologies and identifies promising development initiatives for advanced thermal control

systems. This, in turn, provides a common basis to compare many diverse technologies.

Assessments are presented for five reference missions considered to be likely

candidates for major human space flight initiatives beyond the assembly of International

Space Station. These include:

• International Space Station Evolution

• Space Transportation System Upgrade

• First Lunar Outpost Lander

• Permanent Lunar Base

• Mars Lander

The objective of this study is to estimate potential benefits for various proposed

and under-development thermal control technologies for possible human missions early in

the next century. Twenty advanced thermal technologies currently under various stages of

development by the National Aeronautics and Space Administration (NASA) were applied

to the reference missions. These technologies represent a host of potential advanced

thermal control system architectures. As research and development progress and mission

planning and equipment mature, this study can be revised and extended to include new

data as well as additional technologies and mission scenarios.

STUDY METHODOLOGY

The study format compares all technologies to baseline missions using mass as a

basis. Power consumption is included directly by converting it to an equivalent mass using

a mission appropriate power mass penalty. Qualitative ratings are presented for each

advanced thermal control technology. The qualitative ratings allow other considerations,

besides mass, to be included in the technology assessments. The qualitative rating process

is fully described in Section 1.0 and the qualitative ratings are presented in Section 2.0.

Primary Assessments: Qualitative Assessments:

• equipment mass savings • volume

• power savings • ease of deployment or installation

• power converted to mass • reliability

• overall mass savings • development cost

(including equipment and power) • terrestrial use potential

Table 5.1 within this report gives a complete summary of mass savings while Table 2.1

provides a summary of qualitative ratings for each advanced thermal control technology.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Each baseline architecture uses single-phase flow loops to transport waste heat

from collection sites to rejection devices. More specifically, the baseline architecture

assumes:

• Heat exchangers collect heat and transfer it to flow loops.

• The internal thermal control system collects heat from the cabin occupied by humans.

This heat is rejected to the external thermal control system. The internal thermal

control system working fluid is single-phase water.

• The external thermal control system collects heat from the internal thermal control

system and other payloads outside the cabin. This heat load is rejected to the

environment. The external thermal control system working fluid is a single-phase

refrigerant. For redundancy, there are at least two external thermal control system

flow loops in each architecture.

• The heat-rejection devices assumed in the baseline architecture are aluminum, flowthrough radiators.

This configuration is common to spacecraft currently in service. Advanced technologies

are defined as "not used on previous or current human vehicles." The technologies

considered as advanced include:

• two-phase thermal control systems

• lightweight radiators

• phase-change thermal storage

• rotary fluid coupler

• carbon brush heat exchanger

• heat pipes

• heat pumps

• radiator shades

• plant chamber cooling improvements

These technologies, including several variations, were all compared to the baseline system

for each mission individually without combining them. It is expected that combining some

technologies with others will yield additional savings. Such combinations, however, were

not considered in this initial study. The criteria used to pare the advanced technology list

for each mission were availability of information, practicality and mission suitability, and

engineering judgment. Certain technologies apply to only certain environments or

missions, while other technologies are more generally applicable over a wider range of

missions.

Because the study format compares various architectures with a commonly defined

baseline, it is versatile and expandable. When better information becomes available, the

affected portions of the study can be readily updated without redoing the entire effort.

Further, missions and technologies, including combinations of technologies, can also be

added without disrupting the overall study.

BASELINE MISSION DEFINITION

• International Space Station Evolution

This study assumes that the current Intemational Space Station thermal control

system orbital replacement units will need replacement and augmentation after project year

20, with an overall project design life of 30 years. The current external thermal control

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

system design uses six flow-through radiator orbital replacement units in two clusters.

Due to micrometeoroid impacts, optical property degradation, and other failures, it is

assumed that all six radiator orbital replacement units will be replaced. Further, it is

assumed in year 20 International Space Station will be augmented with two solar dynamic

power modules, increasing the onboard power supply of the U.S.-led portion to provide

an additional 20 kW of user power. This does not include additional power which might

be added for the thermal control system.

The baseline mission adds two additional radiator orbital replacement units to

handle the increased heat load. Therefore, the baseline International Space Station

evolution configuration would replace the original six radiator orbital replacement units

mounted in two clusters of three units with new equipment and add one new cluster of

two units. To mount the third cluster of two units, a hypothetical third truss segment is

added in some unspecified manner.

• Space Transportation System Upgrade

This mission addresses upgrading the current Space Transportation System, or

Shuttle, with new thermal control system components of the same or improved capabilities

as those currently available. Alternatively, one might replace the current vehicles with

new vehicles of similar capabilities. The external thermal control system working fluid is

Freon 21. In addition to flow-through radiators, Shuttle uses evaporative cooling in both

primary and secondary cooling equipment. On orbit, the radiators provide primary

cooling while a flash evaporator provides secondary cooling by evaporating water and

then rejecting the steam. On orbit, Shuttle rejects up to 39.5 kW of heat from its radiators

and flash evaporator. A vehicle life of 140 missions (seven flights per year for 20 years)

following upgrades is assumed. This basis of 140 missions could also refer to the total

number of flights a fleet of vehicles might make over the life of the program. The mass

savings are considered cumulative for the life of the vehicle(s).

• First Lunar Outpost Lander

The baseline First Lunar Outpost Lander thermal control system uses a low solar

absorptivity, horizontal radiator with single-phase liquid ammonia as the working fluid.

No additional cooling devices are presumed for surface operations. The radiator upper

surface coating is silver Teflon, while the lower radiator surface is insulated to reduce

heating by the lunar surface. The user heat load for the thermal control system is 16.0 kW

with a mission length of up to 45 days. Two First Lunar Outpost Landers are assumed.

The mass savings are considered cumulative for both vehicles.

• Permanent Lunar Base

One proposal for a Permanent Lunar Base would bury three modified Space

Station modules beneath the lunar surface to provide living and working space continually

for a crew of three or four. The base elements would include a habitation module, a

laboratory module, and a plant growth module. The plant growth module would be an

integral part of the base environmental control and life support system by replenishing

atmospheric oxygen, removing carbon dioxide, and providing some food for the crew.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Power for the base would be supplied by solar photovoltaic power arrays with

regenerative fuel cells for energy storage. Baseline thermal control system heat rejection

would be accomplished through horizontal radiators with low solar absorptivity surface

coatings. The thermal control working fluid is single-phase liquid ammonia. The overall

thermal user load is 50 kW. Assessments for a single base at the equator with a project

life of 15 years are presented.

• Mars Lander

This study concentrates on Mars Lander while it is situated on the martian surface

for a 30-day stay. The vehicle will have both habitation and laboratory space for a crew of

four. The baseline Mars Lander thermal control system uses low solar absorptivity,

vertical radiators with single-phase liquid ammonia as the working fluid. The vehicle will

land at the martian equator and generate an overall user heat load of 30 kW continuously.

One mission is assumed.

ADVANCED TECHNOLOGIES

The advanced thermal control technologies may be grouped into several general

categories based on functionality. Additionally, some of the technologies considered here

can not be grouped with other technologies. These technologies are placed at the end of

the list.

• Two-Phase Thermal Control Systems:

- Two-Phase Thermal Control System With Mechanical Pump/Separator

- Low-Power Two-Phase Thermal Control System

- Two-Phase Thermal Control System With Electrohydrodynamic Pumping

Capillary Pumped Loops

• Heat Pumps:

- Vapor Compression Heat Pump

- Solar Vapor Compression Heat Pump 1

- Complex Compound Heat Pump

- Zeolite Heat Pump

• Heat Pipe Radiators:

- Arterial Heat Pipe Radiators

- Arterial Heat Pipe Radiators With Electrohydrodynamic Pumping

- Axial-Groove Heat Pipe Radiators

• Lightweight Radiators:

- Composite Flow-Through Radiators

- Composite Reflux Boiler Tube Radiators

- Composite Heat Pipe Radiators

- Unfurlable Radiators

This is also referred to as a vapor compression heat pump with a dedicated solar photovoltaic power

array.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

• Other Heat Rejection Technologies:

- Phase-Change Thermal Storage

- Parabolic Radiator Shade

• Additional Technologies:

- Rotary Fluid Coupler

- Plant Chamber Cooling Improvements

- Carbon Brush Heat Exchanger

Two-phase thermal control systems utilize a vapor and liquid mixture, as opposed

to a liquid, within the thermal control system flow loop. Upon returning from the heatrejection device, the working fluid is a liquid. Upon collecting heat at various sites along

the flow loop, the working fluid vaporizes. At the end of the flow loop the working fluid

passes to the heat-rejection device and reliquefies. Thus, by using the latent heat of the

working fluid more heat can be carded per mass of working fluid circulated.

Heat pumps increase the heat-rejection temperature, and thereby the rate of radiant

heat rejection, at the expense of energy input. For mechanically driven heat pumps, the

energy input is usually in the form of electricity to drive the compressor. For heat-driven

heat pumps, the cycle requires an input of high temperature heat which can come as a by

product of another process or may be generated specifically to drive the heat pump.

Heat pipes are passive, two-phase heat transfer devices. Heat is collected by

evaporating the internal working fluid within the evaporator. The heat pipe rejects heat by

condensing its working fluid in the condenser. The working fluid is pumped back to the

evaporator by capillary forces acting along an internal slit or groove.

Lightweight radiators use composites and advanced material processing techniques

to produced less massive radiator units. These technologies are unified through their

dependence on common materials and processing techniques. The actual heat-rejection

mechanisms vary among the different technologies. In this study, however, lightweight

radiators were assessed a mass savings as a category rather than individually. The

different technologies presented are examples of actual lightweight radiators.

Phase-change thermal storage melts a phase-change material to absorb and store

heat. The stored heat may be rejected at a later time when the primary heat-rejection

device is operating below its capacity.

A radiator shade reduces the effective temperature of the environment to which a

heat-rejection device radiates. In particular, the parabolic radiator shade reflects direct

solar irradiation away from a protected heat-rejection device and blocks infrared radiation

from the planetary surface.

Several additional technologies are unrelated to those in the other categories. The

rotary fluid coupler is a device which passes the thermal control system working fluid from

a habitat or vehicle which is fixed to a radiator which rotates to maintain a particular

thermal environment. Plant chamber cooling improvements include ideas to reduce the

mass of the thermal control equipment within a plant growth chamber. Carbon brush heat

exchangers use numerous carbon fibers to improve conductivity across evacuated gaps.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

RESULTS

• International Space Station Evolution

The solar vapor compression heat pump displayed a mass savings of 32%

compared with the baseline architecture. This option may be potentially the least

disruptive to implement once International Space Station is fully operational. Lightweight

radiators also offer significant mass savings. For this mission, composite flow-through

radiators or composite heat pipe radiators may be appropriate. The two-phase thermal

control systems appear to offer savings of almost 10% for this mission, but that may be

within the range of uncertainty for this type of preliminary analysis. The other

technologies, due to the uncertainty in the estimation procedures, produce systems which,

in terms of mass, are probably similar to the baseline configuration. The enhancing

technologies, listed here as additional technologies, can provide mass savings in their area

of application regardless of the thermal control system architecture selected. Though

these technologies yield small mass savings, they may offer other benefits which improve

system reliability or overall operations.

Thermal Control System Mass Savings for International Space Station Evolution

Basis = 1 Upgrade:

Baseline External Thermal Control System 2

Heat Pumps

Solar Vapor Compression

Vapor Compression

Lightweight Radiators

Two-Phase Thermal Control Systems

Capillary Pumped Loops

Low-Power

With a Mechanical Pump/Separator

Additional Technologies

Rotary Fluid Coupler

Carbon Brush Heat Exchanger

Heat Pipe Radiators

Arterial with Electrohydrodynamic Pumps

Arterial Heat Pipes

Axial-Groove Heat Pipes

Mass [kg]

15,600

Savings Percent

[kg] Savings

5,021 32

-250 -2

1,661 11

1,535 10

1,366 9

1,203 8

444 3

118 1

427 3

-238 -2

-309 -2

This value does not represent the entire thermal control system mass for International Space Station.

However, the omitted items, including the line masses, the line working fluid, and the pump masses,

are similar for both the baseline architecture and each of the advanced technologies.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

• Space Transportation System Upgrade

The Space Transportation System has a relatively small thermal control system

mass. Each vehicle, however, is launched numerous times throughout its life. Thus, even

small savings can be beneficial over the life of a vehicle. Reducing the mass of the

radiators by using lightweight components may reduce overall thermal control system

mass by 11%. Composite flow-through radiators appear to be most appropriate here. A

low-power two-phase thermal control system or a phase-change thermal storage system

will provide thermal control systems with masses similar to that of the baseline

architecture. The phase-change thermal storage system, however, is an enhancing

technology which can be added regardless of other improvements to the thermal control

system.

Thermal Control System Mass Savings for Space Transportation System Upgrade

Basis = 140 Flights:

Baseline External Thermal Control System 3

Lightweight Radiators

Phase-Change Thermal Storage

Low-Power Two-Phase Thermal Control System

Cumulative

Mass [kg]

201,000

Savings

[kg]

22,904

11,875

8,176

Percent

Savings

11

6

4

This value does not represent the entire thermal control system mass for Space Transportation

System. However, the omitted items are equivalent for both the baseline architecture and each of the

advanced technologies.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

• First Lunar Outpost Lander

For First Lunar Outpost Lander, the parabolic radiator shade displayed a mass

savings of roughly 30% over the baseline architecture. Assuming that a lightweight,

remotely-activated deployment scheme is developed, this technology is ideally suited for

the lunar environment near the equator. Lightweight radiators yield a mass savings of

about 28%. Any of the lightweight radiator technologies may be used, although those

which require a vertical configuration may require shading or heat pumping to operate.

Capillary pumped loops also provide a mass savings close to 20%. Both the low-power

two-phase thermal control system and the solar vapor compression heat pump offer mass

reductions of slightly more than 10%.

Thermal Control System Mass Savings for First Lunar Outpost Lander

Basis = 2 Missions:

Baseline External Thermal Control System 4

Parabolic Radiator Shade

Lightweight Radiators

Capillary Pumped Loops

Low-Power Two-Phase Thermal Control System

Solar Vapor Compression Heat Pump

Cumulative

Mass [kg]

2,100

Savings

[kg]

642

584

370

286

259

Percent

Savings

31

28

18

14

12

This value does not represent the entire thermal control system mass for First Lunar Outpost Lander.

However, the omitted items are equivalent for the baseline architecture and each of the advanced

technologies.

xvi
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

• Permanent Lunar Base

For Permanent Lunar Base, the parabolic radiator shade and the solar vapor

compression heat pump are the most promising technologies. Both individually offer mass

savings greater than 65%. These large reductions in thermal control system mass are

possible here because radiator shades and heat pumps both allow large reductions in

radiator surface area during the middle of the lunar day. In fact, shades and heat pumps

are often designated as "enabling technologies" for long-term lunar missions. The heatdriven heat pumps, specifically complex compound and zeolite heat pumps, provide mass

savings between 35 and 40%. Lightweight radiators give a mass savings of 30% for this

mission, while the two-phase thermal control systems yield overall masses similar to the

baseline configuration. Any of the lightweight radiator technologies may be used,

although those which require a vertical configuration may require shading or heat pumping

to operate. The additional technologies are all enhancing technologies which may be

utilized regardless of the thermal control system architecture. Plant chamber cooling

improvements are a collection of ideas which, if implemented in the assumed plant growth

chamber, offer a significant mass savings.

Thermal Control System Mass Savings for Permanent Lunar Base

Basis = 1 Installation:

Baseline External Thermal Control System 5

Parabolic Radiator Shade

Heat Pumps

Solar Vapor Compression

Vapor Compression

Complex Compound

Zeolite

Lightweight Radiators

Additional Technologies

Plant Chamber Cooling Improvements

Carbon Brush Heat Exchanger

Two-Phase Thermal Control Systems

Capillary Pumped Loops

Low-Power

With a Mechanical Pump/Separator

Mass [kg]

10,900

Savings

[kg]

7,416

Percent

Savings

68

7,321 67

7,017 64

4,241 39

3,818 35

3,306 30

2,867 26

154 1

917 8

751 7

456 4

This value does not represent the entire thermal control system mass for Permanent Lunar Base.

However, the omitted items are similar for both the baseline architecture and each of the advanced

technologies.

xvii
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

• Mars Lander

For Mars Lander, lightweight radiators offer the greatest individual mass savings

for any of the technologies considered. Any of the lightweight radiator technologies may

be used. The projected mass savings from the solar vapor compression heat pump

indicate the resulting system mass will be comparable to the baseline architecture.

However, the heat pump will add flexibility to the thermal control system which is

beneficial considering the uncertain climate on the surface of Mars.

Thermal Control System Mass Savings for Mars Lander

Basis = 1 Mission"

Baseline External Thermal Control System 6

Lightweight Radiators

Heat Pumps

Solar Vapor Compression

Vapor Compression

Two-Phase Thermal Control Systems

Capillary Pumped Loops

Low-Power

Mass [kg]

2,170

Savings Percent

[kg] Savings

416 19

127 6

-425 -20

58 3

45 2

This value does not represent the entire thermal control system mass for Mars Lander. However, the

omitted items are equivalent for the baseline architecture and each of the advanced technologies.

,.°

xvln
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

CONCLUSIONS

The solar vapor compression heat pump and lightweight radiators show the

greatest promise as general advanced thermal technologies which can be applied across a

range of missions.

• The vapor compression heat pump suffers from high power requirements.

However, the entire system compares favorably to baseline architectures when the

heat pump is integrated with a dedicated solar photovoltaic power array and

operated only while the vehicle or habitat is in sunlight.

• Lightweight radiators, although not rigorously defined here, offer the hope of

directly reducing the radiator mass, which is the single greatest mass within the

external thermal control system. In this initial study, four lightweight radiator

concepts are considered, allowing this technology to be applied across a wide

range of missions and environments.

• The qualitative assessments, which are separate from the quantifiable mass savings,

additionally identify the rotary fluid coupler, the carbon brush heat exchanger, and

electrohydrodynamic pumping as promising technologies. Even though these

technologies do not yield large reductions in the total thermal control system mass,

they offer other benefits which may warrent consideration.

For certain missions, several other technologies deserve consideration:

Parabolic radiator shades have great utility on the lunar surface. In this airless

planetary environment, they offer the greatest potential mass savings for missions

at low latitudes assuming the shade surface can be maintained free of dust.

Phase-change thermal storage is useful for orbital missions where the time during

which the vehicle must endure its maximum heat load is less than the orbital

period.

Two-phase thermal control systems offer greater promise for larger systems with

long flow lines. With the exception of International Space Station and Permanent

Lunar Base, the vehicles here are too small to make adequate use of this

technology. For Permanent Lunar Base, however, a two-phase thermal control

system may be a wise investment, especially if the base will be expanded sometime

in the future. As flow lines lengthen and heat loads increase, the pump power

saved by a two-phase thermal control system will be significant. Capillary pumped

loops may provide a mass savings even for smaller vehicles assuming the thermal

control system is not expected to operate under high accelerations.

xix
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

• Due to their extra mass, metallic heat pipes appear to be justified for human

missions only when thermal control system flow loop punctures from external

debris are expected to be a serious problem. When a mission flies in such an

environment, mass savings become secondary to the reliability of the heat-rejection

system. Current debris predictions for the missions here do not consider this to be

a problem except possibly for International Space Station evolution, which flies in

low Earth orbit where the debris problem is expected to be greatest. Further,

longer arterial heat pipes are lighter per unit of radiant surface area. Combining

longer units with lighter thermal interfaces and/or electrohydrodynamic pumps also

improves the mass savings.

• Heat-driven heat pumps offer mass savings for Permanent Lunar Base compared

with the baseline architecture. Coupled with a source of high-temperature waste

heat, such as from an industrial process, they may provide economical cooling.

However, heat-driven cycles are not as efficient as mechanically driven cycles and,

therefore, heat-driven heat pumps weigh more than mechanically-driven heat

pumps.

• Several enhancing technologies appear in this study. These include the rotary fluid

coupler, plant chamber cooling improvements, and the carbon brush heat

exchanger. Alone, they do not offer large mass savings. However, they also do

not require a particular thermal control system architecture and, therefore, can be

implemented individually or with other technologies.

This initial study identifies several promising advanced thermal control

technologies which offer both mass savings and other benefits compared to traditional

thermal control systems. Future research and development is expected to more accurately

and precisely def'me these mass savings as various programs proceed to develop flightready hardware. Future technical developments and additional knowledge may change

current understanding on which these results are based. Therefore, this study is expected

to be an ongoing effort which will be updated as needed to allow a common basis for

future programmatic decisions. Subsequent work in this area is currently planned to

include a wider range of technologies as well as architectures which combine various

technologies at the system level.

XX
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

1.0 INTRODUCTION

This study presents advanced active thermal control system (ATCS) architectures

for five proposed National Aeronautics and Space Administration (NASA) vehicles and

habitats. The objective is to define possible alternate thermal control systems using new

or proposed technologies for these vehicles and habitats and determine what mass savings

and other benefits might be available. Non-mass savings are converted to mass using an

appropriate penalty. For example, a power savings is converted to mass using a power

mass penalty. However, many benefits are considered only qualitatively. Because this is

preliminary work, many details of various vehicles and technologies are either assumed or

neglected entirely. To aid the reader, the significant assumptions are listed.

The five reference missions considered below are:

• International Space Station (ISS) Evolution

• Space Transportation System (STS or Shuttle) Upgrade

• First Lunar Outpost (FLO) Lander

• Permanent Lunar Base (PLB)

• Mars Lander (ML)

Twenty different advanced technologies in various stages of maturity are

considered for this study. The main areas collectively addressed by these technologies

include thermal transport systems, heat pumps, and radiators. However, several additional

options, such as the rotary fluid coupler and the carbon brush heat exchanger, fall into

categories by themselves. Because each of the advanced technologies either does not

apply to all of the vehicles or habitats examined, or it is expected that the given technology

would not yield a benefit, the technology list was pared as appropriate for each reference

mission. The advanced technologies considered for each vehicle or habitat are addressed

in Table 1.1.
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Table 1.1 Advanced Technologies Applied to Reference Missions

Advanced Technologies

Reference Vehicles and Habitats

ISS STS FLO PLB ML

Two Phase Thermal Control Systems:

1. Two-Phase TCS With Mechanical Pump/Separator

2. Low-Power Two-Phase TCS

3. Two-Phase TCS with Electrohydrodynamic Pumping

4. Capillary Pumped Loops

X X

X X X X X

X X X

X X X X

Heat Pumps:

5. Vapor Compression Heat Pump

6. Solar Vapor Compression Heat Pump

7. Complex Compound Heat Pump

8. Zeolite Heat Pump

X X X

X X X X

X

X

Heat Pipe Radiators:

9. Arterial Heat Pipe Radiators

10. Axial-Groove Heat Pipe Radiators

11. Arterial Heat Pipes With Electrohydrodynamic Pumping

X

X

X

Lightweight Radiators 7:

12. Composite Flow-Through Radiators

13. Composite Heat Pipe Radiators

14. Composite Reflux Boiler Tube Radiators

15. Unfurlable Radiators

X X X X X

X X X X X

X X X X

X X X

X X X

Other Heat Rejection Technologies:

16. Phase-Change Thermal Storage

17. Parabolic Radiator Shade

X

X X

Additional Technologies:

18. Rotary Fluid Coupler

19. Plant Chamber Cooling Improvements

20. Carbon Brush Heat Exchanger

X

X

X X

Total Number of Applicable Technologies 10 3 5 11 5

It should be noted that the various advanced technologies were assumed to operate as

theorized. Because some technologies are currently not well-developed, the analysis

which follows may seem speculative. However, as stated earlier, the purpose here is to

estimate what mass savings and other benefits might accrue from pursuing various ideas

for active thermal control systems which may be available in the near future.

Each of the technologies considered has desirable and undesirable attributes. To

reduce the time spent on unproductive exercises, the technology list was pared to examine

only technologies which were likely to offer an advantage for a particular mission.

Further, technologies were not combined with each other to optimize the use of a

particular technology in this initial work. In other words, each technology was considered

individually to determine its benefits.

This study evaluates lightweight radiators as a class. The other architectures mentioned in this

category are possible implementations of a lightweight radiator. Those which are possibilities for

each mission are marked with an 'x' (small x) but they are not evaluated separately as a technology.
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Two-phase thermal control systems, according to Ungar (1995), are most likely to

lower pumping requirements for larger systems with larger heat loads. Further,

electrohydrodynamic pumping appears appropriate only for systems not operating while

the vehicle is accelerating rapidly, such as during launch or landing. Capillary pumped

loops employ two-phase heat acquisition and rejection for the thermal control system flow

loop. However, unlike the other two-phase architectures mentioned above, capillary

pumped loops do not use any moving parts. Their main disadvantage is a reliance on

capillary pumping, which typically generates a low pumping head, so that such a system

will not operate under launch or landing accelerations.

Heat pumps are a versatile technology which can be applied to most vehicles and

missions. Their major detracting attribute is the need for a relatively large power source

or a high-temperature heat source. Further, the study here considered heat-pumping for

the entire thermal control system load, and not just for partial loads. Under these

conditions, the vapor compression cycles were deemed feasible for all missions except

Space Transportation System upgrade, and the heat-driven cycles were considered only

for Permanent Lunar Base.

Heat pipes, which are generally heavier than flow-through radiators, are attractive

for environments where high concentrations of debris may strike and puncture the

radiators. Of the missions considered here, only International Space Station evolution is

projected to possibly operate in a high debris environment for long periods of time.

Lightweight radiators represent a general category where the radiators are

constructed using lighter and stronger materials. Due to this category's broad nature, the

implementation of this option for each of the missions is expected to differ. Some initial

speculative calculations are presented which presume lighter components based upon

using flow-through radiator architecture. Specific examples of lightweight radiators

include composite flow-through radiators, composite heat pipe radiators, composite reflux

boiler tube radiators, and unfurlable radiators. Of these examples, composite reflux boiler

tube radiators and unfurlable radiators are restricted to vertical deployments on planetary

surfaces. Further, the composite heat pipe radiator is also deployed vertically when on a

planetary surface, but they may also operate in a microgravity environment. These

technologies are mentioned for each mission when appropriate.

Phase-change thermal storage utilizes the heat of fusion for a homogeneous

material to store cooling or heating capacity. Here such a material is used in conjunction

with flow-through radiators to supplement the radiators' heat-rejection capability by

melting the phase-change material while the vehicle is in a relatively hot thermal

environment. This technology appears to be most attractive for situations where a vehicle

is operated continuously in its hottest environment for short periods of time, such as for

atmospheric braking and on-orbit mission elements.

The parabolic radiator shade requires large surface areas for deployment and

absence of an atmosphere which might allow wind to deposit dust on the shade.

The rotary fluid coupler is appropriate for situations when a tracking radiator is

necessary and less complicated radiator systems would be inadequate. This is a

consideration for orbiting vehicles. On a planetary surface it is generally impossible to

avoid all significant sources of solar radiation on the radiator using selective radiator

positioning.
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Plant chamber cooling improvements are applicable only for missions which

contain plant growth chambers. In the current study only Permanent Lunar Base is

projected to use a plant chamber.

Carbon brush heat exchangers have utility where ever a coldplate exists. Here they

were restricted to coldplates for dc-to-dc converter units for International Space Station

evolution and Permanent Lunar Base.

At the end of each advanced technology assessment, a composite qualitative

assessment is given in addition to the overall mass savings. Assessments are made in five

areas. These assessments rate volume, ease of deployment or installation, reliability,

remaining development cost to bring the technology to a readiness level of at least 6, and

potential for terrestrial applications. In situations where the advanced technology is

expected to have similar performance to the base reference mission, the qualitative

assessment used is "average." In situations where the advanced technology is expected to

differ significantly from the reference architecture, one of the two extreme assessments are

assigned. The development cost is assigned a value of "low" for estimated remaining

development program costs up to one million dollars, "average" for costs between one

and three million dollars, and "high" for costs in excess of three million dollars. More

specifically, the qualitative grading assessments are described in Table 1.2.

Table 1.2 Qualitative Scores for Assessment Areas

Assessment Area

Volume

Deployment (or Installation)

Reliability

Development Cost

Terrestrial Use Potential

-1

large

difficult

degraded

high

none

Qualitative Score

I °

average

average

average

average

possible

+1

compact

easy

improved

low

good

Details of the qualitative assessments are given in Section 2.0 for each advanced

technology.

To aid in assessing the development cost, each advanced technology was assigned

a technology readiness level rating based on NASA (1991) for each reference mission. In

the most basic terms, these technology readiness level ratings are defined as:

Rating Technology Readiness Level Description

Basic principles observed and reported

Technology concept and/or application formulated

Analytical and experimental critical function / characteristic proof-of-concept

Component / breadboard validation in laboratory environment

Component / breadboard validation in relevant environment

System/subsystem model or prototype demonstration in relevant environment

System prototype demonstration in a simulated environment

System completed and flight qualified through test and demonstration

System "Flight proven"

4
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

2.0 ADVANCED THERMAL CONTROL TECHNOLOGIES

This section presents the various advanced thermal control technologies

considered in this study. The technologies are applied to each baseline mission

individually to assess their benefits compared with the baseline architecture. The baseline

architecture is either the present configuration for a vehicle or habitat, if known, or a

defined baseline configuration. The defined baseline configurations assume a thermal

control system (TCS) with fluid flow loops to collect heat loads and flow-through

radiators to reject the heat load to space. More specifically, the defined baseline

configurations use an internal thermal control system (ITCS) flow loop with water as the

working fluid and an external thermal control system (ETCS) flow loop with single-phase,

liquid ammonia as the working fluid. Heat exchange from the ITCS to the ETCS is by

way of interface heat exchangers.

2.1 Two-Phase Thermal Control Systems

A two-phase TCS takes advantage of the working fluid's latent heat of

vaporization to store and release thermal energy. When heat is collected by the ETCS

flow loop, the working fluid vaporizes. The working fluid liquefies when it rejects the

collected heat load to a rejection device. Compared with a single-phase system, the twophase system should use smaller piping, less working fluid, and correspondingly less

pumping power, especially for larger vehicles or habitats with longer ETCS fluid lines.

Ungar (1995) compares single- and two-phase TCSs for space stations of several

sizes. His results indicate that two-phase TCSs are advantageous for larger vehicles,

comparable for medium-sized vehicles, and less efficient for smaller vehicles. In this initial

section on technologies, a larger vehicle, such as ISS after evolution or PLB are assumed

for the qualitative assessments.

2.1.1 Two-Phase Thermal Control System With Mechanical Pump/Separator

The mechanical pump/separator (MP/S) uses centrifugal forces to separate a twophase stream into liquid and vapor before pumping the fluid to the heat exchangers and

radiators. The liquid is extracted from the periphery of the device by a pitot pump and

delivered to the ITCS/ETCS interface heat exchangers while the vapor in the center flows

to the radiators (Figure 2.1). While in the radiators, the vapor condenses, liberating heat

to space, and exits the radiators as a subcooled liquid.

To simplify the analysis for this study, several assumptions are helpful. First, the

masses for two-phase plumbing, fittings, and interface heat exchangers are approximately

the same as the masses of the corresponding single-phase equipment. Second, changes in

the fluid inventory between this two-phase system and the baseline single-phase system are

not considered.

5
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Radiator Orbital Replacement Units

/,, / / / .... /

/,, /

/,,/

/ ,/

Mechanical

Pump/Separator

O I

I |

I I

I I

I

I

|

I

I I I I

ISS

ITCS

(Water)

Figure 2.1 An external thermal control system using a two-phase thermal control system

with mechanical pump/separator. The lines through which two-phase fluid flows are

dashed and the single-phase lines are solid. The example here presents the configuration

for International Space Station.
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Because a two-phase TCS with MP/S was originally proposed for Space Station,

much of the developmental and design work has already been completed for a larger TCS.

For ISS evolution, this option has a technology readiness level (TRL) of 7. For other

reference missions the TRL would be 6. Flight testing is not considered necessary to

complete a system design. Thus, the development cost is expected to be low. The

reliability and deployment of such a TCS should be comparable to current equipment once

development is completed. It is also expected that, due to the high heat-transfer rates and

low power consumption associated with evaporative cooling and condensation in a twophase system, this technology may be attractive to terrestrial users. However, a flash

tank 8 in an environment with gravity is a proven method which requires no input power to

separate liquid and vapor streams.

General Qualitative Assessments:

Technology Readiness Level 9

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

6

average

average

average

low

none

Composite Qualitative Score 0

2.1.2 Low-Power Two-Phase Thermal Control System

The low-power (LP) two-phase TCS is similar to the two-phase TCS with MP/S

with the added benefit that only liquid is pumped. Like the two-phase TCS with MP/S,

the ETCS stream vaporizes while extracting heat from the ITCS. Here a two-phase

liquid/vapor mixture is delivered to the radiators. However, the pump is placed

downstream of the radiators where the fluid is always sub-cooled. This approach reduces

the necessary pumping power requirements and avoids the technical complexities

associated with specialized equipment such as a MP/S. For the study here, the

quantifiable difference between a two-phase TCS with MP/S and a LP two-phase TCS is

an additional reduction in pumping power associated with the latter option.

As with the previous two-phase system, the LP two-phase TCS is still in the

developmental stage. However, because the LP two-phase TCS uses standard

components in an innovative arrangement, the developmental work should be primarily at

the system level. To simplify the analysis for this study, several assumptions are helpful.

First, the masses of two-phase plumbing, fittings, and interface heat exchangers are similar

to the masses of the corresponding single-phase equipment. Second, changes in the fluid

inventory between this two-phase system and the baseline single-phase system are not

considered.

8 A flash tank is a chemical, or flow, processing unit with at least one inlet and two outlets. The tank

itself is typically at a lower pressure than its upstream components, so the process stream "flashes"

or separates into vapor and liquid components using gravity as the driving potential for the

separation.

9 For ISS evolution this technology would have a TRL of 7.
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Technically, this option should be easier to implement than the two-phase TCS

with MP/S. However, the LP two-phase TCS has a lower TRL of 3. Thus, an average

development cost is anticipated. Terrestrial uses for this technology are not expected even

though several possibilities initially suggest themselves. One suggestion would replace a

building's air conditioning cycle with a LP two-phase TCS. However, the coefficient of

performance of this technology is expected to be significantly lower than a standard heat

pump cycle at temperatures common to residential and commercial building cooling

applications. Further, to avoid pumping a two-phase mixture against gravity, the

condenser, which is in contact with a cooler environment, must be located above the

evaporator (cooling unit) in the building 10. Thus, this technology is probably impractical

for cooling buildings. For higher temperature applications, such as industrial process

cooling, liquid sprays or tanks are often used. The additional equipment, such as

coldplates and piping, for a LP two-phase TCS cooling cycle would be a more expensive

alternative as long as the current technology is acceptable. Further, the petrochemical and

other industries already use something similar to the LP two-phase TCS. For transporting

two-phase streams a flash tank is used to separate the components which may then be

pumped individually. When applicable, this approach is generally less expensive than

subcooling a process stream. Thus, while innovative, the LP two-phase TCS and its

associated technology are not expected to inspire terrestrial uses.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

3

average

average

average

average

none

Composite Qualitative Score -1

2.1.3 Two-Phase Thermal Control System With Eleetrohydrodynamie Pumping

To use a two-phase TCS, the difficulty of transporting a mixture of liquid and

vapor must be addressed. The LP two-phase TCS sub-cools the ETCS working fluid in

the radiators, allowing only liquid to enter the ETCS pumps. Electrohydrodynamic

pumping would pump the entire fluid stream, regardless of its phase, by using a distributed

network of electrohydrodynamic pumps. This technology would be appropriate for

thermal control loops which were not subjected to large accelerations. Thus,

10 Radiators, for commercial and residential cooling, are inappropriate because the terrestrial

environmental temperatures are too high when cooling is necessary. However, a cooling tower could

be placed on a building's roof to provide an appropriate sink for a LP two-phase TCS cooling system.

However, cooling towers are, by necessity, extremely heavy. Thus, a LP two-phase TCS cooling

system could probably not be economically installed in an older building. Further, the additional

structural support necessary to implement a LP two-phase TCS in a new building is expected to be

prohibitive.

8
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

electrohydrodynamic pumping would be appropriate for missions in microgravity and on

planetary surfaces but not for vehicles subjected to launch loads.

Electrohydrodynamic pumping should have several advantages for a two-phase

TCS (Bryan, 1995). The mass of any phase separation device could be eliminated. The

standard single-phase pumps could also be removed, but the mass gain associated with the

electrohydrodynamic pumps will mostly offset this benefit. The power for an

electrohydrodynamic pumping system is expected to be similar to that associated with

mechanical pumping. The volume will be comparable to other two-phase TCSs. The

deployment should be comparable to other TCSs. Electrohydrodynamic pumping would

be significantly more reliable because this technology utilizes a greater number of smaller

pumps. Thus, losing a single pump is less damaging than losing a mechanical pump in a

standard TCS. In this application, electrohydrodynamic pumping is immature with a TRL

of 2. Because this technology generates low pumping pressures, a test on orbit is needed

to assess its behavior in a microgravity environment. However, terrestrial testing should

be sufficient for systems deployed on a planetary surface. Thus, the development cost for

a two-phase TCS with electrohydrodynamic pumping is expected to be high for orbital

applications and average for planetary missions. Electrohydrodynamic pumping may have

applications in the medical field and in the petrochemical industry. Medical flows often

contain more than one phase, such as blood or a suspension of bacteria, and

electrohydrodynamic pumps would, in cases where the pumping process does not alter the

fluid in the flow, provide an excellent method of transporting these mixtures. Further,

electrohydrodynamic pumping moves flows over long distances by using numerous

electrohydrodynamic pumps. This pumping configuration avoids introducing extreme

stresses into the flow unlike conventional mechanical pumping. For flows with shear

sensitive components, electrohydrodynamic pumping would be preferable to traditional

mechanical pumping. Unlike the LP two-phase TCS, electrohydrodynamic pumping can

transport two-phase flows without separating them into their components. Thus, in some

industrial applications, electrohydrodynamic pumping may be a more expedient approach

than traditional techniques which separate the flow for transport.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

2

average

average

improved

average

good

Composite Qualitative Score 11 +2

2.1.4 Capillary Pumped Loops

Capillary pumped loops are two-phase thermal control loops which use capillary

forces to drive the flow. More specifically, as illustrated in Figure 2.2, the capillary

11 This assessment is for missions on a planetary surface. For orbital missions the development cost is

"high" and the overall score is +1.

9
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

pumped loop acquires heat by vaporizing the working fluid at the evaporator and rejects

heat to the environment or a heat-rejection interface when the working fluid condenses at

the condenser. The wick, which in American and European implementations 12 is typically

porous high-molecular-weight polyethylene, is normally saturated with liquid. When heat

is applied to the evaporator, the liquid on the outer diameter of the wick evaporates and

moves as a pure vapor to the condenser, where it condenses and is subcooled. The

pumping action of the wick returns the liquid to the evaporator. The reservoir contains

saturated working fluid, liquid and vapor, at the desired loop pressure and, therefore,

temperature. Thus, capillary pumped loops, in their standard mode of operation, are

variable conductance devices 13 which transfer heat proportional to the rate of vapor

formation at the evaporator and operate at some temperature determined by the set-point

pressure of the loop.

Cullimore (1993) provides a broad overview of capillary pumped loop operating

characteristics and traits supported by an extensive literature survey. Several observations

from this work are applicable here. Of greatest importance for rejecting heat from human

vehicles and habitats is the overall capacity of the capillary pumped loop. Studies (see

Cullimore, 1993) demonstrate that capillary pumped loops can be designed to transfer

from 5 W up to 25 kW per loop using aluminum and ammonia systems 14. Further, the

heat load may be acquired using numerous parallel evaporators and rejected from

numerous parallel condensers. Correspondence between the number of evaporators and

the number of condensers is not required. Using multiple evaporators and condensers

allows the capillary pumped loop to function as a thermostat for the loads it services by

heating those loads which are cooler than the loop set-point and cooling only the loads

which are warmer than the loop set-point. However, to prevent the capillary pumped loop

from shutting down 13, the overall sum of the loads serviced by the loop requires cooling.

A capillary pumped loop is not constrained by a fixed geometry but is rather a

combination of components arranged to give a specific TCS architecture 16. Further, they

can pump working fluid 0.3 m to 3.0 m against terrestrial gravity, depending on the

chosen wick structure and flow geometry. Thus, assuming care is taken in the system

design, capillary pumped loops could be used on a planetary surface as well as in space.

The only situation where capillary pumped loops would be completely inappropriate are

12 The experiences from the former Soviet Union, and now Russia, differ in some significant ways.

Cullimore (1993) mentions some of these differences for the interested reader. The remainder of this

description assumes American or European heritage equipment and practice.

13 If the reservoir ever becomes completely filled with liquid, the capillary pumped loop becomes a

constant conductance heat transport device. Traditional practice dictates that capillary pumped

loops are designed and charged so that this cannot occur.

14 Flight-rated hardware so far has concentrated on cooling robotic probes with multiple capillary

pumped loops which transfer no more than hundreds of Watts each. An example of such a vehicle is

the Earth Observing System (EOS), which has three instruments. Each of the instruments, which

have maximum heat loads up to 250 W, uses one of two redundant capillary pumped loops to

transfer those loads to constant conductivity heat pipes (Ku, 1996).

_5 A capillary pumped loop functions as a diode with respect to heat transfer. If the heat load to be

rejected by the loop is cooler than the condenser environment, the liquid line will boil, allowing the

evaporators to deprime and shut down the loop. The capillary pumped loop may be restarted later

(Cullimore, 1993).

16 As such, they are placed here with the two-phase TCSs and not with the heat pipes.

10
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

for systems which must function while enduring high accelerations, such as during launch

or re-entry.

Qin

Va orJ L

/

Wick /

Direction of Fluid Flow

Evaporator

Condenser

Vapor-Liquid Mixture

Liquid

/

Qout

- 1 Reservoir

VaporLiquid

Mixture

Figure 2.2 A simple implementation of a capillary pumped loop (Cullimore, 1993). The

capillary pumped loop acquires heat by vaporizing the working fluid at the evaporator and

rejecting heat to the environment by condensing working fluid at the condenser. The wick

pulls liquid into its pores by capillary action from the liquid volume at its inner diameter.

As the evaporator acquires heat, the liquid on the outer diameter of the wick evaporates

and moves as a slightly superheated vapor to the condenser. The reservoir contains

saturated working fluid liquid and vapor at the desired loop pressure and, therefore,

temperature. Thus, capillary pumped loops, in their standard mode of operation, are

variable conductance devices which transfer heat proportional to the rate of vapor

formation at the evaporator and operate at some temperature determined by the set-point

pressure of the loop. (Capillary pumped loop components are denoted by bold text while

working fluid states are given by italic text.)

As presented in this study, capillary pumped loops are rated assuming they require

no external pumping power and in turn require no additional equipment mass beyond that

for the baseline single-phase TCS architecture. Assuming the capillary pumped loop is

properly sized, this TCS does not require any mechanical pumping except possibly for

starting the system and/or recovering from depriming. The only power input is that for

maintaining the temperature of the reservoir which sets the capillary pumped loop setpoint. This power expenditure is a small fraction of the total heat load carded by the

capillary pumped loop. The actual capillary pumped loop equipment may be significantly

different than that for a comparable single-phase TCS because capillary pumped loop

designs must use components with low flow resistance to minimize the overall loop

11
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

pressure drop 17. However, these components are not necessarily more massive than the

corresponding single-phase components.

Qualitatively, capillary pumped loops integrated into new vehicles and habitats

should have a volume and ease of deployment comparable to the baseline architecture is

The system reliability should be improved relative to the baseline architecture because the

capillary pumped loop has no moving parts and can operate with little external power

input. The development cost for piloted vehicles on orbital missions will be high because

a capillary pumped loop of sufficient capacity with multiple evaporators/condensers has

yet to be tested on orbit. For systems placed on a planetary surface, terrestrial testing

should suffice so an average development cost is anticipated. The TRL for piloted

vehicles is 5. Applications are currently unknown for capillary pumped loops, so no

terrestrial uses are anticipated.

General Qualitative Assessments:

Technology Readiness Level 5

Volume average

Deployment average

Reliability improved

Development Cost average

Terrestrial Use Potential none

Composite Qualitative Score 19

2.2 Heat Pumps

Assuming the radiating area is fLxed, a greater heat load may be rejected by either

lowering the effective environmental, or sink, temperature or by increasing the radiator

surface temperature 20. Heat pumps, as presented in this study for space applications,

effectively cool a vehicle or habitat using a refrigeration cycle and reject the work load and

the original heat load at a higher temperature. Two general classes of heat pumps axe

presented. Vapor compression heat pumps use a mechanically driven cycle built around a

17 Ungar (1996) recommends that the overall loop pressure drop not exceed 3,447 N/m 2 (0.5 lbigin 2)

based on the limited capillary pumping head available from current high-molecular-weight

polyethylene wicks.

18 Retrofitting a vehicle using a single-phase TCS with a capillary pumped loop may be difficult, if not

impossible. Thus, for installation on a pre-existing vehicle, the rating for deployment should be

"difficult."

19 This assessment is for missions on a planetary surface. For orbital missions the development cost is

"high" and the overall score is -1.

2o Heat rejection in a vacuum between a radiator and its environment, q, may be defined as:

q= - T4o ]

where A is the actual radiating surface area, rl is the fin efficiency, c is the surface emissivity, Fij is

the view factor between the radiator surface and the environment, c_ is the Stefan-Boltzmann

constant, and Tr_diatorand Tsmk are the radiator surface and environmental temperatures, respectively.

12
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

compressor 21. The primary utility for a vapor compression cycle is a source of electricity

to run the compressor. Heat-driven beat pumps are built around a chemical or physical

absorption-desorption cycle over a range of temperatures. Here the cycles considered use

solid/vapor adsorption systems with a fixed refrigerant carrier (Yerushalmi, 1992). The

primary utility necessary for heat-driven cycles is a source of high temperature heat.

Ideally, this could be obtained as waste heat from some other process such as a nuclear

power plant or other industrial process. However, solar concentrators can also provide a

high-temperature heat source to operate a heat-driven cycle. The results here for heatdriven heat pumps are based on Ewert (1993) and assume solar concentrators for the heat

source.

Generally, the systems here are sized so that heat pumping is only necessary while

the vehicle or habitat is receiving solar irradiation. Orbital vehicles, however, may receive

continuous heat pumping because their sunlight/shade cycles are relatively short. For

example, ISS will have roughly a 90-minute orbital cycle with about 30 minutes in shade

for a beta angle of 0 degrees. The heat pumping here is also assumed to be centralized as

opposed to distributed.

2.2.1 Vapor Compression Heat Pump

With respect to an active thermal control system (ATCS), heat pumps could be

used to increase the radiator fluid temperature, thereby increasing radiator heat rejection

and reducing the necessary radiator surface area. A typical single-stage heat pump for this

application would accept heat from the ITCS, which is at the cycle's cold temperature, by

evaporating a working fluid. The resulting stream, usually a superheated vapor, is

compressed and passed to the condenser. Within the condenser, which is the cycle's hot

working temperature, the fluid condenses and rejects heat to the ETCS radiators at an

elevated temperature. Upon leaving the condenser, the working fluid flows through an

expansion valve before returning to the evaporator (Figure 2.3). Generally more efficient

high-lift heat pumps use more than a single stage.

The main disadvantage of the vapor compression heat pump is its high power

requirement relative to the baseline TCS architecture using only metal flow-through

radiators. The overall TRL for a vapor compression heat pump is 5. For application on a

planetary surface, the development cost is expected to be low. For orbital applications,

the influence of microgravity on lubricant circulation within the heat pump must be

resolved and tested on orbit. Thus, for orbital applications the development cost is high.

The deployment in general should be comparable to baseline architectures, or average.

Directly, heat pumps can be used for heating or cooling of terrestrial habitats or industrial

equipment. Further, coupled with non-chlorofluorocarbon refrigerants, this technology

may pay commercial dividends. Some work has already been completed for a high-lift,

multistage heat pump designed for use with lunar habitations. This design has fewer

components which are susceptible to micrometeoroid punctures and optical degradation

than the baseline TCS architecture. The vulnerable radiator flow loops are shorter and the

21 Another mechanical heat pump uses the Stirling cycle. The Stirling cycle has many of the same

attributes as a vapor compression cycle. However, to date, Stirling heat pumps have not been

applied to high load or high temperature applications. Thus, they were not considered here.

13
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

heat pump itself is compact enough to take advantage of local shielding. Additional

components with moving parts, however, will degrade the reliability of this technology.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

average

average

degraded

low

good

Composite Qualitative Score 22 +1

2.2.2 Solar Vapor Compression Heat Pump

A solar heat pump uses the same equipment as the heat pump presented in

Section 2.2.1 except that input power is provided by a dedicated solar photovoltaic power

array instead of from the general habitat or vehicle power grid 23. To further reduce mass,

only minimal power conditioning equipment is provided. This is expected to significantly

reduce the mass of the power plant and distribution net for the heat pump. (See Ewert,

Keller, and Hughes, 1996) Further, a bypass line for the heat pump will allow the ETCS

to use the radiators directly when the heat pump is not in use, such as when the base is not

in sunlight. Therefore, a solar heat pump does not require any electrical storage

(Figure 2.4).

The solar heat pump has the same advantages as the standard vapor compression

heat pump presented in the previous section. However, the mass of the required

supporting power systems for the solar heat pump is significantly less. Because the power

system can account for over 90% of the mass added for the vapor compression heat

pump, reducing this mass is highly advantageous.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

compact

average

degraded

low

good

Composite Qualitative Score 24 +2

22 For orbital applications, the development cost is "high" for an overall score of -1.

23 This option is more correctly called a vapor compression heat pump with a dedicated solar

photovoltaic power array.

24 As with the vapor compression heat pump, this option has a "high" development cost for orbital

applications. The overall score for orbital applications is 0.

14
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

I /

\ \

I /

\ \

\ \

/ / /

/ / /

/

/

/

/

/

/

Radiator

ORUs

Condenser

_ 7 Expansion

 .ve

I Evaporator DPump

Module

Compressor

ETCS

Figure 2.3 An external thermal control system using a vapor compression heat pump.

The example here presents the configuration for International Space Station.

15
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

2.2.3 Complex Compound Heat Pump

Heat-driven heat pumps use cycles requiring a source of extremely high

temperature heat to operate. For a complex compound heat pump, the high heat source

requires a temperature of 500 K (Yerushalmi, 1992). Commonly, this heat is transported

to a heat-driven heat cycle using an appropriate working fluid.

In the baseline missions studied here, no sources of waste heat are present. In

order to generate heat to drive a heat-driven heat pump, at least two other methods are

available. One method would generate heat from electricity generated by solar

photovoltaic power arrays through resistive heating. A second method would use a

collector to focus solar radiation on a tube containing a working fluid. This is a more

efficient technique and is assumed for the heat-driven cycles in this study.

The TCS with complex compound heat pumps is similar to the baseline technology

in many areas. This heat pump will allow for more compact radiators, compared with the

baseline TCS, but this benefit will be partially offset by the addition of the heat pumps.

The deployment should be comparable to the baseline mission. The system reliability will

probably be the same as the baseline mission because heat-driven heat pumps contain

fewer components with moving parts than vapor compression heat pumps. The heat

pump itself has few moving parts but the collector system to supply high temperature heat

to the cycle will have a reliability comparable to other technologies. Because current heat

pump research is directed toward an application for space, the development cost for this

system is expected to be average. While orbital missions are not proposed for this heat

pump, the microgravity environment is not expected to affect it. The TRL is assessed at

4. Terrestrial uses could exist for complex compound heat pumps as cooling devices in

heavy industries where a source of high temperature heat is readily available. Examples

might include power generation, metal processing and forging, or possibly use in arid

regions where an uninterrupted source of solar energy can be readily collected.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

4

compact

average

average

average

possible

Composite Qualitative Score +1

2.2.4 Zeolite Heat Pump

Zeolite heat pumps also use a heat-driven cycle and require a high heat source at a

temperature of 700 K (Yerushalmi, 1992). From the perspective of this report, zeolite

heat pumps have many of the same attributes as complex compound heat pumps.

Therefore, the same overall approach as for the preceding technology is employed here.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Radiator

ORUs

\ \

\ \

\ \

/ / / /

/ /

/ /

\ \

/ /

Pump Module

I

t_

o

o

r_

E

Condenser

X 7 Expansion

_E Valve

vaporator

Compres s or

ETCS

Figure 2.4 An external thermal control system using a solar vapor compression heat

pump (a vapor compression heat pump with a dedicated solar photovoltaic power array).

The example here presents the configuration for International Space Station.

17
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

The TCS with zeolite heat pumps is similar to the baseline technology in many

areas. This heat pump will allow for more compact radiators, compared with the baseline

TCS, but again this benefit will be partially offset by the addition of the heat pumps which

are more massive than the complex compound equipment. The deployment should be

comparable to the baseline mission and the system reliability will probably be the same as

the baseline mission. Again, the zeolite heat pump itself has few moving parts but the

collector system to supply high temperature heat to the cycle will have a reliability

comparable to the baseline technologies. Because current heat pump research is directed

toward an application for space, the development cost for this system is expected to again

be average. The zeolite heat pump is slightly behind the complex compound heat pump,

so the TRL is 3. As above, terrestrial uses could exist for the zeolite heat pumps as

cooling devices in heavy industries with a source of waste high-temperature heat.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

3

compact

average

average

average

possible

Composite Qualitative Score +1

2.3 Heat Pipe Radiators

Heat pipes are self-contained, two-phase devices which transfer heat from a hot

source to a cold source. The heat pipe evaporator vaporizes the internal working fluid to

acquire the heat load. The vapor travels out into the condenser and liquefies the working

fluid as the heat pipe rejects the heat load to the cold source. For missions in this study,

the cold source is the environment. The working fluid travels back to the evaporator by

capillary forces, which provide the primary pumping within heat pipes. For a given heat

pipe, the heat transport is limited by the available capillary pumping capability. When the

transport limit is exceeded, the working fluid may collect in the heat pipe's condenser,

rendering the unit unusable. One technique for working around this tendency is to add

extra heat pipe capacity to ensure that the system transport limit is not exceeded. Another

option is to add individual pumps to each heat pipe to increase each unit's transport limit.

The primary advantage of arterial heat pipes, even though equivalent metallic

arterial heat pipe systems are often heavier than corresponding metallic flow-through

radiator systems, is their redundancy which provides added reliability in environments with

micrometeoroids and on-orbit debris. If an ETCS flow line is punctured by a

micrometeoroid, the working fluid is expected to drain quickly to space. Spacecraft and

bases with flow-through radiators are most susceptible to this because their main ETCS

flow lines pass through the exposed and lightly protected radiators. In a similar scenario

with arterial heat pipes, the ETCS flow lines would be protected with only the heat pipe

condensers exposed to on-orbit debris punctures. In this second situation only the

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

working fluid and capacity of a single heat pipe would be lost to a single puncture, not

that of an entire ETCS flow loop. Axial-groove heat pipes replace just the flow-through

tubes within radiator panels (Nguyen, 1992) so they do not yield a significantly more

robust system for applications where micrometeoroid and on-orbit debris are a significant

concern.

2.3.1 Arterial Heat Pipe Radiators

Arterial heat pipe radiators are somewhat different from the current state-of-theart heat pipe 25. In this concept, the current radiators are replaced with a heat exchanger

to which a number of long heat pipes are mounted. The heat exchanger collects thermal

energy from the ETCS loop and passes it to the heat pipe radiators. The heat pipes are

much longer than the axial-groove heat pipes presented in Section 2.3.3 26. As such, the

arterial heat pipes extend individually from the heat exchanger into space to reject heat.

The portion of the heat pipe in contact with the heat exchanger, a base section about

0.91 m (3.0 ft) long on 13.11-m heat pipes, is the evaporator, while the remaining section,

which is connected to a fin and rejects heat to space, is the condenser. Internally, the heat

pipe uses channels shaped liked barbells to convey the fluid. The vapor travels due to a

pressure gradient within the larger channel while the liquid travels in the opposite direction

by a pressure gradient within the liquid channel. The pressure gradients within both

channels are supplied by capillary forces acting along a slit passage which connects the

two channels and runs the length of the heat pipe (Figure 2.5).

Originally Space Station planned to use arterial heat pipes with a two-phase TCS.

In that configuration, the ETCS heat exchangers mentioned above were ETCS

condensers. These ETCS condensers were to be attached to Space Station while the heat

pipes were all individual orbital replacement units (ORUs). The arterial heat pipes could

be individually replaced without disturbing the flow. Because the condensers were

relatively small they could be shielded from projectile impacts. Thus, this design offered

effective resistance to micrometeoroid and orbital debris. However, a major disadvantage

of this design was the relatively heavy interface or clamping mechanism used to attach the

individual heat pipe ORUs to the ETCS condensers.

25

26

Generally the phase "heat pipe," when applied to spacecraft, refers to axial-groove heat pipes or

similar equipment.

Arterial heat pipes of at least four different length specifications have been considered or tested by

NASA. These include a 6.71-m (22-ft) unit [tested on orbit in SHARE II], a 13.11-m (43-ft) unit

[not tested], a 14.63-m (48-ft) unit [tested on the ground in SERS], and a 15.24-m (50-ft) unit [tested

on orbit in SHARE]. Additionally, the internal designs of these different arterial heat pipes varied

both due to the manufacturer and the time period. In general, the longer design specifications

preceded the shorter design specifications.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

ETCS Heat Exchanger J

Z Heat Pipes

Six Arterial Heat Pipes

Plugged into One

ETCS Heat Exchanger

An Arterial Heat Pipe ORU

Heat Pipe Evaporator Cross Section

Figure 2.5

Heat Pipe Condenser Cross Section

An overview of arterial heat pipes installed in a heat exchanger interface.

20
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

The previous comments apply to both of the heat pipe configurations considered

below. The first configuration presented is an installation using 13.11-m (43-ft) long heat

pipes 27. Longer heat pipes during ground testing demonstrated the best heat rejection per

unit mass (Chambliss and Ewert, 1990). Thus, the 13.11-m heat pipes were originally

selected as part of the Space Station baseline ATCS. Tests on orbit using a 15.24-m heat

pipe 28 were less than satisfactory. A shorter heat pipe, 6.71-m (22-ft) long, functioned

well when tested on orbit (Brown, Ungar, and Cornwell, 1992) 29.

13.1 l-meter (43-foot) Long Arterial Heat Pipe Panels

The components for the arterial heat pipe radiator installation using 13.11-m

panels are:

Arterial Heat Pipe Radiator Components

(Chambliss and Ewert, 1990,

Kantara, 1989, and LTV, 1990)

Mass Radiating Area

[kg] [Ibm] [m 2]

ETCS Condenser (per ORU panel) 30

Arterial Heat Pipe ORU Panel (dry) 31

Heat Pipe Working Fluid (ammonia)

Radiator Beam Truss (per ORU panel)

22.90 50.49 --

39.01 86.00 7.43

1.87 4.13 --

1.59 3.50 --

Total Mass and Area (per ORU panel) 65.37 144.12 7.43

The component masses, based on Kantara (1989), for the 13.11-m arterial heat pipes are:

Study Panel

Overall Length [m] 13.11

Evaporator Section: Length [m] 0.91

Mass [kg] 5.20

Condenser Section: Length [m] 12.19

Mass [kg] 33.81

Total Dry Panel Mass [kg] 39.01

The mass per radiating area is 8.80 kg/m 2. Unfortunately, longer heat pipes using a similar

internal configuration to the 13.11-m units did not function properly during the on-orbit

SHARE test. ff this design were modified to incorporate the knowledge gained from the

27 This length is the longest which will fit in the payload bay of a Shuttle (STS) vehicle.

28 This is the SHARE test.

29 This is the SHARE II test. Besides the length, the heat pipe tested during SHARE II used a different

internal arrangement than the heat pipe tested during SHARE.

30 Chambliss and Ewert (1990) mention that the ETCS condenser mass per ORU panel includes

14.0 kg (30.9 Ibm) for the interracial clamping mechanism in the LTV design.

31 These values are for a 13.11-m (43.0-fl) LTV arterial heat pipe. The condenser is 0.30 m (12.0 in.)

wide. The evaporator is 0.91 m (36.0 in.) long and 0.23 m (9.0 in.) wide. The panel is 0.0279 m

(1.10 in.) thick (Oren, 1995). Depending on the thickness, the LTV 13.11-m panels vary in mass

from 32.91 kg to 40.88 kg. The corresponding Grumman panels vary in mass from 41.19 kg to

49.41 kg. The ETCS condenser masses per panel are 22.90 kg and 19.05 kg for the LTV and

Grumman panels, respectively (Kantara, 1989).

21
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

SHARE II test, it is expected that this heat pipe should operate properly (Ungar, 1996).

However, additional testing on orbit would still be needed to flight certify this design.

6. 71-meter (22-foo0 Long Arterial Heat Pipe Panels

The components for the arterial heat pipe radiator installation using 6.71-m panels

are:

Arterial Heat Pipe Radiator Components Mass Radiating Area

(LTV, 1990) [kg] [Ibm] [m E]

ETCS Condenser (per ORU panel) 32 19.73 43.50 --

Arterial Heat Pipe ORU Panel (wet) 33 45.05 99.32 6.94

Radiator Beam Truss (per ORU panel) 3.81 8.39 --

Total Mass and Area (per ORU panel) 68.59 151.21 6.94

The mass per radiating area is 9.88 kg/m 2.

The arterial heat pipes will have high reliability due to using numerous individually

sealed elements and their ease of replacement. Loss of a single heat pipe ORU will have

negligible effect on the overall ATCS heat rejection. Further, any one heat pipe can be

replaced much more easily than a flow-through radiator panel. The overall volume for

either installation of arterial heat pipe radiators is similar to the baseline flow-through

radiators. The development cost of this option is low for the 6.71-m units because

previous work moved them to a TRL of 8. The 13.11-m units have a TRL of 6 but still

require a test on orbit to verify proper internal circulation due to the available capillary

forces. Thus, they will have a high development cost. Arterial heat pipes are flexible to

deploy. They can be individually replaced by astronauts on an extravehicular activity.

Arterial heat pipes can also be manufactured as larger panels by joining several individual

heat pipes together. On a planetary body, arterial heat pipes must be installed level to the

surface so that the local gravity does not supplant the intended working fluid circulation

path with another flow path. However, a horizontal orientation exposes an arterial heat

pipe to full solar irradiation at local noon, which seriously degrades this system's ability to

reject heat. Thus, no terrestrial uses are expected for this technology.

32 The condenser has a mass of 6.80 kg (15.0 Ibm) and the interface mechanism and socket have a total

mass of 12.93 kg (28.5 Ibm).

33 These values are for a 6.71-m (22.0-ft) LTV arterial heat pipe. The condenser is 0.61 m (24.0 in.)

wide and 5.69 m (224.0 in.) long. The evaporator is 0.23 m (9.0 in.) wide. The transition zone

between the evaporator and condenser is 0.33 m (13 in.) long.

22
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

General Qualitative Assessments:

Technology Readiness Level 34 8

Volume average

Deployment average

Reliability improved

Development Cost low

Terrestrial Use Potential none

Composite Qualitative Score 35 +1

2.3.2 Arterial Heat Pipe Radiators With Electrohydrodynamic Pumping

In general, heat pipes depend upon capillary forces to keep the working fluid

circulating. When the transport limit is exceeded, the working fluid may collect in the heat

pipe condenser and render the affected unit unusable. To maintain the capacity of a heat

pipe system, designers add extra units so that the heat transport capacity of all heat pipes

is sufficient for the anticipated load. To augment the heat transfer capacity of individual

heat pipes, Bryan (1995) proposes applying electrohydrodynamic pumping to each

individual unit. This arrangement would guard against the working fluid collecting in the

condenser and ensure that each unit would be able to continually reject its applied heat

load. For the arterial heat pipes presented in Section 2.3.1, Bryan proposes that one of

two condenser extrusions be removed. The fin efficiency is expected to decrease from

0.925 to 0.763. The added mass for electrohydrodynamic pumps is assumed equivalent to

an additional two extrusions in the evaporator section. Currently arterial heat pipes have

two condenser extrusions and six evaporator extrusions. Additionally, the

electrohydrodynamic pumps would require 2 W of power per heat pipe. Further, it is

assumed that electrohydrodynamic pumping will allow the 13.11-m heat pipes to function

as envisioned so they form the basis for the analysis here. Thus, using

electrohydrodynamic pumping for arterial heat pipes:

34 The TRL for the 6.71-m units is 8, while the TRL for the 13.11-m units is only 6.

35 This applies to the 6.71-m units. The 13.11-m units have a "high" development cost and an overall

score of -1.

23
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Arterial Heat Pipe Without Arterial Heat Pipe With

Electrohydrodynamic Electrohydrodynamic

Pumping Pumping 36

Evaporator Section Mass [kg] 5.20 6.93

Condenser Section Mass [kg] 33.81 22.54

Heat Pipe Working Fluid Mass [kg] 1.87 1.87

Electrohydrodynamic Pumping Power [kW] -- 0.002

Eiectrohydrodynamic Pumping Power as

Mass [kg] -- 0.95

Total Panel Mass [kg] 40.88 32.29

The overall mass, including the ETCS interface, per radiating area is 7.64 kg/m 2.

The overall volume for the arterial heat pipe radiators is slightly larger than the

current baseline radiator architecture due to the lower fin efficiency. However, this

difference may not be significant. Like arterial heat pipes, this radiator configuration

should have high reliability. While significant work and some prototype testing have been

completed for arterial heat pipes, a reasonable amount of work is needed to develop

working electrohydrodynamic pumps. Thus, the TRL is currently 4. The development

cost for this option is expected to be high because this assessment is based on the

unproven 13.11-m units which still require on-orbit testing. Finally, there may be some

terrestrial uses for these technologies. (See Sections 2.1.3 and 2.3.1.) However, because

this is an integrated package, no assessment for terrestrial use potential is offered here.

General Qualitative Assessments:

Technology Readiness Level 4

Volume average

Deployment average

Reliability improved

Development Cost high

Terrestrial Use Potential --

Composite Qualitative Score (0)

2.3.3 Axial-Groove Heat Pipe Radiators

Axial-groove heat pipe radiators take advantage of capillary forces to transfer heat

between hot and cold reservoirs. In practice, axial-groove heat pipes are long tubes with

longitudinal fins around the inner wall. Each heat pipe is filled with an appropriate

working fluid and sealed on both ends. Each end is placed in thermal contact with either

36 For purposes of this estimate, the heat pipe with electrohydrodynamic pumping uses an evaporator

mass multiplier of 8/6 and a condenser mass multiplier of 2/3 times the values for the panel without

electrohydrodynamic pumping. The condenser mass multiplier is not one half because the facesheets

are the same for each heat pipe.

24
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

the hot or cold reservoir. At the hot end of the heat pipe, the working fluid absorbs heat

by evaporating. At the cold end of the heat pipe, the working fluid releases heat by

condensing to the periphery. Physically, axial-groove heat pipes operate similarly to

arterial heat pipes. In axial-groove heat pipes, the vapor flows through a central core

while the liquid returns in grooves between internal longitudinal fins running the length of

the heat pipe on the periphery. The grooves also provide the wicking, or the capillary

driving force, that the slit provides in an arterial heat pipe (Figure 2.6).

Heat pipe radiator panels could be used to replace the existing flow-through

radiator panels for active thermal control (Nguyen, 1992). In this application, heat pipes

replace the flow-through tubes which join the fluid manifolds on either edge of the radiator

panels. The hot end of each heat pipe is placed in thermal contact with working fluid from

the ETCS as it enters the radiator assembly while the cold end is placed in contact with the

radiator face sheet. In practice, heat pipes are placed between two separate manifolds

which each contain an entering and returning line for the ETCS working fluid

Heat-pipe panels are expected to be more reliable than conventional flow-through

radiators. A micrometeoroid puncturing any radiator ORU panel destroys only a single

heat pipe. However, because axial-groove heat pipes only replace the flow-through panel

flow tubes, which have a fairly low probability of puncture compared with the overall

probability of a loop puncture, this is not a significant consideration for this technology

(Christiansen, 1992). Finally, axial-groove heat pipes are a proven technology with a TRL

of 9. Axial-groove heat pipes have been used successfully to reject heat on robotic

spacecraft (Nguyen, 1992). However, the TRL for an integrated radiator using axialgroove heat pipes in place of flow-through tubes is only 6. Thus, the developmental cost

for such a system is low. Finally, there may be some terrestrial uses for this technology.

These units resemble reflux boilers or other similar devices.

General Qualitative Assessments:

Technology Readiness Level 6

Volume average

Deployment average

Reliability average

Development Cost low

Terrestrial Use Potential possible

Composite Qualitative Score +1

2.4 Lightweight Radiators

Lightweight radiators replace conventional materials within radiator units with

lighter, high-strength substitutes. The strength and efficiency of the radiator units which

are critical to specific missions are maintained while the overall mass is reduced. Further,

by tailoring the system specifically to its final environment, such as on a planetary surface

versus on orbit, newer radiator designs are optimized for the intended mission.

25
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Radiator

ORUs

Radiator Manifolds

0

Radiator Panel

Heat Extracted from A

Manifolds (Fluid Evaporates

Inside of Heat Pipe)

I -@

Heat Pipe

A_ _ Heat Rejected

to Space (Fluid

Condenses Inside

of Heat Pipe)

)

Liquid Travels by Capillary

Action Between the Fins

Vapor Travels by Pressure

Difference in the Pipe Core

Crosssection A-A

Figure 2.6 An overview

replacement unit.

of axial-groove heat pipes installed in a radiator orbital

26
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Several concepts which are lighter versions of other radiators are examined here.

Thus, this section's uniting technologies are the composites and fabrication techniques

employed to produce each of these radiators rather than a specific heat transfer mechanism

or similar equipment as in other sections. Initially, each radiator is evaluated to determine

its relative mass per surface area. These estimates provide guidance for generic estimates

of the mass savings expected for a particular mission when using lightweight radiators.

Such generic mass savings are employed throughout this study for lightweight radiators.

Once further details for specific lightweight radiator concepts become available, those

concepts might be assessed individually for the various reference missions.

2.4.1 Composite Flow-Through Radiators

This lightweight radiator concept applies composite materials to the traditional

flow-through radiator architecture. This concept is under development to possibly replace

or supplement the water sublimator currently used to reject heat from the portable life

support system of the extravehicular mobility unit (Cross, 1996). The heart of this design

employs flow-through radiator panels utilizing flow passages of rectangular cross section

formed from sheets of carbon composite material. Above the sheet of flow passages is

placed another sheet of carbon composite material separated by a thin gap. The gap is

either evacuated or filled with low-pressure nitrogen gas (10 torr). While the gap is filled

with gas, it will provide conduction paths between the sheet of flow tubes and the

separated sheet. When the gap is evacuated, the only heat exchange mechanism between

the sheet of flow passages and the separated sheet is radiation. In other words, the

evacuated gap effectively acts as insulation for the sheet of flow passages. This prevents

the working fluid from freezing (Figure 2.7). The mass per radiating area for single-sided

heat rejection using this design is 8.21 kg/m 2. This value includes mass for the manifolds

and assumes ammonia as the working fluid. For a double-sided radiator panel, the mass

per radiating area is 4.67 kg/m 2 after accounting for an additional gas gap cover sheet.

There are several issues about this concept which require more work. The most

ingenious aspect of this concept is the use of a gap, which may use either conductive or

radiant heat transfer, to control the rate of heat loss from the radiating surface. This, in

turn, allows the user to protect the radiator from cold environments while making full use

of the radiating surface for hot environments. This is an extremely useful development.

The current design, which is under development for possible use in an extravehicular

mobility unit, is too heavy to be applied directly to vehicle or habitat thermal control

architectures. However, this design for an extravehicular mobility unit uses flow passages

which are spaced so close to each other as to render the radiator fin efficiency as almost

unity. As such, a composite flow-through radiator could possibly have a much lighter

mass per area when optimized for a base or habitat. Further, though the radiator itself is

fairly rigid, work to date has not addressed the structure or deployment necessary to

integrate this technology into a large TCS. Thus, this concept is presented here because

of its general utility even though it is not ready to be applied to a vehicle or habitat TCS.

The composite flow-through radiator is in many ways identical to the baseline

metal flow-through radiators. As such, the volume and deployment for this concept are

expected to be average. The reliability will also be average although the method of

protecting the radiator from episodes of frozen working fluid within the flow passages

differs from that of the baseline architecture. Even though the TRL is 4, the development

27
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

cost for this technology is expected to be average because all remaining testing may be

completed in terrestrial facilities. Terrestrial uses for this technology, or at least for the

material processing techniques used to develop these components, is highly likely.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

4

average

average

average

average

good

Composite Qualitative Score +1

2.4.2 Composite Reflux Boiler Tube Radiators

Composite reflux boiler tube radiators 37 utilize thin-walled reflux tubes encased in

an epoxy sleeve as radiator elements. As such, these devices require a gravitational field

to operate efficiently. The units currently under development have a circular cross-section

with a diameter of 0.02395 m and a length of 1.143 m. The condenser section length is

1.067 m, which implies a radiating surface area of 0.08027 m 2 per unit. Because these

units are not finned, the effective fin efficiency is 1.0. This radiator, however, will "see"

other portions of itself so its view factor to space will not be 1.0 as it is for a planer

radiator. The exact view factor to space will depend upon the selected tube configuration.

Each unit includes a thin-walled titanium tube surrounded by an outer sleeve of epoxy

composite. The composite outer sleeve provides burst support for the titanium tube.

Each unit has a mass of 0.165 kg including 0.020 kg of water which is the working fluid

(Thermacore, 1995 b). An ETCS interface for 320 tubes for a similar deployment has an

estimated mass of 35.62 kg (Blades, 1996) 38

Composite Reflux Boiler Tube Component

Unit Components [kg]

Titanium Liner/Epoxy Sleeve 0.145

Working Fluid (water) 0.020

ETCS Interface per Tube 0.111

Mass per Tube 0.276

37 An earlier reflux boiler tube radiator concept is the ultralight fabric reflux tube radiator (Hurlbert,

Ewert, Graf, Keller, Pauley, Guenther, and Antoniak, 1996). While the composite reflux boiler tube

is a descendent of the previous project, it does not share the shortcomings associated with the

ultralight fabric reflux tube units.

38 The estimate here for the ETCS interface is based on the ultralight fabric reflux tube which is similar

in dimensions to other reflux boiler tube designs.

28
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Gas

Plate

6.25 in

Gas Inlet

24.25 in _lt Port

Manifold

A

v

32.0 inches

22.0 inches I

Water

Outlet

2.25 in

To end

of brass

firing

(0.5 in)

0.75 in

Water

Manifold

Water _ Overall 336in I

Inlet

2.5 in ) of Brass Fitting (0.5 in) _ B

0.5 in OD

Graphite Epoxy Tube

Side View

B

Water Manifold

in OD _h,.._at_r

0.75

Passage

Section A-A

Radiator and Manifold Only

Not to Scale

GasGap

Plate

Continuous I1_ Radiator Top

l Fin \ V / Surface I[

T t "_'' '......... I ....., _.,r .... " "_ * 1/32in

0" 175_in,_ _L-_i_, L_ ",_ _r _ in. 025

T I 0 5 in Water Passage I

View B-B

Rotated 90 ° Right

Not to Scale

Figure 2.7 The configuration for the extravehicular activity composite radiator test article

from Cross (1996). This concept is based on the traditional single-phase, flow-through

radiator architecture. The carbon composites form the unit instead of aluminum as in the

baseline architecture. Additionally, this radiator features a gas gap which insulates the

flow passages when evacuated. The portable life support system for which this concept is

under consideration will use water as its the working fluid. However, a thermal control

system radiator using this approach will probably circulate ammonia or another fluid

besides water.

29
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

The overall mass per surface area for composite reflux boiler tube radiators is

3.44 kg/m:. While planetary surface missions are not expected to operate in high debris

fields, composite reflux boiler tube radiators are very similar, with regard to redundancy,

to arterial heat pipe radiators. If necessary, the ETCS interface could be shielded from

debris punctures while the composite reflux boiler tube radiator units remain exposed to

the environment. A puncture in one or even a few units would not significantly reduce the

heat-rejection ability of the overall system. Based on their architecture, composite reflux

boiler tube radiators have improved reliability. The deployment of the overall system is

also expected to be comparable to the baseline architecture, although the vertical

geometry of the reflux tube units may be advantageous in some cases. The volume should

be no more than that of the baseline architecture. The development cost for this

technology is low because the TRL is already 5 and any remaining testing may be

completed in terrestrial facilities. From this application, both the lightweight reflux boiler

tubes and the associated material processing techniques could have terrestrial uses.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

average

average

improved

low

good

Composite Qualitative Score +3

2.4.3 Composite Heat Pipe Radiators

Composite heat pipe radiators are a lightweight variation on the heat pipes

presented in Section 2.3.3. Based here on the work of Juhasz and Bloomfield (1994) and

Juhasz and Rovang (1995), the composite heat pipe is constructed from a thin niobiumzirconium alloy foil liner covered by a carbon-carbon composite outer shell. The foil liner

has a wall thickness of 0.064 millimeters which is furnace brazed to the composite outer

shell. The evaporator section, which extends 0.076 m beyond the composite liner on one

end of the unit, has a thicker wall of 0.4 millimeters. The heat pipe itself has a circular

cross section with a diameter of 0.025 m. The composite shell, with two opposing 0.05 m

wide fins, surrounds the foil liner for the length of the condenser section, 0.914 m 39. The

composite has a thermal conductivity of roughly 0.575 kW/(m*K) at 300 K 40. The

overall test unit is about a meter in length.

39 The view factor to the environment and the fin efficiency of this radiator are unknown but both are

less than unity. To avoid this problem, numerical analysis, correlation, or experiment are useful.

40 This heat pipe research was funded by NASA Lewis Research Center to develop radiators to reject

high temperature heat from power systems in space. This information is provided by Juhasz (1996).

30
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

The components for the composite heat pipe radiator installation are:

Composite Heat Pipe Radiator Components

(Based on Juhasz and Bloomfield, 1994,

and Juhasz and Rovang, 1995)

Mass Radiating Area

[kg] [m 2]

ETCS Interface (per composite heat pipe) 41 0.111 --

Composite Heat Pipe Panel (dry) 0.322 0.2335

Working Fluid (per composite heat pipe) 0.014 --

Total Mass and Area (per composite heat pipe) 0.447 0.2335

The component masses for the composite heat pipes are (Juhasz and Rovang, 1995):

Masses of 0.99 m Component

Test Panel Components Mass [grams]

Carbon-Carbon Composite Shell 214.0

Liner Including Evaporator 41.2

End Caps 13.1

Fill Tubes 7.2

Braze 22.5

Wick 24.0

Working Fluid (liquid metal) 42 13.5

Mass per Panel 335.5

The overall mass per surface area for composite heat pipe radiators is 1.91 kg/m 2.

This unit is designed to operate vertically on a planetary surface. Because the composite

heat pipe already contains an integral wick, slight modifications should allow them to

function under microgravity also (Juhasz, 1996). The volume for this installation will

probably be comparable to that of the baseline architecture. Significant work and

prototype testing have been completed for composite heat pipes, although flight-testing is

still required for any on-orbit applications. The current TRL is 5. Thus, the development

cost is expected to be low for planetary applications and high for orbital applications.

Like the arterial heat pipes presented in Section 2.3.1, the composite heat pipes detailed

here should provide a high level of radiator redundancy and therefore have high reliability.

Terrestrial uses from this technology include advances in materials processing plus those

presented for axial-groove heat pipes in Section 2.3.3.

41

42

The ETCS interface is assumed to be similar to the unit given by Blades (1996). See Section 2.4.2.

The high temperature (800 K) application investigated by NASA Lewis Research Center required

liquid metal as a working fluid. However, Juhasz and Rovang (1995) note that this technology has

been demonstrated for the 400 K to 450 K range using a stainless steel foil liner and demineralized

water as the working fluid. Both the high temperature panel and this low temperature panel have

roughly the same mass (Juhasz, 1996).

31
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

average

average

improved

low

good

Composite Qualitative Score 43 +3

2.4.4 Unfurlable Radiators

Unfurlable radiators and composite reflux boiler tube radiators both utilize reflux

boiler tubes as their base heat-rejection element. Unfurlable radiators differ from the

previous concept in that the boiler tube condenser is flexible instead of rigid. In fact, the

condenser wall is manufactured from a metal and plastic laminate which collapses under

terrestrial atmospheric pressures when the unit is not operating (Gernert and Donovan,

1994). When the unit is operating, the internal working fluid vapor is sufficient to inflate

the condenser. As with other reflux boiler tube devices, unfurlable radiators require a

gravitational field to operate properly. To package this concept, the unfurlable radiator

tubes would be affixed to an interface module. Upon activating the TCS, working fluid

vapor would inflate the radiator unit condensers.

The prototype test units have a circular cross-section with a diameter of

0.01905 m and an overall length of 1.000 m. The condenser section length is 0.951 m

which implies a radiating surface area of 0.05694m 2 per unit (Thermacore, 1995 a).

These prototype units are not finned, so their effective fin efficiency is 1.0. Again, this

radiator will "see" other portions of itself so its view factor to space will not be unity as it

is for a planer radiator. The exact view factor to space will depend upon the selected unit

configuration. Each unit includes a thin-walled polymeric film and metal laminate

condenser attached to a rigid evaporator base. In the completed radiator, the evaporator

units sit in an ETCS interface assembly which supplies the heat load. An ETCS interface

for 320 reflux boiler tubes has an estimated mass of 35.62 kg (Blades, 1996) 44

Composite Reflux Boiler Tube

Unit Components

(Thermacore, 1996)

Condenser With Fiber and Cap

Evaporator, Epoxy, and Cap

Working Fluid (water)

ETCS Interface per Tube

Component

[kg]

0.051

0.089

0.012

0.111

Mass per Tube 0.263

43 This is for planetary missions. For orbital missions, the development cost is "high" and the overall

score is + 1.

44 The estimate here for the ETCS interface is based on the ultralight fabric reflux tube which is similar

in dimensions to other reflux boiler tube designs.

32
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

The overall mass per surface area for unfurlable radiators is 4.62 kg/m 2. While

planetary surface missions are not expected to operate in high debris fields, unfurlable

radiators are very similar, with regard to redundancy, to arterial heat pipe radiators, ff

necessary, the ETCS interface could be shielded from debris punctures while the radiator

units remain exposed to the environment. A puncture in one or even a few units would

not significantly reduce the heat-rejection ability of the overall system. The deployment of

the overall system should be easy because little setup is required for this radiator. The

vertical geometry of the reflux tube units may be advantageous in some cases. The stored

volume should be less than that of the baseline architecture. The development cost for this

technology is expected to be average because prototype testing is in progress and the

current TRL is 4. As with other lightweight radiators, the primary terrestrial use from this

program will probably be advances in processing composite materials.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

4

compact

easy

improved

average

good

Composite Qualitative Score +4

2.5 Other Heat Rejection Technologies

Several other radiator-related technologies are under consideration for near-term

missions. These include using phase-change materials (PCMs), to provide supplemental

cooling for short periods requiring higher than average heat rejection, and radiator shades,

which shield a radiator from strong irradiation sources.

2.5.1 Phase-Change Thermal Storage

To implement phase-change thermal storage within an ETCS, a portion of the

empty volume within the radiator panels could be filled with packages of PCM, or wax 45

These packages would be positioned around the current flow-through tubes within the

radiator panels. When the vehicle experiences high heat loads, the hot ETCS working

fluid could pass through these radiator panels and melt the previously solidified PCM,

effectively increasing the rate of heat removal from the working fluid while the PCM

melted. When the vehicle experiences low heat loads, the tubes surrounded by PCM

could be bypassed, allowing the PCM to solidify. At least two scenarios are possible to

size a phase-change thermal storage device.

45 A PCM for an application is a solid at the minimum operating temperature, but the material melts at

some temperature below the maximum system operating temperature. Thus, they mimic candle wax,

which is solid at room temperature, but melts easily near a source of flame. In fact, long-chain

alkanes, which are a class of compounds including waxes, are useful as PCMs.

33
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

One case would presume to replace a heat-rejection system which uses expendable

materials with phase-change thermal storage. A system using expendables is sized to

handle an expected heat load for some finite length of time, such as that experienced by a

vehicle during aero-braking. An example of a heat-rejection device using expendable

materials is the ammonia boiler subsystem on STS.

A second case assumes that phase-change thermal storage is used to provide

auxiliary cooling in conjunction with other heat-rejection systems such as radiators and

heat-rejection devices using expendable materials. In particular, phase-change thermal

storage could reduce usage of the expendable heat-rejection device, thereby reducing the

mass consumed by the vehicle 46. The PCM packages are solidified while the radiator

rejects heat above the average orbital value and the PCM is allowed to melt to provide

additional cooling while the radiator is rejecting heat below the average orbital value.

Because the radiator heat load varies around an orbit, even when a vehicle is entirely in

full sun at high beta angles, this approach is generally applicable. Figure 2.8 schematically

illustrates the assumed implementation of phase-change thermal storage for Shuttle. The

PCM packages could be integrated within the aft radiator panels around the existing tube

bundle and panel honeycomb 47 or in a separate unit placed either upstream or

downstream of the radiator panels 48

For this study, two materials were considered. Water has a thermal density 49 of

10.0kg/kW*h for phase-change applications, or 18.0kg/kW*h including packaging 50

Water solidifies at 273.2 K under standard atmospheric pressure. Further, water also

expands upon solidifying. A long-chain alkane, or wax, such as n-dodecane, has a thermal

density of 16.8kg/kW*h for phase-change applications, or an assumed value of

30.2 kg/kW*h including packaging, n-Dodecane solidifies at 263.6 K under standard

atmospheric pressure. While water is the preferred material, n-dodecane is assumed to be

representative of a range of alkanes which, taken collectively, could be tailored to handle

phase-change loads over a wide range of temperatures 51. Water, however, is restricted to

applications near 273.2 K.

From a qualitative perspective, phase-change thermal storage offers some

advantages. The volume consumed by this system could be negligible because it could fill

volume already set aside for radiators. Installation for this system should be relatively easy

because the proposed components should readily fit in available volume. The reliability

should be comparable to the current systems. This technology uses the passive technology

46 For example, as designed, Shuttle uses water generated by the production of electricity in its fuel

ceils as feedwater for the flash evaporator subsystem. Early in the next century, however, that water

generated by Shuttle's fuel cells may be tapped as a source of potable water to be delivered to ISS or

for the crew on extended-duration Shuttle missions. Use of phase-change thermal storage would

save some water which currently is consumed by the flash evaporator subsystem.

,*7 The panel honeycomb is aluminum and provides additional conduction pathways for heat transport

from the radiator tubes to the radiator surface sheet, which is a fin. Thus, the honeycomb is

effectively an extension of the surface fin.

48 The thermal densities used in this study for estimating the mass of the phase-change thermal storage

device include packaging mass and do not assume an in-panel installation.

49 Thermal density is defined here as l/(heat of fusion).

50 Packaging mass is assumed to be 80% of the mass of the PCM.

51 The actual melting points of the various alkanes differ, but their overall thermal densities are

assumed to be roughly similar to that of n-dodecane.

34
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

of phase-change cooling but requires sensor units and automatically controlled valves, so

reliability is expected to be average. Even though phase-change thermal storage has been

tested on orbit historically, the TRL for missions here is 5. Thus, the cost to integrate this

technology into other missions should be no more than average. Terrestrial applications

may exist for the PCMs when packaged within personal protection suits for high

temperature applications. Applications may also exist for cooling equipment, such as

machinery for cutting or forming of hard materials, which can experience extremely high

heat loads for short intervals. PCMs might also be incorporated as packaging materials

for any item which is to be maintained at a temperature other than that of its environment.

Current Shuttle Cooling Schematic

Aft Radiator Panels

: Forward Radiator Panels [I

ETCS Radiator Bypass Line

ETCS Flow Loop

Revised Shuttle Cooling Schematic with Phase-Change Material Packages

___ With Aft Radiator PCM Packages Panels

ETCS Radiator Bypass Line

ETCS Flow Loop

_ Secondary Bypass Line

@.

Foward Radiator Panels

Figure 2.8 To integrate phase-change thermal storage into Shuttle, the phase-change

material packages could be placed within the aft radiator panels or in units outside of the

radiators. To solidify the phase-change material while the radiators are facing a relatively

cold environment, the ETCS flow loop, which is warmer than the melting point of the

phase-change material, bypasses the phase-change material packages by means of the

secondary bypass line. While the vehicle is in a hot environment, the ETCS flow loop can

receive additional cooling by melting the previously solidified phase-change material.

35
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

compact

average

average

average

good

Composite Qualitative Score +2

2.5.2 Parabolic Radiator Shade

Parabolic radiator shades provide a means of lowering the effective sink

temperature around a radiator to allow heat rejection at temperatures associated with

waste heat from environmental control and life support systems for human beings. The

shade, which has a specular upper surface and gray under surface, is designed to redirect

incident solar radiation back to space while directly blocking diffusely scattered and

infrared radiation from the lunar surface (Ewert and Clark, 1991, Keller, 1994, and Ewert,

Graf, and Keller, 1995). More specifically, the shade's parabolic shape focuses the

incoming solar radiation along a line above a radiator placed in the shade's trough. The

surface coatings on the lower side of the shade reflect thermal energy from the planetary

surface (Figure 2.9). The shade must be aligned such that the incident solar vector is

parallel to the radiator throughout the day. Otherwise, the shade itself will direct solar

radiation on to the radiator in addition to any energy directly striking the side of the

radiator. Because Luna only tilts 1.53 degrees on its axis from the orbital plane, a

parabolic radiator shade is ideal at the lunar equator and can be readily adapted for other

lunar sites.

The overall volume is expected to be similar or slightly more compact than a

standard radiator. For a long term mission, deployment of a radiator shade should be

comparable to deployment of a larger flow-through radiator system. For a lander vehicle,

an extravehicular activity may be required to align the shade system properly and this

would be more difficult than deploying radiators mounted on the vehicle. The reliability is

expected to be the same as standard equipment for the short duration mission. For longer

missions, the issue of dust accumulation on the radiator shade, which degrades the shade's

effectiveness, is more significant. Occasional replacement of the shade for longer missions

should alleviate this problem. The program TRL is 5. Some additional testing is needed,

but the development cost remaining is average. Currently, no terrestrial uses are expected

for the parabolic radiator shade.

36
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Incident Solar

Ray Deflected

by Shade

Focal Line for Irradiation

Deflected Ray

4 from Albedo

Parabolic Radiator Shade

Figure 2.9 A parabolic radiator shade around a vertical radiator. The incident solar

irradiation hitting the upper shade surface is reflected to a focal line above the radiator,

while irradiation from the planetary surface is deflected away by the underside of the

shade. Endsheets guard the radiator from radiation enterting through the ends of the

shade.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment 52

Reliability

Development Cost

Terrestrial Use Potential

5

average

average

average

average

none

Composite Qualitative Score -1

2.6 Additional Technologies

The heat transfer technologies presented below are unlike other technologies

presented in this study.

2.6.1 Rotary Fluid Coupler

In the ISS TCS, the radiator ORUs are clustered in two sets of three ORUs each.

Each cluster is fixed to a truss which rotates on a thermal radiator rotary joint (TRRJ) to

maintain the radiator panel faces parallel to the local solar vector. This arrangement

52 For missions where a vehicle mounted system is preferable, the deployment for this option would be

"difficult" yielding a Composite Qualitative Score of "-2" for this technology.

37
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

minimizes solar absorption by the radiators. To allow ATCS fluids to pass from within

ISS to the ETCS radiators, a rotary coupler is part of each TRILl.

A rotary fluid coupler is a fully rotating device which uses liquid seals to contain

and separate the internal fluid flow channels (Harwell, 1992). A fully rotating device

allows the radiators to always be aligned with the panel faces parallel to the solar vector.

The projected rotary fluid coupler mass is 10.9 kg (24 Ibm) with an approximate volume of

0.0127 m 3 (0.45 ft3). Further, the rotary fluid coupler is expected to be more reliable than

the baseline coupler because its components have longer service lives and its smaller size

makes it less susceptible to micrometeoroid puncture. Also because of its size, the rotary

fluid coupler is easier to replace. The rotary fluid coupler is similar in size and mass to a

pump allowing it to be readily handled by an individual astronaut. The rotary fluid coupler

also displays some disadvantages. The liquid seals between flow channels permit low

amounts of leakage between channels when they wear, but this is not a significant

problem. Further, the rotary fluid coupler is geometrically different from the current

coupler. As such, it may be expensive to retrofit ISS to accept rotary fluid couplers.

Even though the previous rotary fluid coupler program was discontinued,

significant work has been completed. The TRL is 7. Therefore, the developmental costs

should be low. This technology provides a generic, low-leakage rotary coupler.

However, the rate of rotation is limited to 45 degrees per minute. Therefore, terrestrial

uses may be possible in some large industrial machines.

General Qualitative Assessments:

Technology Readiness Level 7

Volume compact

Deployment easy

Reliability average

Development Cost low

Terrestrial Use Potential possible

Composite Qualitative Score +3

2.6.2 Plant Chamber Cooling Improvements

Regenerative life support systems are of great importance as an enabling

technology for extended-duration missions with crews. In particular, plant growth

chambers promise to provide future space explorers with both food and oxygen

replenishment. The practical aspect of this system then becomes optimizing the edible

biomass (food) produced by a plant growth module as a function of mass. This study

concentrates on improving an assumed baseline chamber by adding more reliable/less

massive equipment in place of the baseline materials.

The baseline plant growth module for this study is patterned after a test chamber at

the Lyndon B. Johnson Space Center in Houston, Texas, built by the NASA Crew and

Thermal Systems Division. Jerng (1991) and Barta, Dominick, and Kallberg (1995)

provide descriptions of this chamber which is known as the Johnson Space Center 10-foot

Regenerative Life Support System Test Chamber or, more recently, as the Variable

38
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Pressure Growth Chamber (VPGC). The relevant overall parameters of the plant chamber

assumed here are (Jerng, 1991) 53:

Physical Dimensions:

Chamber Volume

Crop Tray Growing Area (per tray)

Number of Crop Trays

Number of Lamps

Total Growing Area

Biomass (wheat, Triticum aestivum) 54

Maximum Power/Heat Loads 55.

Atmospheric Circulation Fan/Blower 56

Lighting:

Lamps (high pressure sodium)

Ballasts

Reheat 57

Total Power/Heat Load

Assumed Power Usage

Power as Mass

13.7 m 3

1.43 m E

4

32

5.72 m 2

21.7 kg

1.9 kW

12.8 kW

2.3 kW

3.0 kW

20.0 kW

Continuous

15,000.0 kg

Figure 2.10 provides a general illustration of a portion of the plant growth

chamber. Overall, the redesigned plant growth chamber will be similar to the baseline

configuration. Because the bulk of the chamber volume is dictated by the volume

requirements of the crop, the volume of the redesigned chamber is expected to be similar

to the baseline. Further, the deployment is expected to be similar to the baseline chamber.

The reliability is expected to be slightly higher due to equipment improvements. The

overall TRL is 5 because some testing and development has been completed. The

development cost for any additional work is expected to be low. Terrestrial uses for these

technologies are less obvious. As applied to plant growth, there may be some interest

within agricultural products companies and research labs, especially for small, high-value

crops which require closely controlled growing conditions. Thus, because few terrestrial

users will, in the foreseeable future, require completely isolated chambers for growing

plants, there is little anticipated terrestrial use for this technology. However, terrestrial

uses for individual components of the plant chamber assembly are possible.

53 The dimensions listed here correspond to one half of those for the overall VPGC.

54 Actual data from the VPGC using high-pressure sodium lamps.

55 Actually, these are power inputs for the listed systems. However, all input power is assumed to

eventually require rejection as heat, so these are also heat loads associated with this equipment.

56 Estimated for 45 Hertz.

57 Pau/ (1995). The reheat increases the temperature of the air stream before it returns to the plant

growth volume. In some cases, to control humidity within the growing volume of the chamber, the

air is cooled below the return air stream set-point temperature. After removing the excess water

vapor, the reheat increases the air temperature before the stream returns to the growing volume.

39
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

average

average

average

low

possible

Composite Qualitative Score +1

2.6.3 Carbon Brush Heat Exchanger

Heat exchange across an evacuated gap between two fiat plates is a common

problem. The gap effectively adds a resistance to heat transfer which, in some cases when

a vacuum fills the gap, is the controlling resistance. One solution, when an extravehicular

activity compatible interface is required, is to mount fins extending perpendicularly from

each surface bounding the gap. When the two fin sets are intermeshed, a radiant fin

interface is formed and heat is transferred primarily by radiation. Knowles (1995) gives

the interfacial conductivity as 60 W/(m2*K) for this arrangement.

Another solution to transfer heat across the gap is to mount fine carbon fibers

extending perpendicularly from the gap boundaries. Heat is again transferred by radiation

but also by contact conductance when the fibers are allowed to intermesh. This

arrangement is a carbon brush interface. Though this arrangement is similar to the radiant

fin interface, the interfacial conductivity increases to 2500 W/(m2*K) (Knowles, 1995) due

to increased area for radiant transfer and the numerous physical contact points for

conductive transfer (Figure 2.11).

Currently the dc-to-dc converter units (DDCUs) are affixed to coldplates

transferring heat using radiant fin interface. The fins, which are made of aluminum, are

0.0508 m tall by 762 micrometers thick and run the length of the DDCU coldplate 58. In

order to save mass and improve DDCU cooling efficiency, the radiant fin interfaces could

be replaced by carbon brush interfaces. A quick one-dimensional heat transfer analysis

reveals that this change will increase the interfacial conductivity by almost two orders of

magnitude. This, in turn, yields more efficient cooling of the DDCUs. What further

advantage or savings this imparts is dependent on the cooling needs of other units on the

ETCS loop.

58 These dimensions apply to units designed for use on ISS. Other units are assumed to be comparable.

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COOLING

WATER

/

AIRFLOW

i

Bypass Air

Figure 2.10 An overall drawing of the Variable Pressure Growth Chamber used by

Ewert, Paul, and Barta (1995). This view shows two of eight plant growth trays plus the

chamber environmental control equipment. The coldplate and heat exchanger use liquid

water as the working fluid while the reheat coil is a resistance heating device. The arrows

indicate the mean airflow pattern within the chamber.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Overall, the carbon brush heat exchanger has excellent properties. The volume of

the carbon brush interface is slightly smaller than a radiant fin interface. Units using a

carbon brush interface should be as easy to deploy as the current design. The carbon

brush should be very reliable and inexpensive. Because the theoretical basis for this idea is

well developed at a TRL of 5, the development cost is low. For terrestrial applications,

this technology could be used to cool high temperature electronics. The major

disadvantage of the carbon brush heat exchanger is the possible generation of small

particle debris as individual fibers detach or break upon handling (Knowles, 1995). This

may be unacceptable near electronic devices such as DDCUs because the particles may

float into the electronics. Depending on other considerations, such as component

packaging, the hazard of such particle debris may be reduced.

General Qualitative Assessments:

Technology Readiness Level

Volume

Deployment

Reliability

Development Cost

Terrestrial Use Potential

5

average

average

improved

low

good

Composite Qualitative Score +3

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DDCU

Radiant Fin

Interface

Coldplate Socket __

I, I I I ..

ColdplateSocket , _ I I , II , I I , I.,l'q

with Carbon Fibers [/_li I II I i !_ I

(for illustrative i./171''1 II ii1"11/,,!" /

purposes only) /171 "1 "11" i i I i i_ /

041111111ill ¢y /

I

DDCU Coldplate

Carbon Brush

Interface _ ....

Coldplate Socket

Figure 2.11 The radiant fin and carbon brush heat exchangers.

to scale.)

(This figure is not drawn

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2.7 Summary

Overall, the qualitative assessments of the various technologies

summarized in the following table:

Table 2.1 Summary of Qualitative Assessments for

Advanced ATCS Architecture Technologies

may be

Qualitative Assessments

,--- O .-- Q

;> o > oo

2.1

2,1.1

2.1.2

2.1.3

2.1.4

Two-Phase Thermal Control Systems

Two-Phase Thermal Control System

With Mechanical Pump/Separator

Low-Power Two-Phase

Thermal Control System

Two-Phase Thermal Control System

With Electrohydrodynamic Pumping

Capillary Pumped Loops

0 0 0 +1 -1 0

0 0 0 0 -! -1

0 0 +1 0 +1 +2

0 0 +1 0 -1 0

2.2

2.2.1

2.2.2

2.2.3

2.2.4

Heat Pumps

Vapor Compression Heat Pump

Solar Vapor Compression Heat Pump

Complex Compound Heat Pump

Zeolite Heat Pump

0 0 -1 +1 +1 +1

+1 0 -1 +1 +1 +2

+1 0 0 0 0 +1

+1 0 0 0 0 +1

2.3

2.3.1

2.3.2

2.3.3

Heat Pipe Radiators

Arterial Heat Pipe Radiators

Arterial Heat Pipe Radiators

With Electrohydrodynamic Pumping s9

Axial-Groove Heat Pipe Radiators

0 0 +1 +1 -1 +1

0 0 +1 -1 -- (0)

0 0 0 +1 0 +1

2.4

2.4.1

2.4.2

2.4.3

2.4.4

Lightweight Radiators

Composite Flow-Through Radiators

Composite Reflux Boiler Tube Radiators

Composite Heat Pipe Radiators

Unfurlable Radiators

0 0 0 0 +1 +1

0 0 +1 +1 +1 +3

0 0 +1 +1 +1 +3

+1 +1 +1 0 +1 +4

2.5

2.5.1

2.5.2

Other Heat Rejection Technologies

Phase-Change Thermal Storage

Parabolic Radiator Shade

+1 0 0 0 +1 +2

0 0 0 0 -1 -1

2.6

2.6.1

2.6.2

2.6.3

Additional Technologies

Rotary Fluid Coupler + 1 + 1 0 + 1 0 +3

Plant Chamber Cooling Improvements 0 0 0 + 1 0 + 1

Carbon Brush Heat Exchanger 0 0 +1 +1 +1 +3

59 This is a combination of technologies, so no assessment is provided for terrestrial use potential,

Such assessments are provided under the individual component technologies.

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Advanced technology radiators exhibit a wide range of masses per radiating area.

The baseline heat-rejection devices for the reference missions below assume flow-through

radiators constructed from aluminum alloys using single-phase working fluids. For the

vehicles and habitats examined, the mass per radiating area, excluding the fin efficiency,

falls between 5.0 kg/m 2 and 8.5 kg/m 2. The table below summarizes the masses per

radiating area for the various radiators presented as advanced technologies. These masses

per radiating area are based on the total radiator mass, including any working fluid,

divided by the radiating surface area.

Table 2.2 Summary of Masses per Radiating Area

for Advanced Technology Radiators

Mass per Effective Mass

Radiating Area Fin per Radiating

[kg/m 2] Efficiency Area 60 [kg/m 1]

2.3

2.3.1

2.3.2

Heat Pipe Radiators

13.11-m Arterial Heat Pipe Radiators

6.71-m Arterial Heat Pipe Radiators

Arterial Heat Pipe Radiators

With Electrohydrodynamic Pumping

8.80 0.925 9.51

9.88 0.925 10.68

7.64 0.763 10.01

2.4

2.4.1

Lightweight Radiators

Composite Flow-Through Radiators

(Single-Sided Rejection)

Composite Flow-Through Radiators

(Double-Sided Rejection) 4.67 ~ 1.00 4.67

2.4.2 Composite Reflux Boiler Tube Radiators 3.44 1.00 3.44

2.4.3 Composite Heat Pipe Radiators 61 1.91 0.97 1.96

2.4.4 Unfurlable Radiators 4.62 1.00 4.62

8.21 ~1.00 8.21

60 The effective mass per radiating area is the mass per radiating area divided by the fin efficiency.

61 The fin efficiency listed for the composite heat pipe radiator is estimated based on an analysis using

the Thermal Synthesizer System (TSS) and SINDA/FLUINT. The actual value is unknown.

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3.0 ORBITAL MISSIONS

Several aspects of orbital operations are unique. Microgravity, generated by the

motion of the vehicle itself, is the only body force which applies to orbital operations. The

time constant associated with an orbital cycle is on the order of a couple of hours or less

for a low altitude orbit. This allows a frequent transition between sunlight and shade.

Finally, by the nature of the environment itself, there are few additional natural resources

or atmosphere of any type associated with an orbital mission.

3.1 INTERNATIONAL SPACE STATION EVOLUTION

3.1.1 Reference Mission

This study assumes that the current International Space Station (ISS) TCS ORUs

will need replacement and augmentation after year 20 (year 2017, assuming a project

genesis in 1997). The current radiator configuration is six flow-through radiator ORUs in

two clusters. Due to micrometeoroid impacts, optical property degradation, and other

failures, it is assumed that all six radiator ORUs will be replaced. Further, it is assumed in

year 20 ISS will be augmented with two solar dynamic power modules increasing the

onboard power supply of the U.S.-led portion of ISS from 75 kW to 95 kW. This will

require an additional 20 kW of heat-rejection capability. An alternate reference mission

for the advanced thermal technologies below would be to construct an entirely new space

station.

3.1.2 Baseline Case

ATCS requirements for ISS (Howell, et al., 1994):

• Loop A and Loop B are independent flow passages with Loop A on

segment S1 and Loop B on segment P1. Each loop will use liquid

ammonia (NH3) to carry heat from the ITCS to the heat-rejection devices

in the ETCS. (See Figure 3.1 for a schematic of the ISS ATCS.)

• Steady-state heat rejection must be greater than or equal to 11.67 kW per

radiator ORU for flowrates of 0.1591 kg/s (1263 Ibm/hr) per loop with an

entrance temperature Tin of 283.3 K (50.2 °F). The set-point temperature

Tsp is 275.4 K (36 °F). This assumes an equal heat load on each ETCS

loop.

• The temperature for ammonia leaving the ETCS heat-rejection device

should fall between 199.8 K (-100 °F) and 273.7 K (33 °F), inclusive.

• The temperature of ammonia entering the ETCS heat-rejection device falls

between 280.8 K (45.8 °F) and 285.9 K (54.9 °F), inclusive.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

v

!"1 I

41k

E

Figure3.1 International Space Station external thermal control system baseline

configuration.

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• Total mass for each of the six ETCS radiator ORUs on ISS should be no

more than 1,080 kg (2,380 Ibm) per ORU. With fluids, the current radiator

ORU mass is 1,104.9 kg (2,435.9 Ibm) 62.

• Each radiator ORU has 129.8 m 2 (1397.2 ft 2) of radiating surface area for

heat transfer. The radiator mass per radiating area is 8.51 kg/m 2.

• The two ammonia loops have mass flowrates of 1.121 kg/s (8,900 lbJhr)

and 0.958 kg/s (7,600 lbJhr). The corresponding total pressure drops for

each loop are 386,100N/m 2 (561bf/in 2) and 296,500N/m 2 (43 lbf/in2),

respectively. Assuming a fluid density of 635.9 kg/m 3 (which is for

saturated liquid ammonia at 275.4 K or 36 °F), the necessary power

delivered to the fluid is 1.130 kW. The total power set aside for both

pumps, which are identical, is 2.650 kW. Therefore, the overall efficiency

of the pumping system must be at least 42.6%.

• The pressure drop across any ORU radiator panel set is at least

34,470 N/m 2 (5.0 lbf/in 2) and not more than 48,260 N/m 2 (7.01bf/in/).

Assuming a fluid temperature equal to the loop temperature, 275.4 K

(36 °F), and maximum mass flowrate through the radiators, 0.9546 kg/s

(7578 lbm/hr), the mechanical power dissipated within the radiator panel

flow passages is 72.5 W. Thus, assuming a pump efficiency of 0.45,

161.1 W of pumping power must be available for flow through the

radiators in addition to other needs.

The above criteria apply when the heat load is split evenly between Loop A and Loop B.

As of March 30, 1994, Loral Vought Systems, the prime ETCS radiator subcontractor, presented the following design information (Howell, et al., 1994):

Current Radiator Design: Each radiator ORU will use eight radiator panels.

Each radiator panel has 22 tubes in spread-spacing. In spread-spacing, the

effective tube fin area at the panel center is less than for the tubes at the

extremities. Based on analysis, the estimated maximum steady-state heat

rejection is 11.67 kW per ORU for a flowrate of 0.1591 kg/s (1263 lbm/hr)

per path with an entrance temperature Tin of 283.3 K (50.2 °F). The ORU

design has a mass of 1,088.38 kg (2,399.49 Ibm) dry and 1,123.25 kg

(2,476.35 Ibm) with all fluids. The mass breakdown is:

62 See Table 3.1 for details.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Table 3.1 ATCS Radiator ORU Hardware for ISS

Radiator ORU Hardware

Quantity Total Item Mass

per ORU kg Ibm

Base Structure

Scissors Beam and Hinges

Torque Panels and Arms

Cinching Mechanism

Deployment Mechanism

Deployment Motor

Flex Hose and Manifold Set

Radiator Panel Set

Radiator Panel Set

Electrical

Assembly Hardware

Boeing Furnished Equipment

(1) 200.81 442.72

(1) 92.76 204.50

( 1) 47.26 104.20

(6) 23.26 51.27

(1) 34.40 75.85

(1) 11.39 25.12

(8) 246.68 543.84

(7) 358.37 790.07

(1) 50.83 112.06

(1) 9.71 21.40

(1) 8.92 19.66

(1) 3.99 8.80

Total Dry Mass as of 29-Mar-94 1088.38 2399.49

(per ORU radiator)

Total Dry Mass as of 28-Sep-94 1070.0 2359.0

(per ORU radiator) (NASA, 1994)

Fluids - NH3 (estimated) -- 34.86 76.86

Single-phase ammonia (NH3) is the working fluid. This is the baseline

radiator ORU design.

Table 3.2 through Table 3.8 show the overall ISS ITCS and ETCS

hardware masses (NASA, 1994).

Table 3.2 Significant Structural ATCS Hardware per ORU Cluster

Significant Structural ATCS Hardware

Masses per Radiator ORU Cluster

Total Item Mass

kg Ibm

Attachment Hardware 552.9 1219.0

Keel Structure 95.3 210.0

Primary Truss Structure 3286.2 7244.8

Radiator Beam 1210.0 2667.6

Tank and Pump Structures 298.0 657.0

TCS Structural Support 68.0 150.0

Thermal Radiator Rotary Joint:

TRRJ 306.0 674.6

Flex Hose Coupler 158.8 350.0

TRRJ Torque Box 138.3 305.0

Total Structural TCS Hardware

per Radiator ORU Cluster 6113.5 13478.0

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Table 3.3 Total Active ETCS Mass in ISS Segments

Total Active ETCS Hardware Mass

in each Module (excluding structural Total Item Mass

members, and wiring) kg Ibm

Integrated Truss Segment P 1

DDCU Coldplate 51.1 112.6

Fluids (NH3 and N2) 322.0 710.0

Instruments and Sensors 352.2 776.4

Pump Module 332.0 731.9

Radiator Beam Valves 104.2 229.8

Radiator ORUs (3) 3210.1 7077.0

Other 570.5 1257.8

Integrated Truss Segment SO

Coldplates

Plumbing and Instruments

469.9 1036.0

249.7 550.6

Integrated Truss Segment S 1

DDCU Coldplate 51.1 112.6

Fluids (NH3 and N2) 322.0 710.0

Instruments and Sensors 352.2 776.4

Pump Module 332.0 731.9

Radiator Beam Valves 104.2 229.8

Radiator ORUs (3) 3210.1 7077.0

Other 581.9 1282.8

Node 2

Heat Exchangers 209.6 462.0

Heaters 12.5 27.5

U. S. Habitation Module

Heat Exchangers 45.4 100.0

Other 25.9 57.1

U. S. Laboratory

Heat Exchangers 55.5 122.3

Other 4.8 10.5

Total 10968.9 24182.0

Table 3.4 Other Significant ATCS Mass in lSS

Other Significant Total Item Mass

ATCS Hardware Masses kg Ibm

Piping and Plumbing 63 3363.2 7414.6

Ammonia per ETCS Loop 64 191.8 422.9

Science Power Platform Radiator 65 1000.2 2205.0

63 Estimated from Ungar (1995).

64 From Wuestling (1994).

65 Russian hardware.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Table 3.5 Total Active ITCS Mass in ISS Segments

Total Active ITCS Hardware Mass in each Total Item Mass

Module (excluding structural members) kg Ibm

Common US/RSA Airlock 169.4 373.5

Node I 269.0 593. I

Node 2 416.9 919.1

U. S. Habitation Module 849.4 1872.7

U. S. Laboratory 1037.0 2286. I

Total 2741.7 6044.5

Table 3.6 Percentage of ISS ETCS Mass as a Function of Category

Excluding Structural Elements

Category Percentage of ETCS Mass

1. Coldplates 4.1

2. Heat Exchangers 2.5

3. Radiator ORUs 50.8

4. Pump Modules 5.3

5. Instruments and Sensors 6.1

6. Fluids (NH3 and N2) 5.1

7. Plumbing 16.5

8. TRILl Assemblies 9.6

Table 3.7 Percentage of ISS ETCS Mass as a Function of Category

Category Percentage of ETCS Mass

1. Coldplates 2.1

2. Heat Exchangers 1.3

3. Radiator ORUs 25.9

4. Pump Modules 2.7

5. Instruments and Sensors 3.1

6. Fluids (NH3 and N2) 2.6

7. Plumbing 8.4

8. Primary Structure 31.6

9. TRRJ Assemblies 4.9

10. Radiator Beams 9.7

11. Other TCS Structure 2.9

12. Other 4.8

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Table 3.8 Breakdown of Mass for ISS Truss Segment P1

Total Mass for Segment P1 Module Total Item Mass

(including structural members) kg Ibm

Additional Mechanical Equipment

Additional Wiring

Antennas and Cameras

Attachment Hardware

Crew Equip. Trans. Assembly Cart B

Command and Data Handling

DDCU and Connector Box

112.9 248.9

1283.7 2830.1

375.3 827.4

552.9 1219.0

319.8 705.1

289.7 638.6

129.2 284.9

Dedicated ETCS Equipment

DDCU Coldplate 51.1 112.6

Fluids (NH3 and N2) 322.0 710.0

Instruments and Sensors 352.2 776.4

Pump Module 332.0 731.9

Radiator Beam Valves 104.2 229.8

Radiator ORUs (3) 3210.1 7077.0

Other 570.5 1257.8

Handholds and Worksites 253.6 559.1

P1 Keel Structure 95.3 210.0

P1 Primary Truss Structure 3286.2 7244.8

Passive TCS Equipment 129.0 284.3

Radiator Beam 1210.0 2667.6

Remote Power Controller Modules 72.2 159.2

RJMC 70.9 156.2

Support Radiator 4.5 10.0

Tank and Pump Structures 298.0 657.0

TCS Structural Support 68.0 150.0

Thermal Radiator Rotary Joint,

Torque Box, and Coupler 679.9 1498.9

Utility Distribution System

Support Structure 428.2 944.0

Total 14601.4 32190.6

Elements listed as structural or mechanical members are included in the

listings for truss P1 for reference purposes. These, after all, are not

designed to be easily replaced once ISS is actually on orbit.

The photovoltaic power array thermal control system (PV-TCS), though

separate from the central ATCS, is very similar. It is therefore appropriate

to examine mass savings for this system should replacement be necessary.

When fully operational, ISS will derive electricity from four solar power

modules. Each module mounts two external dc-to-dc converter units

(DDCU-E) and a PV-TCS radiator ORU. The PV-TCS radiator ORUs are

smaller versions of the flow-through radiator ORUs used in the ATCS.

This study will briefly examine replacing the PV-TCS radiators and the

DDCU-E interfaces. Table 3.9 presents the relevant baseline masses.

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Table 3.9 Masses of Photovoltaic Power Array Thermal Control System Members

Masses for Some PV-TCS Equipment

Found on Segments $4, $6, P4, and P6)

Quantity

per

Segment

Item Mass

kg Ibm

dc-to-dc Converter Unit-External

PV-TCS Radiator ORU

(2) 63.5 139.9

(1) 618.0 1362.5

3.1.3 Implementing the Reference Mission

If one replaces and augments the ISS TCS with additional flow-through radiator

ORUs, the TCS configuration following year 20 would include eight radiator ORUs.

These replacement radiators would be mounted in two clusters of three ORUs each and

one cluster of two ORUs. In addition to the two radiator ORU clusters on segments S 1

and P1, a hypothetical third truss segment, designated here as P0, would be added to ISS

to support the radiator cluster with two ORUs. Table 3.10 shows the assumed

configuration for P0.

Table 3.10 Breakdown of Mass for Proposed Truss Segment P0

Total Mass for Segment P0 Module

(including structural members) 66

Total Item Mass

kg Ibm

Attachment Hardware

DDCU, Coldplate, and Connector Box

Dedicated ETCS Equipment for

Two Radiator ORUs

P0 Primary Truss & Keel Structure

Passive TCS Equipment

Radiator Beam

Tank and Pump Structures

TCS Structural Support

Thermal Radiator Rotary Joint,

Torque Box, and Coupler

Other

414.7 914.2

180.3 397.5

3260.7 7188.6

2254.3 4970.0

129.0 284.3

806.7 1778.4

198.7 438.0

45.4 100.0

679.9 1498.9

2408.1 5309.0

Total 10377.8 22878.9

66 This construction for segment P0 is based primarily on the values given in Table 3.8 for segment P1.

Some of the component masses for segment P0 are lower to reflect attachment of only two radiator

ORUs instead of three. Specifically, the masses for "Attachment Hardware" and "Other" are threequarters of the corresponding masses listed for segment P1, which are more conservative estimates to

account for equipment which must be installed regardless of the number of ORUs included. The

masses for "Primary Truss & Keel Structure," "Radiator Beam," "Tank and Pump Structures," and

"TCS Structural Support" are two-thirds of the corresponding masses listed for segment P1.

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Table 3.11 Summary of Active ETCS Mass to be Added in Year 20

Total Active ETCS Hardware Mass Added Total Item Mass

in Year 20 (excluding structural members) kg Ibm

Integrated Truss Segment P0

DDCU Coldplate 51.1 112.6

Fluids (NH3 and N2) 214.7 473.3

Instruments and Sensors 234.8 517.6

Pump Module 221.3 487.9

Radiator Beam Valves 69.5 153.2

Radiator ORUs (2) 2140.0 4718.0

Other 380.4 838.5

Total 3311.8 7301.1

This would lead to an overall ETCS equipment mass in year 20, excluding structural

components and piping, of 14,280.7 kg (31,483.7 Ibm) for the reference ISS evolution

mission. (See Table 3.3 and Table 3.11) Overall, this vehicle would circulate 453.3 kg

(999.5 Ibm) of ammonia, or 226.7 kg (499.7 Ibm) per ETCS loop 67. It is presumed that

clusters with four radiator ORUs in the present radiator locations are not feasible for

dynamic or structural reasons. Assuming a pump efficiency of 0.45, the power necessary

just to pump ammonia through the eight radiator ORUs is 215 W. Thus, the total

pumping power will be 2,704 W. The assumed power mass penalty is 476 kg/kW while

using continuous power. Thus, the power as mass is 1,287.1 kg (2,837.6 lbm). The power

mass penalty (Table 3.12) may be broken down as:

Table 3.12 Breakdown of ISS Power Mass Penalty

Mass Penalty

Source [kg/kWsupplied]

Photovoltaic Array (generation) 50

Storage

Batteries 157

Thermal Control System 80

Power Management and Distribution 54

Structure 135

Total for Continuous Power 476

Thus, the overall baseline ETCS mass for ISS evolution is 15,567.8 kg (34,321.3 Ibm).

A major disadvantage of this reference mission is the need to add additional truss

segments to accommodate the additional radiator ORUs. Because these additional trusses

need to be fixed with respect to the SO truss segment, they need to eventually be located in

67 Actually, the total working fluid mass will probably be greater than the values listed here because the

ETCS fluid mass in the lines to and from the two additional radiator ORUs has not been included. If

any additional habitable segments are added in the evolution, the ETCS fluid mass will further

increase.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

board of the solar alpha rotary joints on segments $3 and P3. It is presumed that this

augmentation operation would require major reconstruction of the baseline ISS structure

to properly locate the additional radiator ORUs. This would probably require lengthy and

unacceptable periods during which ISS is not fully available for use.

3.1.4 Parametric Study Using the Baseline Case

Parametric studies are often used to determine the sensitivity of a design to a set of

prescribed variables. For those variables to which the design shows the greatest

sensitivity, the greatest care must be exercised to accurately determine these values. This

study provides bounds on the extent of ISS ETCS heat-rejection variation based on some

of the more common ETCS parameters.

A simple model for the Baseline Case can be developed by considering an overall

heat balance for a single radiator ORU. Basically, the heat lost by the ammonia flowing

through the radiator tubes must be equal to the net emission by the radiator panels. Thus:

dE • + m- l out

d---_= E generated

Assumptions and Restrictions:

The heat transfer for a radiator panel is steady without internal generation.

The difference between the simple average temperature of the panel

surface, Tp, and the simple average temperature of the ammonia, T, is a

known linear function of the sink temperature, Te. Here:

T - Tp = --0.0500 T e + 13.88 K

The base values for this model (Andish, 1995) are:

Te

K

244.3

210.9

T - Tp

°F K oF

-20 1.7 3

-80 3.3 6

The radiator ORU fin efficiency, r I (Te), is a linear function of the sink

temperature, Te. Here:

TI(T e ) = 0.00204 K T e + 0.380

The base values for this model (Farner, 1995, and Ungar, 1995) are:

We

K

240.0

249.8

aq

oF

-27.7 0.87

-10.0 0.89

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

An average specific heat, computed at the simple average temperature of

the ammonia, Ep (T), may be used for the fluid side heat transfer. Here:

D

Ep(T) = 0.008286

kW*s - kW*s

T + 2.424--

kg * K2 kg * K

The parametric study computes the heat rejection for a radiator ORU containing

eight individual panels. Six radiator ORUs are present for the TCS on ISS. The LVS

Base Case presented below uses the values presented by Loral Vought Systems (LVS)

(Howell, et al., 1994). They report a heat rejection of 11.67 kW which the current simple

model reproduces at a sink temperature of 245.4 K (-18.0 °F).

Input study constants:

Stefan-Boltzmann Constant

Radiating Area

Infrared Emissivity 68

Calculation Tolerance

5.670 x 10 -11 kW/(m2*K 4)

129.8 m2

0.90

0.0001

The initial cases examine the LVS Case (design case) and combinations of the minimum

and maximum allowable ammonia mass flowrates and inlet temperatures. These cases

provide bounds on the current abilities of the radiator ORU design.

LVS Case Case AI 1 Case A12

Mass Flowrate of Ammonia [kg/s] 0.1591 0.1591 0.0031

Ammonia Inlet Temperature [K] 283.3 283.3 283.3

Sink Temperature [K] 245.4 249.8 249.8

Fin Efficiency 0.88 0.89 0.89

Average Ammonia Temperature [K] 69 275.5 276.2 253.4

Average Panel Surface Temperature [K] 273.9 274.8 252.1

Average Ammonia Specific Heat [kW*s/kg*K] 4.71 4.71 4.52

Outlet Temperature of Ammonia [K] 267.7 269.1 223.6

Total Heat Rejection per Radiator ORU [kW] 11.68 10.66 0.84

Heat Rejection per Unit Area [kW/m 2] 0.0900 0.0821 6.45x10 3

Case Bll Case B12 Case Cll Case C12

Mass Flowrate of Ammonia [kg/s] 0.1591 0.0031 0.1591 0.0031

Ammonia Inlet Temperature [K] 280.8 280.8 285.9 285.9

Sink Temperature [K] 249.8 249.8 249.8 249.8

Fin Efficiency 0.89 0.89 0.89 0.89

Average Ammonia Temperature [K] 274.3 253.3 278.2 253.6

Average Panel Surface Temperature [K] 272.9 251.9 276.7 252.2

Average Ammonia Specific Heat [kW*s/kg*K] 4.70 4.52 4.73 4.53

Outlet Temperature of Ammonia [K] 267.7 225.7 270.4 221.3

Total Heat Rejection per Radiator ORU [kW] 9.74 0.77 11.64 0.91

Heat Rejection per Unit Area [kW/m 2] 0.0750 5.95x10 3 0.0897 6.98x10 3

68 This value was suggested by Keller (1995 a) and is consistent with other sources for ISS.

69 This is the simple arithmetic average of ammonia inlet and outlet temperatures.

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Case A21 Case A22

Mass Flowrate of Ammonia [kg/s] 0.1591 0.0031

Ammonia Inlet Temperature [K] 283.3 283.3

Sink Temperature [K] 238.7 238.7

Fin Efficiency 0.87 0.87

Average Ammonia Temperature [K] 274.6 244.0

Average Panel Surface Temperature [K] 272.6 242.1

Average Ammonia Specific Heat [kW*s/kg*K] 4.70 4.45

Outlet Temperature of Ammonia [K] 265.8 204.8

Total Heat Rejection per Radiator ORU [kW] 13.08 1.08

Heat Rejection per Unit Area [kW/m 2] 0.101 8.34x10 -3

Case B21 Case B22 Case C21 Case C22

Mass Flowrate of Ammonia [kg/s] 0.1591 0.0031 0.1591 0.0031

Ammonia Inlet Temperature [K] 280.8 280.8 285.9 285.9

Sink Temperature [K] 238.7 238.7 238.7 238.7

Fin Efficiency 0.87 0.87 0.87 0.87

Average Ammonia Temperature [K] 272.6 243.9 276.6 244.3

Average Panel Surface Temperature [K] 270.7 241.9 274.6 242.3

Average Ammonia Specific Heat [kW*s/kg*K] 4.68 4.44 4.72 4.45

Outlet Temperature of Ammonia [K] 264.4 206.9 267.2 202.6

Total Heat Rejection per Radiator ORU [kW] 12.19 1.02 14.02 1.15

Heat Rejection per Unit Area [kW/m 2] 0.0939 7.84x10 -3 0.108 8.85×10 -3

The following studies increased the inlet ammonia temperature and varied the sink

temperature while holding the mass flowrate constant. The purpose here was to observe

how an increased thermal bus temperature affected heat rejection.

Case DI 1 Case D21 Case D31 Case D41

Mass Flowrate of Ammonia [kg/s] 0.1591 0.1591 0.1591 0.1591

Ammonia Inlet Temperature [K] 290.9 290.9 290.9 290.9

Sink Temperature [K] 249.8 238.7 227.6 210.9

Fin Efficiency 0.89 0.87 0.84 0.81

Average Ammonia Temperature [K] 281.9 280.4 279.1 277.7

Average Panel Surface Temperature [K] 280.6 278.4 276.6 274.4

Average Ammonia Specific Heat [kW*s/kg*K] 4.76 4.75 4.74 4.73

Outlet Temperature of Ammonia [K] 273.0 269.9 267.4 264.5

Total Heat Rejection per Radiator ORU [kW] 13.57 15.88 17.75 19.81

Heat Rejection per Unit Area [kW/m 2] 0.105 0.122 0.137 0.153

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Case D 12 Case D22 Case D32

Mass Flowrate of Ammonia [kg/s] 0.0031 0.0031 0.0031

Ammonia Inlet Temperature [K] 290.9 290.9 290.9

Sink Temperature [K] 249.8 238.7 227.6

Fin Efficiency 0.89 0.87 0.84

Average Ammonia Temperature [K] 254.0 244.6 235.6

Average Panel Surface Temperature [K] 252.6 242.7 233.1

Average Ammonia Specific Heat [kW*s/kg*K] 4.53 4.45 4.38

Outlet Temperature of Ammonia [K] 217.0 198.4 180.3

Total Heat Rejection per Radiator ORU [kW] 1.04 1.28 1.51

Heat Rejection per Unit Area [kW/m 2] 7.99x10 3 9.84x10 -3 0.0116

The final studies are similar to the previous set except that the inlet ammonia

temperature was varied until the outlet ammonia temperature reached either the maximum

or minimum set-point temperature [273.7 K (33 °F) or 199.8 K (-100 °F)].

Case E 11 Case F21 Case G31 Case H41

Mass Flowrate of Ammonia [kg/s] 0.1591 0.1591 0.1591 0.1591

Ammonia Inlet Temperature [K] 292.3 298.3 302.8 307.5

Sink Temperature [K] 249.8 238.7 227.6 210.9

Fin Efficiency 0.89 0.87 0.84 0.81

Average Ammonia Temperature [K] 283.0 286.0 288.3 290.6

Average Panel Surface Temperature [K] 281.6 284.1 285.8 287.3

Average Ammonia Specific Heat [kW*s/kg*K] 4.77 4.79 4.81 4.83

Outlet Temperature of Ammonia [K] 273.7 273.7 273.7 273.7

Total Heat Rejection per Radiator ORU [kW] 14.12 18.75 22.30 25.94

Heat Rejection per Unit Area [kW/m 2] 0.109 0.144 0.172 0.200

Case 112 Case J22

Mass Flowrate of Ammonia [kg/s] 0.0031 0.0031

Ammonia Inlet Temperature [K] 310.9 289.2

Sink Temperature [K] 249.8 238.7

Fin Efficiency 0.89 0.87

Average Ammonia Temperature [K] 255.3 244.5

Average Panel Surface Temperature [K] 254.0 242.6

Average Ammonia Specific Heat [kW*s/kg*K] 4.54 4.45

Outlet Temperature of Ammonia [K] 199.8 199.8

Total Heat Rejection per Radiator ORU [kW] 1.56 1.23

Heat Rejection per Unit Area [kW/m 2] 0.0120 9.50x10 3

These results are summarized graphically by Figure 3.2 through Figure 3.7.

59
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

0.014

0.012

0.010

0.OO8

_ 0.(106

0.004

280

o

J Sink Temperature

r-_ - ..o-- - 238.7 K

249.8 K

290 300 310

An-monia Inlet Temperature [K]

320

Figure 3.2 Baseline Case Parameter Study Results: Heat rejected per unit radiator area as

a function of ammonia inlet temperature for a flowrate of 0.0031 kg/s, which is the

minimum radiator flowrate.

0.25

E

e_

0.20

0.15

•_ 0.10

0

.o

_°

oO" 0

t," ,- ,_" Sink Temperature

P

zr -. ¢,- - 210.9 K

... t,... 227.6 K

._ - -.o- - 238.7 K

_ 249.8 K

0.05 L I

280 290 300 310

Anm_nia Inlet Temperature [K]

32O

Figure 3.3 Baseline Case Parameter Study Results: Heat rejected per unit radiator area as

a function of ammonia inlet temperature for a flowrate of 0.1591 kg/s, which is the

maximum radiator flowrate.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

1.6

1.4

0

!

"_ 1.2

_ 1.0

._,

0.8

D

J Sink Temperature

/,/- - -zx- - 238.7 K

---o_ 249.8 K

0.6 i i

280 290 300 310

Arrraonia Inlet Temperature [K]

320

Figure 3.4 Baseline Case Parameter Study Results: Total heat rejected per radiator

orbital replacement unit as a function of ammonia inlet temperature for a flowrate of

0.0031 kg/s, which is the minimum radiator flowrate.

30

O

t_

20

i.

N 15

_ 10

.S

°S

o_ oo°°

S

A" ss

SS _S Sink Temperature

-. o- - 210.9 K

---zx- - - 227.6K

- -o- - 238.7 K

----o--- 249.8 K

5 ! I I

280 290 300 310 320

Arranonia Inlet Temperature [K]

Figure 3.5 Baseline Case Parameter Study Results: Total heat rejected per radiator

orbital replacement unit as a function of ammonia inlet temperature for a flowrate of

0.1591 kg/s, which is the maximum radiator flowrate. The LVS Case has a sink

temperature of 245.4 K and an ammonia inlet temperature of 283.3 K.

61
Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

230

225

E

220

E 215

[..,

= 210

©

=

205

E

<

200

195

Sink Temperature

- -o- - 238.7 K

a, -----z3-----249.8 K

_ck

280 290 300 310 320

Arrrnonia Inlet Temperature [K]

Figure 3.6 Baseline Case Parameter Study Results: Radiator ammonia outlet temperature

as a function of ammonia inlet temperature for a flowrate of 0.0031 kg/s, which is the

minimum radiator flowrate.

274

272

_" 270

268

_ 266

<

264

280

II " I

iI .' I"

J j.... /:'

/ I ," S

[ / /1 ,' I' Sink Temperature

/el / ; t' --_-- 210.9K

¢ /' -..zx- - - 227.6K

I s

i - --o-- - 238.7 K

/ i _ 249.8 K

I 4 °

I

290 300 310

Arrrnonia Inlet Temperature [K]

320

Figure 3.7 Baseline Case Parameter Study Results: Radiator ammonia outlet temperature

as a function of ammonia inlet temperature for a flowrate of 0.1591 kg/s, which is the

maximum radiator flowrate.

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3.1.5 Advanced ATCS Architecture for International Space Station Evolution

Currently, the assumed ISS project life is 30 years from the date construction

begins on orbit. As noted above, the TCS radiator ORUs are assumed to need

replacement in year 20. Because the original architecture and the advanced architectures

which follow have service lives in excess of 10 years, the computed mass savings is for a

single radiator-ORU replacement. Qualitative assessments for these advanced

technologies are presented in Section 2.0.

3.1.5.1 Two-Phase Thermal Control System With Mechanical Pump/Separator

For the specific ISS evolution to a two-phase TCS with MP/S 70, several

assumptions are helpful. Overall, the current single-phase plumbing and lines will not be

removed. The single-phase radiator and heat exchanger ORUs will be replaced with an

equivalent number of two-phase units. Therefore, changes in the ORU mass, fluid mass,

and pumping power need to be considered.

Ungar (1995) presents a study comparing single-phase and two-phase TCS designs

for space stations of differing sizes. His example for a large station is the previous U.S.-

led Space Station which contained two more crew modules than the current ISS design.

However, with the additional power modules in year 20, ISS may add some additional

modules for human activities leaving it to closely resemble Ungar's large station. Ungar

(1995) gives the pumping power for the large station with a two-phase TCS with MP/S 71

as 0.697 kW. In year 20, ISS will use 2.7 kW for its single-phase cascade system. From

Ungar (1995), the projected radiator areas are roughly equivalent for the two-phase TCS

with MP/S design and the corresponding single-phase cascade design for the large station.

(The two-phase TCS with MP/S design uses 9% less radiator area.) Assuming the two

TCSs use the same radiating area, the mass for a comparable two-phase radiator is

available. Howell, et al. (1994), presents the final two-phase radiator ORU design before

the ISS TCS evolved from a two-phase system to a single-phase cascade system. This

two-phase radiator ORU design (designated Revision K) uses the same surface area and

roughly the same tube arrangement as the current ISS single-phase radiator ORUs.

Dry ORU Mass ORU Fluids Pumping Power as Mass

Thermal Control System [kg] Mass [kg] Power [kW] [kg]

Single-Phase Cascade 1070.0 kg 34.9 kg 2.700 1285.2 kg

Two-Phase With Mechanical

Pump/Seperator 1051.0 kg 22.7 kg 0.697 331.8 kg

Total Mass Savings 19.0 kg 12.2 kg 2.00 953.4 kg

For a system of eight radiator ORUs, the total mass savings is 249.6 kg (550.3 Ibm) for

hardware deleted and 953.4 kg (2101.9 Ibm) for the reduction in required pumping power.

70 See Section 2.1.1 for details of a two-phase TCS with MP/S.

71 Ungar (1995) designates this TCS as a rotary fluid management device (RFMD) type two-phase

TCS.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

250kg

2.00kW

953 kg

Overall Mass Savings 1,203 kg

Composite Qualitative Score

3.1.5.2 Low-Power Two-Phase Thermal Control System

0

As in the previous section, several assumptions are necessary to analyze a LP twophase TCS 72 for ISS evolution. Overall, the current single-phase plumbing and lines will

not be removed. The single-phase radiator and heat exchanger ORUs will be replaced

with an equivalent number of two-phase units. Therefore, changes in the ORU mass, fluid

mass, and pumping power need to be considered. These are identical to the assumptions

in the previous section. Thus, with respect to the analysis here, the only quantifiable

difference between a two-phase TCS with MP/S and a LP two-phase TCS is the additional

savings in pumping power associated with the latter option. Ungar (1995) presents a

pumping power value of 355 W for the LP two-phase TCS on a large space station.

Thus:

Dry ORU Mass ORU Fluids Pumping Power as Mass

Thermal Control System [kg] Mass [kg] Power [kW] [kg]

Single-Phase Cascade 1070.0 kg 34.9 kg 2.700 1285.2 kg

Low-Power Two-Phase 1051.0 kg 22.7 kg 0.355 169.0 kg

Total Mass Savings 19.0 kg 12.2 kg 2.35 1116.2 kg

For a system of eight radiator ORUs, the total mass savings is 249.6 kg (550.3 Ibm) for

lighter hardware and 1116.2 kg (2460.8 Ibm) for the reduction in required pumping power.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

250 kg

2.35 kW

1,116 kg

Overall Mass Savings 1,366kg

Composite Qualitative Score -1

3.1.5.3 Capillary Pumped Loops

Based on Section 3.1.3, capillary pumped loops 73 will save 2.7 kW, which is the

estimated ETCS pumping power following ISS evolution. Upon converting this to an

72 See Section 2.1.2 for details of a LP two-phase TCS.

73 See Section 2.1.4 for background on capillary pumped loops.

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equivalent mass, the savings in power is 1,285.2 kg (2,833.4 Ibm). As above, most of the

capillary pumped loop equipment mass is assumed to be similar to the corresponding mass

of the equipment for the single-phase TCS with mechanical pumps. However, the

capillary pumped loop will use the two-phase radiator panels which yield a savings of

249.6 kg (550.3 Ibm) in lighter hardware for a system of eight radiator ORUs. Thus, the

overall savings for this option is 1,534.8 kg (3,383.7 Ibm).

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

250 kg

2.70 kW

1,285 kg

Overall Mass Savings 1,535 kg

Composite Qualitative Score 74 -2

3.1.5.4 Vapor Compression Heat Pump

To incorporate a vapor compression heat pump 75 system for ISS evolution several

assumptions apply. The heat pump is generic and ammonia is the working fluid within the

radiators. The NASA specification used by LVS to size the radiator ORUs, given above

by the LVS Case in the parametric study, is the sizing criterion for the heat pump. Finally,

the design here is assumed to have six radiator ORUs like the current ISS ATCS

configuration.

Cold Source Temperature, Tc (average ETCS temperature)

Temperature Lift, TH - Tc

Hot Source Temperature, Tri (condenser temperature)

Radiator Inlet Temperature, Tin

Environmental Temperature, Tsink

Total Cooling Load, Qc

(which is equivalent to the load from 8 LVS Case ORUs)

275.5 K

30.6 K

306.1 K

302.0 K

245.4 K

93.36 kW

For a Carnot efficient heat pump, the coefficient of performance (COP) may be expressed

as:

Qc Tc

COPcarnot -

W TH-T c

74 Because ISS evolution assumes that an existing vehicle would be retrofit to use a capillary pumped

loop, the assessment for deployment for this mission is "difficult." Further, because this is an orbital

mission, the development cost is "high."

75 See Section 2.2.1 for details of vapor compression heat pumps.

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For a nonideal heat pump, the required work input is:

Wreal =

Qc

rl COPcarnot

From these equations,

Ideal Coefficient of Performance, COPcarnot

Heat Pump Efficiency, I"i 76 (Ewert, 1991)

Necessary Input Power, Wreal

9.00

0.50

20.7 kW

The ETCS radiators are modeled by the model developed for the parametric study.

Radiating Area per radiator ORU

Fin Efficiency

Emissivity

Radiator Mass Flowrate of Ammonia per Radiator ORU

Average Panel Surface Temperature

Heat Rejection per Unit Area

129.8 m 2

0.88

0.90

0.1591 kg/s

288.1 K

0.147 kW/m 2

By the first law of thermodynamics for a heat pump,

QH, real = Wreal + Qc

Total Heat Rejected by the Radiators, QH,real

Necessary Radiator Surface Area

Number of Radiator ORUs Required

114.1 kW

778.0 m 2

6

To check that this configuration was indeed a minimum, a parametric study was

conducted using the model developed above (Figure 3.8). From this study, the mass for

the heat pump power requirements increases more quickly than the mass of the radiator

ORUs decreases. Therefore, the heat pump saves mass overall because the truss segment

P0 may be deleted once a sufficiently high temperature lift is used. In this case that

temperature lift is 30.6 K.

To account for pumping fluid through the heat pump condenser and the heat pump

evaporator, pumping power equal to 150% of that required for the radiators is added.

After deducting the radiator pumping power for the baseline case, this results in a net

pumping power increase of 187.1 W. However, this heat pump will save mass in other

areas because the truss segment P0 may be deleted entirely.

76 Percentage of Carnot coefficient of performance (COP).

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

30000

20000

0 10 20 30 40 50

Temperature Lift [K]

Heat Pump Power _' Pumping Power

Heat Pump # P0 Truss

A Radiator (wet) ---" Total

Figure 3.8 Variation of the ISS ETCS mass as a function of heat pump temperature lift

for the LVS Case.

Net Mass Decrease Due to Vapor Total Item

Compression Heat Pump Installation Mass [kg]

Heat Pump 77 -699.1

Segment P0 78 10,377.8

Mass of Power for Heat Pump, 20.7 kW

Mass of Pump Power Due to Added

Equipment, 0.1871 kW

Total -250.0

77 The heat pump mass is the sum of its component masses. The evaporator and condenser masses are

2.72 kg per kW of heat exchanger capacity (Swanson, Sridhar, and Gottman, 1993). The total of the

heat pump compressor and driving motor masses is taken as 31.83 (Input Work) °476 (Green, 1991).

78 See Table 3.10 for details.

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This heat pump's main disadvantage is an extremely high power requirement 79 but this

option does not require adding additional truss segments for radiator ORUs. If the heat

pump supports could be affixed to existing structures, this would minimize disruptions to

ISS operations.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

9,679 kg

-20.9 kW

-9,929 kg

Overall Mass Savings -250 kg

Composite Qualitative Score 8o 0

3.1.5.5 Solar Vapor Compression Heat Pump

A solar vapor compression heat pump 8_ uses the same equipment as the heat

pump presented in Section 3.1.5.4 except that input power is provided by a dedicated

solar photovoltaic array instead of by the general station resources. Further, a bypass for

the heat pump will allow the ETCS to use the radiators directly when the heat pump is not

in use on the dark side of the orbit. Therefore, a solar heat pump does not require any

electrical storage. Using the assumption that the power management and distribution

system mass can be cut by a third and the power storage mass can be entirely eliminated,

the appropriate power mass penalty from Table 3.12 becomes 221 kg/kW. The pumping

power is computed using a penalty factor of 476 kg/kW because continuous use is

assumed. The net mass decrease for a solar vapor compression heat pump is:

Net Mass Decrease Due to Solar Vapor Total Item

Compression Heat Pump Installation Mass [kg]

Heat Pump 82 -699.1

Segment P0 83 10,377.8

Mass of Power for Heat Pump, 20.7 kW

Mass of Pump Power Due to Added

Equipment, 0.1871 kW

Total 5021.2

79 The power consumption here is equivalent to the entire output of the solar power modules added as

part of the upgrade to ISS in year 20. While additional power modules may be added to offset usage

by the heat pump, and this is the assumed scenario, this comparison hopefully puts some perspective

on the heat pump's power consumption.

80 The deployment for this option is "easy" because this technology does not require additional

supporting trusses. However, the development cost is "high" because this is an orbital mission.

81 See Section 2.2.2 for details on solar vapor compression heat pumps.

82 See the footnotes for Section 3.1.5.4 for the heat pump sizing correlation.

83 See Table 3.10 for details.

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The solar heat pump has the same advantages as the standard heat pump presented in the

previous section. However, the mass of the required supporting power systems for the

solar heat pump is about half of the corresponding power mass for the standard heat

pump. Because the power system accounts for over 90% of the mass added for the vapor

compression heat pump, reducing this mass by over half is highly advantageous.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

9,679 kg

-2O.9 kW

-4,658 kg

Overall Mass Savings 5,021 kg

Composite Qualitative Score 84 +1

3.1.5.6 Arterial Heat Pipe Radiators

For the mission of ISS evolution, the first arterial heat pipe 85 configuration

presented uses 13.11-m (43-ft) units. These units are the longest that will fit in the

payload bay of a Shuttle vehicle. Longer heat pipes during ground testing demonstrated

the best heat rejection per unit system mass (Chambliss and Ewert, 1990). Thus, the

13.11-m heat pipes were originally selected as part of the Space Station baseline ATCS.

Tests on orbit using 15.24-m heat pipes 86 were less than satisfactory. A shorter heat pipe,

6.71-m (22-ft) long, functioned well when tested on orbit (Brown, Ungar, and Comwell,

1992) 87. The second configuration presented employs this shorter heat pipe.

13.11-meter (43-foot) Long Arterial Heat Pipe Panels

After the ISS evolution mission in year 20, the total load rejected by the eight

ETCS flow-through radiator ORUs at a sink temperature of 245.4 K would be 93.39 kW.

Assuming an equivalent load for arterial heat pipes, analysis yields:

Heat Load Rejected 93.36 kW

Average ETCS Ammonia Temperature 275.5 K

Fin Efficiency (Pekrul, et al., 1989) 0.925

Surface Emissivity 0.90

Temperature Drop Between the ETCS

Heat Exchanger and the Arterial

Heat Pipe (Chambliss and Ewert, 1990) 2.4 K

Average Radiator Temperature 273.1 K

Sink Temperature 245.4 K

84 The deployment for this option is "easy" because this technology does not require additional

supporting trusses. However, the development cost is "high" because this is an orbital mission.

85 See Section 2.3.1 for a description of arterial heat pipes.

86 This is the SHARE test.

87 This is the SHARE II test. Besides the length, the heat pipe tested during SHARE II used a different

internal arrangement than the heat pipe tested during SHARE.

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Arterial Heat Pipe ORU Surface Area

Total Heat Pipe Radiating Area

Total Number of Heat Pipe ORUs

7.43 m 2

1,021.55 m 2

137.45

In terms of actual hardware, it is assumed that each ETCS heat exchanger will hold six

arterial heat pipe ORUs. The ETCS heat exchangers are affixed to the radiator beam truss

in a line with the arterial heat pipes extending from one side. The center of the radiator

beam truss attaches to the TRRJ interface.

Rounding the previous estimate upward, the final configuration will use 138 heat

pipe ORUs and 23 ETCS heat exchangers. They will be deployed in two clusters of eight

ETCS heat exchangers plus one cluster of seven ETCS heat exchangers, to remain

consistent with the assumed ISS reference mission. Further, the mass for the ETCS

condenser in the original two-phase system given in Section 2.3.1 will be assumed for the

ETCS heat exchangers. The power for pumping flow through the ETCS heat exchangers

is assumed to equal the power required for pumping fluid through the current ISS flowthrough radiator ORUs. Additionally, to install the 13.11-m arterial heat pipe radiators on

ISS, new radiator beam trusses will be needed. The baseline radiator beam trusses will be

removed and the radiator beam truss for segment P0 88 will be omitted. Thus, for the

13.11-m arterial heat pipes:

Mass Savings Using Item Mass Quantity in Total Mass

13.11-m Arterial Heat Pipes [kg] Year 20 [kg]

Arterial Heat Pipe ORU

ETCS Heat Exchanger and

Interfacial Mechanism

Radiator Beam Truss (per ORU panel)

-40.88 (138) -5641.4

- 137.40 (23) -3160.2

-1.59 (138) -219.4

Total Mass for Arterial

Heat Pipe Radiators -9021.0

ISS Flow-Through Radiator ORU With

Fluids (28-Sep-94) 1104.9 (8) 8838.9

Segment P0 Radiator Beam 806.7 (1) 806.7

Total Baseline Configuration Mass 9645.6

Equipment Mass Savings 624.6

The 13.11-m arterial heat pipe radiators will yield a mass savings of 624.6kg

(1,377.0 Ibm) in hardware compared with the ISS reference mission using flow-through

radiator ORUs.

6. 71-meter (22-foot) Long Arterial Heat Pipe Panels

After the ISS evolution mission in year 20, the total load rejected by the eight

ETCS flow-through radiator ORUs at a sink temperature of 245.4 K would be 93.39 kW.

Assuming an equivalent load for arterial heat pipes, analysis yields:

88 See Table 3.10 for a listing of truss segment P0.

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Heat Load Rejected

Average ETCS Ammonia Temperature

Fin Efficiency 89

Surface Emissivity

Temperature Drop Between the ETCS

Heat Exchanger and the Arterial

Heat Pipe (Chambliss and Ewert, 1990)

Average Radiator Temperature

Sink Temperature

Arterial Heat Pipe ORU Surface Area

Total Heat Pipe Radiating Area

Total Number of Heat Pipe ORUs

93.36 kW

275.5 K

0.925

0.90

2.4 K

273.1 K

245.4 K

6.94 m z

1021.55 m 2

147.20

In terms of actual hardware, it is assumed that each ETCS heat exchanger will hold four

arterial heat pipe ORUs. The ETCS heat exchangers are affixed to the radiator beam truss

in a line with the arterial heat pipes extending from both sides. In fact, each ETCS heat

exchanger has two sockets on each side. The end of the radiator beam truss attaches to

the TRRJ interface.

Rounding the previous estimate upward, the final configuration will use 148 heat

pipe ORUs and 37 ETCS heat exchangers. They will be deployed in two clusters of

13 ETCS heat exchangers plus one cluster of ten ETCS heat exchangers, to remain

consistent with the assumed ISS reference mission. Further, the mass for the ETCS

condenser in the original two-phase system given in Section 2.3.1 will be assumed for the

ETCS heat exchangers. The power for pumping flow through the ETCS heat exchangers

is assumed to equal the power required for pumping fluid through the current ISS flowthrough radiator ORUs. Additionally, to install the 6.71-m arterial heat pipe radiators on

ISS, new radiator beam trusses will be needed. Again, the baseline radiator beam trusses

will be removed and the radiator beam truss for segment P0 will be deleted. Thus:

Mass Savings Using Item Mass Quantity in Total Mass

6.71-m Arterial Heat Pipes [kg] Year 20 [kg]

Arterial Heat Pipe ORU

ETCS Heat Exchanger and

Interfacial Mechanism

Radiator Beam Truss (per ORU panel)

-45.05 (148) -6667.4

-78.92 (37) -2920.0

-3.81 (148) -563.9

Total Mass for Arterial

Heat Pipe Radiators -10151.3

ISS Flow-Through Radiator ORU With

Fluids (28-Sep-94) 1104.9 (8) 8838.9

Segment P0 Radiator Beam 806.7 (1) 806.7

Total Baseline Configuration Mass 9645.6

Equipment Mass Savings -505.7

89 Internally, 6.71-m arterial heat pipes have four extrusions in their evaporators and condensers.

Thus, even though these panels are wider, they have the same fin efficiency as the 13.1 l-m heat

pipes. The fin efficiency for the 13.11-m heat pipes is given by Pekrul, et al. (1989).

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The 6.71-m arterial heat pipe radiators will require 505.7 kg (1,114.9 ibm) more in

hardware compared with the ISS reference mission using flow-through ,:adiator ORUs.

Arterial Heat Pipe Reliability

The arterial heat pipes will have high reliability due to using numerous individually

sealed elements and their ease of replacement. Loss of a single heat pipe ORU will have

negligible effect on ISS's overall heat rejection. Further, any one heat pipe ORU can be

replaced much more easily than the flow-through ORUs. Christiansen (1992) predicts that

micrometeoroids and orbital debris will puncture the flow-through radiator ORUs 1.55

times during the final ten years of the original 30-year ISS project life. His results are

presented below and graphically in Figure 3.9. Here the arterial heat pipe installation

using 6.71-m heat pipes is selected. While the 13.11-m option is less massive, the shorter

heat pipes are already proven on orbit.

Time [Years]

10 20 30

Cumulative Number of Flow-Through Radiator

Perforations Due to On-Orbit Debris 90 0.616 1.6 3.15

It is assumed that any debris puncture to the ISS ETCS will drain the

corresponding flow loop in about two hours (Keller, 1995 a). After the hole is patched, or

the damaged section of the radiator is closed off from the rest of the flow loop, the

appropriate ETCS flow loop can be refilled using new ammonia, 226.7 kg 91, brought on a

resupply flight. Further, this process will be acutely necessary because one loop

represents half of the heat-rejection capability of the U.S.-led portion of ISS. In contrast,

assuming the ETCS heat exchangers for the arterial heat pipes are well shielded such that

an ETCS loop puncture is unlikely, puncturing debris will render only a single heat pipe

unusable. Because a single arterial heat pipe represents only 1/148th of the ATCS

capacity for the U.S.-led portion of ISS, replacing this loss is not immediately critical.

Assuming:

• Any puncture of an ETCS flow-through radiator for the U.S.-led portion of

ISS will be repaired or closed off from the rest of the flow loop. (It is

assumed the radiator ORU itself will not require complete replacement.)

• The ETCS heat exchangers for the arterial heat pipes are well shielded

from impacting space debris.

• Any punctured arterial heat pipe ORU will also not be replaced. (Its loss

will be considered insignificant.)

Then, a mass savings for the arterial heat pipe ORUs compared with the baseline flowthrough radiator ORUs is equal to the mass of ammonia lost using the flow-through panels

over the final ten years of the original life of ISS.

90 From Christiansen (1992). The value for 20 years is estimated using the curve fit to the data found

in Figure 3.9.

91 This value assumes that the ISS ATCS grows to include eight flow-through radiator ORUs. There is

191.8 kg of ammonia per ETCS loop when using six flow-through radiator ORUs (Wuestling, 1994).

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3.5

3.0 E

© 2.5

O

E

2.0

o

_ 1.5

_a 1.0

_D

_0.5

b.

0.0

-----o--- Both ETCS Flow Loops

-- _- Flow Panel Tubes (Only)

0 5 10 15 20 25 30

ISS Project Life [years]

Figure 3.9 Perforations in the ISS radiator ORUs as a function of time. The flow loop

segments within the remainder of ISS are sufficiently well-shielded that punctures due to

on-orbit debris are not expected during the life of the project. The data points are from

Christiansen (1992) while the curves are second-order polynomial fits to the data. The

upper curve represents the number of ETCS loop punctures as a function of time for the

flow-through radiator ORUs. The lower curve gives the number of punctures expected

for just the panel flow tubes, which corresponds to the number of punctures expected in

axial-groove heat pipes.

Mass of Ammonia per ETCS Loop 226.7 kg

Number of Flow-Through Radiator ORU

Punctures Expected in the Final 10 Years × 1.55

Mass of Ammonia Lost in the Final 10 Years

Using Flow-Through Radiators 351.3 kg

Thus, any arterial heat pipe installation will yield a mass savings of 351.3 kg over ten years

by saving ETCS loop ammonia which might be lost to space.

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The 6.71-m arterial heat pipes require 505.7 kg more for hardware while saving

351.3 kg in ETCS working fluid when compared with the baseline ETCS. Overall, this

option is 154 kg heavier than the baseline ETCS.

Specific Assessments:

Equipment Mass Savings

Replacement Mass Savings

Power Savings as Mass

-505 kg

351 kg

neghgible

Overall Mass Savings -154 kg

Composite Qualitative Score +1

3.1.5.7 Axial-Groove Heat Pipe Radiators

The table below presents axial-groove heat pipe 92 and flow-through tube radiator

panels to reject comparable quantities of heat for ISS evolution. Both radiator ORUs

have the same surface area and use all the same equipment, including manifolds and

radiator panels, except for the flow tubes/heat pipes themselves. The first design, by OAO

Corporation (Nguyen, 1982, and Swerdling, 1993), employs 18 heat pipes (outside

diameter of 0.015 m or 0.59 inch) spaced evenly in each radiator panel. The second

design, by Loral Vought Systems (Howell, et al., 1994) uses 22 flow-through tubes in

spread-spacing. The overall fin efficiencies are similar.

Radiator ORU Hardware

Tubes

Mass of Flow Tubes per Fin

/ Heat Pipes [kg] Panel Efficiency

OAO Corporation Heat Pipe

ORU Radiator Panel Set 93

Loral Vought Systems Flow-Through

ORU Radiator Panel Set 94

143.92 18 0.872

105.27 22 0.88

The radiator ORU mass increase using axial-groove heat pipes is 38.65 kg (85.21 lbm).

This translates to an equipment mass increase of 309.2 kg (681.7 Ibm) overall in year 20.

Because the volume devoted to the ammonia circulating within the ETCS will decrease, a

mass savings can be obtained by deleting some ammonia. The necessary power for

pumping ammonia through the radiator manifolds will increase slightly (Nguyen, 1992)

which will roughly offset the mass savings associated with using less ammonia.

A similar mass increase will apply to the PV-TCS radiator ORUs. Assuming the

panel set mass versus the total radiator ORU mass is the same for the ETCS radiator

92 See Section 2.3.3 for details of axial-groove heat pipes.

93 The mass for this option includes both heat pipe and heat exchanger masses (Nguyen, 1992). An

extrusion is assumed to be unnecessary because it could be affixed to the heat pipe during

fabrication.

94 The mass for this option includes the flow tubes, the extrusion used to position the tube within the

radiator panel, and the silver epoxy (Oren, 1995).

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ORUs and the PV-TCS radiator ORUs, an axial-groove heat pipe PV-TCS radiator ORU

will increase by 38.65 kg (85.21 Ibm). This is a mass increase of 154.6 kg (340.8 Ibm)

upon replacing all four PV-TCS radiator ORUs.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

-309 kg

negligible

Overall Mass Savings -309 kg

Composite Qualitative Score +1

PV-TCS Assessments:

PV-TCS Equipment Mass Savings

PV-TCS Power Savings as Mass

-155 kg

negligible

PV-TCS Mass Savings -155 kg

3.1.5.8 Arterial Heat Pipe Radiators With Electrohydrodynamic Pumping

As noted in Section 2.3.2, when the heat transport capacity of an arterial heat pipe

is exceeded, the working fluid collects in the condenser, leaving the unit unusable. To

prevent this, Bryan (1995) proposes applying electrohydrodynamic pumping to each

individual unit. This arrangement would guard against the working fluid collecting in the

condenser and ensure that each unit would be able to continually reject its applied heat

load. In terms of actual hardware, the arterial heat pipes with electrohydrodynamic

pumping will replace the 13.11-m arterial heat pipes from Section 3.1.5.6 without

electrohydrodynamic pumping. Because the heat pipes with electrohydrodynamic

pumping have a lower fin efficiency, the new configuration will be:

13.11-m Arterial Heat Pipes

Without Electrohydrodynamic

Pumping 95

13.11-m Arterial Heat Pipes

With Electrohydrodynamic

Pumping

Heat Pipe Panel Mass [kg] 40.88 32.29

Fin Efficiency 0.925 0.763

Heat Pipe ORU 138 167

ETCS Heat Exchanger and

Interfacial Mechanism 23 28

Thus, this configuration will use 167 heat pipe ORUs and 28 ETCS heat exchangers.

They will be deployed in two clusters of ten ETCS heat exchangers plus one cluster of

eight ETCS heat exchangers. As above, the radiator beam trusses will require

replacement, to be compatible with the arterial heat pipe architecture, and the baseline

segment P0 radiator beam truss can be omitted. The power for pumping flow through the

95 From Section 3.1.5.6.

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ETCS heat exchangers is again assumed to equal the power required for pumping fluid

through the current ISS flow-through radiator ORUs.

Mass Savings Using Arterial Heat Pipes Item Mass Quantity in Total Mass

With Electrohydrodynamic Pumping [kg] Year 20 [kg]

13.11-m Arterial Heat Pipe ORU

ETCS Heat Exchanger and

Interracial Mechanism

Radiator Beam Truss (per ORU panel)

-31.34 (167) -5233.8

- 137.40 (28) -3847.2

- 1.59 (168) -267.1

Equipment Mass for Arterial Heat Pipe

Radiators With Electrohydrodynamic -9348.1

Pumping

Electrohydrodynamic Pumping Power as -0.95 (167) - 158.7

Mass

Total Mass for Arterial Heat Pipe

Radiators With Electrohydrodynamic -9506.8

Pumping

ISS Flow-Through Radiator ORU With

Fluids (28-Sep-94) 1104.9 (8) 8838.9

Segment P0 Radiator Beam Truss 806.7 (1) 806.7

Total Baseline Configuration Mass 9645.6

Equipment Mass Savings 138.8

The medal heat pipe radiators with electrohydrodynamic pumping will yield a mass

savings of 138.8 kg (306.0 Ibm) in hardware compared with the baseline ISS reference

mission. However, as outlined in Section 3.1.5.6, arterial heat pipes will save an

additional 351.3 kg by not losing ETCS loop ammonia over the final ten years of ISS.

Therefore, the total overall mass savings for this option is 490.1 kg (1,080.5 Ibm) 96

Specific Assessments:

Equipment Mass Savings

Replacement Mass Savings

Power Savings

Power Savings as Mass

298 kg

351 kg

-0.334 kW

-159 kg

Overall Mass Savings 490 kg

Composite Qualitative Score 97 (o)

96 Because electrohydrodynamic pumping is an immature technology, the mass estimates are extremely

tentative and may be conservative. In fact, recent measurements by Bryan (1995) indicate that

electrohydrodynamic pumping power consumption per arterial heat pipe ORU may be less than the

2.0 W assumed here.

97 This is a combined technology so no assessment is provided for terrestrial use potential.

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3.1.5.9 Lightweight Radiators

A parametric study here examines the potential savings from using lighter materials

for various portions of the current radiator ORU assembly. Table 3.1 gives a fairly

specific mass breakdown for a radiator ORU as of 29 March 1994. While this design has

changed slightly since then, these values reflect the general mass distribution within an ISS

radiator ORU. Further, to simplify and structure this study, the major subassemblies can

be grouped into four categories.

Percentage of

Category Mass [kg] ORU Mass

Base Structure, Deployment, and Panel Support: 35.5

Base Structure 200.81

Scissors Beam and Hinges 92.76

Torque Panels and Arms 47.26

Cinching Mechanism 23.26

Deployment Mechanism 34.40

Flex Hose and Manifold Set 246.68 22.0

Radiator Panel Set 409.20 36.4

Additional Equipment and Fluids: 6.1

Deployment Motor 11.39

Electrical 9.71

Assembly Hardware 8.92

Fluids 34.86

Other 3.99

Total 1123.24 100.0

This study varies the component radiator ORU masses linearly based on the

original total mass for that category. Here the radiator panel mass is reduced up to 50%

and the flex hose and manifold set masses are reduced up to 25%. This study assumes up

to 20% mass savings for the base structure, deployment and panel support. Finally, no

mass savings for the fluids and additional equipment is presumed 9s.

Lightweight radiators, as presented here, are purely speculative. Two factors

which will heavily influence ATCS component mass is the heat-rejection system design

and the component materials. Here both the design and the component materials are, out

of necessity, vague. The designs are vague because new radiator configurations which are

under development may have significantly different mass requirements than current

radiator technology. Further, lighter materials will yield additional mass savings which are

currently not quantified. Rather, this section attempts to show the overall mass savings

that might be realized if certain component radiator masses can be reduced. Some actual

lightweight radiator concepts for ISS evolution are mentioned below.

The radiator panel set mass is the largest single component within the radiator

ORU. Any technologies or materials which reduce this mass would form the basis for a

lightweight radiator design. Assuming the design optimization would target mass

98 See Section 2.4 for additional general background on lightweight radiators plus specific examples of

proposed lightweight radiators.

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reductions in the radiator panel set, the corresponding flex hoses and manifolds are

assumed to decrease in mass at half the rate of the radiator panels. The flex hoses are

already probably as light as possible for the appropriate component parts. Thus, any

reduction in flex hose mass would require using smaller channel passages. Further, the

constraint of using the lightest and strongest materials for the radiator panel set will

probably dictate that the flex hose and manifold materials can not be completely optimized

in order to maintain compatibility with the radiator panel set materials. (See Howell, et aI.

(1992), for an example of some material interaction problems associated with a radiator

design.) The base structure, deployment, and panel support designs are driven more by

the overall volume of the radiator ORU than by the mass of the other components. Thus,

here it is assumed that the structural component mass can be reduced by only 20% for a

reduction in the panel mass of 50%. Finally, the additional equipment and fluids are fixed

masses. The additional equipment items are "off the shelf', while the fluid mass is a

function of radiator volume and fluid density, both of which are constant. This approach

gives an overall radiator ORU mass reduction of 30.8% when the radiator panel set mass

is reduced by 50% (Figure 3.10).

Percent Reduction in Radiator Panel Mass

Category 10% 30% 50%

Base Structure, Deployment, and Panel Support [kg]

Flex Hose and Manifold Set [kg]

Radiator Panel Set [kg]

Additional Equipment [kg]

382.55 350.67 318.80

234.35 209.68 185.01

368.28 286.44 204.60

68.87 68.87 68.87

Overall Mass per Radiator ORU [kg] 1054.05 915.66 777.27

Overall Mass Reduction per Radiator ORU [kg] 69.19 207.58 345.97

Mass Reduction as a Percentage

of the Original Radiator ORU Mass [%] 6.2 18.5 30.8

Radiator Mass Per Surface Area [kg]m 2] 99 8.12 7.05 5.99

Considering the available lightweight radiators presented in Section 2.4, an overall mass

reduction of 18.5% was selected as a representative value. Thus, the mass savings for

eight radiator ORUs is 1,660.64 kg (3,661.10 Ibm). Because the flow geometry should be

the same, the required pumping power is unchanged.

Two lightweight radiator concepts are appropriate for ISS evolution. The first

would be to substitute composite flow-through radiator panels for the baseline

architecture. This is the case most accurately represented by the analysis above. The

second concept would be to use composite heat pipe radiators.

A similar mass savings may be achieved for the PV-TCS radiator ORUs.

Assuming the ratio of the components within the PV-TCS radiator ORUs is similar to

those in the ETCS radiator ORU, a lightweight PV-TCS radiator ORU will have a mass of

503.7 kg (1,110.4 Ibm). This is a savings of 114.3 kg (252.0 Ibm) per PV-TCS radiator

ORU or 457.2 kg (1,008.0 Ibm) upon replacing all four PV-TCS radiator ORUs.

99 These values are based on a radiating area of 129.8 m 2 per radiator ORU.

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Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

Overall Mass Savings

PV- TCS Assessments:

PV-TCS Equipment Mass Savings

PV-TCS Power Savings as Mass

PV-TCS Mass Savings

1,661 kg

none

1,661 kg

457 kg

none

457 kg

1200-

1000

800-

_0

600

400

2OO

0 0 0 0

0 I I I I

0 10 20 30 40

Percent Reduction in Radiator ORU Panel Mass

Flex Hose & Manifold Set _ Additional Equipment & Fluids

Radiator Panel Set _ Base Structure, Deployment & Panel Support

..L

-.- Total Radiator ORU Mass

50

Figure 3.10 Radiator ORU mass as a function of mass reduction within the radiator panel

set. Study Assumption: For each 10% mass reduction in the radiator panel set, the mass

for the flex hoses and manifolds decreases by 5% and the mass of the base structure,

deployment, and panel support decreases by 4%.

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3.1.5.10 Rotary Fluid Coupler

The current ISS TCS uses a flex hose coupler with a mass of 159 kg (350 lbm) and

an approximate volume of 0.634 m 3 (22.4 ft3) (Harwell, 1992). Flexible hoses within the

flex hose coupler allow ATCS fluid streams to transfer to and from the radiator ORUs

irrespective of the TRRJ's orientation. Because the flexible hoses wrap around a central

hub as the TRRJ traverses, the entire radiator structure may rotate no more than

+_105 degrees before a counter-rotation is required to unwind the flexible hoses.

Additionally, the hoses themselves are subject to fatigue. Because ISS can usually

perform the required counter-rotation while it is in Earth's shadow, the flex hose coupler

has only a small performance penalty. (At some beta angles the rotation limit is reached

while ISS is still in sunlight.)

A rotary fluid coupler 10o is a fully rotating device which uses liquid seals to

contain and separate the internal fluid flow channels (Harwell, 1992). A fully rotating

device allows the radiators to always be aligned with the panel faces parallel to the solar

vector. Because the flexible hoses of the flex hose coupler restrict rotation of the

radiators, there are some orbital positions in which the best allowable radiator ORU

alignment is not optimal. The projected rotary fluid coupler mass is 10.9 kg (24 Ibm) with

an approximate volume of 0.0127 m 3 (0.45 ft3). Therefore, the rotary fluid coupler would

save 444 kg (978 Ibm) and 1.86 m 3 (65.9 ft 3) in three TRRJs when new couplers are

necessary.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

444 kg

negligible

Overall Mass Savings 444 kg

Composite Qualitative Score +3

3.1.5.11 Carbon Brush Heat Exchanger

In order to save mass and improve DDCU cooling efficiency for ISS evolution, the

current radiant fin interfaces could be replaced by carbon brush interfaces. As

Section 2.6.3 implies, a quick one-dimensional heat transfer analysis reveals that this

change will increase the interfacial conductivity by almost two orders of magnitude. This,

in turn, yields more efficient cooling of the DDCUs. What further advantage or savings

this imparts is dependent on the cooling needs of other units on the ETCS loop 101. The

equipment mass savings is summarized below 102.

10o See Section 2.6.1 for details.

101 Assuming the DDCU baseplate will remain at its original temperature, then a more efficient, lower

resistance interface will allow a lower ETCS flowrate past the DDCUs.

102 Values followed by a question mark, ?, are assumed values; the actual masses were not identified.

The values for the DDCUs on S1 and P1, which are known, were assumed here.

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Location Coldplate Mass Fin Mass Total Equipment

(Type of DDCU Coldplate) Quantity (each coldplate) (each coldplate) Mass Savings

SO (DDCU) 4 62.4 kg 9.86 kg? 78.9 kg?

137.6 Ibm 21.73 Ibm? 173.9 Ibm?

S1, P1 (DDCU) 2 51.1 kg 9.86 kg 39.4 kg

112.6 Ibm 21.73 Ibm 86.9 Ibm

$4, $6, P4, P6 8 63.5 kg 4.29 kg 68.6 kg

for PV-TCS (DDCU-E) 139.9 Ibm 9.46 Ibm 151.3 Ibm

l

The total mass savings is twice the fin mass to account for the corresponding fin set on the

coldplate socket on which the DDCU sits. Aluminum 6061-T6 is the assumed coldplate

material with a density of 2,713 kg/m 3 (0.098 Ibm/in3). Further, while it is assumed that

the carbon brush heat exchangers will have negligible mass, the rest of the baseplate and

coldplate mass will remain unchanged.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

118 kg

none directly

Overall Mass Savings 118 kg

Composite Qualitative Score +3

PV-TCS Assessments:

PV-TCS Equipment Mass Savings

PV-TCS Power Savings as Mass

69 kg

none directly

PV-TCS Mass Savings 69 kg

3.1.6 Summary

The various advanced technologies and their estimated benefits are summarized in

the table below for the evolution of ISS. From Section 3.1.3, the mass of the baseline

ETCS, excluding structural components, is 15,567.8kg. Assuming the mass

determinations throughout this study have associated uncertainties on the order of 10%,

an overall TCS with an advanced technology would need to show a savings of at least

1,557 kg to ensure a mass savings. Further, because design and development costs are not

trivial, a mass savings of at least 25%, or 3,892 kg, is desirable. Using these criteria, the

TCSs with advanced technologies proposed for ISS evolution may be divided into five

categories:

• TCSs using advanced technologies requiring a mass penalty greater than

10% of the overall baseline ETCS mass: none.

• TCSs using advanced technologies requiring a mass penalty less than 10%

of the overall baseline ETCS mass: vapor compression heat pump

(continuously operated), arterial heat pipe radiators, and axial-groove heat

pipe radiators.

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• TCSs using advanced technologies with a mass savings less than 10% of

the overall baseline ETCS mass: two-phase TCS with MP/S, LP twophase TCS, arterial heat pipe radiators with electrohydrodynamic pumping,

and capillary pumped loops.

• TCSs using advanced technologies with a mass savings between 10 and

25% of the overall baseline ETCS mass: lightweight radiators.

• TCSs using advanced technologies with a mass savings greater than 25%

of the overall baseline ETCS mass: solar vapor compression heat pump.

The second and third categories include those technologies which show a mass

savings or deficit less than 10% of the baseline ETCS mass. These technologies will

produce an ETCS which is comparable to the baseline system. The technology in the

fourth category is promising but not outstanding for this mission. The technology in the

fifth category shows significant promise for this mission.

For ISS evolution, another significant consideration is the time necessary to install

any new TCS equipment on orbit. One criterion for grouping options is whether they

require installation of an additional truss segment. Of the technologies discussed above,

the continuously operated vapor compression heat pump and the solar vapor compression

heat pump do not require any additional truss segments. While additional power systems,

and their associated supporting trusses, are required for both of these options, it is

presumed that these power systems can be readily mounted on either or both ends of the

U.S. truss outboard of the alpha joints while the heat pumps are located inboard near the

current ATCS heat-rejection equipment.

Several technologies addressed in this section are more correctly identified as

enhancing technologies. Enhancing technologies are advanced technologies which will

uniformly deliver a mass savings or penalty for a specified reference mission regardless of

the type of TCS selected. These technologies include the rotary fluid coupler and the

carbon brush heat exchanger. Thus:

• Enhancing technologies which require a mass penalty: none.

• Enhancing technologies which yield a mass savings: rotary fluid coupler

and carbon brush heat exchanger.

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Table 3.13 Advanced Active Thermal Control System Architecture

for International Space Station Evolution

Summary of Advanced Active Thermal Control

System Architecture for International Space

Station Evolution

ETCS PV-TCS

Overall Mass Qual. Overall Mass Qual.

Savings [kg] Score Savings [kg] Score

3.1.5.1 Two-Phase Thermal Control System

With Mechanical Pump/Separator 1,203

3.1.5.2 Low-Power Two-Phase

Thermal Control System 1,366 -1

3.1.5.3 Capillary Pumped Loops 1,535 -2

3.1.5.4 Vapor Compression Heat Pump -250 0

3.1.5.5 Solar Vapor Compression Heat Pump 5,021 +1

3.1.5.6 Arterial Heat Pipe Radiators 103 -154 + 1

3.1.5.7 Axial-Groove Heat Pipe Radiators -309 + 1

3.1.5.8 Arterial Heat Pipe Radiators

With Electrohydrodynamic Pumping 490 (0)

3.1.5.9 Lightweight Radiators 1,661 --

3.1.5.10 Rotary Fluid Coupler 444 +3

3.1.5.11 Carbon Brush Heat Exchanger 118 +3

-155 +1

457 --

69 +2

103 This value assumes 6.71-m heat pipe units. Assuming 13.11-m heat pipe units instead gives an

overall mass savings of 625 kg and an overall qualitative score of - 1.

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3.2 SPACE TRANSPORTATION SYSTEM UPGRADE

3.2.1 Reference Mission

This study assumes that the current Space Transportation System (STS), or

Shuttle, may be upgraded or refit with new TCS components as part of a program to

extend the life of the fleet. Though less likely, this study could also apply to a new

replacement vehicle with similar capabilities. The assumed power mass penalty is

100 kg/kW. This value assumes 56 kg/kW for power generation 104 plus 44 kg/kW for

power management and distribution. A vehicle (or program) life of 140 missions (seven

flights per year for 20 years) following upgrades is assumed. The mass savings are

considered cumulative for the life of the vehicle or fleet of vehicles.

3.2.2 Baseline Case

Shuttle uses several devices in its ATCS to cover its wide range of operating

environments. Listed here are the major components of the Shuttle ATCS which reject

heat from the vehicle (Figure 3.11 ):

• Ammonia boiler subsystem (ABS): This device was designed for use on

descent below 36,600 m (120,000 ft). In practice, except during a launch

abort, the ABS provides cooling after Shuttle lands until the ground

support personnel complete hookup of the ground support equipment heat

exchanger. Simon (1994) lists the ABS heat-rejection ability as 33.2 kW

(ll3,200Btu/hr). The maximum energy capacity, using available

ammonia, is 63,300 kW*s (60,000 Btu).

• Flash evaporator subsystem (FES): This device provides primary cooling

during ascent while Shuttle is above 42,700 m (140,000 ft), and during

descent while Shuttle is above 30,500 m (100,000 ft). In this high load

mode, the FES provides heat rejection up to 43.4 kW (148,000 Btu/hr)

(Simon, 1994). On orbit, while the payload bay doors are open, the FES

provides supplementary cooling of up to 11.4 kW (39,000 Btu/hr) in its

topping mode. When the payload doors are closed on orbit, the FES

provides primary cooling for Shuttle in its high load mode. The FES

exhaust lines, nozzles, and feedwater lines use up to 1435.5 W for internal

heaters to prevent in-line ice blockage.

• Ground support equipment heat exchanger (GSEHX): This device

provides cooling while Shuttle is on the ground, both before launch and

after ground personnel hook up a portable cooling cart after landing.

Simon (1994) lists the maximum heat-rejection capability for the GSE HX

as 31.4 kW (107,000 Btu/hr). Jaax (1978) specifies the GSE HX volume

as 7.04 x 10 -3 m3 (0.249 ft 3) with adry mass of 6.1 kg (13.4 Ibm).

104 The power generation value corresponds to using three sets of hydrogen and oxygen tanks during six

days on orbit. This translates to a power output of 15 kW.

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Radiator panels: Shuttle uses six or eight radiator panels mounted inside

the payload bay on orbit to reject heat. While the payload doors are open,

this is the primary heat-rejection system. To increase the effective radiator

heat transfer area, the forward two panels on each side may be deployed

away from the payload doors while the aft radiators, plus the kit radiators if

they are installed, are fixed to the payload bay doors. These same panels

are used, after cold-soaking, as a vehicle heat sink while Shuttle is

transiting between altitudes serviced by the other heat-rejection equipment

on descent. Rotter (1987) notes that the radiator heat-rejection capability

is a function of the orbit altitude and attitude. According to Jaax (1978),

the six-panel configuration has a maximum heat rejection of 22.0 kW

(75,000 Btu/hr). Physically, this configuration has a total mass of 573.2 kg

(1,263.81bm), including the necessary Freon21 fluid. When fi_y

deployed, the panels display an effective surface area of 114.3 m 2

(1,229.8 ft2). The mass per surface area is 5.01 kg/m 2. In the eight-panel

configuration, the maximum heat rejection is 28.1 kW (96,000 Btu/hr).

Physically, the total panel mass is 744.7 kg (1,641.8 Ibm) with an effective

radiator surface area of 140.4 m 2 (1,511.6 ft2). The mass per surface area

is 5.30 kg/m 2. Because all working Shuttle flights utilize the eight panel

configuration (Rotter, 1996), this will be assumed standard for this

analysis. The specifications for the radiator panels are (Jaax, 1978):

Overall Panel Measurements:

Length (All): 4.60 m (15.1 ft)

Width (All): 3.20 m (10.5 ft)

Thickness (Forward or Mid-Forward Panels): 0.0229 m (0.900 in.)

(Mid-Aft or Aft Panels):

Facesheet (Aluminum 2024):

Thickness:

Panel Honeycomb Core (Aluminum 5056-H39):

Density:

Panel Manifold Lines (Aluminum 5083):

Outer Diameter:

Wall Thickness:

Panel Flow-Through Tubes (Aluminum 6061-T6):

Forward and Mid-Forward Panels:

Number of Tubes per Panel:

Outside Diameter:

Wall Thickness:

Mid-Aft and Aft Panels:

Number of Tubes per Panel:

Outside Diameter:

Wall Thickness:

0.0127 m (0.500 in.)

279 × 10 -6 m (0.011 in.)

49.7 kg/m 3 (3.1 lbm/ft 3)

0.0222 m (0.875 in.)

889 x 10 -6 m (0.035 in.)

68

0.00475 m (0.187 in.)

508 × 10 -6 m (0.020 in.)

26

0.00599 m (0.236 in.)

508 × 10 -6 m (0.020 in.)

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

13a

13b

11

10

13d

13c

2b 4c

(NH 3) (H20)

2a

4a

!

Figure 3.11 An overall schematic of the Space Transportation System active thermal

control system. The labels refer to: (1) aft coldplates; (2) ammonia boiler subsystem,

(a) boiler, (b) ammonia storage; (3) cabin interchanger; (4) flash evaporator subsystem,

(a) high load unit, (b) topping unit, (c)water storage; (5) flow proportioning module;

(6) fuel cell heat exchanger; (7) ground support equipment heat exchanger; (8) hydraulics

heat exchanger; (9) midbody coldplates; (10) oxygen restrictor; (11) payload heat

exchanger; (12) pump package (Freon 21); (13) flow-through radiators, (a) port radiator

panels, (b) port flow control assembly, (c) starboard radiator panels, (d) starboard flow

control assembly. Only one Freon 21 fluid loop, Loop 1, is shown. The second loop is

like the first except that it serves the starboard radiators and not the port radiators.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

The breakdown of area and mass within the radiator panels is (Jaax, 1978):

Table 3.14 Effective Area Breakdown of Shuttle Radiator Panels

Shuttle Side Forward Panels Mid-Forward Mid-Aft Panels Aft Panels

(Servicing Freon Loop) [m z] Panels [m 2] [m2] [m 2]

Left (Loop 1) 21.93 22.58 12.89 13.15

Right (Loop 2) 21.73 22.32 12.79 13.03

Average Effective Area 1o5 21.83 22.45 12.84 13.09

Table 3.15 Mass Breakdown of Shuttle Radiator Panels

Forward Panel Mid-Forward Mid-Aft Panel Aft Panel

Item [kg] Panel [kg] [kg] [kg]

Facesheets 22.8 22.8 22.8 22.8

Honeycomb 15.5 15.5 8.3 8.3

Manifolds 2.0 2.0 2.0 2.0

Flow Tubes 11.4 11.4 5.6 5.6

Other Items 40.2 43.6 38.8 40.2

Fluids (Freon 21) 8.5 7.7 5.7 6.8

Total Mass per Panel 100.4 103.0 83.2 85.7

Aluminum Densities:

2024:

5083 (Assumed):

6061-T6:

Freon 21 pump package:

2770 kg/m 3 (0.100 Ibm/in 3)

2660 kg/m 3 (0.096 Ibm/in 3)

2710 kg/m 3 (0.098 lbm/in 3)

Shuttle has four pumps housed in two pump

modules to circulate Freon 21 within the vehicle. The second pump in each

module is a spare and is normally not used. Each package has a mass of

20.3 kg (44.7 Ibm) and a volume of 0.180 m 3 (6.34 ft3). The Freon 21

loops, excluding the accumulator, have a total volume of 0.159 m 3

(5.60 ft3). The loop volumes are not identical with Loop 2 accounting for

51.8% of the total system volume. From the pump curves in Mistrot

(1994), the current pumps have the following characteristics:

Output Mass Flowrate 106 Pressure Rise Across Pump Required Input Power

[kg/s] [kN/mEl [kW]

0.283 501 0.340

0.321 469 0.360

0.340 455 0.370

0.378 416 0.389

105 The total effective area, or sum, for all radiator panels is 140.42 m 2. The effective area here

accounts for the surface available to exchange heat with the environment (space). This does not

include a fin efficiency, which is currently unknown, so a value of 1.0 is assumed.

106 These flowrates include the minimum and maximum loop flowrates.

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The Shuttle ATCS component breakdowns are presented in Table 3.16 and Table 3.17.

Table 3.16 Heat-Rejection ETCS Hardware for Shuttle

28.6 63.0

45.8 100.9

4.5 10.0

299.4 660

90.7 200

Qty Maximum Heatper Total Item Mass Rejection Capability

STS ETCS Hardware STS kg Ibm kW kBtu/hr

Ammonia Boiler: (1) 33.2 113.2

Equipment 27.6 60.8

Ammonia (NH3) 44.3 97.6

Vent Line 107 ........

Flash Evaporator: (1)

FES

Ducting & Nozzles

H20 Accumulators

Feedwater Supply:

Max. Capacity

Min. Requirement

High Load Mode 43.4 148.0

Topping Only Mode 11.4 39.0

How Cont. Assembly (2)

Dry FCA 22.1 48.8

FCA Fluid 1.4 3.0

Ground Support Equipment

Heat Exchanger (1) 6.1 13.4 31.4 107.0

Radiator Panels 108.

Plumbing 8.3 18.2

Huid in Plumbing 2.0 4.5

Dry Panels (6) 529.5 1167.4

Panel Huid 43.7 96.4

Dry Panels + Kit (8) 687.3 1515.2

Panel + Kit Huid 57.4 126.6

22.0 75.0

28.1 96.0

Total for ETCS Heat

Rejection Hardware:

Using 6 Rad Panels 672.5 1482.5

Using 8 Rad Panels 830.3 1830.3

Other Fluids lO9 44.3 + 97.6 +

90.7 200.0

Max Heat Rejection:

Ascent or Descent 43.4 148.0

On Orbit (8 Panels) 39.5 135.0

On Ground - ABS 33.2 113.2

On Ground - GSE 31.4 107.0

107 The references consulted did not give a value for the vent line, although this mass is not

insignificant.

108 The heat rejection values here assume the best spacecraft attitude for radiant heat rejection

(Jaax, 1978).

109 This total is expressed as ABS ammonia plus, "+", FES feedwater. The total system Freon 21 is

included in Table 3.17 below.

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Table 3.17 Other ETCS Hardware for Shuttle

Maximum HeatQty Rejection

per Total Item Mass Capability 110

STS ETCS Hardware STS kg lbm kW kBtu/hr

Aft Coldplates (7) 44.0 96.9

Max. Heat Load - 2.9 - 10.0

On-Orbit Heat Load - 1.8 - 6.3

O2 Restrictors (2) 1.4 3.0

Interchanger (1) 14.5 32.0 - 14.2 - 48.3

Payload Bay Heat

Exchanger:

Max. Heat Load

Bay Doors Closed

Bay Doors Open:

w/6 Rad Panels

w/8 Rad Panels

(1) 14.9 32.8

- 1.5 - 5.2

- 6.3 - 21.5

- 8.5 - 29.0

Fuel Cell Heat Exchangers (2) 14.6 32.2 - 13.2 - 44.9

Midbody Coldplates (2) 28.8 63.6 - 1.4 - 4.7

Hydraulics Heat Exchanger (1) 10.6 23.4 4.4 15.0

Flow Distribution:

Flow Proportioning

Module (1) 1.7 3.7

Pump Package (2) 40.6 89.5

Plumbing 111 ........

Total Fluids

(Freon 21 only) 112 229.5 505.9

Total for Other ETCS

Hardware:

Equipment (Dry) 171.1 377.1

Fluids 229.5 505.9

To be consistent with the definitions used for ISS, all of the equipment

listed in these tables are ETCS while the label ITCS is reserved to describe

the water loop in the cabin ATCS.

Based on Table 3.16 and Table 3.17 the baseline ETCS mass for STS upgrade

includes 1365.9kg for equipment and 364.5 kg for fluids. An additonal 72.0kg

represents the mass of the ETCS power systems assuming the nominal pumping power

consumption is 0.720 kW. Thus, the overall baseline ETCS mass is 1437.9 kg per vehicle,

or 201,306 kg for 140 flights.

110 A negative heat transfer capability denotes hardware which contributes heat to the TCS instead of

rejecting heat.

111 The plumbing mass is significant for a vehicle the size of STS, but a value of the plumbing mass is

not given in the references consulted.

112 Based on a total fluid volume for Freon 21 of 0.167 m3 (5.9 fi3) at 294.3 K (70°F) and an orbiter

configuration using 8 radiator panels. This is the total Freon 21 circulating within the Freon loops.

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3.2.3 Parametric Study Using the Baseline Case

Like the study presented in Section 3.1.4 for ISS, a simple model for Shuttle

radiator heat rejection can be constructed based on an overall energy balance for the

radiators. Actual STS flight data is used for an assumed baseline case. For this study, the

chosen mission is STS-41 (1990). To get an accurate estimate of the heat rejected from

the radiators, a time index was selected on orbit when the radiator panel exit temperatures

were above the Freon loop set-point of 276.5±1.1 K (38-*-2 °F). This condition implies

that the entire Freon loop flow passes through the radiator panels and none uses the

radiator bypass line. Such a situation is recorded in the STS-41 data at 14 hours after

liftoff.

Assumptions and Restrictions:

• All of the Freon in both loops passes through its corresponding radiator

panel set.

• The effective radiator surface area for either loop is haft of the total

radiator surface area, 70.2 m 2. This assumes the forward two panels on

either side are in the deployed position.

• The difference between the simple average Freon 21 loop temperature

within the panels and the average radiator surface temperature is 1.5 K.

• The average specific heat, computed from the known properties for

Freon 21 (Mistrot, 1994), may be used for the fluid side heat transfer.

Here:

_p (q') = I0.9335 + 0.001231 exp(0.01555 1 T)I _gkW_'*_

This parametric study computes the heat rejection for both sets of Shuttle

radiators. The data for the STS-41 Case (Case A) come directly from that flight. At 14

hours after lift off, Discovery's radiators were rejecting 15.84 kW.

Input study constants:

Stefan-Boltzmann Constant

Radiator Area per Freon Loop

Infrared Emissivity 113

Calculation Tolerance

5.670 x 10 -ll kW/(mE*K 4)

70.2 m 2

0.76

0.00001

The initial case examines the STS-41 Case (Case A). The following series of cases

present combinations of the minimum and maximum allowable Freon 21 mass flowrates

and inlet temperatures. These cases provide bounds on the current abilities of the Shuttle

radiator design.

113 Jaax (1978).

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case A

Loop 1/Loop 2

0.3340/0.3513

301.1 / 300.4

258.9 / 252.5

290.4 / 289.0

288.9 / 287.5

1.05 / 1.04

279.7 / 277.6

7.48 / 8.37

15.85

0.113

Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case BIA

Loop 1/Loop 2

0.2835/0.2835

283.2 / 283.2

199.8 / 199.8

265.8 / 265.8

264.3 / 264.3

1.01 / 1.01

248.5 / 248.5

9.95 / 9.95

19.90

0.142

Case B1B

Loop 1/Loop 2

0.2835/0.3780

283.2 / 283.2

199.8 / 199.8

265.8 / 269.2

264.3 / 267.7

1.01 / 1.01

248.5 / 255.3

9.95 / 10.72

20.67

0.147

Case B1C

Loop 1/Loop 2

0.3780/0.3780

283.2 / 283.2

199.8 / 199.8

269.2 / 269.2

267.7 / 267.7

1.01 / 1.01

255.3 / 255.3

10.72 / 10.72

21.44

0.153

Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case C1A

Loop 1/Loop 2

0.2835/0.2835

322.0 / 322.0

199.8 / 199.8

293.0 / 293.0

291.8 / 291.8

1.05 / 1.05

264.6 / 264.6

17.11 / 17.11

34.22

0.244

Case C1B

Loop 1/Loop 2

0.2835/0.3780

322.0 / 322.0

199.8 / 199.8

293.0 / 298.6

291.8 / 297.1

1.05 / 1.06

264.6 / 275.3

17.11 / 18.76

35.87

0.255

Case C1C

Loop I/Loop 2

0.3780/0.3780

322.0 / 322.0

199.8 / 199.8

298.6 / 298.6

297.1 / 297.1

1.06 / 1.06

275.3 / 275.3

18.76 / 18.76

37.51

0.267

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case B2A

Loop 1/I.xmp 2

0.2835/0.2835

283.2 / 283.2

227.6 / 227.6

270.0 / 270.0

268.5 / 268.5

1.02 / 1.02

256.8 / 256.8

7.60 / 7.60

15.21

0.108

Case B2B

Loop l/Loop 2

0.2835/0.3780

283.2 / 283.2

227.6 / 227.6

270.0 / 272.5

268.5 / 271.0

1.02 / 1.02

256.8 / 261.9

7.60 / 8.21

15.81

0.113

Case B2C

Loop 1/Loop 2

0.3780/0.3780

283.2 / 283.2

227.6 / 227.6

272.5 / 272.5

271.0 / 271.0

1.02 / 1.02

261.9 / 261.9

8.21 / 8.21

16.42

0.117

Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case C2A

Loop 1/Loop 2

0.2835/0.2835

322.0 / 322.0

227.6 / 227.6

297.1 / 297.1

295.6 / 295.6

1.06 / 1.06

272.1 / 272.1

14.97 / 14.97

29.93

0.213

Case C2B

Loop 1/Loop 2

0.2835/0.3780

322.0 / 322.0

227.6 / 227.6

297.1 / 301.6

295.6 / 300.1

1.06 / 1.07

272.1 / 281.3

14.97 / 16.43

31.40

0.224

Case C2C

Loop 1/Loop 2

0.3780/0.3780

322.0 / 322.0

227.6 / 227.6

301.6 / 301.6

300.1 / 300.1

1.07 / 1.07

281.3 / 281.3

16.43 / 16.43

32.86

0.234

Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case B3A

Ix)op 1/l_x_p 2

0.2835/0.2835

283.2 / 283.2

249.8 / 249.8

274.5 / 274.5

273.0 / 273.0

1.02 / 1.02

265.8 / 265.8

5.03 / 5.03

10.06

0.0716

Case B3B

Loop 1/Ix_p 2

0.2835/0.3780

283.2 / 283.2

249.8 / 249.8

274.5 / 276.2

273.0 / 274.7

1.02 / 1.02

265.8 / 269.1

5.03 / 5.44

10.47

0.0746

Case B3C

Loop I/Loop 2

0.3780/0.3780

283.2 / 283.2

249.8 / 249.8

276.2 / 276.2

274.7 / 274.7

1.02 / 1.02

269.1 / 269.1

5.44 / 5.44

10.88

0.0775

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Mass Flowrate of Freon 21 [kg/s]

Freon 21 Inlet Temperature [K]

Sink Temperature [K]

Average Freon 21 Temperature [K]

Average Panel Temperature [K]

Average Freon 21 Specific Heat

[kW*s/kg*K]

Outlet Temperature of Freon 21 [K]

Loop Heat Rejection [kW]

Total Heat Rejection

per Radiator Panel Set [kW]

Heat Rejection per Area [kW/m 2]

Case C3A Case C3B Case C3C

Loop 1/l__op 2 Loop 1/I.x_p 2 Loop 1/Loop 2

0.2835/0.2835 0.2835/0.3780 0.3780/0.3780

322.0 / 322.0 322.0 / 322.0 322.0 / 322.0

249.8 / 249.8 249.8 / 249.8 249.8 / 249.8

301.2 / 301.2 301.2 / 304.9 304.9 / 304.9

299.7 / 299.7 299.7 / 303.4 303.4 / 303.4

1.07 / 1.07 1.07 / 1.07 1.07 / 1.07

280.3 / 280.3 280.3 / 287.9 287.9 / 287.9

12.61 / 12.61 12.61 / 13.87 13.87 / 13.87

25.22 26.47 27.73

0.180 0.189 0.198

3.2.4 Advanced ATCS Architecture for Space Transportation System Upgrade

The upgraded STS vehicles are projected to serve for 20 additional years with an

average of seven flights per year, or 140 missions overall. It is expected that the advanced

architectures outlined below will, once installed, last for the life of the upgraded vehicle.

However, regular vehicle ground maintenance should allow any failing components to be

identified and replaced while the vehicle is not in use. As with the section on ISS above,

each advanced architecture for Shuttle is assessed numerically for overall mass savings

(for 140 missions). Qualitative assessments for these advanced technologies are presented

in Section 2.0.

3.2.4.1 Low-Power Two-Phase Thermal Control System

To estimate the mass of a LP two-phase TCS 114 for STS upgrade this study

extrapolates predictions from Ungar (1995) which gives estimates for space stations.

Ungar (1995) includes estimates for three different vehicles using four different TCS

architectures. Of the options presented, Shuttle most closely resembles a small system

with a single-phase cascade TCS. Ungar also presents a LP two-phase TCS. The LP

two-phage TCS requires tess pumping power and smaller diameter flow lines than the

single-phase cascade TCS. The required radiator area is roughly the same for the two

systems. Thus, neglecting the flow line differences, the systems may be summarized by:

Pump Power [kW]

Radiator Area [m 2]

Loop Set-Point [K]

Baseline

Shuttle Thermal

Control System

Loop 1 Loop 2

0.360 0.360

70.2 70.2

276.5 276.5

Small System (Ungar, 1995)

Single-Phase Cascade

Thermal Control System

LTL LTL

0.320 0.320

197 197

275.2 275.2

Low-Power Two-Phase

Thermal Control System

LTL MTL

0.068 0.080

195 195

275.2 287.2

LTL refers to the low temperature loop while MTL refers to the moderate temperature

loop. Shuttle uses a single-phase TCS which is similar to the single-phase cascade TCS

114 See Section 2.1.2 for details.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

given by Ungar. Comparing the Shuttle TCS with the simple single-phase TCS for a small

system reveals that the pumping power per loop is almost the same and the Shuttle TCS

loop set-points are approximately that of the LTL for the single-phase cascade TCS. A

comparable LP two-phase TCS for Shuttle then would use two LTL loops. The pumping

_ower requirement for the revised Shuttle LP two-phase TCS then is 136 W. Thus:

Current Shuttle Low-Power TwoThermal Control Phase Shuttle Thermal Change Total Change

System Control System (for 1 Mission) (140 Missions)

Pump Power [kW] 0.720 0.136 -0.584

Mass Due to

Pump Power [kg] 72.0 13.6 -58.4 -8176.0

Because the overall vehicle mass decreases only slightly and approximations for these

computations contain significant uncertainty, the two systems effectively have equivalent

masses.

Specific Assessments (for 140 missions):

Equipment Mass Savings

Power Savings

Power Savings as Mass

negligible

81.76 kW

8,176 kg

Overall Mass Savings 8,176 kg

Composite Qualitative Score -1

3.2.4.2 Lightweight Radiators

A parametric study here examines the potential savings from using lighter materials

for various portions of the current Shuttle radiator assembly. Table 3.15 gives a fairly

specific mass breakdown for the Shuttle radiators. To simplify this analysis, the major

subassemblies within the Shuttle radiators can be grouped into four categories.

Mass per Side Percentage of

Category of an STS [kg] Radiator Mass

Facesheet and Honeycomb: 37.3

Facesheet 91.2

Honeycomb 47.6

Flow Tubes and Manifolds: 11.3

Flow Tubes 34.0

Manifolds 8.0

Other Items 162.8 43.7

Fluids 28.7 7.7

Total 372.3 100.0

This study varies the component Shuttle radiator masses linearly based on the

original total mass for that category. Here the facesheet and honeycomb masses are

reduced up to 60% and the flow tube and manifold set masses are reduced up to 24%.

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This study assumes up to 18% mass savings for other items, including the panel support

and deployment mechanisms. Finally, no mass savings for the fluids is expected 115

Lightweight radiators, as presented in this report, are purely speculative. Two

factors which will heavily influence ATCS component mass are the heat-rejection system

design and the component materials. Here both the design and the component materials

are, out of necessity, vague. The designs are vague because new radiator configurations

currently under development may have significantly different mass requirements than

current radiator technology. Further, lighter materials will yield additional mass savings

which are currently not quantified. Rather, this section attempts to show the overall mass

savings that might be realized if certain component radiator masses can be reduced.

As with the radiator ORUs for ISS, advanced materials are most likely to

significantly reduce the mass of the panel facesheets and honeycomb. Some mass

reduction can also be expected for the flow tubes and manifolds, but the dimensions, and

therefore mass, of these components are primarily determined by their function of

containing the fluid within a specified volume. The category "Other Items" includes the

deployment and support structures which are sized based upon the overall volume of the

radiators and less on their absolute mass. Still, lighter radiators should allow some mass

reductions in these supporting structures. Finally, the fluid mass is a function of radiator

volume and fluid density, both of which are constant here. Overall, this study gives a

radiator mass reduction of 33.0% when the facesheet and honeycomb masses are reduced

by 60% (Figure 3.12).

Percent Reduction

in Facesheet and Honeycomb Mass

Category 20% 40% 60%

Facesheet and Honeycomb [kg] 111.0 83.3 55.5

Flow Tubes and Manifolds [kg] 38.6 35.3 31.9

Other Items [kg] 153.0 143.3 133.5

Fluids [kg] 28.7 28.7 28.7

Overall STS Radiator Mass per Vehicle Side [kg] 331.4 290.5 249.6

Overall Mass Reduction per Vehicle Side [kg] 40.9 81.8 122.7

Mass Reduction as a Percentage

of the Original Shuttle Radiator Mass [%] 11.0 22.0 33.0

Radiator Mass Per Surface Area [kg/m 2] 116 4.72 4.14 3.56

Considering the available lightweight radiators presented in Section 2.4, an overall mass

reduction of 22.0% was selected as a representative value. Thus, the mass savings for a

complete Shuttle vehicle is 163.6 kg. For the life of the vehicle, which is 140 missions,

this is a savings of 22,904.0 kg. Because the flow geometry should be the same, the

required pumping power is unchanged. For actual equipment, the composite flow-through

radiators are the only concept in the current study which are appropriate for STS upgrade.

115 See Section 2.4 for additional general background on lightweight radiators plus specific examples of

proposed lightweight radiators.

116 These values assume a radiating area of 70.2 m 2 as found in the deployed 8 panel configuration.

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Specific Assessments (for 140 missions):

Equipment Mass Savings

Power Savings as Mass

Overall Mass Savings

22,904 kg

none

22,904 kg

400 -

_0

300

200-

100

0

13 13 8 o

I I I I I

0 10 20 30 40 50 60

Percent Reduction in Facesheet & Honeycomb Mass

A Other Items _ Facesheets & Honeycomb

Fluids _ Flow Tubes & Manifolds

Total Radiator Mass per Vehicle Side

Figure 3.12 Shuttle radiator mass as a function of mass reduction within the radiator

facing and honeycomb. The masses here are for one of two equivalent sets of radiator

panels on a Shuttle vehicle. Study Assumption: For each 10% mass reduction in the

facesheets and honeycomb the mass for the flow tubes and manifolds decreases by 4% and

the mass of the other items decreases by 3%.

3.2.4.3 Phase-Change Thermal Storage

At least two scenarios are possible to size a phase-change thermal storage device

for Shuttle. One case would presume to replace the ABS with phase-change thermal

storage. This would require a system sized for 63,300 kW*s which is the current design

capacity for the ABS. Preliminary calculations indicate that a corresponding phase-change

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thermal storage system sized to replace the ABS would require 175.8 kg of PCM

alone 117. The mass of a fully charged ABS, however, is only 71.9 kg. Thus, to replace

the ABS with phase-change thermal storage would lead to a substantial increase in vehicle

launch mass.

A second case assumes that phase-change thermal storage is used to provide

auxiliary cooling in conjunction with the radiators and FES. In particular, phase-change

thermal storage could be used to reduce FES usage, thereby reducing the mass of cooling

water rejected by the vehicle 118. The PCM packages are solidified while the radiator

rejects heat above the average orbital value and the PCM is allowed to melt to provide

additional cooling while the radiator is rejecting heat below the average orbital value.

Because the radiator heat load varies around an orbit, even when Shuttle is entirely in full

sun at high beta angles, this approach is generally applicable. Figure 2.8 schematically

illustrates an implementation of phase-change thermal storage for Shuttle 119

Shuttle is an extremely versatile craft with a correspondingly complex heatrejection profile. However, any mission will fly somewhere around planet Earth. The

most advantageous heat-rejection environment would be at a beta angle of 50 degrees,

while the most disadvantageous would be an orbit which is continually in full sun. For this

second extreme an orbit at a beta angle of 75 degrees was selected. A constant altitude of

407,400 m (220 nautical miles) and an attitude of payload-bay-to-Earth complete the

description of the vehicle thermal environment for this analysis. The radiator inlet

temperature was held constant at 308.2 K (95.0 °F) and the working fluid mass flowrate

was maintained at a value of 0.352 kg/s (2795 lbm/hr) per loop. For these parameters,

numerical analysis 120yields:

Average Orbital Radiator

Heat Rejection Rate [kW]

Energy Rejected Above the

Average Rate / Orbit [kW*s]

Equivalent Mass of

FES Water / Orbit [kg] 121

Vehicle Orbital Beta Angle [degrees]

I

0 50 I 75

20.52

3117.9

1.3963

18.91 16.22

3859.3 1204.1

1.7283 0.5392

117 The assumed PCM is water with a thermal density of 10.0 kg/kW*h.

118 As designed, Shuttle uses water, generated by producing electricity in its fuel cells, as feedwater for

the FES. Early in the next century, however, water generated by Shuttle's fuel cells may be tapped

as a source of potable water to be delivered to Space Station or for the crew on extended duration

Shuttle missions. Use of phase-change thermal storage would save some water which currently is

consumed by the FES.

119 See Section 2.5.1 for more information on phase-change thermal storage.

120 The analysis here used TSS with SINDA/FLUINT to compute radiator heat rejection profiles for the

310 node LTV model of Shuttle. After determining the average orbital radiator heat rejection, the

energy above and below the average heat rejection were determined. The approach includes thermal

capacitance for the radiator components.

12t Based on assuming a heat of vaporization for water within the FES of 2,233 kW*s/kg (960 Btu/lbm)

(Lucas, 1996).

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For this study, two materials were considered. Water has a thermal density 122 of

18.0 kg/kW*h including packaging. A long-chain alkane, or wax, such as n-dodecane has

a thermal density of 30.2 kg/kW*h including packaging.

Phase-Change Thermal Storage Devices Phase-Chan ge Materials

Time Basis: 1 orbit Water n-Dodecane

Thermal Density [kg/kW*h] 10.0 16.8

Thermal Density + Packaging [kg/kW*s] 123 0.00500 0.00840

Beta Angle = 0 degrees:

Design PCM Load [kW*s] 3117.9 3117.9

Mass of PCM Device [kg] 15.59 26.19

Beta Angle = 50 degrees:

Design PCM Load [kW*s] 3859.3 3859.3

Mass of PCM Device [kg] 19.30 32.42

Beta Angle = 75 degrees:

Design PCM Load [kW*s] 1204.1 1204.1

Mass of PCM Device [kg] 6.02 10.11

From the definition of the reference mission, the assumed mission length is 6 days.

Assuming an average orbital period of 90 minutes, the standard mission is 96 orbits. The

mass savings using n-dodecane as the PCM is:

Mass Savings Using Water for Vehicle Orbital Beta Angle [degrees]

PCM Device per Mission 0 50 75

Mass of PCM Device [kg] 26.19 32.42 10.11

Mass of FES Water Saved [kg] 134.04 165.92 51.77

Overall Mass Savings [kg] 107.85 133.50 41.66

The worst-case scenario for this technology would be to fly a mission in full sun at

a beta angle of 75 degrees using a vehicle with a PCM device sized for a beta angle of

50 degrees. For this case the FES water savings would be only 51.77 kg while the PCM

device penalty would be 26.19 kg, yielding a mass savings per mission of 25.58 kg, or

3,581.2 kg for the life of the vehicle. In general, Shuttle will fly at a variety of beta angles

from 0 up to 75 degrees or even higher in some cases. Assuming this reflects the

spectrum of Shuttle operating conditions, then the average amount of FES water saved

per mission would be 117.24 kg. Further, assuming the PCM device is sized for a beta

angle of 50 degrees, the average mass savings per mission is 84.82 kg, or 11,875.3 kg for

the life of the vehicle.

Specific Assessments (for 140 missions):

Equipment Mass Savings

Power Savings as Mass

11,875 kg

negligible

Overall Mass Savings 11,875 kg

Composite Qualitative Score +2

122 Thermal density is defined here as 1/(heat of fusion).

123 Including mass for packaging which is assumed as an additional 80% of the PCM mass.

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3.2.5 Summary

The various advanced technologies and their estimated benefits are summarized in

the table below for STS upgrade. From Section 3.:2.2, the mass of the baseline ETCS is

201,306 kg for 140 missions. Again, assuming the mass determinations throughout this

study have associated uncertainties on the order of 10%, a complete TCS with an

advanced technology would need to show a savings of at least 20,131 kg to ensure a mass

savings. Further, because design and development costs are not trivial, a mass savings of

25%, or 50,327 kg, is desirable. Using these criteria, the TCSs with advanced

technologies proposed for STS upgrade may be divided among five categories:

• TCSs using advanced technologies requiring a mass penalty greater than

10% of the overall baseline ETCS mass: none.

• TCSs using advanced technologies requiring a mass penalty less than 10%

of the overall baseline ETCS mass: none.

• TCSs using advanced technologies with a mass savings less than 10% of

the overall baseline ETCS mass: LP two-phase TCS.

• TCSs using advanced technologies with a mass savings between 10 and

25% of the overall baseline ETCS mass: lightweight radiators.

• TCSs using advanced technologies with a mass savings greater than 25%

of the overall baseline ETCS mass: none.

The technology in the third category will produce an ETCS which is comparable to

the baseline system. The technology in the fourth category is promising for this mission.

One technology presented above is really an enhancing technology. In other

words, this technology will deliver a mass savings or penalty regardless of the TCS

selected. Thus:

• Enhancing technologies which require a mass penalty: none.

• Enhancing technologies which yield a mass savings: phase-change thermal

storage.

Table 3.18 Advanced Active Thermal Control System Architecture

for Space Transportation System Upgrade

Summary of Advanced Active Thermal Control System

Architecture for Space Transportation System Upgrade

Overall Mass Qualitative

Savings [kg] Score

3.2.4.1 Low-Power Two-Phase Thermal Control System

3.2.4.2 Lightweight Radiators

3.2.4.3 Phase-Change Thermal Storage

8,176 -1

22,904 --

11,875 +2

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4.0 PLANETARY MISSIONS

Several aspects of surface operations distinguish planetary missions from other

missions. Most obvious is the presence of a nontfivial gravity field. Further, the time

constant associated with the sunlight/shade cycle on a planetary surface is on the order of

tens to hundreds of hours. The planet itself provides an additional support surface for

equipment including TCSs. Finally, some planets also have an atmosphere which can

deposit material on TCS surfaces and buffet fragile structures which are best suited to a

vacuum.

4.1 FIRST LUNAR OUTPOST LANDER

4.1.1 Reference Mission

One possible approach to re-establish a human presence on Luna is to send a series

of expendable vehicles similar to those used by the Apollo program. One proposal, known

as First Lunar Outpost (FLO), would utilize two vehicles to place a crew of up to four on

the lunar surface for up to 45 days (including one lunar day and one lunar night.)

(Woodcock, 1993) The fh-st transfer vehicle would carry the crew from Earth to the lunar

surface and then back to Earth. The second vehicle, which would actually arrive on the

lunar surface before the crew, would be a pilotless lander which would serve as a habitat

while the crew is on the lunar surface. This second lander is to be built around a Space

Station habitation module with appropriate modifications to make it an independent

vehicle. This will be FLO Lander. In addition to the habitation module, FLO Lander will

include a base containing landing gear and a descent motor, external tankage and ladders,

an airlock, and a complete ATCS designed for both lunar day and lunar night. For this

study, the continuous power mass penalty is 616 kg/kW while the daytime penalty, which

presumes solar power generation, is 42.2 kg/kW (Woodcock, 1993). Here, two FLO

Landers (one for each of two human missions) are assumed, and mass savings are

considered cumulative for both missions. The crew transfer vehicle and the Earth to Luna

transfer of FLO Lander are not considered in this study.

4.1.2 Baseline Case

The baseline FLO Lander ETCS will use a low solar absorptivity, horizontal

radiator 124 with single-phase ammonia as the ETCS fluid. The ITCS loop is assumed to

use water. No additional cooling devices are presumed, although some may be required

for mission phases other than surface operations. The radiator upper surface coating is

silver Teflon with an assumed surface emissivity of 0.80 and a solar absorptivity of 0.10.

The lower radiator surface is insulated to reduce heating from solar irradiation reflected by

the lunar surface. The radiator fin efficiency and mass per surface area are 0.85 and

124 A study by Cross (1995) indicates that radiators are significantly less massive than evaporative

systems for spacecraft in near-Earth space on missions lasting more than a week.

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6.00 kg/m 2 125, respectively. Woodcock (1993), assuming an ETCS with a heat pump,

sized the FLO Lander radiator at 63 m2 and placed it directly above the habitation module.

Here it is assumed that any radiator area in excess of 63 m2 will initially be folded on top

of a base radiator of 63 m 2. The additional radiating area will be deployed once FLO

Lander is in place on the lunar surface (Figure 4.1). The deployable radiator panels have

an assumed mass per surface area of 9.00 kg/m 2. This higher penalty accounts for the

additional deployment mechanism and structure associated with the deployable panel

segments. The heat-rejection requirement of the ETCS is 16.0 kW and the total pumping

power is 0.30 kW (Woodcock, 1993). In this study a heat pump is not included in the

baseline ETCS because heat pumps are considered an advanced technology. Thus, the

baseline FLO Lander ETCS design is:

• Overall ETCS Performance:

Heat Rejected

ETCS Mass Flowrate

ETCS Working Fluid

16.0 kW

0.1453 kg/s

single-phase ammonia

• Low Solar Absorptivity, Horizontal Radiators:

Surface Emissivity

Solar Absorptivity

Fin Efficiency

0.80

0.10

0.85

Average Radiator Surface Temperature

Average ETCS Fluid Temperature

Radiator Outlet Temperature

Radiator Inlet Temperature

285.0 K

286.5 K

274.82 K

298.18 K

Radiator Pressure Drop (as per ISS)

Pump Efficiency

Pumping Power for Radiators

48.26 kNlm z

0.45

0.0248 kW

125 This penalty includes 5.20 kg/m 2 for radiator panel mass and 0.80 kg/m 2 for base structure and

support (Woodcock, 1993).

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

i , illll ,i,

Figure 4.1 A sketch of a typical deployable radiator, as proposed for the baseline,

mounted on top of the modified Space Station habitation module of First Lunar Outpost

Lander. The top figure illustrates the radiators stowed for flight and to protect the

radiating surface while the vehicle external thermal control system is not in use. The

lower figure illustrates how the radiators might look when deployed. The large radiator

directly affixed to the habitation module has a radiating area of 63 m 2 while the two panels

deployed on either end provide any additional radiating area.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Summary of ETCS Masses for a Single FLO Lander Vehicle:

Radiator Surface Area 126

Base Radiator Support and Structure

Radiator Panel Mass

Support and Deployment for Added Panels

Total Radiator Mass

Deployment Motor (assumed)

Overall Pumping Power

Pumping Power as Mass

116.08 m 2

50.40 kg

603.62 kg

201.70 kg

855.72 kg

10 kg

0.300 kW

184.80 kg

Total ETCS Heat-Rejection System Mass 127 1050.5 kg

The average radiator mass per surface area is 7.37 kg/m z for the baseline configuration.

Figure 4.2 presents an overall sketch of the FLO Lander ETCS reference or baseline.

4.1.3 Parametric Study Using the Baseline Case

Because FLO Lander is still in the design stage, the baseline ETCS for this study

was determined through a parametric analysis. Density and specific heat for single-phase

ammonia were determined based on the average ETCS loop temperature using curve fits

to existing thermodynamic data.

P('T) = [- 0"002 _22 T2 - 0" 2331 K T + 858" 966_-_ 3

Ep(T)= [0.0061T + 2.9957 k--w-*s /kg *K

In addition to the constants above, other assumed study constants are:

Stefan-Boltzmann Constant

Solar Irradiation at Lunar Noon

Environmental Temperature (space)

Maximum Radiator Outlet Temperature

(ETCS set-point temperature)

ETCS Loop Temp. - Radiator Surface Temp.

5.670 × 10 -l 1 kW/(m2,K 4)

1.371 kW/m 2

3K

274.82 K

1.5 K

126

127

The radiator was sized using the study presented in Section 4.1.3 below by assuming an average

radiator surface temperature of 285.0 K.

This total excludes the mass of the ETCS working fluid, pump package, and piping and fittings.

While these masses are significant, their total is not expected to change significantly between the

baseline and the advanced technology options discussed below. For comparison, if an Apollo era

sublimator were used to reject a heat load of 16 kW for 45 days it would consume 25,500 kg of

water, not including tankage and other equipment masses.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

,a, _-

'.:7

< 5.,_

j;

,_ -

J

!

i

_ m m

1

!

.I

??

I' I

:+' ii_

|

ITCS Interchanger

I

I

I

I

l

I

!"

Module

Loop B

}i -,i: N Radiators

)

Z,

,x

I

I

l

I

) ,

I

I

I

Loop A

/ \

/ \

ITCS

Figure 4.2 An overall sketch of the baseline external thermal control system for First

Lunar Outpost Lander. Two flow loops are shown because the actual vehicle will

probably use two loops as a safety precaution. However, to simplify calculations and

analysis, the heat load and pumping power for both Ioops has been lumped into single

values for the First Lunar Outpost Lander vehicle as a whole.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

Because the FLO Lander radiator was not sized when this study was initiated, both the

surface solar absorptivity and the average radiator temperature were initially restricted to a

range of values. Both of these parameters were varied to determine their effect on the

required radiating area. While silver Teflon has a solar absorptivity when new of 0.08 to

0.09 (Peck, 1990), the landing of the vehicle with the crew and actual radiator usage over

45 days will degrade this coating slightly, so a value of 0.10 was assumed. The radiator

area was expected to be about twice that quoted by Woodcock (1993) because the

baseline would not include a heat pump. Further, a higher ETCS loop temperature than

used by either ISS or Shuttle would be permitted. An analysis of the required radiating

area is presented in Table 4.1 and Table 4.2, and graphically in Figure 4.3 and Figure 4.4.

Table 4.1 Various ETCS Flow Loop Values for First Lunar Outpost Lander

as a Function of Radiator Panel Temperature

Average Radiator Panel Temperature

275.0 K 280.0 K 285.0 K 290.0 K 295.0 K 300.0 K

Average ETCS Loop

Temperature [K] 276.5 281.5 286.5 291.5 296.5 301.5

Average NH3 Specific

Heat [kW*s/(kg*K)] 4.65 4.68 4.71 4.74 4.77 4.80

Average NH3

Density [kg/m 3] 641.6 634.9 628.1 621.1 614.1 606.9

NH3 Fiowrate [kg/s] 1.0239 0.2557 0.1453 0.1011 0.0773 0.0624

Required Radiator

Pumping Power [kW] 0.1711 0.0432 0.0248 0.0175 0.0135 0.0110

Required Radiator

Area at Night [m 2] 72.56 67.52 62.90 58.67 54.80 51.23

Table 4.2 Daytime First Lunar Outpost Lander Radiator Area [m 2]

as a Function of Both Solar Absorptivity and Radiator Panel Temperature

Solar

Absorptivity

0.080

0.085

0.090

0.095

0.10

0.11

0.12

0.13

0.14

0.15

0.16

0.17

0.18

0.19

0.20

Average Radiator Panel Temperature

275.0 K 280.0 K 285.0 K 290.0 K 295.0 K 300.0 K

125.71 111.30 99.29 89.15 80.50 73.04

131.74 116.00 103.02 92.14 82.93 75.03

138.38 121.12 107.03 95.34 85.51 77.14

145.73 126.71 111.37 98.77 88.26 79.37

153.89 132.84 116.08 102.46 91.19 81.73

173.32 147.07 126.80 110.72 97.68 86.90

198.36 164.71 139.70 120.43 105.16 92.78

231.86 187.16 155.53 132.01 113.88 99.50

278.97 216.70 175.40 146.06 124.18 107.27

350.10 257.32 201.09 163.44 136.53 116.37

469.93 316.66 235.59 185.53 151.60 127.14

714.47 411.59 284.40 214.52 170.42 140.12

1489.67 587.80 358.69 254.24 194.57 156.04

-- 1027.85 485.54 312.02 226.70 176.05

-- 4088.93 751.20 403.78 271.54 201.95

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400

350

300

250

<

L

200

150

100 -

50

270

Solar

Absorptivity

0.08

0.10

I I I I I

275 280 285 290 295 _00

Average Radiator Surface Temperature [K]

Figure 4.3 Required radiator surface area to reject a 16.0 kW heat load during lunar

noon at the lunar equator as a function of the average radiator surface temperature for a

horizontal, low solar absorptivity radiator on First Lunar Outpost Lander. A surface

emissivity of 0.8 and a fin efficiency of 0.85 are assumed. Silver Teflon coating, the

assumed surface coating, has a solar absorptivity of 0.08 when new and a value of 0.15 at

end-of-life.

4.1.4 Advanced ATCS Architecture for First Lunar Outpost Lander

FLO Lander vehicles are assumed here to be expendable. They will be placed on

the lunar surface and left behind once the objectives at their landing site have been

accomplished. This is not to say that the vehicle may not be used by more than one crew.

In fact, FLO Lander may be used multiple times by crews to carry out studies at a single

lunar site, especially if the perishable items are replenished between visits. However, to

bound this study, it is assumed that the vehicle usage does not permit significant

degradation of the ETCS components either by solar irradiation, debris impact, or

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

abrasion due to lunar soil propelled by the thrust of other vehicles landing nearby 128. As

noted earlier, only two FLO Lander vehicles are presumed. Each advanced architecture is,

where applicable, assessed numerically for the overall mass savings when compared with

the FLO Lander baseline. Qualitative assessments for these advanced technologies are

presented in Section 2.0.

q_

t_,

4

1000

900

800

700

600

500

400

300

Average Radiator Surface Temperature

275.0 K I_ 290.0 K

---'-'<)"--- 280.0 K "* 295.0 K

A 285.0 K _ 300.0 K

100

I I I I I

0.08 0.1 0.12 0.14 0.16 0.18 0.2

Solar Absorptivity

Figure 4.4 Required radiator surface area to reject a 16.0 kW heat load during lunar

noon at the lunar equator as a function of the solar absorptivity for a horizontal, low solar

absorptivity radiator on First Lunar Outpost Lander. A surface emissivity of 0.8 and a fin

efficiency of 0.85 are assumed. The curves represent performance for various average

radiator surface temperatures. For comparison, the STS-41 case for Shuttle has an

average radiator surface temperature of 288.3 K and the LVS Base Case for International

Space Station has an average radiator surface temperature of 273.9 K.

128 By assuming that the radiators will be stowed while other vehicles are landing or taking off, there

should be little erosion of the surface coatings due to lunar soil propelled at the radiators by the

thrust of other vehicles nearby.

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4.1.4.1 Low-Power Two-Phase Thermal Control System

To estimate the mass of a LP two-phase TCS 129 for FLO Lander, this study

extrapolates predictions from Ungar (1995) which gives estimates for space stations.

Ungar (1995) includes estimates for three different vehicles using four different TCS

architectures. Of the options presented, FLO Lander most closely resembles a small

system with a single-phase cascade TCS. Ungar also presents a LP two-phase TCS. The

LP two-phase TCS requires less pumping power and smaller diameter flow lines than the

single-phase cascade TCS. The required radiator area is roughly the same for both the

single-phase cascade TCS and the LP two-phase TCS. Thus, neglecting the differences in

the flow line mass, the systems may be summarized by:

Pump Power [kW]

Radiator Area [m z]

Loop Set-Point [K]

First Lunar Outpost

Lander Baseline

Thermal Control System i

0.300

116.1

274.8

Small System (Ungar, 1995)

Single-Phase Cascade

Thermal Control System

LTL LTL

0.320 0.320

197 197

275.2 275.2

Low-Power Two-Phase

Thermal Control System

LTL MTL

0.068 0.080

195 195

275.2 287.2

LTL refers to the low temperature loop while MTL refers to the moderate temperature

loop. Actually, the baseline TCS for FLO Lander uses a pair of single-phase cascade, low

temperature loops. The listed values in the table above, however, are for the entire

baseline TCS for FLO Lander. Thus, comparing the FLO Lander TCS with the singlephase cascade TCS for a small system reveals that the pumping power per loop is about

half, the FLO Lander TCS loop set-points are approximately that of the LTL for the

single-phase cascade TCS, and FLO Lander uses 37% of the effective radiator area of the

single-phase cascade TCS. A comparable LP two-phase TCS for FLO Lander, then,

would use two LTL loops. The total pumping power requirement for the revised FLO

Lander LP two-phase TCS is 68 W, which is half of the total pumping power for two LTL

loops on Ungar's LP two-phase TCS for a small station. Thus:

Mass Gained Using a

Low-Power Two-Phase

Thermal Control System

Baseline First Lunar

Outpost Lander TCS

Low-Power TwoPhase First Lunar

Outpost Lander TCS

Change Total Change

(for 1 Vehicle) (2 Vehicles)

Pump Power [kW] 0.300 0.068 -0.232

Mass due to

Pump Power [kg] 184.8 41.9 - 142.9 -285.8

Overall, the LP two-phase TCS is lighter than the baseline single-phase cascade TCS.

However, because FLO Lander is a very small vehicle, like Shuttle, a two-phase TCS does

not provide a large savings.

129 See Section 2.1.2 for details.

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Specific Assessments (for 2 missions):

Equipment Mass Savings

Power Savings

Power Savings as Mass

negligible

0.464 kW

+286 kg

Overall Mass Savings 286 kg

Composite Qualitative Score -1

4.1.4.2 Two-Phase Thermal Control System With Eiectrohydrodynamic Pumping

Based on the discussion in Section 2.1.3, FLO Lander could utilize a two-phase

TCS with electrohydrodynamic pumping. A two-phase TCS with electrohydrodynamic

pumping may yield a slight mass savings over a LP two-phase TCS. A significant mass

savings over the baseline TCS would also be expected. However, without more data on

this technology, no mass savings estimates may be determined.

Specific Assessments (for 2 missions):

Equipment Mass Savings

Power Savings as Mass

unknown

unknown

Overall Mass Savings unknown

Composite Qualitative Score +2

4.1.4.3 Capillary Pumped Loops

Based on the baseline architecture given in Section 4.1.2, a capillary pumped

loop 130 could save 0.30 kW, which is the estimated ETCS pumping power for a FLO

Lander. A single capillary pumped loop might be used for the heat load rejected by this

vehicle. However, such considerations do not affect the mass estimates for this study

because the capillary pumped loop equipment mass is assumed to be comparable to the

equipment mass for the baseline single-phase TCS architecture. The overall savings for

this option is 369.6 kg for two FLO Landers based on saving power for pumping and a

power mass penalty of 616 kg/kW.

Specific Assessments (for 2 missions):

Equipment Mass Savings

Power Savings

Power Savings as Mass

none

0.30 kW

370 kg

Overall Mass Savings 370 kg

Composite Qualitative Score 0

130 See Section 2.1.4 for background on capillary pumped loops.

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4.1.4.4 Solar Vapor Compression Heat Pump

A vapor compression heat pump could be used to reject heat at higher

temperatures so as to reduce the required radiator area. In fact, Woodcock (1993)

decided against using just a radiator, as in the baseline for FLO Lander here, and sized the

ETCS assuming a vapor compression heat pump. As will be shown below, a continuous

heat pump is not necessary on Luna and, therefore, is inefficient because it consumes

power during the night. Further, insufficient detail is present in this study to consider a

non-solar heat pump separately from a solar heat pump. The analysis here follows that in

Section 3.1.5.4 using the baseline from Section 4.1.2 to size the heat pump for FLO

Lander. Some assumptions are:

• Because this is a solar heat pump, the appropriate power mass penalty for

the heat pump power is 42.2 kg/kW.

• Because FLO Lander is compact and even smaller than Shuttle, the heat

pump evaporator will take heat directly from the ITCS through a heat

exchanger (Figure 4.5).

• The ITCS pumping power is assumed to be unchanged.

• The baseline ETCS pumping power is retained and added to the ITCS

pumping power 131

A two-parameter parametric study (Figure 4.6) indicates that the ETCS mass is minimized

for FLO Lander when the heat pump temperature lift is 27.5 K and the ETCS fluid mass

flowrate is increased to 0.368 kg/s. Specifically then:

Cold Source Temperature, Tc (evaporator temperature) 286.5 K

Temperature Lift, TH - Tc 27.5 K

Hot Source Temperature, TH (condenser temperature) 314.0 K

Radiator Inlet Temperature, Tin 310.0 K

Lunar Surface Solar Irradiation 1.371 kW/m 2

Environmental Temperature, Tspace 3.0 K

Total Cooling Load, Qc 16.00 kW

Ideal Coefficient of Performance, COPcarnot 10.05

Heat Pump Efficiency, 1"1132 (Ewert, 1991) 0.50

Necessary Input Power, Wreal 3.18 kW

The ETCS radiators are represented by the model developed for the parametric study in

Section 4.1.3.

Emissivity

Solar Absorptivity

Fin Efficiency

Radiator Mass Flowrate of Ammonia

0.80

0.10

0.85

0.368 kg/s

131 This accounts for the longer ITCS lines and increased ITCS pumping loads.

132 Percentage of Carnot coefficient of performance (COP).

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Radiators

L

il

i _ _ ill

if'

¢

ilJ i_

Heat Pump Bypass Loop ¢'-

-- ndln "'- -_ - I

Compressor I

 :i°n ',

I

Evaporator I I

/ I

_ _..........................

Figure 4.5 An overall sketch of the First Lunar Outpost Lander external thermal control

system using a solar vapor compression heat pump. Because the heat pump evaporator

receives heat directly from the internal thermal control system loop, the external thermal

control system flow loop exists only to carry heat from the heat pump condenser to the

radiators and to bypass the heat pump during the lunar night.

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Average Ammonia Temperature

Average Panel Surface Temperature

Heat Rejection per Unit Area

Total Heat Rejected by the Radiators, QH,real

Radiator Mass Per Unit Area (< 63 m 2)

Radiator Mass Per Unit Area (> 63 m 2)

Necessary Radiator Surface Area During the Day

300.0 K

298.5 K

0.304 kW/m 2

19.18 kW

6.00 kg/m 2

9.00 kg/m 2

62.82 m 2

Mass Gained by Using a

Solar Vapor Corn- Baseline FLO Solar Heat Pump Change Total Change

pression Heat Pump Lander TCS FLO Lander TCS (for 1 Vehicle) (2 Vehicles)

Radiator and

Support Mass [kg] 855.7 376.9 -478.8 -957.6

Heat Pump Mass 133 [kg] -- 149.7 +149.7 +299.4

Radiator Plus

Heat Pump Mass [kg] 855.7 526.6 -329.1 -658.2

Heat Pump Power [kW] -- 3.07 +3.07

Additional Pumping

Power [kW] -- 0.11 +0.11

Power as Mass[kg] -- 199.5 +199.5 +399.0

Total Mass [kg] 855.7 726.1 - 129.6 -259.2

As Woodcock (1993) and the tabulation suggest, a solar vapor compression heat pump

offers a mass savings for FLO Lander. While the radiator area decreases, the added solar

array will offset this savings. The deployment should be easier because the deployable

radiator wings have been omitted. The deployment motor is retained for the solar heat

pump configuration to deploy and retract a thin protective covering over the radiator

surface while other vehicles are taking off or landing near FLO Lander. Though the motor

can be lighter, it is assumed that the covering mass will offset any savings.

Specific Assessments (for 2 missions):

Equipment Mass Savings

Power Savings

Power Savings as Mass

658 kg

-6.36 kW

-399 kg

Overall Mass Savings 259 kg

Composite Qualitative Score 134 +1

133 See the footnotes for Section 3.1.5.4 for the heat pump sizing correlation.

134 Volume is rated as "average" for this mission because this attribute is expected to be comparable to

the baseline architecture for the solar vapor compression heat pump.

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800

600

400

0

0 10 20 30 40 50

Temperature Lift [K]

r_

t_

800

600

400

200

w

I I I I

}.0 0.2 0.4 0.6 0.8

Radiator Loop Mass Flowrate [kg/s]

Heat Pump Power _ Pumping Power

Heat Pump _ Total

Radiator

.0

Figure 4.6 Variation of First Lunar Outpost Lander external thermal control system mass

as a function of the vapor compression heat pump temperature lift and the radiator loop

mass flowrate. The baseline from Section 4.1.2 forms the basis for this study. This study

optimized external thermal control system mass with respect to both of the variables

mentioned.

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4.1.4.5 Lightweight Radiators

A parametric study here is employed to identify potential mass savings through

using lighter materials for portions of the FLO Lander ETCS. Unlike ISS and Shuttle, the

FLO Lander ETCS is not well defined, so the components and categories listed here are,

out of necessity, vague.

Mass per First Mass Percentage of

Category Within Lunar Outpost External Thermal

External Thermal Control System Lander Vehicle [kg] Control System

Structure and Deployment: 24.0

Base Radiator Support and Structure 50.4

Support and Deployment for Added Panels 201.7

Radiator Panels 603.6 57.5

Other Items: 18.5

Deployment Motor 10.0

Pumping Power Mass 184.8

Total 1050.5 100.0

This study varies the component FLO Lander ETCS masses linearly based on the

original mass in each category. The structure and deployment masses are reduced by up

to 20% when the radiator panel mass is reduced by up to 40%. The ETCS working fluid

is not listed, but it is not expected to change as a result of lighter material components

throughout the rest of the ETCS 135

The underlying assumption is that composites and other advanced materials are

most likely to offer a mass reduction for some components of the radiator panels such as

the honeycomb and the facesheets. Because the radiator panels are the single most

massive item of the FLO Lander ETCS, such savings would constrain the remainder of the

design. A less significant mass reduction is assumed for the structures and deployment

because the sizes of these components are primarily dictated by the dimensions of the

radiator array. However, composites should allow comparable components to replace

some of the structure and deployment with lighter parts. In summary, this approach yields

an overall ETCS mass reduction of 27.8% for a 40% reduction in the radiator panel mass

(Figure 4.7).

135 See Section 2.4 for additional general background on lightweight radiators plus specific examples of

proposed lightweight radiators.

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1000

800

400-

200_

A

o o

I I I

0 10 20 30 40

Percent Reduction in Radiator Panel Mass

Structure and Deployment _ Other Items

Radiator Panels _ Total ETCS Mass

Figure 4.7 First Lunar Outpost Lander radiator mass as a function of mass reduction

within the radiator panel set. Study Assumption: For each 10% mass reduction in the

radiators the mass for the structure and deployment decreases by 5%.

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Percent Reduction in

Category Within Radiator Panel Mass

External Thermal Control System 20% 40%

Structure and Deployment [kg] 226.9 201.7

Radiator Panels [kg] 482.9 362.2

Other Items [kg] 194.8 194.8

Total ETCS Mass per Vehicle [kg] 904.6 758.7

Overall Mass Reduction per Vehicle [kg] 145.9 291.9

Mass Reduction as a Percentage

of the Original FLO Lander ETCS 13.9 27.8

Radiator Mass Per Surface Area [kg/m 2] 136 6.11 4.86

Considering the available lightweight radiators presented in Section 2.4, an overall mass

reduction of 27.8% was selected as a representative value. Thus, the mass savings for one

FLO Lander vehicle is 291.9 kg, or 583.8 kg for the reference mission of two vehicles.

The power requirements will be unchanged because the internal fluid dynamics and,

therefore, the pumping power are functions of the working fluid material properties.

Because FLO Lander is a planetary mission, any of the lightweight radiator

concepts may be used. All of the radiators presented in Section 2.4 can be deployed

vertically. However, vertical radiators away from polar regions require integration with

either a radiator shade or a heat pump to reject life support heat loads at lunar noon.

Additionally, the composite flow-through radiators could also be oriented horizontally like

the radiator in the baseline architecture.

Specific Assessments (for 2 missions):

Equipment Mass Savings

Power Savings as Mass

584 kg

none

Overall Mass Savings

4.1.4.6 Parabolic Radiator Shade

584 kg

Parabolic radiator shades provide a means of lowering the effective sink

temperature around a radiator to allow rejection at temperatures associated with waste

heat from environmental control and life support systems for human beings 137.

Two possible deployments using a parabolic radiator shade for FLO Lander at an

equatorial landing site are presented in Figure 4.8. The first would place the shade on top

of the modified Space Station habitation module with a vertical radiator. This option is

difficult to employ for at least two reasons. Preliminary calculations indicate that the top

of the FLO Lander vehicle has insufficient room for a radiator shade. Further, because

FLO Lander is pilotless, computer guidance would need to flawlessly position the vehicle

on touchdown to align the radiator's main axis with the local incident solar vector.

136 These values are based on a radiating area of 116.08 m 2 as per the baseline architecture for FLO

Lander. Further, these values include only the radiator mass and not the Other Items in the ETCS.

137 See Section 2.5.2 for more information on parabolic radiator shades.

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Without allowing for additional equipment to adjust the radiator's alignment on top of the

vehicle, this approach is unworkable.

Figure 4.8 Two deployments for a parabolic radiator shade. The first deploys the shade

on top of First Lunar Outpost Lander. This option could be deployed remotely but would

rely on the First Lunar Outpost Lander vehicle positioning itself in the proper position

relative to the solar vector. The second deploys the radiator with parabolic shade on the

lunar surface. This option would be carried in a box and deployed by the crew after

arrival. The positioning of the First Lunar Outpost Lander vehicle itself would not be

critical because the crew themselves could adjust the parabolic shade and radiator

assembly to align with the solar vector.

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A second approach would use the crew to set up the radiator and shade on the

lunar surface after arrival. This approach allows the shade to be aligned properly with

respect to the local solar vector. Further, preliminary calculations indicate that this option

requires significantly less mass than a vehicle-mounted installation 138. This implies that

the habitat will not be ready for the crew when they reach the lunar surface. Because the

mission profile calls for an extravehicular activity only sufficient for the crew to transfer

from the piloted vehicle to the habitat, the habitat should be remotely readied (Woodcock,

1993). However, assuming that this readiness issue can be solved, a surface installation of

a vertical radiator with a parabolic shade is the most reasonable approach 139. The

radiator and shade combination could be set up as illustrated in the lower part of

Figure 4.8 or as a single continuous unit. In either case, endsheets (not pictured) should

be used to further minimize heating from surface irradiation.

For a parabolic radiator shade on the lunar surface (Ewert and Clark, 1991):

Heat Rejected 16.0 kW

Radiator Surface Temperature 285.0 K

Lunar Surface

Emissivity 0.93

Absorptivity 0.93

Albedo 0.07

Solar Irradiation 1.371 kW/m 2

Incident Angle 90.0 °

Orbital Inclination 1.53 °

Radiator (with silver Teflon)

Fin Efficiency 0.85

Infrared Emissivity 0.80

Solar Absorptivity 0.10

Shade

Specular Upper Surface

Emissivity 0.03

Absorptivity 0.04

Diffuse Lower Surface

Emissivity 0.81

Absorptivity 0.371

Sink Temperature 140 155.0 K

138 The preliminary mass calculations include only a modest estimate for a shade deployment based on a

surface installation. A vehicle-mounted installation would probably require more massive pointing

mechanisms.

139 To allow the habitat to be remotely readied, radiators could be vehicle-mounted when the crew

arrives. Assuming the crew arrives at lunar sunrise, the unshaded radiators should be sufficient to

cool the vehicle until the crew can set up the parabolic shade on a subsequent extravehicular activity.

Finally, the vehicle-mounted radiators might be transferred to the lunar surface for use with the

parabolic shade.

140 Keller (1995 a) computed a sink temperature of 153.58 K for a parabolic shade and radiator

assembly mounted on a lander vehicle 10 m off of the lunar surface. For a shade on the lunar

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Radiator

Area 68.9 m 2

Mass Penalty (Keller, 1994) 4.5 kg/m 2

Mass 310.2 kg

Dimensions

Height 1.00 m

Length 34.47 m

Shade

Parabolic Arc Length 141 4.59 m

Length 34.47 m

Area 15 8.2 m E

Mass Penalty 142 I. 1 kg/m 2

Mass 174.0 kg

Total System Mass 143 484.2 kg

Comparing the radiator and flexible parabolic shade with the baseline presented in

Section 4.1.2 yields:

Mass Gained by Using a

Parabolic Radiator Shade

Baseline First First Lunar

Lunar Outpost Outpost Lander

Lander Thermal With Parabolic Change Total Change

Control System Radiator Shade (for 1 Vehicle) (2 Vehicles)

Radiator Area [m 2] 116.1 68.9 -47.2

Radiator Mass [kg] 603.6 310.2 -293.4 -586.8

Added Panel Support &

Deployment [k_] 201.7 -- -201.7 -403.4

Shade Area [m E] -- 158.2 +158.2

Shade Mass [kg] -- 174.0 +174.0 +348.0

Total Mass [kg] 805.3 484.2 -321.1 -642.2

The deployment motor mass is retained in this configuration to account for the mass of a

remotely removable covering over the radiators while they are vehicle-mounted. The

overall savings for using a parabolic radiator shade is 642.2 kg.

surface the effective sink temperature would be slightly higher. The shade properties assumed here

are beginning-of-life values which are appropriate for this mission because dust accumulation will be

negligible. This explains the difference between the lower sink temperature here and the higher

value listed for the PLB shade which is sized using end-of-life shade properties. Keller used TSS.

141 A full parabolic shade is assumed. In other words, the shade top is even with the radiator top and

the shade focus is at the height of the radiator.

142 This penalty includes mass for a flexible parabolic shade plus appropriate support structure. This

penalty is twice that used by Keller (1994). The additional mass accounts for more extensive

deployment structure which will allow for the radiator and shade to be deployed quickly.

143 This value does not include mass for piping, fittings, and working fluid to and from the radiators.

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Specific Assessments (for 2 missions):

Equipment Mass Savings

Power Savings as Mass

642 kg

negligible

Overall Mass Savings 642 kg

Composite Qualitative Score 144 -2

4.1.5 Summary

The various advanced technologies and their estimated benefits are summarized in

the table below for FLO Lander. From Section 4.1.2, the mass of the baseline ETCS is

2,101.0 kg for two missions. Assuming the mass determinations throughout this study

have associated uncertainties on the order of 10%, an overall TCS with an advanced

technology would need to show a savings of at least 210 kg to ensure a mass savings.

Further, because design and development costs are not trivial, a mass savings of 25%, or

525 kg, is desirable. Using these criteria, the TCSs with advanced technologies proposed

for FLO Lander may be divided among five categories:

• TCSs using advanced technologies requiring a mass penalty greater than

10% of the overall baseline ETCS mass: none.

• TCSs using advanced technologies requiring a mass penalty less than 10%

of the overall baseline ETCS mass: none.

• TCSs using advanced technologies with a mass savings less than 10% of

the overall baseline ETCS mass: none.

• TCSs using advanced technologies with a mass savings between 10 and

25% of the overall baseline ETCS mass: LP two-phase TCS, capillary

pumped loops, and solar vapor compression heat pump.

• TCSs using advanced technologies with a mass savings greater than 25%

of the overall baseline ETCS mass: lightweight radiators and parabolic

radiator shade.

The technologies in the fourth category are promising but not outstanding for this

mission. The technologies in the final category show significant promise for this mission.

However, as discussed above, the parabolic radiator shade needs to be properly deployed

on the lunar surface to function properly and this may require an extravehicular activity.

144 The deployment is rated as "difficult" for this mission because the mass assessment assumes a crew

extravehicular activity to place the parabolic shade on the lunar surface.

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Table 4.3 Advanced Active Thermal Control System Architecture

for First Lunar Outpost Lander

Summary of Advanced Active Thermal Control System

Architecture for First Lunar Outpost Lander

Overall Mass

Savings [kg]

Qualitative

Score

4.1.4.1 Low-Power Two-Phase Thermal Control System

4.1.4.2 Two-Phase TCS with Electrohydrodynamic Pumping

4.1.4.3 Capillary Pumped Loops

4.1.4.4 Solar Vapor Compression Heat Pump

4.1.4.5 Lightweight Radiators

4.1.4.6 Parabolic Radiator Shade

286

unknown

370

259

584

642

-1

+2

0

+1

-2

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4.2 PERMANENT LUNAR BASE

4.2.1 Reference Mission

In order to intensively investigate and utilize Luna, a Permanent Lunar Base (PLB)

is, for expedience, necessary. It is conceivable that PLB could become a reality sometime

early in the next century. One proposal, as presented in Figure 4.9, would bury three

modified Space Station modules under two meters of lunar regolith and provide living and

working space continually for a crew of three or four. The regolith acts as a radiation

shield for typical levels of incident irradiation. The base elements would include a

habitation module, a laboratory module, and a plant growth module. The plant growth

module would be an integral part of the base environmental control and life support

system by replenishing atmospheric oxygen and removing carbon dioxide. The plants are

also expected to provide food for the crew. Power for the base would be supplied by

photovoltaic solar arrays located near the crew's quarters on the lunar surface. Base

access would be through Space Station-type nodes connected to airlocks on the surface.

Baseline ATCS heat rejection would be accomplished through horizontal radiators with

low surface solar absorptivity 145. These radiators would be insulated from below, using

lunar regolith, and deployed on the lunar surface above the living areas. The overall

nominal ATCS heat load is projected at 50 kW with peak individual module loads of

25 kW. The ETCS fluid is single-phase liquid ammonia. The continuous power mass

penalty is 750kg/kW which assumes solar photovoltaic power generation with

regenerative fuel cell energy storage. The power mass penalty for "daytime only" power

usage is 35.3 kg/kW for electricity taken from the main PLB power grid and 20.2 kg/kW

for electricity taken from a dedicated photovoltaic array (Hughes, 1995). The project life

is 15 years and the total mass savings computed below are for a single base.

4.2.2 Baseline Case

The baseline PLB ETCS uses low solar absorptivity, horizontal radiators with

single-phase liquid ammonia as the working fluid. As illustrated in Figure 4.10, each

module is serviced by two of the three ETCS loops. The TCS utilizes a single-phase

cascade similar to the TCS for ISS. Each module is supplied with both a low-temperature

ETCS loop interface (LTL) and a moderate-temperature ETCS loop interface (MTL).

Radiator bypass valves provide ETCS loop set-point temperature control. Further, each

PLB module has two ITCS loops using liquid water as the working fluid. Additional

fittings in each PLB module allow either ITCS loop to service all of the module's

coldplates and heat exchangers. This arrangement provides extra flexibility if any one of

the ITCS or ETCS loops fail. Each ETCS loop also has a dedicated pump module. The

overall base ATCS heat load is 50 kW, but each module may have a peak load up to

25 kW. A value of 7.50 kg/m 2 is assumed for the radiator mass per radiating area for

horizontal radiators. Vertical radiators are assumed to use 5.625 kg/m 2 146. This base will

145 These are also known as low alpha, horizontal radiators.

146 The radiators on ISS, which are completely vehicle-mounted and use an extensive deployment

mechanism, have a mass of 8.24 kg per m2 of radiating area. The ISS radiators are two-sided. The

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be located at the lunar equator which has the most severe thermal environment on Luna.

Here surface temperatures range from 102 K at night to 384 K at lunar noon (Ewert,

Petete, and Dzenitis, 1990). The solar irradiation at noon is 1.371 kW/m 2 with a surface

albedo on average of 0.07 147. The lunar surface has the properties of a diffuse gray

surface with an emissivity of 0.93 to 0.94 148

Overall ETCS Performance:

Heat Rejected

ETCS Working Fluid

50.O kW

single-phase ammonia

Low Solar Absorptivity, Horizontal Radiators 149.

Service Life (each set of panels)

Surface Coating

Surface Emissivity (end-of-life)

Solar Absorptivity (end-of-life)

Fin Efficiency

Average Radiator Surface Temperature

Radiator Surface Area (day load)

Radiator Surface Area (night load)

Mass Penalty for Dry Radiators 150

Radiator Panel Mass

Number of Radiator Panel Sets per Project Life

Total Project Radiator Mass

Radiator Working Fluid Mass Penalty (ISS)

Radiator Working Fluid Mass 151

7.5 years

10 mil silver Teflon

0.86

0.14

0.85

275.0 K

650.81 m 2

217.26 m 2

7.50 kg/m 2

4881.08 kg

2

9765.16 kg

0.358 kg/m 2

232.99 kg

147

148

149

150

151

radiator on FLO Lander is also vehicle-mounted, but single-sided (it is insulated on its underside),

and has a base mass of 6.00 kg per m 2 of radiating area when not considering the heavier deployable

panel sections. For PLB, the radiators will probably be packaged in some manner such that they can

be deployed with minimal extravehicular activity time while using the lunar surface for support.

Thus, some integral mass for a deployment mechanism is appropriate. However, the overall package

is expected to be less complicated and lighter than the ISS radiators, yet heavier and possibly more

complex than the FLO Lander radiators.

This value is characteristic of a lunar mare. Surface albedo may vary between 0.07 and 0.24

according to Binder (1990).

Keller (1995 a) indicates that surface solar absorptivity can vary between 0.7 and 1.0. Thus, a

diffuse gray surface is an approximation for the lunar surface.

The radiators were sized based on the parametric study given in Section 4.2.3 below.

This mass penalty includes structure and deployment mass as well as actual radiator panel mass.

The ETCS working fluid is assumed to not require replacement during the project life except if the

original fluid is somehow lost (such as through a puncture in the radiators). This mass estimate is

based on data for the working fluid mass used in the ISS radiator ORUs as a function of radiating

area.

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<

tJ

.a

"6

0

f.

ID

"6

F.

z

,..1

o

ID

Z

Figure 4.9 A side view of the Permanent Lunar Base baseline case. Shown are the three

Space Station modules buried under 2 m of regolith. The external thermal control system

uses three flow loops (two-fault tolerant) and low solar absorptivity, horizontal radiators

with 10 mil silver Teflon surface coating. The power grid is also two-fault tolerant. The

flow loops and the power distribution lines are for illustrative purposes only. (See

Figure 4.10 for details of the external thermal control system plumbing.)

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Temperature

Control

|

|

Bypass Valve

f I. LTL

|

|

ET Loop B ,

ETCS Loop A | ...................... '

Habitation Module Laboratory Module

ETCS

Loop C

Plant Growth Module

Figure 4.10 Plumbing for the baseline PLB ETCS. LTL is low-temperature ETCS loop

and MTL is moderate-temperature ETCS loop. The system uses three ETCS loops to

give two-fault tolerance with single-phase liquid ammonia as the working fluid. The

bypass valves allow for temperature control. Each module has two ITCS water loops

which can be interconnected to allow all of a module's heat load to be removed by either

the LTL heat exchanger or MTL heat exchanger. A pump module (not pictured) is

assumed for each of the three ETCS loops.

ETCS Working Fluid Loop Mass Flowrates:

Average ETCS Fluid Temperature

Radiator Outlet Temperature

Radiator Inlet Temperature

Average Ammonia Specific Heat

Total Mass Flowrate

(full load/equal distribution)

Average Flowrate per Loop

(full load/equal distribution)

Maximum Flowrate per Loop 152

278.0 K

274.82 K

281.18 K

4.62 (kW*s)/(kg*K)

1.702 kg/s

0.5672 kg/s

0.6807 kg/s

152 This assumes a module loading of: habitation module: 10 kW, laboratory module: 25 kW, and plant

growth module: 15 kW. This arrangement includes the maximum loading for any one module plus

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ETCS Pump Module: Two cases were considered for pumping power

calculations. They may be summarized as:

Case Heat Load [kW] Flowrate [kg/s] Pump Efficiency

Nominal Load:

Loop A 16.67 0.5672 0.40

Loop B 16.67 0.5672 0.40

Loop C 16.67 0.5672 0.40

Maximum Load:

Loop A 12.50 0.4254 0.35

Loop B 20.00 0.6807 0.45

Loop C 17.50 0.5956 0.40

Average Ammonia Density

Radiator Pressure Drop per Radiating Area 153

Radiator Pressure Drop

Pumping Power for Radiators (day load) 154

Pumping Power for Radiators (night load) 155

Estimated Effective Line Lengths 156:

Loop A

Loop B

Loop C

Line Diameter 157

Maximum Line Pumping Power 158

Overall Pumping Power 159

Continuous Power Mass Penalty

Daytime Only Power Mass Penalty

Pumping Power as Mass 160

Mass of Ammonia in Lines

631.96 kg/m 3

0.3718 kN/m 4

242.0 kN/m 2

1.629 kW

0.544 kW

25.65 m

71.10m

57.90 m

0.0236 m

0.628 kW

2.26 kW

750 kg/kW

35.3 kg/kW

917.3 kg

32.76 kg

reserves 15 kW, as specified, for the plant growth module. The maximum per loop load is 20 kW

assuming each loop picks up half of a module's heat load.

153 This penalty is based on the values for ISS which specify a maximum radiator pressure drop of

48.26 kN/m 2 for 129.8 m 2of radiating area.

154 The nominal load yielded the greatest pumping power requirement for the radiators.

155 The night load pumping power is the day load pumping power multiplied by the ratio of the night

load and day load radiator areas.

156 These values are the sum of the estimated line lengths based on Figures 28 and 29 plus an additional

50% to account for fittings and valves.

157 This is the inside diameter of the ETCS lines within ISS modules. Specifically, this value is 0.93 in.

158 The maximum load yielded the greatest pumping power requirement for the ETCS lines.

159 This value includes 0.63 kW to pump ammonia through the ETCS loops and 1.63 kW to pump

ammonia through the radiators.

160 Assumption: The ETCS flowrate is constant with portions of the radiator either freezing and/or

being closed off at night when the heat rejection requirement is less. In pratical terms, this implies

that the pumping power is less at night than during daylight. Thus, the continuous power mass

penalty is applied only to the power associated with running the pumps continuously. The daytime

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• Summary of Permanent Lunar Base Active Thermal Control System:

Heat Load Rejected

Total Dry Radiator Mass (for 15 years)

Working Fluid Mass

Pumping Power

Pumping Power as Mass

50.0 kW

9765.2 kg

265.8 kg

2.26 kW

917.3 kg

Total ETCS Heat-Rejection System Mass 161 10,948.3 kg

The baseline radiator mass per area, based on a single set of radiator panels, is 7.86 kg/m 2.

Upon considering both sets of radiator panels, this value increases to 15.36 kg/m 2.

4.2.3 Parametric Study Using the Baseline Case

PLB is still undergoing preliminary design work, so the baseline for the ETCS

assumes only the values mentioned above. The actual baseline hardware includes low

solar absorptivity, horizontal radiators fabricated from materials like those specified for

ISS. Several radiator surface coatings are available, including silver Teflon and Z-93.

Additionally, the radiators could be designed for the entire project life of 15 years, or they

could be designed for a shorter time and then replaced as appropriate during scheduled

maintenance. The possible benefit of using radiators with a design life less than 15 years is

that the end-of-life surface coating properties would not be as severely degraded due to

environmental effects so that the overall radiator size could be smaller. For this study, the

masses of the surface coatings themselves were assumed negligible compared with the

masses for the other radiator components. A comparison of radiator designs for a

50.0kW heat load using three surface coatings and two radiator panel surface

temperatures is presented below and in Figure 4.11 and Figure 4.12. The radiators are

deployed horizontally on the lunar surface.

Table 4.4 Total Masses for the Lunar Base Radiator [kg]

Using Various Surface Coatings and Various Total Design Lives

Design Average Radiator Panel Surface Temperature (Horizontal Units)

Radiator 275.0 K 280.0 K

Panel Set 5.5 mil 10 mil 5.5 mil 10 mil

Life [years] Silver Teflon Silver Teflon Z-93 Silver Teflon Silver Teflon Z-93

15.0 12,258.9 52,765.9 7,846.8

7.5 16,025.6 9,762.1 12,080.7 7,935.0 52,662.3

5.0 16,189.6 11,706.6 89,466.0 13,231.5 9,869.6 36,157.2

2.5 23,815.1 19,181.4 73,383.3 20,403.5 16,617.0 45,731.3

1.0 50,870.4 42,955.8 133,625.2 44,454.6 37,704.3 92,768.4

161

only power mass penalty is applied to the additional power used to run the pumps during the day

when the entire radiator is in use.

This value does not include masses for the pumps, lines, and fittings, but these are expected to be

roughly the same for all PLB ATCS configurations.

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lO0

tD

'¢ 80- t_

t_

¢ff

o 60

_ 40

N 2o

o

ERadiator Surface Coating

[] 5.5 mil Silver Teflon

O 10 nail Silver Teflon

A Z-93

El

0 I I

0 5 10

Design Radiator Panel Set Life [years]

Figure 4.11 Total mass of horizontal radiator panels for the life of Permanent Lunar Base

(15 years) as a function of the expected design life for the panel sets. The study

assumptions are: Rejected heat load is 50 kW, the radiator mass per rejection area for a

dry panel is 7.50 kg/m 2, the fin efficiency is 0.85, and the radiator surface temperature is

275.0 K. The units on the vertical axis are Megagrams [Mg].

Table 4.5 Surface Coatings as a Function of Time (Peck, 1990)

Time 5.5 mil Silver Teflon 10 mil Silver Teflon Z-93

[:,,ears] E _ E _ E ot

0 0.78 0.08 0.89 0.09 0.90 O. 17

10 0.73 O. 15 0.85 0.15 0.90 0.23

30 0.65 0.20 0.81 0.20 0.90 0.30

Surface properties were estimated with second order polynomials using data from Peck

(1990) to set the function constants.

In all cases the 10 mil silver Teflon yields design estimates which are significantly

less massive than those for the other two coatings. Thus, barfing excessive cost or mass

contingencies, this is the recommended coating for this application.

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100

o,0

o 80

_"O 60

_ 40

°.=,,i

_ 20

O

0

Radiator Surface Coating

[] 5.5 mil Silver Teflon

O 10 mil Silver Teflon

A Z-93

0 5 10 15

Design Radiator Panel Set Life [years]

Figure 4.12 Total mass of horizontal radiator panels for the life of Permanent Lunar Base

(15 years) as a function of the expected design life for the panel sets. The study

assumptions are: Rejected heat load is 50 kW, the radiator mass per rejection area for a

dry panel is 7.50 kg/m 2, the fin efficiency is 0.85, and the radiator surface temperature is

280.0 K. The units on the vertical axis are Megagrams [Mg].

For the baseline, which uses a panel temperature of 275.0 K, the least massive

radiators have a design life of 7.5 years. Thus, the radiators will be replaced once as

scheduled maintenance. In addition to surface properties, the radiator design for PLB is

sensitive to the radiator surface temperature. Increasing the surface temperature by five

degrees to 280.0 K leads to a different optimal design. For this hotter case, a single

radiator with a design life of 15 years is the least massive option.

4.2.4 Advanced ATCS Architecture for Permanent Lunar Base

The PLB will be a single base with a projected life of 15 years. As stated above,

the radiators will be replaced once after 7.5 years. Unless otherwise noted, all advanced

technologies are expected to last at least for the entire project life of 15 years. Each

advanced architecture is, where applicable, assessed numerically for the overall mass

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savings when compared with the baseline. Qualitative assessments for these advanced

technologies are presented in Section 2.0.

4.2.4.1 Two-Phase Thermal Control System With Mechanical Pump/Separator

To convert PLB to use a two-phase TCS this study extrapolates predictions from

Ungar (1995) which provides estimates for space stations. Ungar (1995) includes a

medium space station which is similar to PLB. Ungar's medium station with a singlephase cascade TCS is like the baseline PLB TCS. Ungar (1995) also presents a

comparable two-phase TCS with MP/S 162

Pump Power [kW]

Radiator Area [m 2]

Loop Set-Point [K]

PLB Total

Baseline PLB

TCS Baseline

(1 loop) TCS

0.753 2.260

216.9 650.8

274.8 274.8

Medium Space Station (Ungar, 1995)

Single-Phase Cascade Two-Phase TCS With

TCS MP/S

LTL LTL LTL MTL

0.615 0.615 0.299 0.330

395 395 390 390

275.2 275.2 275.2 287.2

LTL refers to the low temperature loop while MTL refers to the moderate temperature

loop. The baseline ETCS for PLB uses three single-phase cascade, low temperature

loops. Comparing the PLB TCS with the single-phase cascade TCS for a medium station

reveals that the pumping power per loop is about 22% more for PLB and the PLB TCS

loop set-points are approximately that of the LTL for the single-phase cascade TCS. A

comparable two-phase TCS with MP/S for PLB would use two LTL loops and one MTL

loop. The radiator area will remain approximately the same. The pumping power is

estimated to be 0.366 kW for each LTL and 0.404 kW for the MTL 163. ThUS, the

significant savings for this option compared with the baseline TCS is a reduction in

9umping power.

Mass Saved by Using a Two-Phase Thermal Control Baseline TCS With Total Change

System With Mechanical Pump/Separator PLB TCS MP/S for PLB for TCS

Pump Power [kW] 2.260 l. 136

Mass Due to Pump Power [kg] 164 917.3 461.1 456.2

The two-phase TCS with MP/S uses smaller fluid lines than the baseline single-phase

cascade TCS which translates into an additional mass savings. However, this savings is

not included here.

A more important issue related to this technology is whether PLB will end after its

initial project life of 15 years or will it be an initial phase for a much larger installation?

From Ungar (1995), the two-phase TCSs offer the greatest advantage for large systems

with long ETCS transport lines. If PLB will end after its initial project life, then the mass

savings associated with reduction in pumping power is not as great as some of the other

162 Ungar (1995) refers to this option as a rotary fluid management device (RFMD) type two-phase

TCS. See Section 2.1.1 for additional details about the MP/S.

163 These values are 22% more than Ungar's values for the two-phase TCS with MP/S.

164 See Section 4.2.2 for details.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

advanced technologies presented below. However, if PLB will eventually expand into a

much larger facility, a two-phase TCS will yield substantial mass savings due to reductions

in pumping power, line mass, and radiator mass for that larger facility. As such, a twophase TCS for PLB would be a sound investment on which to build a larger TCS at a later

date.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

negligible

1.124 kW

456 kg

Overall Mass Savings 456 kg

Composite Qualitative Score

4.2.4.2 Low-Power Two-Phase Thermal Control System

0

To convert PLB to use a two-phase TCS this study extrapolates predictions from

Ungar (1995) which gives estimates for space stations. Ungar (1995) includes a medium

space station which is similar to PLB. Ungar's medium station with a single-phase cascade

TCS is comparable to the baseline PLB TCS. A comparable LP two-phase TCS 16s is also

presented.

Pump Power [kW]

Radiator Area [m2]

Loop Set-Point [K]

PLB Total

Baseline PLB

TCS Baseline

(1 loop) TCS

0.753 2.260

216.9 650.8

274.8 274.8

Medium Space Station (Ungar, 1995)

Single-Phase Cascade

TCS

LTL LTL

0.615 0.615

395 395

275.2 275.2

LP Two-Phase TCS

LTL MTL

0.094 0.147

390 390

275.2 287.2

LTL and MTL refer to low and moderate temperature loops, respectively. The Baseline

ETCS for PLB uses three single-phase cascade, low temperature loops. Comparing the

TCS for PLB with the single-phase cascade TCS for a medium station reveals that the

pumping power per loop is about 22% more for PLB and the PLB TCS loop set-points

are approximately that of the LTL for the single-phase cascade TCS. A comparable LP

two-phase TCS for PLB would use two LTL loops and one MTL loop. The radiator area

will remain approximately the same. The pumping power is estimated to be 0.115 kW for

each LTL and 0.180 kW for the MTL 166. THUS, the significant savings for this option

compared with the baseline TCS is a reduction in pumping power.

165 See Section 2.1.2 for additional details on LP two-phase TCSs.

166 These values are 22% more than Ungar's values for a LP two-phase TCS.

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Mass Saved by Using a Baseline LP Two-Phase Total Change

Low-Power Two-Phase Thermal Control System PLB TCS TCS for PLB for TCS

Pump Power [kW] 2.260 0.410

Mass Due to Pump Power [kg] 167 917.3 166.4 750.9

The LP two-phase TCS uses smaller fluid lines than the baseline single-phase cascade TCS

which translates into an additional mass savings. However, this savings is not included

here.

As in the previous section, a more important issue related to this technology is

whether PLB will end after its initial project life of 15 years or will it be an initial phase for

a much larger installation? From Ungar (1995), the two-phase TCSs offer the greatest

advantage for large systems with long ETCS transport lines. If PLB will end after its

initial project life, then the mass savings associated with reduction in pumping power is

not as great as some of the other advanced technologies presented below. However, if

PLB will eventually expand into a larger facility, a two-phase TCS will yield substantial

mass savings due to reductions in pumping power, line mass, and radiator mass for that

larger facility. As such, a two-phase TCS for PLB would be a sound investment on which

to build a larger TCS at a later date.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

negligible

1.850 kW

751 kg

Overall Mass Savings 751 kg

Composite Qualitative Score -1

4.2.4.3 Two-Phase Thermal Control System With Electrohydrodynamic Pumping

Based on the discussion in Section 2.1.3, PLB could utilize a two-phase TCS with

electrohydrodynamic pumping. A two-phase TCS with electrohydrodynamic pumping

may yield a mass savings over a LP two-phase TCS. A significant mass savings over the

baseline TCS would also be expected. However, without more data on this technology,

no mass savings estimates may be determined.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

unknown

unknown

Overall Mass Savings unknown

Composite Qualitative Score +2

167 See Section 4.2.2 for details.

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4.2.4.4 Capillary Pumped Loops

Based on Section 4.2.2, capillary pumped loops 168 will save 2.26 kW, which is the

estimated ETCS pumping power for PLB. As above, the capillary pumped loop

equipment mass is assumed to be comparable to the equipment mass for the baseline

single-phase TCS architecture. Thus, the overall savings for this option is 917.3 kg based

on saving power for pumping 169

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

none

2.26 kW

917 kg

Overall Mass Savings 917 kg

Composite Qualitative Score 0

4.2.4.5 Vapor Compression Heat Pump

A vapor compression heat pump 170 could be used to reject heat at higher

temperatures so as to reduce the required radiator area. In general, as the effective

environmental temperature and, correspondingly, the ETCS mass increases, heat pumps

are more likely to offer a mass savings compared with an ETCS using only radiators.

While PLB can operate without a heat pump, an ETCS incorporating a heat pump is

expected to yield a mass savings compared with the PLB baseline. Because the PLB

ETCS is expected to be two-fault tolerant, this study assumes a heat pump rejecting a

nominal load of 16.67 kW for each ETCS loop. Further, to reject a full load of 50 kW

even after a single failure, the actual heat pumps must each have a 25 kW capacity. The

analysis is for a single ETCS loop. Here the heat pumps will operate only while PLB is in

sunlight drawing power from the main PLB power grid. The appropriate power mass

penalty for "daytime only" usage for this scenario is 35.3 kg/kW (Hughes, 1995). The

continuous power mass penalty of 750 kg/kW applies for any continuous use.

The radiators may be oriented in numerous ways. The confgurafions of greatest

interest are the horizontal radiator, as included in the baseline TCS, and a vertical radiator.

Because PLB will have an equatorial site, the vertical radiator is positioned so its radiating

areas face north and south with its length parallel to the lunar equator. Further, because

the radiator rejection temperature with a heat pump will be above 280 K, as Section 4.2.3

indicates, a single set of radiator panels designed for 15 years will be less massive than two

sets of radiator panels designed for 7.5 years each. Thus, the assumed surface properties

are those for 10 mil silver Teflon after 15 years of service.

168 See Section 2.1.4 for background on capillary pumped loops.

169 See Section 4.2.2 for details.

170 See Section 2.2.1 for background on vapor compression heat pumps.

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Horizontal Radiators

A two parameter parametric study (Figure 4.13) indicates that the ETCS mass is

minimized for PLB using horizontal radiators when the heat pump temperature lift is

52.3 K and the radiator fluid mass flowrate is increased to 1.27 kg/s per ETCS loop.

Actually, as Figure 4.14 shows, this heat pump design is constrained by the minimum

radiating area for rejecting a full heat load at night without a heat pump at an average

panel temperature of 275.0 K. Specifically then (for one of three ETCS loops):

Cold Source Temperature, Tc (evaporator temperature) 274.0 K

Temperature Lift, TH - Tc 52.3 K

Hot Source Temperature, TH (condenser temperature) 326.3 K

Radiator Inlet Temperature, Tin 322.3 K

Lunar Surface Solar Irradiation 1.371 kW/m 2

Environmental Temperature, Tspace 3.0 K

Effective Sink Temperature, Tsink 171 265.3 K

Total Nominal Cooling Load per Loop, Qc 16.67 kW

Ideal Coefficient of Performance, COPcarnot 5.24

Heat Pump Efficiency, I"1172(Ewert, 1991) 0.50

Necessary Input Power, Wreal 6.36 kW

The low solar absorptivity, horizontal radiators are simulated by the model developed for

the parametric study.

Emissivity (end-of-life)

Solar Absorptivity (end-of-life)

Fin Efficiency

Radiator Mass Flowrate of Ammonia (per ETCS loop)

Average Ammonia Temperature

Average Panel Surface Temperature

Heat Rejection per Unit Area

Total Heat Rejected per Loop by the Radiators, QH,real

Radiator Mass Per Unit Area

Number of Radiator Panel Sets per Project Life

Radiator Working Fluid Mass Per Unit Area

Necessary Radiator Surface Area During the Day

Necessary Radiator Surface Area at Night

0.83

0.17

0.85

1.17 kg/s

316.3 K

314.8 K

O.318 kW/m 2

23.03 kW

7.50 kg/m 2

1

0.36 kg/m 2

72.44m2

72.42 m 2

Here the radiator design is constrained by the radiating area necessary to reject the

nominal load at night (Figure 4.14). The power may be grouped according to its

applicable power mass penalty.

171 The effective environmental sink temperature is provided here for comparison. These computations

used the solar irradiation and the temperature of space, not the effective sink temperature.

172 Percentage of Carnot coefficient of performance (COP).

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3000

2000

1000

0

3000 -

10 20 30 40 50

Temperature Lift [K]

60

r_

2000

1000

T ¢- 41' m

0

0.50

A

0.75 1.00 1.25 1.50

Radiator Working Fluid Mass Flowrate [kg/s]

Heat Pump Power _" Working Fluid

Heat Pump _, Pumping Power

Radiator "-.- Total

Figure 4.13 Variation of Permanent Lunar Base external thermal control system mass per

loop as a function of the vapor compression heat pump temperature lift and the radiator

loop mass flowrate. This study used horizontal radiators. The baseline mission from

Section 4.2.2 forms the basis for this study. The external thermal control system mass was

optimized with respect to both variables mentioned above.

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Power Mass Penalties:

Heat Pump Power

Evaporator Pumping Power

Condenser Pumping Power

Radiator Pumping Power

Daytime Only

Daytime Only

Daytime Only

Continuous

Mass Gained by Using

a Vapor Compression Heat Pump

PLB ETCS Total PLB

Loop With ETCS With Baseline PLB Change for

Heat Pump Heat Pumps TCS PLB ETCS

Radiator Mass [kg]

Working Fluid [kg]

Heat Pump Mass 173 [kg]

543.3 1629.9 9765.2 -8135.3

25.9 77.7 233.0 -155.3

255.1 765.3 -- +765.3

Radiator, Fluid, and

Heat Pump Mass [kg] 824.3 2472.9 9998.2 -7525.3

Power:

Heat Pump [kWl 6.36

Evaporator Pumping Power [kW] 0.27

Condenser Pumping Power [kW] 0.06

Radiator Pumping Power [kW] 0.12

19.08 --

0.81 --

0.18 --

0.36 1.63

Total Power [kW] 6.81 20.43 1.63

Mass due to Power [kg] 326.2 978.6 471.0 +507.6

Total Mass [kg] 1150.5 3451.5 10469.2 -7017.7

The ETCS line pumping power is unchanged. The power values include penalties for the

pressure drops associated with the condenser and the evaporator. The condenser power

penalty is 50% of the power to pump fluid through the radiators designed for use with the

heat pump (or 0.5 x 0.12 kW). The evaporator power penalty is 50% of the power to

pump fluid through the baseline radiator configuration (or 0.5 x 1.63 kW/3). The radiator

pumping power is proportional to the radiator size, in terms of radiating area, and is based

on values from ISS.

Vertical Radiators

A two-parameter parametric study (Figure 4.15) indicates that the ETCS mass is

minimized for PLB using vertical radiators when the heat pump temperature lift is 122.3 K

and the radiator fluid mass flowrate is increased to 0.88 kg/s per ETCS loop. Again, as

Figure 4.16 shows, this heat pump design is constrained by the minimum radiating area for

rejecting a full heat load at night without a heat pump at an average panel temperature of

275.0 K. Specifically then (for one of three ETCS loops):

Cold Source Temperature, Tc (evaporator temperature)

Temperature Lift, TH - Tc

Hot Source Temperature, TH (condenser temperature)

274.0 K

121.6 K

395.6 K

173 See the footnotes in Section 3.1.5.4 for the heat pump sizing correlation.

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-o

200

180

160

140

120

100

80

60

Radiator Area.

Required at Night __

• " " I " " " I I " " " I I

l0 20 30 40 50

Temperature Lift [K]

60

200

180

°1,,_

160

140

120

100-

801 60 Radiator Area Required at Night

I I I

0.50 0.75 1.00 1.25 1.50

Radiator Working Fluid Mass Flowrate [kg/s]

Figure 4.14 Variation of radiator area for Permanent Lunar Base per external thermal

control system loop using horizontal radiators as a function of the vapor compression heat

pump temperature lift and the radiator loop mass flowrate. This heat pump option is

constrained by the radiator area required for rejection of the active thermal control system

heat load at night without the heat pump operating.

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Radiator Inlet Temperature, Tin

Effective Sink Temperature, Tsink

Total Nominal Cooling Load per Loop, Qc

Ideal Coefficient of Performance, COPcarnot

Heat Pump Efficiency, rl 174(Ewert, 1991)

Necessary Input Power, Wreal

391.6 K

317.1 K

16.67 kW

2.25

0.50

14.80 kW

The low solar absorptivity, vertical radiators are simulated by the model developed for the

parametric study using the appropriate higher effective sink temperature.

Emissivity (end-of-life)

Solar Absorptivity (end-of-life)

Fin Efficiency

Radiator Mass Flowrate of Ammonia (per ETCS loop)

Average Ammonia Temperature

Average Panel Surface Temperature

Heat Rejection per Unit Area

Total Heat Rejected per Loop by the Radiators, QH,real

Radiator Mass Per Unit Area

Number of Radiator Panel Sets per Project Life

Radiator Working Fluid Mass Per Unit Area

Necessary Radiator Surface Area During the Day

Necessary Radiator Surface Area at Night

0.83

0.17

0.85

0.88 kg/s

381.6 K

380.1 K

0.430 kW/m 2

31.46 kW

5.625 kg/m 2

1

0.27 kg/m 2

73.11 m 2

73.07 m 2

Mass Gained by Using

a Vapor Compression Heat Pump

PLB ETCS Total PLB

Loop With ETCS With Baseline PLB Change for

Heat Pump Heat Pump TCS PLB ETCS

Radiator Mass [kg]

Working Fluid [kg]

Heat Pump Mass 175 [kg]

411.2 1233.6 9765.2 -8531.6

19.7 59.1 233.0 -173.9

335.6 1006.8 -- +1006.8

Radiator, Fluid, and Heat Pump

Mass [kg] 766.5 2299.5 9998.2 -7698.7

Power:

Heat Pump [kW] 14.80

Evaporator Pumping Power [kW] 0.27

Condenser Pumping Power [kW] 0.06

Radiator Pumping Power [kW] 0.12

44.40 --

0.81 --

0.18 --

0.36 1.63

Total Power [kW] 15.25 45.75 1.63

Mass due to Power [kg] 624.1 1872.3 471.0 +1401.3

Total Mass [kg] 1390.6 4171.8 10469.2 -6297.4

The ETCS fine pumping power is unchanged. As above, the power values include

penalties for the pressure drops associated with the condenser and the evaporator.

174 Percentage of Carnot coefficient of performance (COP).

175 See the footnotes in Section 3.1.5.4 for the heat pump sizing correlation.

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3000

2000

10130

0

70

1

+ _+..._..__...___+ __+

"A. °

80 90 100 110 120 130

Temperature Lift [K]

1600

1400

1200

1000

800

600

400

200

0

0.0

+ -''".-.-.44444.+ _+

r"l

°'AQ.°

o- .-o-- - _ - - -- -_- .......

t: : _:: _.2..**_**-. - •

.... r . . AAAAAA D _ , _ . _ . .

0.5 1.0 1.5 2.0

Radiator Working Fluid Mass Flowrate [kg/s]

Heat Pump Power - -o- - Heat Pump

... A... Radiator Panel _. A.. _ Working Fluid

-. _--. Pumping Power ---+-- Total

Figure 4.15 Variation of Permanent Lunar Base external thermal control system mass per

loop as a function of the vapor compression heat pump temperature lift and the radiator

loop mass flowrate. This study uses vertical radiators. The external thermal control

system mass was optimized with respect to both variables mentioned above.

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190

170

150

130

110

90

70

50

70

\

b

o

1:1°

o

oo

Radiator Area Required at Night

I .... ]

80 90 100 110

Temperature Lift [KI

i ,

120 130

190

170

150

E

_ 130

_ 110

90

70

50

0.0

I-I

_'['1.

Radiator Area Required at Night

i .... I !

0.5 1.0 1.5 2.0

Radiator Working Fluid Mass Flowrate [kg/s]

Figure 4.16 Variation of radiator area for PLB per external thermal control system loop

using vertical radiators as a function of the vapor compression heat pump temperature lift

and the radiator loop mass flowrate. This heat pump option is constrained by the radiator

area required for rejection of the active thermal control system heat load at night without

the heat pump operating (solid line).

Both designs using vapor compression heat pumps are less massive than the

baseline ETCS configuration. However, the horizontal radiators provided a less massive

system than using vertical radiators. The vertical radiator units are less massive simply

because they have a lower mass per radiating area than the horizontal radiator units.

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However, the vertical radiators partially view the lunar surface while horizontal radiators

view only the Sun and space. The lunar surface radiates its energy in the same frequency

band that the radiator uses to reject heat. This rejection frequency band is associated with

the radiating surface's infrared emissivity. The Sun irradiates at frequencies which are

much greater than the heat-rejection frequency band. The solar irradiation frequency band

is associated with the radiating surface's solar absorptivity. Because the emissivity is large

compared with the solar absorptivity, the overall effective sink temperature for the vertical

radiators is significantly greater than for the horizontal radiators. Therefore, the heat

pumps using vertical radiators use a greater temperature lift which requires more power

input and a larger capacity heat pump. In summary, the configuration using vertical

radiators has lighter radiators while the configuration using horizontal radiators sees a

lower effective sink temperature at lunar noon and uses less power and lower capacity

heat pumps. Thus, assuming the loss of either power generation capacity or heat pump

units, the configuration with horizontal radiators would be the more suitable system

because it uses less power to operate and it experiences the lower effective sink

temperature at lunar noon which allows this system to reject more heat without heat

pumping than the corresponding system using vertical radiators 176

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

Using Horizontal Using Vertical

Radiators: Radiators:

7,525 kg 7,708 kg

- 18.80 kW -44.36 kW

-508 kg -1401 kg

Overall Mass Savings 7,017 kg 6,297 kg

Composite Qualitative Score 177 +2

4.2.4.6 Solar Vapor Compression Heat Pump

As with ISS, the single greatest drawback of a TCS using a vapor compression

heat pump is the power requirement. The configuration in the previous section using

horizontal radiators consumed power equivalent to 41% of the nominal load for PLB.

The configuration using vertical radiators consumed power equivalent to 92% of the

nominal load for PLB. To minimize the mass of this commodity, dedicated photovoltaic

solar arrays can be installed with the heat pumps. This option is called the solar vapor

compression heat pump 178. The appropriate mass penalty for equipment which will

176

177

178

At an effective sink temperature of 265.3 K, which applies for horizontal radiators, heat can be

rejected without a heat pump at a panel surface temperature of 275 K. This is the basis for the

baseline TCS design. However, at an effective sink temperature of 317.1 K, the corresponding value

for vertical radiators, no heat can be rejected without a heat pump. Therefore, on Luna horizontal

radiators display an added factor of safety compared with vertical radiators.

For this application the volume is "compact" compared with the baseline architecture.

Technically, both heat pumps presented for PLB are "solar heat pumps" because they only operate

while PLB is in sunlight and their assumed power mass does not include any mass for power storage.

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require power only during the day is 20.2 kg/kW (Hughes, 1995). The continuous power

mass penalty is still 750 kg/kW.

From an analysis standpoint, this option differs from the previous section only in

the value of the "daytime only" power mass penalty assigned. As such, the equipment

masses, capacities, and set-points are the same for these two options. The masses for the

power, however, will differ. Thus, assuming the configuration from the previous section

using horizontal radiators:

Mass Gained by Using a

Solar Vapor Compression Heat Pump 179

PLB ETCS Loop Total PLB ETCS

With Solar With Solar Baseline PLB Change for PLB

Heat Pump Heat Pump TCS ETCS

Radiator, Fluid, and

Heat Pump Mass [kg] 824.3 2472.9 9998.2 -7525.3

Power:

Heat Pump [kW] 6.36 19.08 --

Evaporator Pumping

Power [kW] 0.27 0.81 --

Condenser Pumping

Power [kW] 0.06 0.18 --

Radiator Pumping Power

[kW] 0.12 0.36 1.63

Total Power [kW] 6.81 20.43 1.63

Mass Due to Power [kg] 225.1 675.3 471.0 +204.3

Total Mass [kg] 1049.4 3148.2 10469.2 -7321.0

The ETCS line pumping power is unchanged. As in the previous section, the power

values include penalties for the pressure drops associated with the condenser and the

evaporator. It is assumed that the pumping power for flow through the heat pump

condensers and evaporators is supplied by the arrays dedicated to the heat pumps.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

7,525 kg

-18.80 kW

-204 kg

Overall Mass Savings 7,321 kg

Composite Qualitative Score +2

4.2.4.7 Complex Compound Heat Pump

A complex compound heat pump 180 requires a source of high temperature heat to

drive its operating cycle. The assumed method within this study uses a collector to focus

179

However, Section 4.2.4.5 assumes power comes from the main PLB power grid while Section 4.2.4.6

assumes power comes from a dedicated solar photovoltaic power array. See also Section 2.2.2.

Please see Section 4.2.4.5 (Horizontal Radiators) for details supporting these results.

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solar radiation on a tube containing a working fluid. This is the method assumed while

updating the work of Ewert (1993) for the complex compound heat pump. Ewert also

includes systems using the vapor compression heat pump with vertical and horizontal

radiators. Because Ewert's work differs from the other heat pump studies presented in

this report, his systems are all presented below.

Overall Values:

Radiator Surface Properties:

Emissivity (end-of-life)

Solar Absorptivity (end-of-life)

Fin Efficiency

Radiator Effective Sink Temperature, Tsink:

Horizontal Radiator

Vertical Radiator

Radiator Mass (dry) per Radiating Area:

Horizontal

Vertical

Heat Source Mass (dry) per Heat Energy Required 181

Total Nominal Cooling Load, Qc

Power Mass Penalty

0.83

0.17

0.85

265 K

317K

7.50 kg/m 2

5.63 kg/m 2

2.4 kg/kW

50 kW

35.3 kg/kW

Cold Source (minimum ITCS) Temperature [K]

Temperature Lift [K]

Hot Source (radiator) Temperature [K]

Heat Input per Cooling Load [kWheat input/kWcooling]

Heat Source Temperature [K]

Peak Power Input per Cooling Load

[kWpower/kWcooling]

Radiator Orientation

Heat Rejection per Cooling Load by Primary Radiator

[kWrejected/kWcooling]

Heat Pump Mass per Cooling Load [kg/kWcooling]

Complex

Compound Vapor Compression

Heat Pump Heat Pumps

282 275 275

78 85 85

360 360 360

3.34 ....

500 ....

-- 0.667 0.667

vertical vertical horizontal

4.34 1.67 1.67

4.7 11.0 11.0

180

181

See Section 2.2.3 for additional background material.

This mass penalty is for a solar collector which will direct solar energy into a working fluid passing

through an enclosed channel. This value is "optimistic" and does not include mass for the piping,

pumps, pumping power, or working fluid needed to transfer the collected energy to a heat-driven

heat pump.

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Radiator Orientation

Heat Rejection per Unit Area [kW/m 2]

Required Radiator Area During the Day [m2]

Radiator Mass (dry) [kg]

Power as Mass [kg]

Heat Pump Mass [kg]

Mass of Solar Collector for Heat Source [kg]

Total Heat Pump System Mass [kg]

For Comparison (from Section 4.2.4.5):

Total System Mass [kg]

Mass Savings Compared to PLB Baseline [kg]

Complex

Compound Vapor Compression

Heat Pump Heat Pumps

vertical vertical horizontal

0.268 0.268 0.475

810 312 176

4556 1750 1317

-- 1177 1177

235 550 550

401 ....

5192 3477 3044

4171 3452

6297 7017

Based on the system masses given above, a complex compound heat pump using the

analytical process from Section 4.2.4.5 would have a total mass of (4171 kg/3477 kg) x

5192 kg, or 6,228 kg. Thus, compared with the mass of the baseline ETCS for PLB of

10,469 kg, the complex compound heat pump will save 4,241 kg.

Specific Assessments:

Overall Mass Savings 182 4,241 kg

Composite Qualitative Score +1

4.2.4.8 Zeolite Heat Pump

Zeolite heat pumps also use a heat-driven cycle and require a high-temperature

heat source 183. Ewert (1993) also examined this system. From the perspective of this

report, zeolite heat pumps have many of the same attributes as complex compound heat

pumps. Therefore, the same overall approach as in the preceding section is employed

here. The previous work of Ewert (1993) may be revised for use here for the zeolite heat

pump by assuming the high-temperature heat source is supplied using a solar energy

focusing collector instead of waste heat from a nuclear power plant. Again, Ewert also

includes systems using the vapor compression heat pump with vertical and horizontal

radiators. Because Ewert's work differs from the other heat pump studies presented in

this report, his systems are all presented below.

Overall Values:

Radiator Surface Properties:

Emissivity (end-of-life)

Solar Absorptivity (end-of-life)

Fin Efficiency

0.83

0.17

0.85

182 Only an overall mass savings is given because this estimate is lacking in detail.

183 See Section 2.2.4 for additional background details.

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Radiator Effective Sink Temperature, Tsink:

Horizontal Radiator

Vertical Radiator

Radiator Mass (dry) per Radiating Area:

Horizontal

Vertical

Heat Source Mass (dry) per Heat Energy Required 184

Total Nominal Cooling Load, Qc

Power Mass Penalty

265 K

317K

7.50 kg/m 2

5.63 kg/m 2

2.4 kg/kW

50 kW

35.3 kg/kW

Cold Source (minimum ITCS) Temperature [K]

Temperature Lift [K]

Hot Source (radiator) Temperature [K]

Heat Input per Cooling Load [kWheat input/kWcooling]

Heat Source Temperature [K]

Peak Power Input per Cooling Load

[kWpower/kWcooling]

Radiator Orientation

Heat Rejection per Cooling Load by Primary Radiator

[kWrejected/kWcooling]

Heat Pump Mass per Cooling Load [kg/kWcooling]

Zeolite Heat Vapor Compression

Pump Heat Pumps

275 275 275

85 85 85

360 360 360

2.5 ....

700 ....

-- 0.667 0.667

vertical vertical horizontal

3.5 1.67 1.67

31.4 11.0 11.0

Radiator Orientation

Heat Rejection per Unit Area [kW/m 2]

Required Radiator Area during the Day [m 2]

Radiator Mass (dry) [kg]

Power as Mass [kg]

Heat Pump Mass [kg]

Mass of Solar Collector for Heat Source [kg]

Total Heat Pump System Mass [kg]

For Comparison (from Section 4.2.4.5):

Total System Mass [kg]

Mass Savings Compared to PLB Baseline [kg]

Zeolite Heat Vapor Compression

Pump Heat Pumps

vertical vertical horizontal

0.268 0.268 0.475

652 312 176

3674 1750 1317

-- 1177 1177

1570 550 550

300 ....

5544 3477 3044

4171 3452

6297 7017

Based on the system masses given above, a zeolite heat pump using the analytical process

from Section4.2.4.5 would have a total mass of (4171 kg/3477 kg) x 5544kg, or

6,651 kg. Compared with the mass of the baseline ETCS for PLB of 10,469 kg, the

zeolite heat pump will save 3,818 kg.

184 This mass penalty is for a solar collector which will direct solar energy into a working fluid passing

through an enclosed channel. This value is "optimistic" and does not include mass for the piping,

pumps, pumping power, or working fluid needed to transfer the collected energy to a heat-driven

heat pump.

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Specific Assessments:

Overall Mass Savings 185 3,818 kg

Composite Qualitative Score +1

4.2.4.9 Lightweight Radiators

A parametric study may be used to identify potential mass savings by using lighter

materials for portions of the PLB radiator assembly. As with FLO Lander, the PLB ETCS

is not well defined, so again the components and categories listed here are out of necessity

vague.

Category Within

Permanent Lunar Base Radiator Assembly 186 Mass [kg] Mass Percentage

Structure and Deployment: 30.0

Base Radiator Support Structure 1041.3

Deployment Mechanism 1952.4

Radiator Panels 6768.4 67.7

Radiator Panel Working Fluid (ammonia) 233.0 2.3

Total 9995.1 100.0

This study varies the component PLB radiator assembly masses linearly based on

the original total mass in each category. The structure and deployment masses are

reduced by up to 20% when the radiator panel mass is reduced by up to 40%. The

radiator panel working fluid volume will be unchanged, so the working fluid mass will be

unchanged 187

The most important underlying assumption is that composites and other advanced

materials are most likely to offer significant mass reduction for only some components of

the radiator panels such as the honeycomb and the facesheets. Because the radiator panels

are the single most massive item of the PLB radiator assembly, such savings would

constrain the remainder of the design. A less significant mass reduction is assumed for the

structures and deployment because the sizes of these components are primarily dictated by

the dimensions of the radiator array and the deployment scheme. However, composites

should allow comparable components to replace some of the structure and deployment

with lighter parts. In summary, this approach yields an overall ETCS mass reduction of

185 Again, only an overall mass savings is given because this estimate is lacking in detail.

186 The radiators here will be changed after 7.5 years as in the baseline case. Further, as a worst case,

the structure and deployment will be replaced at the same time. Thus, the masses include two sets of

radiator panels, two support structures, and two deployment mechanisms. As in the baseline

mission, the radiator working fluid will not require replacement. Here the overall mass per radiating

area is 7.5 kg/m 2 for the dry radiator panel assembly. The assumed component composition is:

5.2 kg/m 2 for radiator panels, 0.8 kg/m 2 for the base support structure, and 1.5 kg/m 2 for

deployment. As presented above, the radiator panel working fluid mass is 0.358 kg/m 2 of radiating

area.

187 See Section 2.4 for additional general background on lightweight radiators plus specific examples of

proposed lightweight radiators.

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33% for a 40% reduction in the radiator panel mass (Figure 4.17). It is further assumed

that, as in the baseline case, PLB will use two sets of radiator panels during the 15 years

of the initial project. As a worst case, the radiator support structure and deployment

mechanisms will be replaced when the radiator panels are replaced after 7.5 years.

Percent Reduction in

Category Within Radiator Panel Mass

Permanent Lunar Base Radiator Assembly 20% 40%

Structure and Deployment [kg] 2694.4 2395.0

Radiator Panels [kg] 5414.7 4061.1

Radiator Working Fluid [kg] 233.0 233.0

Overall Radiator Mass [kg] 8342.1 6689.1

Overall Mass Reduction in Radiators [kg] 1653.0 3306.0

Overall Mass Reduction as a Percentage of the

Baseline Permanent Lunar Base Radiator Assembly 16.5 33.1

Radiator Mass Per Surface Area [kg/m 2] 12.82 10.28

Considering the available lightweight radiators presented in Section 2.4, an overall mass

reduction of 33.1% was selected as a representative value. Thus, the mass savings for

PLB is 3306.0 kg for the reference mission. The power requirements will be unchanged

because the internal fluid dynamics and, therefore, the pumping power are functions of the

working fluid material properties.

Any of the lightweight radiator concepts may be used for a base on the lunar

surface. All of the radiators presented in Section 2.4 can be deployed vertically. The

vertical radiators will require integration with either a radiator shade or a heat pump to

reject life support heat loads at lunar noon away from the polar regions. However, the

composite flow-through radiators could also be oriented horizontally like the radiators in

the baseline architecture.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

1,999 kg

none

Overall Mass Savings 1,999 kg

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10000

8000

6000

4000-

2000 A

/k A

0 0 0

T • V

I I I

0 10 20 30 40

Percent Reduction in Radiator Panel Mass

Radiator Panels T Working Fluid

Support Structure _ Total

A Deployment

Figure 4.17 Permanent Lunar Base radiator mass as a function of mass reduction within

the radiator panel set. Study Assumptions: 1) For each 10% mass reduction in the

radiators the masses for the support structure and the deployment decrease by 5%, and 2)

Permanent Lunar Base will use two panel sets during the initial 15-year project life.

4.2.4.10 Parabolic Radiator Shade

The parabolic radiator shade 188 is an excellent option for lowering the effective

sink temperature that a vertical radiator experiences during lunar noon at an equatorial

site is9. For PLB, the radiator and shade would be deployed on the lunar surface near the

buried habitation and laboratory modules. Alignment of the shade and radiator is not

difficult because the entire system can be deployed under human supervision. The final

system might look something like the lower portion of Figure 4.8 or it might be deployed

188 See Section 2.5.2 for additional background information.

189 For a vertical radiator with the end-of-life surface properties assumed above for the baseline mission

(emissivity of 0.86 and solar absorptivity of 0.14) the effective sink temperature is 319.6 K.

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with the radiator panels in a single line. In either case endsheets (not pictured) should be

included to prevent infrared surface emissions from impinging on the radiator and to

inhibit contamination of the shade by lunar dust.

For the mission using a parabolic radiator shade, a single set of vertical radiators

coated with silver Teflon is assumed. However, the shade and its deployment mechanism

will be replaced once every five years 190. Thus, for a parabolic radiator shade at PLB:

Heat Rejected 50.0 kW

Radiator Surface Temperature 275.0 K

Lunar Surface

Emissivity 0.93

Absorptivity 0.93

Albedo 0.07

Solar Irradiation 1.371 kW/m 2

Incident Angle 90.0 °

Orbital Inclination 1.53 °

Radiator 191

Fin Efficiency 0.85

Infrared Emissivity 0.83

Solar Absorptivity 0.17

Shade (End-of-Life) 192

Specular Upper Surface

Emissivity 0.06

Absorptivity 0.14

Diffuse Lower Surface

Emissivity 0.90

Absorptivity 0.90

Sink Temperature 193 195 K

190 The thermal protection value of the shade is highly sensitive to its surface properties. While the

actual material and its surface coatings are expected to withstand continuous use for the life of PLB,

dust accumulation is a significant concern. It is expected that possible dust accumulation due to

human activities near PLB can be anticipated and controlled. However, dust accumulation due to

other aspects of the lunar environment are currently unknown. Thus, a relatively high rate of

replacement is assumed here for the parabolic shade to ensure its reliability.

191 The radiator is assumed to have a 15-year service life.

192 Keller (1995 a). The shade is assumed to have a 5-year service life.

193 Keller (1995 a) computed a sink temperature of 188.8 K for a vertical radiator with a surface

emissivity of 0.86 and a solar absorptivity of 0.14 using TSS.

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Radiator (with silver Teflon)

Area

Mass Penalty

Total Dry Radiator Panel Mass (one set)

Dimensions

Height

Length

Working Fluid Mass Penalty

Working Fluid Mass

Shade

Parabolic Arc Length 194

Length

Area

Mass Penalty 195

Mass (per parabolic radiator shade)

Shades per Project Life

Total Shade Mass

Total System Mass 196

292.5 m z

5.63 kg/m 2

1645.3 kg

1.0 m

146.3 m

0.269 kg/m 2

78.7 kg

4.59 m

146.3 m

671.5 m z

0.56 kg/m 2

376.0 kg

3

1128.0 kg

2852.0 kg

Comparing the radiator and flexible parabolic shade with the baseline mission presented

above yields:

Mass Gained by Using Baseline PLB PLB TCS

a Parabolic Radiator Shade TCS With Shade Total Change

Radiator Area [m 2]

Radiator Mass per Panel Set [kg]

Panel Sets for Life of PLB

Total Dry Radiator Mass [kg]

Radiator Pressure Drop [kN/m 2]

Radiator Pumping Power [kW]

Radiator Pumping Power Mass [kg]

Radiator Working Fluid [kg]

Shade Area [m 2]

Shade Mass per Shade [kg]

Number of Shades for Life of PLB

Total Shade Mass [kg]

Total Mass [kg]

650.8 292.5

4881.1 1645.3

2 1

9765.2 1645.3 -8119.9

242.0 108.8

1.629 0.733

471.0 200.8 -270.2

233.0 78.7 - 154.3

-- 671.5

-- 376.0

3

-- 1128.0 +1128.0

10469.2 3052.8 -7416.4

194 A full parabolic shade is assumed. In other words, the shade top is even with the radiator top and

the shade focus is at the height of the radiator.

195 This penalty includes mass for a flexible parabolic shade plus appropriate support structure

(Keller, 1994).

196 This value includes mass for one set of radiator panels and three parabolic radiator shades plus one

charge of working fluid. Masses for piping, fittings, and working fluid to and from the radiators are

not included. The pumping power mass is also omitted from this value.

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The overall savings for using a parabolic radiator shade is 7,416.4 kg. Because PLB is a

longer mission, the issue of dust accumulation on the radiator shade, which degrades the

shade's effectiveness, is a significant concern. Further research in this area should indicate

how to maintain a parabolic shade on the lunar surface for extended missions. Several

ideas to either decontaminate or guard the shade are under consideration (Keller, 1995 b).

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

7,146 kg

0.896 kW

270 kg

Overall Mass Savings 7,416 kg

Composite Qualitative Score -1

4.2.4.11 Plant Chamber Cooling Improvements

Regenerative life support systems are of great importance as an enabling

technology for extended-duration missions with crews. In particular, plant growth

chambers promise to provide future space explorers with both food and air revitalization.

Therefore, a plant growth module will be included in PLB. The practical aspect of this

system then becomes optimizing the biomass (food) produced by a plant growth module

as a function of system mass. The baseline plant growth module is in Section 2.6.2 which

provides details for a unit based on half of the Johnson Space Center Variable Pressure

Growth Chamber.

Several design changes are expected to reduce the mass of the plant module over

its baseline configuration. Further, each design change is assumed to be completely

independent of any other design change. As such, the savings for all options here are

assumed to be cumulative. Figure 2.10 provides a general illustration of a portion of the

plant growth chamber.

1) Heaters: With proper management of the air stream dehumidification process, it should be possible to completely eliminate any

reheat. The mass of the heaters (36.5 kg total) plus the associated

power (3.0 kW which is 2250.0 kg as mass) may be eliminated.

Total Savings: 2,286.5 kg

2) Ducting and Vents: The current ducting and vents, which are

fabricated of metal, have a mass of 125.7 kg. Using lighter

materials, such as plastic and composites, it is estimated that this

mass may be cut in half. Total Savings: 62.9 kg

3) Lightbox Barriers: The baseline lightbox barriers between the plant

growth lamps and the plants are tempered glass (4.76 mm thick)

and have a total mass of 63.8 kg. The glass barriers can be

replaced with Teflon FEP fluorocarbon film (0.127 mm thick)

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which removes approximately all of the mass associated with the

lightbox barriers (Ewert, Paul, and Barta, 1995). Total Savings:

4) Coldplate Lamp Cooling: A more efficient method of removing

heat generated by plant growth lamps from the plant chamber is to

mount the lamps on coldplates. It is assumed that 40% of the

lighting heat energy can be removed by the coldplates. Thus, this

energy will not reach the plant chamber atmosphere. To maintain

the same temperature distribution across the plant chamber, 40%

less volumetric flowrate of air is required which corresponds to a

40% savings in blower power. This corresponds to a savings of

570 kg (1.9 kW × 0.4 × 750 kg/kW). The coldplates have a mass

of 41.9 kg each or 167.6 kg total (four coldplates are required).

Total Savings:

5) Ballasts: The lamp ballasts have a total mass of 103.4 kg. Using

lighter materials it is assumed that this mass may be cut in half.

Total Savings:

The overall savings for all of these options is 2,867.3 kg.

63.8 kg

402.4 kg

51.7 kg

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

47 kg

3.76 kW

2,820 kg

Overall Mass Savings 2,867 kg

Composite Qualitative Score +1

4.2.4.12 Carbon Brush Heat Exchanger

It is assumed that the baseline design for PLB will include DDCUs affixed to

coldplates transferring heat to the ETCS working fluid using radiant fin interfaces. As

presented above, radiant fin interfaces are the current standard for power converter units

aboard ISS. Here the same dimensions and masses are assumed for the coldplates

associated with the converter units for PLB. Replacement of the coldplates and the

coldplate sockets is again the reference mission for this technology. According to Hughes

(1995), a representative power grid for PLB will use twelve DDCUs, three main bus

switching units (MBSU), and three dc switching units (DCSU). While these latter two

units are functionally different from DDCUs, they are power conditioning devices which

depend on the same coldplate/radiant-fin structure used by the DDCUs for proper cooling.

Thus, it is assumed here that the MBSUs and DCSUs will use the same coldplate design as

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the DDCUs, and that this coldplate design is equivalent to an external DDCU (DDCU-E)

coldplate design designated for ISS units 197

In order to save mass and improve DDCU cooling efficiency, the radiant fin

interfaces could be replaced by carbon brush interfaces. The equipment mass savings

assuming 18 equivalent DDCU coldplate fin masses are replaced is:

Power Conversion and/or

Conditioning Coldplate

Coldplate Mass Fin Mass (each Total Mass

Quantity [kg] coldplate) [kg] Savings [kg]

DDCU 12 63.5 4.29 102.96

MBSU 3 63.5 4.29 25.74

DCSU 3 63.5 4.29 25.74

The total mass savings is twice the fin mass to account for the corresponding fin set on the

coldplate socket on which the DDCU sits. Aluminum 6061-T6 is the assumed coldplate

material with a density of 2,713 kg/m 3. Further, while it is assumed that the carbon brush

heat exchangers will have negligible mass, the rest of the baseplate and coldplate mass will

remain unchanged.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

154 kg

none directly

Overall Mass Savings 154 kg

Composite Qualitative Score +3

4.2.5 Summary

The various advanced technologies and their estimated benefits are summarized in

the table below for PLB. From Section 4.2.2, the mass of the baseline ETCS is

10,948.3 kg. Assuming the mass determinations throughout this study have associated

uncertainties on the order of 10%, a complete TCS with an advanced technology would

need to show a savings of at least 1,095 kg to ensure a mass savings. Further, because

design and development costs are not a'ivial, a mass savings of 25%, or 2,737 kg, is

desirable. Using these criteria, the TCSs with advanced technologies proposed for PLB

may be divided into five categories:

• TCSs using advanced technologies requiring a mass penalty greater than

10% of the overall baseline ETCS mass: none.

• TCSs using advanced technologies requiring a mass penalty less than 10%

of the overall baseline ETCS mass: none.

197 See Section 2.6.3 for details of the radiant fin design. These assumptions, though simple, should

give reasonable estimates of the fin mass for power conversion and conditioning unit coldplates if

the units are mounted externally. The DDCU-E from ISS was selected because this unit's intended

usage most nearly reflects the mission for which the units here are envisioned. Further, the

DDCU-E is the most numerous DDCU on ISS.

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• TCSs using advanced technologies with a mass savings less than 10% of

the overall baseline ETCS mass: two-phase TCS with MP/S, LP twophase TCS, and capillary pumped loops.

• TCSs using advanced technologies with a mass savings between 10 and

25% of the overall baseline ETCS mass: none.

• TCSs using advanced technologies with a mass savings greater than 25%

of the overall baseline ETCS mass: vapor compression heat pump, solar

vapor compression heat pump, complex compound heat pump, zeolite heat

pump, lightweight radiators, and parabolic radiator shade.

Overall, the baseline PLB ETCS is massive and unwieldy though it uses familiar

and dependable technology. As such, most of the advanced technologies considered

provide a significant mass savings compared to the baseline ETCS. The technology in the

third category will produce an ETCS which is comparable to the baseline system. This

category includes the two-phase TCSs. The technologies in the f'mal category show

significant promise. These include the various heat pumps, the parabolic radiator shade,

and lightweight radiators. Within this category, the two vapor compression heat pumps

and the parabolic radiator shade offer the greatest mass savings at roughly 65%. The two

heat pumps using heat-driven cycles fall into a second tier with mass savings around 40%.

Several technologies addressed for PLB are enhancing technologies. As before,

enhancing technologies are advanced technologies which will uniformly deliver a mass

savings or penalty for a specified reference mission regardless of the type of TCS selected.

These technologies include plant chamber cooling improvements and the carbon brush

heat exchanger. Thus:

• Enhancing technologies which require a mass penalty: none.

• Enhancing technologies which yield a mass savings: plant chamber cooling

improvements and carbon brush heat exchanger.

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Table 4.6 Advanced Active Thermal Control System Architecture

for Permanent Lunar Base

Summary of Advanced Active Thermal Control System

Architecture for Permanent Lunar Base

Overall Mass Qualitative

Savings [kg] Score

4.2.4.1 Two-Phase Thermal Control System

With Mechanical Pump/Separator

Low-Power Two-Phase Thermal Control System

Two-Phase Thermal Control System

With Electrohydrodynamic Pumping unknown +2

4.2.4.4 Capillary Pumped Loops 917 0

4.2.4.5 Vapor Compression Heat Pump 198 7,017 +2

4.2.4.6 Solar Vapor Compression Heat Pump 7,321 +2

4.2.4.7 Complex Compound Heat Pump 4,241 + 1

4.2.4.8 Zeolite Heat Pump 3,818 + 1

4.2.4.9 Lightweight Radiators 3,306 --

4.2.4.10 Parabolic Radiator Shade 7,416 - 1

4.2.4.11 Plant Chamber Cooling Improvements 2,867 +1

4.2.4.12 Carbon Brush Heat Exchanger 154 +3

456 0

751 -1

198 The value here is for a configuration with horizontal radiators. A configuration using vertical

radiators saves 6297 kg.

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4.3 MARS LANDER (ML)

4.3.1 Reference Mission

Mars is, like Luna, a planetary body of extremes. Surface temperatures vary

between 130 K at the height of the martian winter to 300 K during the martian summer

(Reich and Scoon, 1993). The planetary inclination is similar to Earth's at 24.8 degrees.

Further, the martian days correspond to 24.66 hours. However, unlike Earth, the martian

orbit is much more elliptical. At its warmest, during summer in the southern hemisphere,

Mars is 1.381 astronomical units [AU] from the Sun. During the much cooler Northern

summer, Mars reaches 1.666 AU from the Sun (Dzenitis, 1992). Additionally, Mars has a

thin, yet significant, atmosphere comprised of 96% carbon dioxide and 3% nitrogen.

Atmospheric pressures vary between 600 and 1000 N/m 2 (Dzenitis, 1992). As such, Mars

can support wind velocities up to 30 rn/s which permit huge dust storms to spread over

the entire planetary surface during the southern summer (Reich and Scoon, 1993).

For an initial human exploration of Mars a current design philosophy is to use an

expendable craft which, like Apollo, is comprised of a crew transport vehicle and a landing

vehicle. The crew transport vehicle would house the mission personnel while in transit

from Earth orbit to an orbit around Mars. An aero-braking maneuver would be used to

decelerate from interplanetary speeds into a stable martian orbit. The crew would then

descend to an equatorial site on the martian surface in a landing vehicle, Mars Lander

(ML). After a 30-day stay, the crew would return to the crew transport vehicle and return

to Earth. A capsule with an independent heat shield would allow the crew to return

directly to the Earth's surface following an interplanetary voyage from Mars.

This study concentrates on ML while it is situated on the martian surface. The

transport of ML to the martian surface is not considered, nor is the crew transit vehicle

considered. The ML itself will have both habitation and laboratory space for a crew of

four. The overall architecture will probably be similar to FLO Lander except that this

vehicle is not necessarily expected to fit entirely within a standard 10-m launch shroud. In

fact, the combined Mars expedition vehicle will probably be assembled in Earth orbit. The

ML ATCS will be designed to effectively reject a nominal load of 30 kW regardless of

environmental conditions at an equatorial landing site. The assumed continuous power

mass penalty is 111 kg/kW while the "daytime" only power mass penalty is 25 kg/kW.

These power mass penalties are not currently attainable and assume some improvements in

power systems 199. One mission is assumed.

4.3.2 Baseline Case

The baseline ML ETCS uses low solar absorptivity, vertical radiators with singlephase liquid ammonia as the working fluid. For redundancy, three ETCS loops are

provided. Each loop will service two module zones in cascade. Physically, each zone may

be an individual module or a single module may be subdivided into more than one zone.

Each loop will be sized to handle half of the total nominal heat load, 15 kW, so as to

199 For purposes of comparison, the power mass penalties assumed for PLB are based on current

technology or anticipated technology from current research.

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provide complete capacity even after one fault. The TCS utilizes a single-phase cascade

similar to the TCS for ISS. Each ETCS loop functions as both low temperature and

moderate temperature service in series. In short, those areas requiring the lowest

temperature cooling are placed upstream of those items with higher temperature needs.

Radiator bypass valves provide ETCS loop set-point temperature control. Further, each

ML zone has two ITCS loops using liquid water as the working fluid. Additional fittings

will allow either ITCS loop to service all of ML's coldplates and heat exchangers. This

arrangement provides extra flexibility if any one of the ITCS or ETCS loops fail. Each

ETCS loop also has a dedicated pump module (Figure 4.18).

The overall base ATCS heat load is 30 kW. A value of 7.50 kg/m 2 is assumed for

the radiator mass per radiating area for horizontal radiators. Vertical radiators are

assumed to use 5.625 kg/m 2 200. The ML will land at the martian equator which has

surface temperatures ranging from 185 K to 190 K at night to 220 K to 260 K during the

day (Reich and Scoon, 1993). The solar irradiation at noon varies between 0.493 kW/m 2

at aphelion to 0.718 kW/m 2 at perihelion. The martian surface has the properties of a

diffuse gray surface with an emissivity of 0.90 to 0.98 (Dzenitis, 1992).

Overall ETCS Performance:

Heat Rejected

ETCS Working Fluid

30.0 kW

single-phase ammonia

• Low Solar Absorptivity, Vertical Radiators 201.

Surface Emissivity

Solar Absorptivity

Fin Efficiency

Average Radiator Surface Temperature

0.8

0.2

0.85

275.0 K

Heat Rejection per Area (worst/design case)

Radiator Surface Area

Mass Penalty for Dry Radiators 202

Radiator Panel Mass

Radiator Working Fluid Mass Penalty 203

Radiator Working Fluid Mass

Total Radiator Mass (wet) 204

O. 10 kW/m 2

352.94 m z

5.625 kg/m 2

1985.3 kg

0.269 kg/m 2

94.9 kg

2080.2 kg

200 These radiator masses are identical to those assumed for PLB. They are not as frugal as the values

used for a similar vehicle, FLO Lander, but rather reflect a desire to include more robustness.

201 The radiators were sized based on the parametric study given in Section 4.3.3 below.

202 This mass penalty includes structure and deployment mass as well as actual radiator panel mass.

203 This mass penalty estimate is based on data for the working fluid mass used in the ISS radiator

ORUs as a function of radiating area.

204 A comparable horizontal radiator design, based on heat rejection during a dust storm, has 392.59 m 2

of radiating area, with a mass of 2944.4 kg for the radiator plus 140.6 kg for the working fluid.

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Two-Sided, Vertical Radiator

!

ETCS LoopA _ ,. i

Zone I

!i!i ETCS Loop Bj

an I1_ i IIm i

Zone II

ETCS i!i!i

Loop C

i!i_i

Zone HI

Figure 4.18 Plumbing for the baseline Mars Lander external thermal control system.

LTL is low temperature loop and MTL is moderate temperature loop. The system uses

three external thermal control system loops to give two-fault tolerance with single-phase

liquid ammonia as the working fluid. The bypass valves allow for temperature control.

Overall, both the radiator and the volume occupied by the crew are subdivided into three

cooling zones. Each zone may correspond to a separate physical device, such as a

module, or one larger device may be subdivided. Each cooling zone has two water

internal thermal control system loops which are internally interconnected so that any heat

load can be serviced by either loop. A pump module is assumed for each loop.

ETCS Working Fluid Loop Mass Flowrates:

Average ETCS Fluid Temperature

Radiator Outlet Temperature

Radiator Inlet Temperature

Average Ammonia Specific Heat

278.0 K

274.82 K

281.18 K

4.62 (kW*s)/(kg*K)

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Total Mass Flowrate

(full load/equal d _ribution)

Nominal Flowrate per Loop

(full load/equal distribution)

Maximum Flowrate per Loop

1.021 kg/s

0.340 kg/s

0.510 kg/s

• ETCS Pump Module:

Pump Efficiency

Average Ammonia Density

Radiator Pressure Drop per Radiating Area 205

Radiator Pressure Drop

Pumping Power for Radiators

Assumed Effective Line Length per Loop

Number of ETCS Loops

Line Diameter 206

Nominal Pressure Drop Line (per loop)

Nominal Line Pumping Power (per loop)

Overall Pumping Power

Continuous Power Mass Penalty

Pumping Power as Mass

Mass of Ammonia in Lines

0.45

631.96 kg/m 3

0.3718 kN/m 4

131.2 kN/m 2

0.471 kW

40m

3

0.0236 m

14.58 kN/m 2

0.0174 kW

0.523 kW

111 kg/kW

58.1 kg

33.2 kg

• Summary of the Mars Lander Active Thermal Control System:

Heat Load Rejected

Radiator Mass (dry)

Working Fluid Mass

Pumping Power

Pumping Power as Mass

30.0 kW

1985.3 kg

128.1 kg

0.523 kW

58.1 kg

Total ETCS Heat-Rejection System Mass 207 2171.5 kg

The baseline radiator mass per area, including fluids, is 5.89 kg/m 2.

4.3.3 Parametric Study Using the Baseline Case

Dzenitis (1992) presents a package of models for heat transfer with the martian

environment. For the study here, the ML radiators were sized considering only radiation.

The radiation case is, assuming no wind, a worst-case scenario because any wind will

provide some convection cooling which will in turn lower the effective environmental

temperature and aid heat rejection by the radiator. Further, as found by Dzenitis (1992),

205 This penalty is based on the values for ISS which specifies a maximum radiator pressure drop of

48.26 kN/m 2 for 129.8 mz of radiating area.

206 This the inside diameter of the ETCS lines within ISS modules. Specifically this value is 0.93 in.

207 This value does not include masses for the pumps, lines, and fittings, but these are expected to be

roughly the same for all ML ATCS configurations.

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convection cooling is expected to only account for about 10% of the heat rejected by a

radiator in a fairly strong martian surface wind of 10 m/s. Thus, heat rejection from

human spacecraft on Mars will primarily be by radiant heat transfer.

Dzenitis (1992) divides the irradiation in a martian environment for any flat surface

into solar fluxes and infrared fluxes. This model, from Section 2.0 of Dzenitis (1992), was

implemented, with the exception of Equation (2.18) 208, in a spreadsheet program. Heat

rejection results for both vertical and horizontal surfaces were obtained for a typical "hot"

equatorial site 209. Data for this study come primarily from Dzenitis (1992) and Reich and

Scoon (1993). Specifically:

Surface and Orbital Location:

Orbital Position, Ls 210

Surface Latitude, _.

270 degrees

0 degrees

Time:

Local Midnight

Local Sunrise, tsr (also assumed as to)

Local Noon, tnoon

Time of Local High Temperature, th

Local Sunset, tss

Length of Local Day, tday

-6.16 hours

0.00 hours

6.16 hours

7.19 hours

12.33 hours

24.66 hours

Local Environmental Parameters Under Clear Skies (optical depth, 'r, of 0.1):

Solar Absorptivity of Martian Surface, CtG 0.7

Infrared Emissivity of Martian Surface, e.G 0.94

Average Surface Temperature, TG 225 K

Diurnal Surface Temperature Variation, ATG 70 K

Infrared Emissivity of Martian Atmosphere, eA 0.225

Average Atmospheric Temperature, TA 205 K

Diurnal Atmospheric Temperature Variation, ATA 6.05 K

208

209

210

Equation (2.18) is a correction for the optical depth of the atmosphere when the surface of interest is

not horizontal to the mean horizon. This effect, when included in the model, gives questionable

results for vertical surfaces, something Dzenitis noted in his paper. While the effective optical depth

for a vertical surface may differ from that of a horizontal surface, this adjustment should probably be

omitted until more work in this area can be completed.

Environmentally, the hottest site on Mars is located 40 degrees south of the equator during summer

in the southern hemisphere. This corresponds to the time when the most severe dust storms form

beginning within the southern hemisphere (Reich and Scoon, 1993).

The variables listed in this section correspond to the nomenclature employed by Dzenitis (1992).

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Local Environmental Parameters in a Dust Storm (optical depth, x, of 2.8 21 l):

Solar Absorptivity of Martian Surface, 0tG 0.7

Infrared Emissivity of Martian Surface, _3 0.94

Average Surface Temperature, TG 220 K

Diurnal Surface Temperature Variation, ATG 30.25 K

Infrared Emissivity of Martian Atmosphere, eA 0.83

Average Atmospheric Temperature, TA 219 K

Diurnal Atmospheric Temperature Variation, ATA 25.93 K

Radiator Properties:

Surface Infrared Emissivity

Surface Solar Absorptivity

Average Surface Temperature

0.8

0.2

275 K

Here two types of radiators are examined. The first is a single-sided, horizontal

unit mounted on top of ML similar to the radiator for FLO Lander. The second is a twosided, vertical unit which is also assumed to be mounted on top of ML. Table 4.7 and

Table 4.8 below present the radiator heat fluxes as a function of time for a typical clear or

dusty day at a hot equatorial landing site.

Table 4.7 Daytime Heat Rejection From Mars Lander Radiators

for an Equatorial Site on a Clear Day (x = 0.1)

Vertical Radiator Vertical Radiator

Time of Day Horizontal Radiator Facing North/South Facing East/West

[hours] [kW/m 2] [kW/m 2] [kW/m 2]

0.00

(Sunrise)

1.23

2.47

3.70

4.93

(Noon) 6.16

7.40

8.63

9.86

11.10

12.33

(Sunset)

0.2424 0.2231 0.2231

0.2074 0.1907 0.1693

0.1712 0.1710 0.1530

0.1429 0.1491 0.1423

0.1246 0.1259 0.1344

0.1181 0.1082 0.1341

0.1241 0.1042 0.1127

0.1419 0.1091 0.1023

0.1697 0.1189 0.1009

0.2057 0.1350 0.1136

0.2407 0.1719 0.1719

Minimum

[Rejection 0.1181 0.1042 0.1009

211 Reich and Scoon (1993) note that the most severe dust storm observed at a Viking landing site

corresponded to an optical thickness of roughly 3.0. Pollack, Colburn, Flasar, Kahn, Carlston, and

Pidek (1979) list the landing sites for both Viking spacecraft. Viking Lander 1 is at 22.3 degrees

north latitude and 47.9 degrees west longitude, while Viking Lander 2 is at 47.7 degrees north

latitude and 225.7 degrees west longitude.

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Table 4.8 Daytime Heat Rejection From Mars Lander Radiators

for an Equatorial Site During a Dust Storm (x = 2.8)

Vertical Radiator Vertical Radiator

Time of Day Horizontal Radiator Facing North/South Facing East/West

[hours] [kW/m 2] [kW/m 2] [kw/m2]

0.00

(Sunrise)

1.23

2.47

3.70

4.93

(Noon) 6.16

7.40

8.63

9.86

11.10

12.33

(Sunset)

0.1916 0.1879 0.1879

0.1788 0.1784 0.1784

0.1568 0.1606 0.1605

0.1283 0.1373 0.1371

0.1044 0.1167 0.1170

0.0899 0.1033 0.1046

0.0939 0.1042 0.1046

0.1077 0.1129 0.1127

0.1278 0.1266 0.1265

0.1457 0.1398 0.1398

0.1598 0.1509 0.1509

Minimum

Rejection 0.0899 0.1033 0.1046

The fluxes for the vertical radiator facing north and south are the average of values for a

single surface facing north and a second surface facing south. Likewise the fluxes for the

vertical radiator facing east and west are the average of values for a single surface facing

east and a second surface facing west.

The performance of the horizontal radiator has several noteworthy features. In

both environments the horizontal radiator is least effective at rejecting heat at local noon,

which corresponds to the greatest direct solar irradiation load. Further, the horizontal

radiator rejected less heat when the direct solar load was mostly blocked by a severe dust

storm. This reduced performance results from receiving a much higher infrared irradiation

from the atmosphere which, while filled with dust, absorbs a greater percentage of the

direct solar irradiation and then reradiates the energy as infrared radiation. Because the

radiator's surface coatings are designed to most readily transfer heat in the infrared range,

the radiator's heat-rejection capability degrades during a dust storm 212

Vertical radiators displayed a variety of results. Because of ML's position on the

equator during summer in the southern hemisphere, a surface facing north never receives

any direct solar irradiation while a surface facing south receives direct solar irradiation

while the Sun is up. A surface facing east receives direct solar irradiation before local

noon, and a surface facing west receives direct solar irradiation after local noon. The local

high temperature occurs shortly after local noon. This time is assumed in this study to

correspond to 1/24th of a day after local noon. As the results above indicate, a vertical

radiator will reject at least 0.1 kW/m 2 regardless of the radiator's orientation. The radiator

facing north and south was more effective on a clear day, while the radiator facing east

212 Another explanation is that the dust storm blocks a radiator's view of its coldest sink, space. As

such, the radiator's performance decreases.

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and west was slightly more effective during a dust storm. From the individual surface

results (not presented), the heat rejection for a radiator surface not receiving direct solar

irradiation degraded during a dust storm because of an increase in atmospheric infrared

irradiation. However, this degradation was partially offset by an increase in heat rejection

during a dust storm for surfaces normally exposed to higher levels of direct solar

irradiation. The dust storm acted as a shield against the Sun, especially for radiators

facing east and west which normally view both the early morning and late evening

sunlight.

The variation in performance between the vertical and horizontal radiators arises

from the difference in the effective view factor for each surface. The horizontal radiator

views only sky, so its performance is especially susceptible to variations in the atmospheric

infrared irradiation level. However, half of the view of the vertical radiator is the martian

surface, which has a more constant infrared irradiation flux. Further, the infrared flux of

the ground actually decreases during a dust storm because the maximum surface

temperature is not quite as great (Reich and Scoon, 1993).

The radiator performance as a function of surface radiator temperature was also

investigated. The results are presented in tabular form in Table 4.9 through Table 4.12,

and graphically in Figure 4.19 and Figure 4.20. The first figure presents radiator area as a

function of orientation and radiator surface temperature. The second figure presents the

dry radiator mass as a function of the radiator orientation and radiator surface

temperature. The assumed fin efficiency for this study is 1.00.

Table 4.9 Required Daytime Radiator Area

for an Equatorial Site on a Clear Day (x = 0.1)

Radiator Surface

Temperature [K]

Vertical Radiator Vertical Radiator

Horizontal Radiator Facing North/South Facing East/West

[m2] [mz] [m2]

265.0

267.5

270.0

272.5

275.0

277.5

280.0

282.5

285.0

364.1 438.4 460.4

329.8 389.6 406.9

300.7 349.7 363.5

275.7 316.3 327.6

254.0 288.0 297.4

235.0 263.8 271.6

218.2 242.8 249.4

203.3 224.5 230.1

189.9 208.3 213.2

The required radiator area is a strong function of the radiator surface temperature

for the range of temperatures normally associated with the rejection of metabolic heat.

For a radiator facing east and west in a dust storm, the required radiator area decreases by

44% if the radiator temperature increases from 265 K to 275 K. Increasing the

temperature from 275 K to 285 K decreases the required radiator area by a third for the

same radiator orientation. Henson (1995) provides similar results.

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8OO

6OO

200

0

265 270 275 280 285

Radiator Temperature [K]

Clear Sky, x = 0.1

Horizontal Radiator

Vertical Radiator Facing North/South

Vertical Radiator Facing EastY_est

Dusty Sky, x = 2.8

". Horizontal Radiator

¢ Vertical Radiator Facing North/South

Vertical Radiator Facing East/West

Figure 4.19 Mars Lander radiator area as a function of radiator surface temperature for

various radiator orientations and environmental conditions. The assumed fin efficiency is

1.00 for these results.

Table 4.10 Daytime Dry Radiator Mass

for an Equatorial Site on a Clear Day ('_ = 0.1)

Radiator Surface

Temperature [K]

Vertical Radiator

Horizontal Radiator Facing North/South

[kg] [kg]

Vertical Radiator

Facing East/West [kg]

265.0

267.5

270.0

272.5

275.0

277.5

280.0

282.5

285.0

2730.8 2465.9 2589.6

2473.7 2191.7 2288.9

2255.4 1966.8 2044.7

2067.8 1779.1 1842.6

1904.9 1620.1 1672.6

1762.2 1483.9 1527.8

1636.3 1365.9 1403.0

1524.4 1262.8 1294.4

1424.4 1171.9 1199.1

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7000

6000

5000

_-" 4000

r_

3000

.w.._

2000

1000-

.

265 270 275 280 285

Radiator Temperature [K]

Clear Sky, x = 0.1

-----O----- Horizontal Radiator

Vertical Radiator Facing North/South

Vertical Radiator Facing East/West

¢,

.liT

Dusty Sky, x = 2.8

Horizontal Radiator

Vertical Radiator Facing North/South

Vertical Radiator Facing East/West

Figure 4.20 Mars Lander dry radiator mass as a function of radiator surface temperature

for various radiator orientations and environmental conditions. The assumed fin efficiency

is 1.00 for these results. The horizontal radiator uses a mass penalty of 7.50 kg/m 2 while

the vertical radiators use a penalty of 5.625 kg/m 2.

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Table 4.11 Required Daytime Radiator Area

for an Equatorial Site During a Dust Storm (x = 2.8)

Vertical Radiator Vertical Radiator

Radiator Surface Horizontal Radiator Facing North/South Facing East/West

Temperature [K] [m 2] [m 2] [m 2]

265.0

267.5

270.0

272.5

275.0

277.5

280.0

282.5

285.0

843.4 724.6 647.8

679.8 600.5 546.7

566.7 510.5 471.1

483.9 442.4 412.5

420.8 389.0 365.7

371.0 346.1 327.5

330.8 310.8 295.8

297.7 281.4 269.0

269.9 256.5 246.1

Table 4.12 Daytime Dry Radiator Mass

for an Equatorial Site During a Dust Storm (x = 2.8)

Vertical Radiator

Radiator Surface Horizontal Radiator Facing North/South Vertical Radiator

Temperature [K] [kg] [kg] Facing East/West [kg]

265.0

267.5

270.0

272.5

275.0

277.5

280.0

282.5

285.0

6325.1 4076.1 3643.7

5098.1 3377.6 3075.2

4250.2 2871.5 2650.0

3629.5 2488.2 2320.2

3155.8 2188.0 2057.0

2782.6 1946.6 1842.2

2481.1 1748.5 1663.8

2232.7 1583.0 1513.2

2024.5 1442.7 1384.6

4.3.4 Advanced ATCS Architecture for Mars Lander

ML is designed for a single 30-day stay on the martian surface at an equatorial site.

The baseline mission assumes a vertical, double-sided radiator with single-phase ammonia

as the working fluid. Each advanced architecture is, where applicable, assessed

numerically for the overall mass savings when compared with the baseline design.

Qualitative assessments for these advanced technologies are presented in Section 2.0.

4.3.4.1 Low-Power Two-Phase Thermal Control System

To size a LP two-phase TCS 213 for ML this study extrapolates predictions from

Ungar (1995) which is based on work for space stations. Ungar (1995) includes a small

space station which, of the cases presented, is closest in size to the TCS for ML. Ungar's

small station with a single-phase cascade TCS compares well with the baseline TCS for

ML. Ungar (1995) also presents a comparable LP two-phase TCS for the small station.

213 See Section 2.1.2 for additional material on LP two-phase TCSs.

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Pump Power [kW]

Radiator Area [m 2]

Loop Set-Point [K]

ML Total

Baseline ML

TCS Baseline

(1 loop) TCS

0.174 0.523

117.6 352.9

274.8 274.8

Small Space Station (Ungar, 1995)

Single-Phase Cascade

Thermal Control System

LTL LTL

0.320 0.320

197 197

275.2 275.2

Low-Power Two-Phase

Thermal Control System

LTL MTL

0.068 0.080

195 195

275.2 287.2

LTL refers to the low temperature loop while MTL refers to the moderate temperature

loop. The baseline TCS for ML uses three single-phase cascade, low temperature loops.

Upon comparing the ML baseline TCS with the single-phase cascade TCS for the small

station several things are apparent. The ML TCS uses about 45% less pumping power

than the small station with a single-phase cascade TCS. Further, the ETCS loop set-point

temperatures are about the same for the two systems. A comparable LP two-phase TCS

for ML then should use only about 55% of the pumping power listed in Ungar (1995) for

a small station. Thus, the pumping power for a LTL in a LP two-phase TCS for ML

would use 0.037 kW, while a corresponding MTL loop would use 0.044 kW. A complete

LP two-phase TCS for ML will use two LTL loops and one MTL loop with a total ETCS

pumping power requirement of 0.118 kW. Based on Ungar (1995), the radiator area is

more dependent on the heat load rejected than on the TCS employed, so the radiator area

for ML with a LP two-phase TCS will be similar to that for the baseline TCS. Thus, LP

two-phase TCS will lead to a reduction in pumping power.

Mass Saved by Using a Low-Power

Two-Phase Thermal Control System

Baseline Mars LP Two-Phase Total Change

Lander TCS TCS for ML for TCS

Pump Power [kW] 0.523 0.118

Mass due to

Pump Power [kg] 58.1 13.1 45.0

Finally, the LP two-phase TCS uses smaller fluid lines than the baseline single-phase

cascade TCS which translates into an additional mass savings. However, these savings are

not included here.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

negligible

0.405 kW

45 kg

Overall Mass Savings 45 kg

Composite Qualitative Score -1

4.3.4.2 Two-Phase Thermal Control System With Electrohydrodynamic Pumping

Based on the discussion in Section 2.1.3, ML could utilize a two-phase TCS with

electrohydrodynamic pumping. A two-phase TCS with electrohydrodynamic pumping

may yield a slight mass savings over a LP two-phase TCS. A significant mass savings

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over the baseline TCS would also be expected. However, without more data on this

technology, no mass savings estimates may be determined.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

unknown

unknown

Overall Mass Savings unknown

Composite Qualitative Score +2

4.3.4.3 Capillary Pumped Loops

Based on the baseline architecture given in Section 4.3.2, a capillary pumped

loop 214 could save 0.523 kW, which is the estimated ETCS pumping power for ML.

While it is possible that a single capillary pumped loop might be employed, that does not

affect the mass estimates for this study because the capillary pumped loop equipment mass

is assumed to be comparable to the equipment mass for the baseline single-phase TCS

architecture. Thus, the overall savings for this option is 58.1 kg based on saving power

for pumping.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

none

0.523 kW

58 kg

Overall Mass Savings 58 kg

Composite Qualitative Score

4.3.4.4 Vapor Compression Heat Pump

In contrast to Luna, Mars offers a somewhat less extreme thermal environment.

The days are only slightly longer than terrestrial days and the diurnal temperature

variations are less extreme than on Luna. However, using a heat pump to increase the

radiator temperature is still a reasonable idea. This section examines using a vapor

compression heat pump 215 continuously throughout the martian day and night.

As in the baseline mission, the ML ETCS here is two-fault tolerant and each ETCS

loop has its own heat pump. During nominal operation, each heat pump handles one third

of the nominal load, or 10 kW. However, each heat pump is sized to handle a load 50%

greater than the nominal load, or 15 kW, to allow ML to still reject its full heat load after

losing one heat pump. The analysis uses a single ETCS loop for a basis.

A two parameter parametric study assuming a vertical radiator (Figure 4.21)

indicates that the ETCS mass is minimized for ML when the heat pump temperature lift is

214 See Section 2.1.4 for background on capillary pumped loops.

215 Section 2.2.1 provides additional background on vapor compression heat pumps.

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26.0 K and the radiator fluid mass flowrate is increased to 1.13 kg/s per ETCS loop.

Specifically then (for one of three ETCS loops):

Cold Source Temperature, Tc (evaporator temperature)

Temperature Lift, TH - Tc

Hot Source Temperature, TH (condenser temperature)

Radiator Inlet Temperature, Tin

Environmental Temperature, Tsink 216

Total Cooling Load per Loop, Qc

Ideal Coefficient of Performance, COPcarnot

Heat Pump Efficiency, rl 217 (Ewert, 1991)

Necessary Input Power, Wreal

274.0 K

26.0 K

300.0 K

296.0 K

243.14 K

10.0 kW

10.54

0.50

1.90 kW

The vertical radiators are simulated by the model developed for the parametric study.

Vertical radiators, compared to comparable horizontal units, have better heat rejection

under their worst-case conditions 218

Emissivity

Solar Absorptivity

Fin Efficiency

Radiator Mass Flowrate of Ammonia (per ETCS loop)

Average Ammonia Temperature

Average Panel Surface Temperature

Heat Rejection per Unit Area

Total Heat Rejected per Loop by the Radiators, QH,,_d

Radiator Mass Per Unit Area

Radiator Working Fluid Mass Per Unit Area

Necessary Radiator Surface Area during the Day

0.80

0.20

0.85

1.13 kg/s

294.4 K

292.9 K

0.149 kW/m 2

11.90 kW

5.625 kg/m 2

0.269 kg/m 2

79.90 m E

216 The sink temperature is computed for the poorest heat rejection by a vertical radiator facing east and

west on a completely clear day. Using the poorest heat rejection for a horizontal radiator, which

occurs during a dust storm, the corresponding effective sink temperature is 247.23 K.

217 Percentage of Carnot coefficient of performance (COP).

218 This occurs for a vertical radiator facing east and west on a clear day during the late afternoon.

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1000

800

600

400

200

0

1000

0 10 20 30 40 50

Temperature Li_ [K]

800

600

400

200

O--------O

0 .... i ---- - -7 .... , .... T- -'_.vJ ....

0.00 0.25 0.50 0.75 1.00 1.25 1.50

Radiator Loop Mass Flowrate [kg/s]

Heat Pump Power _ Working Fluid

Heat Pump • Pumping Power

A Radiator _ Total

Figure 4.21 Variation of the Mars Lander ETCS mass per ETCS loop as a function of

the vapor compression heat pump temperature lift and the radiator loop mass flowrate.

The baseline mission from Section 4.3.2 forms the basis for this study. The ETCS mass

was optimized with respect to both variables mentioned above.

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Mass Gained by Using Mars Lander Total Mars Total Baseline Change for

a Vapor Compression ETCS Loop With Lander ETCS Mars Lander Mars Lander

Heat Pump Heat Pump With Heat Pumps TCS ETCS

Radiator Mass [kg]

Working Fluid [kg]

Heat Pump Mass 219 [kg]

449.4 1348.2 1985.3 -637.1

21.5 64.5 94.9 -30.4

141.7 425.1 -- +425.1

Radiator, Fluid, and

Heat Pump Mass [kg] 612.6 1837.8 2080.2 -242.4

Power:

Heat Pump [kW] 1.90 5.70

Evaporator Pumping

Power [kW] 0.08 0.24

Condenser Pumping

Power [kW] 0.06 0.18

Radiator Pumping

Power [kW] 0.12 0.36 0.47

Total Power [kW] 2.16 6.48 0.47

Mass due to Power [kg] 239.8 719.4 52.2 +667.2

Total Mass [kg] 852.4 2557.2 2132.4 +424.8

The ETCS line pumping power is unchanged. The power values include penalties for the

pressure drops associated with the condenser and the evaporator. The condenser power

penalty is 50% of the power to pump fluid through the radiators designed for use with the

heat pump (or 0.5 x 0.12 kW). The evaporator power penalty is 50% of the power to

pump fluid through the baseline radiator configuration (or 0.5 x 0.47 kW/3). As the

tabulation above indicates, a continuously operating vapor compression heat pump will be

more massive than the baseline ETCS for ML.

Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

242 kg

-6.01 kW

-667 kg

Overall Mass Savings -425 kg

Composite Qualitative Score +1

4.3.4.5 Solar Vapor Compression Heat Pump

The single greatest drawback of a TCS continuously using a vapor compression

heat pump is the power requirement. For the vapor compression heat pump presented in

the previous section, the power accounts for about a quarter of the total ETCS mass. On

Mars it is only necessary to operate a vapor compression heat pump while the vehicle or

219 See the footnotes in Section 3.1.5.4 for the heat pump sizing correlation.

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habitat is in sunlight. Heat can be rejected at night using only radiators. Thus, to reduce

overall system mass for ML using a vapor compression heat pump, the heat pump is

bypassed after the Sun sets to allow the ETCS working fluid to flow directly to the

radiators. During the day, the heat pump radiator loop operates normally with the radiator

working fluid flowing only between the heat pump condenser and the radiators. Here the

heat pumps will only operate while sunlight is available using dedicated photovoltaic solar

arrays and the assumed power mass penalty is 25 kg/kW for energy consumed only in this

mode 220. As above, the continuous power mass penalty is 111 kg/kW.

A two parameter parametric study (Figure 4.22) indicates that the ETCS mass per

loop is minimized for this option when the heat pump temperature lift is 28.1 K and the

radiator fluid mass flowrate is increased to 0.82 kg/s per ETCS loop. The controlling

limitation is the radiator area required while the heat pump is not available, which is

78.0 m 2 per loop. This is the radiating area required to reject the nominal heat load at

sunset under the least favorable environmental conditions. (See Section 4.3.3) Therefore:

Cold Source Temperature, Tc (evaporator temperature)

Temperature Lift, TH - Tc

Hot Source Temperature, TH (condenser temperature)

Radiator Inlet Temperature, Tin

Environmental Temperature, Tsink

Total Cooling Load per Loop, Qc

Ideal Coefficient of Performance, COPcarnot

Heat Pump Efficiency, rl 221 (Ewert, 1991)

Necessary Input Power, Wreal

274.0 K

28.1 K

302.1 K

298.1 K

243.14 K

10.0 kW

9.75

0.50

2.05 kW

The vertical radiators are simulated by the model developed for the parametric study.

Vertical radiators, compared to comparable horizontal units, have better heat rejection

under their worst-case conditions 222

Emissivity

Solar Absorptivity

Fin Efficiency

Radiator Mass Flowrate of Ammonia (per ETCS loop)

Average Ammonia Temperature

Average Panel Surface Temperature

Heat Rejection per Unit Area

0.80

0.20

0.85

0.82 kg/s

295.8 K

294.3 K

0.154 kW/m 2

220 This study has not addressed the issue of whether sufficient power can be generated by a photovoltaic

solar array if a dust storm is present. This is significant because midday on Mars is thermally the

hottest diurnal environment even when a dust storm is present. As such, if heat pumping is not

available, the radiators alone may not be able to reject the full heat load by themselves during the

entire Martian day. See Section 2.2.2 for additional information on solar vapor compression heat

pumps.

221 Percentage of Carnot coefficient of performance (COP).

222 This occurs for a vertical radiator facing east and west on a clear day during the late afternoon.

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800

6OO

e:_

_ 400-

200-

°

0 10 20 30 40

Temperature Lift [K]

50

r_

t_

800-

600

400

200

A

0

0.00 0.25 0.50 0.75 1.00 1.25

Radiator Loop Mass Flowrate [kg/s]

Heat Pump Power --- Working Fluid

"----'<>--- Heat Pump # Pumping Power

A Radiator _ Total

1.50

Figure 4.22 Variation of the Mars Lander external thermal control system mass per loop

as a function of the solar vapor compression heat pump temperature lift and the radiator

loop mass flowrate. The baseline mission from Section 4.3.2 forms the basis for this

study. The external thermal control system mass was optimized with respect to both

variables mentioned above.

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Total Heat Rejected per Loop by the Radiators, QH,real

Radiator Mass Per Unit Area

Radiator Working Fluid Mass Per Unit Area

Necessary Radiator Surface Area during the Day (per loop)

Necessary Radiator Surface Area at Night (per loop)

12.05 kW

5.625 kg/m 2

0.269 kg/m 2

78.05 m 2

78.oo m2

Here the radiator design is constrained by the radiating area necessary to reject the

nominal load at sunset (Figure 4.23). The power may be grouped according to its

applicable power mass penalty.

Power Mass Penalties:

Heat Pump Power

Evaporator Pumping Power

Condenser Pumping Power

Radiator Pumping Power

Daytime Only

Daytime Only

Daytime Only

Continuous

Vlass Gained by Using

a Solar Vapor

Compression Heat Pump

Mars Lander Total ML ETCS Total Baseline Change for

ETCS Loop With With Solar Heat Mars Lander Mars Lander

Solar Heat Pump Pump TCS ETCS

Radiator Mass [kg]

Working Fluid [kg]

Heat Pump Mass 223 [kg]

439.0 1317.0 1985.3 -668.3

21.0 63.0 94.9 -31.9

144.3 432.9 -- +432.9

Radiator, Fluid, and

Heat Pump Mass [kg] 604.3 1812.9 2080.2 -267.3

Power:

Heat Pump [kW] 2.05 6.15 --

Evaporator Pumping

Power [kW] 0.08 0.24 --

Condenser Pumping

Power [kW] 0.04 0.12 --

Radiator Pumping

Power [kW] 0.09 0.27 0.47

Total Power [kW] 2.26 6.78 0.47

Mass due to Power [kg] 64.2 192.6 52.2 +140.4

Total Mass [kg] 668.5 2005.5 2132.4 - 126.9

The ETCS line pumping power is unchanged. As in the previous section, the power

values include penalties for the pressure drops associated with the condenser and the

evaporator.

As anticipated, the heat pump using solar power generation is less massive than a

continually operating heat pump. Further, the reduction in mass also allows the solar heat

pump to be less massive than the baseline ETCS configuration.

223 See the footnotes for Section 3.1.5.4 for the heat pump sizing correlation.

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L.

2

o_

120.

100-

80-

60-

40

0

Raditator

Required at Night

I I I I

10 20 30 40 50

Temperature Lift [K]

120 -

L

°,,,,4

100-

80-

60-

40 .

0.00

Radiator Area

Required at Night

" I .... I .... I .... I .... I " " "

0.25 0.50 0.75 1.00 1.25 1.50

Radiator Loop Mass Flowrate [kg/s]

Figure 4.23 Variation of radiator area per ETCS loop for the Mars Lander ETCS as a

function of the solar vapor compression heat pump temperature lift and the radiator loop

mass flowrate. The solar heat pump option is constrained by the radiator area required for

rejection of the ATCS heat load at night without the heat pump operating.

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Specific Assessments:

Equipment Mass Savings

Power Savings

Power Savings as Mass

267 kg

-6.31 kW

-140 kg

Overall Mass Savings 127 kg

Composite Qualitative Score +2

4.3.4.6 Lightweight Radiators

To identify potential mass savings from using lighter materials for portions of the

ML radiator assembly, a parametric study is presented below. As with several previous

vehicles, the ML ETCS is in an early design stage and is not well defined, so the

components and categories listed here are estimates.

Category Within Mars Lander Radiator Assembly 224 Mass [kg] Mass Percentage

Structure and Deployment: 29.3

Base Radiator Support Structure 211.8

Deployment Mechanism 397.1

Radiator Panels 1376.4 66.1

Radiator Panel Working Fluid (ammonia) 94.9 4.6

Total 2080.2 100.0

This study varies the component ML radiator assembly masses linearly based on

the original total mass in each category. The structure and deployment masses are

reduced by up to 20% when the radiator panel mass is reduced by up to 40%. The

radiator panel working fluid volume will be unchanged, so the working fluid mass will also

be unchanged 225

The most important underlying assumption is that composites and other advanced

materials are most likely to offer a significant mass reduction for only some components of

the radiator panels such as the honeycomb and the facesheets. Because the radiator panels

are the single most massive item of the ML radiator assembly, such savings would

constrain the remainder of the design. A less significant mass reduction is assumed for the

structures and deployment because the sizes of these components are primarily dictated by

the dimensions of the radiator array and the deployment scheme. However, composites

should allow comparable components to replace some of the structure and deployment

with lighter parts.

224 The overall mass per radiating area excluding fluids is 5.625 kg/m 2 for the dry radiator panel

assembly. The assumed component composition is: 3.9 kg/m 2 for radiator panels, 0.6 kg/m / for the

base support structure, and 1.125 kg/m 2 for deployment. As presented above, the radiator panel

working fluid mass is 0.269 kg/m 2 of radiating area.

225 See Section 2.4 for additional general background on lightweight radiators plus specific examples of

proposed lightweight radiators.

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Z_

2000

1500

looO

5OO

0

0 0 0 0 0 0 0

v • • v

I I I

10 20 30

Percent Reduction in Radiator Panel Mass

Radiator Panels _' Working Fluid

Deployment _ Total

Support Structure

40

Figure 4.24 ML radiator mass as a function of mass reduction within the radiator panel

set. Study Assumption: For each 10% mass reduction in the radiators the masses for the

support structure and the deployment decrease by 5%.

Percent Reduction in

Radiator Panel Mass

Category Within Mars Lander Radiator Assembly 20% 40%

Structure and Deployment [kg] 548.0 487.1

Radiator Panels [kg] 1l01.1 825.8

Radiator Working Fluid [kg] 94.9 94.9

Overall Radiator Mass [kg] 1744.0 1407.8

Overall Mass Reduction in Radiators [kg] 336.2 672.4

Mass Reduction as a Percentage

of the Original ML Radiator Assembly 16.2 32.3

Radiator Mass Per Surface Area [kg/m 2] 4.94 3.99

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Considering the available lightweight radiators presented in Section 2.4, an overall mass

reduction of 20.0% was selected as a representative value. Thus, the mass savings for ML

is 416.0 kg for the reference mission (Figure 4.24). The power requirements will be

unchanged because the internal fluid dynamics and, therefore, the pumping power are

functions of the working fluid material properties.

Because ML is a planetary mission, any of the lightweight radiator concepts may

be used. All of the radiators presented in Section 2.4 can be deployed vertically. Unlike

the lunar missions, vertical radiators may be used without additional equipment. Again,

the composite flow-through radiators could also be oriented horizontally.

Specific Assessments:

Equipment Mass Savings

Power Savings as Mass

416 kg

none

Overall Mass Savings 416 kg

4.3.5 Summary

The various advanced technologies and their estimated benefits are summarized in

the table below for ML. From Section 4.3.2, the mass of the baseline ETCS is 2,171.5 kg.

Again, assuming the mass determinations throughout this study have associated

uncertainties on the order of 10%, a complete TCS with an advanced technology would

need to show a savings of at least 217 kg to ensure a mass savings. Further, because

design and development costs are not trivial, a mass savings of 25%, or 543 kg, is

desirable. Using these criteria, the TCSs with advanced technologies proposed for ML

may be divided into five categories:

• TCSs using advanced technologies requiting a mass penalty greater than

10% of the overall baseline ETCS mass: vapor compression heat pump.

• TCSs using advanced technologies requiring a mass penalty less than 10%

of the overall baseline ETCS mass: none.

• TCSs using advanced technologies with a mass savings less than 10% of

the overall baseline ETCS mass: LP two-phase TCS, capillary pumped

loops, and solar vapor compression heat pump.

• TCSs using advanced technologies with a mass savings between 10 and

25% of the overall baseline ETCS mass: lightweight radiators.

• TCSs using advanced technologies with a mass savings greater than 25%

of the overall baseline ETCS mass: none.

A continuously operated vapor compression heat pump is not appropriate for ML.

The technologies in the third category will produce an ETCS which is comparable to the

baseline ETCS. The technology in the fourth category is promising for this mission.

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Table 4.13 Advanced Active Thermal Control System Architecture for Mars Lander

Summary of Advanced Active Thermal Control System

Architecture for Mars Lander

Overall Mass Qualitative

Savings [kg] Score

4.3.4.1 Low-Power Two-Phase Thermal Control System

4.3.4.2 Two-Phase Thermal Control System

With Electrohydrodynamic Pumping

4.3.4.3 Capillary Pumped Loops

4.3.4.4 Vapor Compression Heat Pump

4.3.4.5 Solar Vapor Compression Heat Pump

4.3.4.6 Lightweight Radiators

45 -1

unknown +2

58 0

-425 +1

127 +2

416 --

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5.0 CONCLUSIONS

Each section above provides conclusions within the summary highlighting the most

promising set of thermal control technologies for each reference mission. This section

presents conclusions concerning the mass savings, or general appropriateness, of each

advanced thermal technology across the spectrum of missions examined in this study.

Table 5.1 summarizes the mass savings for each technology as it applies to the various

missions.

Table 5.1 Mass Savings [kg] of Advanced Technologies

Applied to Reference Missions

Reference Vehicles and Habitats

ISS STS FLO PLB ML

Baseline ETCS Architecture Masses 226 15,568 201,306 2,101 10,948 2,172

Advanced Technologies

Two Phase Thermal Control Systems (TCSs):

1. Two-Phase TCS With Mechanical Pump/Separator 1,203

2. Low-Power Two-Phase TCS 1,366

3. Two-Phase TCS with Electrohydrodynamic Pumping

4. Capillary Pumped Loops 1,535

8,176

456

286 751 45

unknown unknown unknown

370 917 58

Heat Pumps:

5. Vapor Compression Heat Pump

6. Solar Vapor Compression Heat Pump

7. Complex Compound Heat Pump

8. Zeolite Heat Pump

-250

5,021

7,017 -425

259 7,321 127

4,241

3,818

Heat Pipe Radiators:

9. Arterial Heat Pipe Radiators -154

10. Axial-Groove Heat Pipe Radiators -309

11. Arterial Heat Pipes With Electrohydrodynamic Pumping 490

Lightweight Radiators 227.

12. Composite Flow-Through Radiators

13. Composite Heat Pipe Radiators

14. Composite Reflux Boiler Tube Radiators

15. Unfurlable Radiators

1,661 22,904 584 3,306 416

X X X X X

X X X X

X X X

X X X

Other Heat Rejection Technologies:

16. Phase-Change Thermal Storage

17. Parabolic Radiator Shade

11,875

642 7,416

Additional Technologies:

18. Rotary Fluid Coupler

19. Plant Chamber Cooling Improvements

20. Carbon Brush Heat Exchanger

444

2,867

118 154

226

227

See reference missions in Sections 3.0 and 4.0 for details.

The listed mass savings comes from the generic lightweight radiator assessment while the

lightweight radiator technologies which appear appropriate are marked with an 'x'. Actual

lightweight radiators may offer a greater or lesser mass savings than the figure listed here.

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Qualitative assessments for these technologies are presented with discussion in Section 2.0

and the assessments are summarized in Table 2.1.

From the information above, the radiator masses per radiating area for the

individual reference mission architectures are summarized in Table 5.2. These values are

based on the total radiator mass, including any working fluid, divided by the radiating

surface area.

Table 5.2 Summary of Masses per Radiating Area for Reference Mission Baseline

Thermal Control System Architectures

Reference Mission

Mass per Effective Mass

Radiating Area Fin per Radiating

[kg/m 2] Efficiency Area 228 [kg/m 2]

3.1.2 International Space Station Evolution

3.2.2 Space Transportation System Upgrade

(6-Panel Configuration) 229

Space Transportation System Upgrade

(8-Panel Configuration)

4.1.2 First Lunar Outpost Lander

4.2.2 Permanent Lunar Base 230

4.3.2 Mars Lander

8.51 0.88 9.67

5.01

5.30

7.37 0.85 8.67

7.86 0.85 9.25

5.89 0.85 6.93

5.1 General

Over the spectrum of missions presented above, solar vapor compression heat

pumps and lightweight radiators are the most generally applicable advanced thermal

control technologies. The vapor compression heat pump suffers from high power

consumption. Offsetting this drawback, vapor compression heat pumps drastically reduce

the required thermal control radiator area for thermal environments where rejection of life

support heat loads would normally be difficult. The solar vapor compression heat pump

concept attempts to deal with the compressor's power requirement using a systems

engineering approach to integrate the heat pump with its power source in one package. In

addition to comparing favorably in terms of mass to the baseline architecture, the solar

vapor compression heat pump is applicable to a wide range of duties, including both

planetary and orbital missions. Likewise, lightweight radiators also apply to a wide range

of planetary and orbital missions. Specific lightweight radiators may be mission-specific or

mission-type-specific. For example, composite reflux boiler tube radiators require a

gravitational force to operate and are inappropriate for orbital missions.

The qualitative assessments are independent of the quantified mass savings. The

qualitative assessments represent a conglomeration of expert opinions dealing with

volume, ease of deployment, reliability, development cost, and terrestrial use potential of

the various technologies. Several technologies which did not offer large mass savings

228 The effective mass per radiating area is the mass per radiating area divided by the fin efficiency.

229 Jaax (1978) provides the effective radiator area, not the actual radiator area.

230 These values reflect a single set of radiator panels. For the entire 15-year reference mission,

assuming two sets of radiator panels, the mass per radiating area is 15.36 kg/m 2.

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scored high in the qualitative assessments. These include the rotary fluid coupler, the

carbon brush heat exchanger, and electrohydrodynamic pumping. Thus, these assessments

indicate that these technologies may offer other benefits and, therefore, reasons for

developing them other than their potential for saving mass.

5.2 Two-Phase Thermal Control Systems

Two-phase thermal control systems offer greater promise for larger systems with

long flow lines. With the exception of International Space Station and Permanent Lunar

Base, the vehicles here are too small to make adequate use of this technology.

Additionally, retrofitting International Space Station with a two-phase thermal control

system may be difficult. For Permanent Lunar Base, however, a two-phase thermal

control system may be a wise investment, especially if the base will be expanded sometime

in the future. As flow lines lengthen and heat loads increase, the mass saved by a twophase thermal control system could be significant. Capillary pumped loops provide a twophase thermal control architecture which can be economical even for smaller vehicles.

The only major disadvantage of this specific two-phase thermal control system is the

relatively low pressure rise generated by capillary forces. As such, care is required when

designing a capillary pumped loop for use on a planetary surface and this technology is

probably inappropriate for a thermal control system while a vehicle is accelerating during a

launch or re-entry maneuver.

5.3 Heat Pumps

In general, heat pumps are extremely versatile. They depend on elevating waste

heat to relatively high temperatures. High rejection temperatures in turn reduce the

thermal control system's sensitivity to radiator surface properties by increasing the driving

temperature difference between the radiator and the environment. Further, heat pumps are

not sensitive to planetary surface conditions except as those conditions affect the radiator

surface properties. Design of heat pumps for use in microgravity environments is an

unsettled issue.

Vapor compression heat pumps suffer from high power consumption. However,

provided with a sufficiently inexpensive power source, such as a dedicated solar

photovoltaic power array, vapor compression heat pumps offer one of the least massive

options for many thermal control systems.

For some missions, heat-driven heat pumps also appear to offer mass savings

compared with the baseline architectures using only metal flow-through radiators.

Coupled with a source of high-temperature waste heat, such as from an industrial process,

they may provide economical cooling. However, they are not as efficient as the vapor

compression heat pumps.

5.4 Heat Pipe Radiators

Due to their extra mass, metal heat pipes appear to be justified for human missions

only when thermal control system flow loop punctures from external debris are expected

to be a serious problem. When a mission flies in such an environment, mass savings

become secondary to the reliability of the heat-rejection system. Longer heat pipe units

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exhibit lower mass per radiating area ratios than shorter units. As such, longer heat pipes

are preferred. Mass predictions for systems augmented with electrohydrodynamic pumps

are lower than for systems which are not augmented because the pumps increase the heat

pipe's transport limit over that derived from capillary forces alone. Current predictions for

the missions here do not consider flow loop punctures from micrometeroids and on-orbit

debris to be a significant problem except possibly for the International Space Station

evolution mission. For International Space Station evolution, the thermal control system

mass assuming metallic heat pipe radiators will be comparable to that of the baseline flowthrough radiator technology. Additionally, the operational benefit of greater resistance to

thermal control flow loop punctures is gained with this architecture. Table 5.3 presents

the mass per radiating area for the arterial heat pipe concepts along with a similar value

from the baseline architecture for International Space Station evolution.

Table 5.3 Summary of Masses per Radiating Area

for Arterial Heat Pipe Radiators

Mass per Effective Mass

Radiating Area Fin per Radiating

[kg/m 2] Efficiency Area [kg/m 2]

3.1.2 International Space Station Evolution 8.51 0.88 9.67

2.3 Heat Pipe Radiators

2.3.1 13.11-m Arterial Heat Pipe Radiators

6.71-m Arterial Heat Pipe Radiators

2.3.2 Arterial Heat Pipe Radiators

With Electrohydrodynamic Pumping

8.80 0.925 9.51

9.88 0.925 10.68

7.64 0.763 10.01

5.5 Lightweight Radiators

Lightweight radiators, though not rigorously defined here, offer the hope of

directly reducing the radiator mass, which is the single greatest mass within the external

thermal control system. Additionally, lightweight flow-through radiators can be used

regardless of the other external thermal control system elements.

To add a more concrete basis to this approach, four different lightweight radiator

concepts are presented under this category. These include composite flow-through

radiators and composite heat pipe radiators which may be applied for either planetary or

orbital missions. Composite reflux boiler tube radiators and unfurlable radiators both

depend on working fluid density differences for internal heat transport and therefore

require gravity to operate. Table 5.4 presents the mass per radiating area for each of these

concepts along with the corresponding values from the baseline architectures of the

reference missions for comparison.

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Table 5.4 Summary of Masses per Radiating Area

for Lightweight Radiators

Mass per Effective Mass

Radiating Area Fin per Radiating

[kg/m 2] Efficiency Area [kg/m 2]

Reference Missions

3.1.2 International Space Station Evolution 8.51 0.88 9.67

3.2.2 Space Transportation System Upgrade

(6-Panel Configuration) 231 5.01

Space Transportation System Upgrade

(8-Panel Configuration) 5.30

4.1.2 First Lunar Outpost Lander 7.37 0.85 8.67

4.2.2 Permanent Lunar Base 7.86 0.85 9.25

4.3.2 Mars Lander 5.89 0.85 6.93

2.4

2.4.1

Lightweight Radiators

Composite Flow-Through Radiators

(Single-Sided Rejection)

Composite Flow-Through Radiators

(Double-Sided Rejection) 4.67 ~ 1.00 4.67

2.4.2 Composite Reflux Boiler Tube Radiators 3.44 1.00 3.44

2.4.3 Composite Heat Pipe Radiators 232 1.91 0.97 1.96

2.4.4 Unfurlable Radiators 4.62 1.00 4.62

8.21 -1.00 8.21

5.6 Other Heat Rejection Technologies

The technologies listed under this category have more limited applicability than

some of the other technologies presented above. Phase-change thermal storage is useful

for orbital missions where the time during which the vehicle must endure its maximum

heat load is less than the orbital period. This allows the phase-change thermal storage

device to provide supplemental cooling while the vehicle is under its maximum heat

loading and to solidify the phase-change material while the vehicle experiences its

minimum heat loading.

The parabolic radiator shade operates well on airless planetary bodies with small

planetary inclinations. For a lunar mission, these devices offer the greatest mass savings of

the technologies analyzed. While a parabolic shade could be adapted for other

environments, this would require additional equipment, including a rigid shade assembly

and tracking equipment. For a planetary body with an atmosphere, additional equipment

or techniques would be needed to keep dust and other debris off of the shade surface.

231 Jaax (1978) provides the effective radiator area, not the actual radiator area.

232 The fin efficiency listed for the composite heat pipe radiator is estimated based on an analysis using

the Thermal Synthesizer System (TSS) and SINDA/FLUINT. The actual value is unknown.

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5.7 Additional Technologies

The heat transfer technologies presented under this heading are all enhancing

technologies. While individually they are not usually expected to yield large mass savings

for their respective missions, the listed mass savings is independent of the architecture of

the remainder of the thermal control system. Thus, these technologies are extremely

useful in a supporting role to reduce the overall mass of the thermal control system.

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6.0 RECOMMENDATIONS

While this study is a good start, it falls to answer all questions one might have

about advanced thermal control system technology. Some answers are unknowable using

the preliminary analysis format presented here. However, other areas could be more

adequately addressed by additional study. The recommendations which follow focus on

additional analysis which could extend and improve the current study without relying on

extensive input from future research and development. As technologies mature and data

become available, the study estimates could be updated and refined. Further, additional

reference mission scenarios could be added and examined as future missions are

contemplated.

6.1 Additional Mission Scenarios

• Human Lunar Return - Consider a small lunar habitat with room for a crew of

two to four and a very limited mission length of 3 to 14 days.

• Permanent Lunar Base at a Lunar Pole - Consider Permanent Lunar Base

located at a lunar pole instead of an equatorial site. Some scientific missions, such

as a telescope located on Luna, would prefer a polar site to an equatorial site

(Burke, 1989). Further, finding water on Luna, if it exists, is deemed more likely

at one of the poles. Such a discovery could reduce the mass of supplies sent from

Earth for human operations.

• Extended Presence on Mars - Consider a base at an equatorial latitude on the

martian surface built up over several missions from remotely placed vehicles. A

single thermal control system will be used for both the transit and surface phases of

the mission. A common thermal control architecture should be employed for all

vehicles.

• Generic Transfer Vehicles 233,

- Interplanetary Vehicle for Transfer to Luna - Consider a vehicle with a

capacity of three or four crew members for a transfer mission from Earth to

Luna.

- Orbit to Surface Transfer Vehicle for Luna - Consider a vehicle for the

transfer of two to three crew members to the surface of Luna and back again

from lunar orbit.

- Combined Transfer Vehicle to Luna - Consider a vehicle which would carry

a crew of three or four to the surface of Luna from Earth orbit and back again.

- Interplanetary Vehicle for Transfer to Mars - Consider a vehicle with a

capacity of three or four crew members for a transfer mission from Earth to

Mars.

233 A basic study of thermal control technologies for transfer vehicles might identify heat rejection

technologies which are more generally applicable to not only the surface or orbital stages of a

mission, but also to the transit stages.

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6.2

- Orbit to Surface Transfer Vehicle for Mars - Consider a vehicle for the

transfer of two to three crew members to the surface of Mars and back again

from martian orbit.

- Combined Transfer Vehicle to Mars - Consider a vehicle which would carry

a crew of three or four to the surface of Mars from Earth orbit and back again.

Additional Thermal Control Technology Studies

General:

- Perform conceptual design studies for aluminum, flow-through radiators to

determine more accurately the associated equipment mass per radiating area

for various radiator deployments. In particular, look at vertical and horizontal

radiator deployments on planetary surfaces such as Luna and Mars.

Two-Phase Thermal Control Systems:

- Determine the variation in mass savings as a function of assumed power mass

penalty for two-phase thermal control systems. Because the two-phase

thermal control system's greatest overall benefit is to reduce thermal control

system power requirements, the mass savings for these systems are dependent

on the price of power for a mission.

- Conduct a more in-depth analysis for the two-phase thermal control systems

similar to the analysis in Ungar (1995) instead of the more simplistic approach

used. These studies, like Ungar (1995), would assume and analyze a base or

vehicle architecture including flow line lengths, mass flowrates, and pumping

power. This, in turn, would permit estimates for the ETCS lines and fittings.

Heat Pumps:

- Determine the variation in mass savings as a function of assumed power mass

penalty for the vapor compression heat pumps.

- Develop estimates based on more in-depth heat pump models.

- Consider distributed heat pumping versus centralized heat pumping.

Lightweight Radiators:

- As more detailed information becomes available for individual lightweight

radiator concepts, assess these concepts individually, when appropriate, for

each of the reference missions.

Combining Technologies:

- Combine a parabolic radiator shade with a solar vapor compression heat pump

and determine the least massive system. This technology would apply to

missions on the lunar surface.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

6.3

- Combine a parabolic radiator shade with a solar vapor compression heat pump

using lightweight radiators in place of vertical flow-through radiators and

determine the least massive system. This technology would apply to missions

on the lunar surface.

- Combine the carbon brush heat exchanger and arterial heat pipes with

electrohydrodynamic pumping.

- Combine the carbon brush heat exchanger with composite heat pipes.

Additional Mission Specific Work for Current Study

International Space Station Evolution:

- Determine the variation in mass savings for various advanced heat-rejection

technologies as a function of assumed environmental temperature.

- Conduct a transient analysis for the option using a solar vapor compression

heat pump. Check to ensure that adequate heat rejection can be maintained

while ISS is within the planetary shadow.

- Add phase-change thermal storage.

Lunar Planetary Missions:

- Determine the thermal control system mass as a function of radiator surface

temperature.

- Determine the thermal control system mass as a function of latitude on the

lunar surface.

- Consider the effect of vertical radiators in place of the horizontal radiators.

- Measure the frequency and distribution of micrometeoroid impacts and dust

accumulation on the lunar surface.

Permanent Lunar Base:

- Conduct an in-depth parametric study for a system using a solar vapor

compression heat pump. Consider vertical and horizontal radiators and vary

both the surface properties and the assumed radiator mass per radiating area.

- Determine the variation in heat pump mass as a function of heat pump

efficiency and power mass penalty.

- Determine the variation in thermal control system mass as a function of surface

latitude. Use a combined system with a parabolic radiator shade and a solar

vapor compression heat pump.

• Mars Lander:

Replace the assumed external and internal thermal control systems with a

single thermal control loop using a non-toxic working fluid. Possible working

fluids might include air, carbon dioxide, or water. The system might use a

Brayton gas cycle heat pump or a standard flow loop.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

- Consider the case where the radiator rejection temperature is allowed to rise to

that assumed for First Lunar Outpost Lander.

- Measure the frequency and distribution of micrometeoroid impacts and dust

accumulation on the martian surface.

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Advanced Active Thermal Control Systems Architecture Study NASA TM 104822

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196

Form Approved

REPORT DOCUMENTATION PAGE OMB No. 0704-0188

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1. AGENCY USE ONLY (Leave Blank) 2. REPORT DATE

October 1996

4. TITLE AND SUBTITLE

Advanced Active Thermal Control Systems Architecture Study

6. AUTHOR(S)

Michael K. Ewert; Anthony J. Hanford*

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

Lyndon B. Johnson Space Center

Crew and Thermal Systems Division

Houston, Texas 77058

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, D. C. 20546-0001

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NASA Technical Memorandum

5. FUNDING NUMBERS

8. PERFORMING ORGANIZATION

REPORT NUMBERS

TM-104822

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

S-817

11. SUPPLEMENTARY NOTES

*Lockheed Martin Engineering and Science Services, Houston, Texas

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13. ABSTRACT (Maximum 200 words)

The Johnson Space Center (JSC) initiated a dynamic study to determine possible improvements available through advanced

technologies (not used on previous or current human vehicles), identify promising development initiatives for advanced active

thermal control systems (ATCSs), and help prioritize funding and personnel distribution among many research projects by

providing a common basis to compare several diverse technologies. Some technologies included were two-phase thermal control

systems, light-weight radiators, phase-change thermal storage, rotary fluid coupler, and heat pumps. JSC designed the study to

estimate potential benefits from these various proposed and under-development thermal control technologies for five possible

human missions early in the next century. The study compared all the technologies to a baseline mission using mass as a basis.

Each baseline mission assumed an internal thermal control system; an external thermal control system; and aluminum,

flow-through radiators.

=Solar vapor compression heat pumps and light-weight radiators showed the greatest promise as general advanced thermal

technologies which can be applied across a range of missions. This initial study identified several other promising ATCS

technologies which offer mass savings and other savings compared to traditional thermal control systems. Because the study

format compares various architectures with a commonly defined baseline, it is versatile and expandable, and is expected to be

updated as needed.

14. SUBJECT TERMS

two-phase systems, heat pumps, heat pipes, heat radiators, radiators, heat exchangers,

missions, lunar bases, space transportation system, thermal analysis

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OF REPORT

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216

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