NASA Contractor Report 178208

(_ASA-CR-]782G8) A_ ADVANCE._ _EC_JGLOGY

SPACE S_ATIGM FOR IBE YEAR 20_.5, STUDX A_D

C£NCEPTS Contractor _eport, Eay-_ov.: 1986

(Eionetics Corp.) 191 p CSC_L 22B

N87-2 0340

Unclas

G3/15 _5406

An Advanced Technology Space Station

for the Year 2025, Study and Concepts

M.J.Queijo, A.J.Buttorfleld, W.F.Cuddihy, C.B.King,

and P.A.Garn

The Bionetics Corporation

Hampton, VA. 23666

Contract NA61-18267

March 1987

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23665
TABLE OF O0_

5'_oooooOoooooooooooooooooooooo°°°

AB_qATiONS ............................

ooooooooooooooooooooooooooeoooo

1.0 INTRCD[K_ION ..........................

2.0 ADVANO_ T_ SPACE STATION CAPABILITY REDUIREMENTS ....

3.0 SPACE STATION SJBSYSTSMS, TECSN[XEGY ASSESSM_TgS, ASD _IS .

4.0 _JBSY_ STATE-OF-THE-ART, T_ TRENDS, AM) CAM)IDATE

ADVANCID SCBSYSTERS .......................

4.1

4.1.1

4.1.2

4.1.3

Electrical Power Subsystem ..............

State-of-the-Art Electrical Power Subsystem ......

EPS Technological Trends . .

Candidate _vanced Electrical _er _tJ _ _ [ _ _

4.2

4.2.1

4.2.2

4.2.3

Guidance, Navigation, and Control Subsystem ......

State-of-the-Art ...................

Technology Trends ...................

Candidate Advanced Subsystem .............

4.3

4.3 .i

4.3.2

4.3.3

Structures and Mechanisms Subsystems . • •

State-of-the-Art

Advanced Space Station .......... Materials Techn ogy

°°

Candidate _dvanced Structures and Materials Subsystem .

4.4

4.4.1

4.4.2

4.4.3

On-board Propulsion Subsystem .............

State-of-the-Art Review of Status _d •T_ol _ for Prop_s ion .....

Subsystem Candidates • • •

Candidate Propulsion S_t_'for'_v_c_ _ce

Station and Reccmmenc]ations ..............

4.5

4.5.1

4.5.2

4.5.3

Thermal Control Subsystem ...............

State-of-the-Art ...................

Technology Trends ...................

Candidate _vanced Subsystem .............

viii

ix

I

5

7

ii

15

15

15

16

28

33

33

37

39

39

39

42

47

51

52

58

76

80

80

83

86

i
TABLE OF OONT_qI_ (Concluded)

4.6

4.6.1

4.6.2

4.6.3

4.7

4.7.1

4.7.2

4.7.3

4.8

4.8.1

4.8.2

4.8.3

4.9

4.9.1

4.9.2

4.9.3

Environmental Control and Life Support Subsystem . . .

State-of-the-Art ..........

Technology Trends ..........

Candidate _dvanced Subsystem ....

Extravehicular Activity Subsystem . .

State-of-the-Art ..........

Technology Trends ..........

Candidate Advanced Subsystem ....

Communications and Tracking Subsystem

State-of-the-Art ..........

Technology Trends ..........

Recon_ended Subsystem ........

Data Management Subsystem ......

State-of-the-Art ..........

Technology Trends ..........

Candidate Advanced Subsystem ....

5.0 S/BSYSTEM SYNERSIES FOR SPACE STATION ..............

5.1

5.2

Assessment of Synergies ................

Operating Synergy from O2-H 2 On-board Generation

by Water Electrolysis .................

6.0 R]VANCED TECHNCXEGY SPACE STATION OO_IGURATION CONCEPT .....

6.1

6.2

6.3

6.4

6.5

Approach ...................

Estimates of Power Requirements and Crew Size . [ [ [ [

Considerations for Artiflcial Gravity and Space

Station Orientation . . .

Sizing the Space Station _ _ _ _ _ _ _ _ [ _ _ _ [ _ _

Space Station Configuration ..............

7.0 _ TOPICS FOR _ RESEARCH ..............

_Soooooooooooooooooooooooooooooo

APPEI_IX ...............................

94

94

i00

i00

106

107

IIi

113

113

115

120

125

128

128

135

135

137

137

140

150

150

151

151

153

155

163

165

169

ii
LIST OF TABLES

TAIK,E 2.0-1 STATISTICS FROM NASA MISSION MODEL .......... 8

TABLE 4.1.1-1 _ SPACE STATION ELECI_C PfRER _ ...... 18

TABLE 4.1.2-1 POWER SYSTEMS TECHNOLOGY FO_ .......... 23

TA_E 4.1.3-1 PERSONNEL A_D POWER RECE_ ........... 29

TABLE 4.2.1-1 SPACE STATION GN&C EQUIPMENT LIST ........... 36

TAI_E 4.2.2-1 NAVIGATION, _3IDANCE, A_D CONTRfL TEf_ FORECASTS. 38

TABLE 4.3.2-1 DAMAGE MECHANISMS FROM THE SPACE E_ ..... 45

TABLE 4.3.2-2 ASSEMBLY AIDS AND HUMAN INVOLVEM_2_ .......... 49

TABLE 4.4.1-i SPACE STATION BAS_.INE _ON SYSTEM UTILIZING

MDNOPROPELLANT HYDRAZINE, AT A SPECIFIC IMPULSE

OF 220 SEC ...................... 53

TABLE 4.4.2-1 SUMMARY OF _IIP/E PROPULSION _ FOR

ADVANCED SPACE STATION APPLICATION .......... 59

TA_K,E 4.4.2-2 PROIWY-SION SYSTEM RA_INS CRITERIA .......... 61

TABLE 4.4.2-3 C0MPARISON OF ELBCgRICAL PROPULSION SYSTEMS ...... 69

TABLE 4.5.1-1 S24MARY OF PRfkTECTIONS ................ 82

TABLE 4.5.2-1 THEI_MAL O0NTRfL TECHNOLOGY FORECASPS ......... 85

TABLE 4.6.1-1 IOC SPACE STATION _ OF TE_ STA_--_JSFOR

THE ECLSS FUNCTIONS A_D EQUIPMENT ........... 95

TABLE 4.6.1-2 ECLSS DESIGN IDADS .................. 102

TABLE 4.6.1-3 ATMOSPHERE A_D WATER R__ ........... 104

TABLE 4.7 .I-i CHARACTERISTICS OF T_E EVA SUBSYST_4 ......... 109

TABLE 4.8.1-1 _CATION LINK CHARACTERISTICS RE_tTIVE TO

THE SPACE STATION ................... 117

TABLE 4.8.1-2 SPACE STATION TRACKING LINK CHARACTERISTICS ...... 118

iii
LIST OF TABLES (Concluded)

TABLE 4.9.2-I ON-BOAFD PROCESSING TE_EE_GY FORECASTS .......

TABLE 5.2.5-1

TABLE 6.2-1

ASSESSMENT OF DAn.Y WATER P_DUIR_NTS

FOR ON-BOArD _ION OF O2-H 2 ...........

ESTIMATED ELECTRICAL POWER Pa_UIR_4ENTS FOR THE

SPACE STATION FUNCTIONS LISTED IN SECTION 2.0 .....

TA_E 6.5-1 ADVANCED T_ SPACE RATION,

_OFPERTINENTFEA_JRES .............

TABLE A-I MIRRORS FOR SPACE OPTICS ...............

TABLE A-2 MIRRORS A_D SUB_ ...............

TABLE A-3 PHYSICAL PROPERTIES OF MIRROR SUBSTRATE MATERIALS

AT RDOM T_4_ ..................

136

147

152

158

171

171

172

iv
LIST OF FIGURES

Figure 3.2-1 Baseline Configuration Dual-Keel Space Station,

Principal Features ..................

Figure 3.3-1 Advanced Technology Space Station Concept,

Principal Features ..................

Figure 4.1.1-1 Simplified Features of the EPS Architecture ......

Figure 4.1.2-1 Specific Mass and Area of Several Space

Station Power b-_]stems.................

Figure 4.1.2-2 Specific Mass and Area of Several Space

Station Power Systems .................

Figure 4.1.2-3 Total System Mass for Potential Space

Station Power Systems .................

Figure 4.1.2-4 Initial Launch Mass and 10-Year Propulsion Requirements

for Several Space Station Power systems ........

Figure 4.1.3-1 Point Design Calculations ...............

Figure 4.1.3-2 Temperature and Area Versus Efficiency ........

Figure 4.1.3-3 Blackbody Radiation ..................

Figure 4.2.1-1 Space Station GN&C Architecture ............

Figure 4.3.1-1 Erectable Strut Attachment Method ...........

Figure 4.3.1-2 Preloaded Erectable Strut Joint--Exploded View ....

Figure 4.3.1-3 Mobile Remote Manipulator System ...........

Figure 4.3.1-4 Platform Attachment Method ..............

Figure 4.3.2-1 Space Environment Lifetime of Resin Matrix Composites .

Figure 4.3.2-2 Maximum Usable Non-Metallic Service Temperature ....

Figure 4.3.2-3 Specific Stiffness, E/p , Versus Thermal

Deformation Resistance, KT/_L, for Spacecraft

Structural Materials .................

12

14

17

24

25

26

27

31

32

32

34

41

41

43

43

46

46

48

.v
LIST OF FIGURES(Continued)

Figure 4.3.2-4 Structural Weight Reduction for Space Systems ..... 48

Figure 4.3.2-5 Space Spider Crane .................. 50

Figure 4.3.2-6 Mobile Rail System Concept .............. 50

Figure 4.4.1-1 Reference Configuration Propulsion System with

Approximate Bquipment Locations ............ 55

Figure 4.5.2-1 Radiator Specific Mass and Specific Power ....... 84

Figure 4.5.3-1 Basic Heat Pipe .................... 87

Figure 4.5.3-2 Schematic of Capillary Pumped Loop Engineering Model . 87

Figure 4.5.3-3 Monogroove Heat Pipe Configuration ........... 89

Figure 4.5.3-4 Typical Variable Conductance Heat Pipe ........ 89

Figure 4.5.3-5 Circuit Board Heat Pipe ................ 90

Figure 4.5.3-6 Liquid Droplet Radiator ................ 90

Figure 4.5.3-7 Extrapolated End-of-Life Absorptance ......... 92

Figure 4.5.3-8 Fairchild and Northrop Louver Assembly Schematic . . . 92

Figure 4.5.3-9 Multilayer Insulation Performance Data ........ 93

Figure 4.6.1-1 Space Station ECLSS Architecture ........... 101

Figure 4.7.1-1 Extravehicular Mobility Unit Components ........ 108

Figure 4.7.1-2 Manned Maneuvering Unit ................ Ii0

Figure 4.7.1-3 Astronaut in _4U and MMU ............... 110

Figure 4.7.2-1 EVA Cycle Time .................... 112

Figure 4.7.2-2 EVA Bubble Concept .................. 114

Figure 4.7.2-3 EVA Hard Suit Concept ................. 114

Figure 4.8-1 Atmospheric Absorption for Microwaves ......... 116

Figure 4.9.1-1 DMS Configuration .................... 130

Figure 4.9.1-2 DMS Functional Operations ............... 131

vi
LIST Ce FIGtU_S (Concluded)

Figure 5.2.3-1 Mass Balance for 60-Ferson Crew Using Wet Air

Oxidation Reclamation and O2-_ 2 Generation ......

Figure 6.3-1

Figure 6.5-1

Artificial Gravity Parameters .............

A_vanced Technology Space Station Concept, Features

as Described in Table 6.5-1, Solar Facing Side ....

Figure 6.5-2 Advanced Technology Space Station Concept, Features

as Described in Table 6.5-1, Side View ........

Figure A-I Sources of S_acecraft Contamination ..........

143

154

156

157

177

vii
ABSTRACT

A survey has been made of potential Space Station missions that might

exist in the 2020-2030 time period. Also, a brief study of the current

state-of-the-art of the major subsystems was undertaken, and trends in

technologies that could impact the subsystems were reviewed. The results of

the survey and study were then used to arrive at a conceptual design of a

Space Station for the time frame mentioned above. Factors taken into

account in the conceptual design included the possible requirements for

artificial gravity, synergies between subsystems, and the use of robotics.

Suggestions are made relative to more in-depth Studies concerning the

conceptual design and alternative configurations.

viii
ABBREVIATIONS

AC

ACD

ACS

"AR

BCD

BR

C8C

CFU

CMG

CPS

DC

DMS

DOD

DOF

E

ECLSS

EDP

EMU

EOL

EPS

EVA

FEP

FOV

Alternating Current

Architectural Control Document

Atmosphere Pressure and Composition Control

Atmosphere Revitalization

Bridge Control Unit

Bridge (Data)

Closed Brayton Cycle

Colony Forming Units

Control-mument

Control Point System

Direct Current

Data Management System

Department of Defense

Degree of Freedom

Modulus of Elasticity

Environmental Control and Life Support System

Embedded Data Processor

Extravehicular Mobility Unit

End of Life

 oxy

Electrical Power System

Extravehicular Activity

Fluorinated Ethylene Propylene

Field of View

ix
ABBREVIATIONS (Continued)

GC_

GE0

GN&C

_GE

IrEAS

I/0

IOC

IPACS

IR

I_J

ISA

Isp

LEO

IR

MBPS

_M

MFR

Microrad

Mipe

MMC

_RU

MSD

Gas Chromatograph/Mass Spectrometer

Geosynchronous Equatorial Orbit

Guidance, Navigation, and Control

Graphite

Gateway (Data)

Integrated Design & Evaluation of A_vanced Spacecraft

InputlOutput:

Initial Operational Capebility

Integrated Power and Control System

Infrared

Inertial Reference Unit

Inertial Sensor Assembly

Specific I_se

Low Earth Orbit

Lunar Module

Megabyte

Megabits Per Second

Multiplexer/De-multiplexer

Mobile Foot Restraint

Microradian

Million Instructions per Second

Metal Matrix Composite

Manned Maneuvering Unit

Module

x
ABBREVIATIONS (Continued)

Mops

MPD

MPAC

Mrad

MR_

MSU

NASA

NIU

NM

NOS

OAO

OHV

ORC

ORU

06

OTV

PDCA

PLSS

PROP

PV

0

RCA

RCS

RF

Million Operations per Secor_

Magnetoplam_ic

Multi-Purpose Apglication Console

Milliradian

Mobile Remote Manipulator System

Mass Storage Unit

National Aeronautics and Space A_mlnistration

Network Interface Unit

Nautical Miles

Network Operating System

Orbiting Astronomical Observatory

Orbital Maneuvering Vehicle

Organic Rankine Cycle

Orbital l_=placeable Unit

Operating System

Orbital Transfer Vehicle

Power Distribution Control Assembly

Portable Life Support System

Propulsion

Photovoltatc

Heat Flux

Radio Corporation of America

Rsaction Control System

Radio Frequency

xi
ABBREg"IATIONS(Concluded)

RFC

SAB

SAT

SD

SDP

SOTA

SPAR

S/S

SSE

TDAS

TDRSS

TGS

THC

W

_0

P

6

Regenerative Fuel Cell

Spacecraft Analysis Branch

Satellite

Solar Dynamic

Standard Data Processor

State-of-the-Art

Space Power Advanced Reactor

Space Station

Subsystem Sensor and Effector

High Temperature

Low Temperature

Tracking and Data Acquisition Satellite

Tracking and Data Relay Satellite System

Time Generation System

Temperature and Hmnidity Control (Module)

View Factor

Wet Air Oxidation

Density

Emissivity

xii
AN _ _ SPACE STATION

FOR THE YEAR 2025, SILDY AND CONCEPTS

M. J. Queijo, A. O. Butterfield, W. F. Cuddihy,

C. B. King, and P. A. Gain

The Bionetics Corporation

."*heBionetics Corporation was tasked by the NASA Langley Spacecraft

Analysis Branch to perform a study, consisting of four parts, relating to an

advanced technology Space Station for the time period around 2025. The four

parts were:

i. Examine the capabilities (functions) that a Space Station

might be required to perform.

2. Review the Space Station major subsystems to determine stateof-the-art and the predicted trends in subsystaa technology.

3. Synthesize candidate advanced subsystem configurations.

4. Create an integrated Space Station configuration using

advanced subsystem configurations.

The first part of the study _ms a review of literature concerned with

proposed and suggested future space missions. The review resulted in a list

of seventeen possible functions which could be performed on or by a manned

Space Station in support of the missions:

I. A permanent observatory to look down upon the earth and out

into the universe

1
2. An orbiting science, medical, materials, and new technologies

laboratory

3. A service and repair facility for payloads, spacecraft, and

platforms

4. An assembly facility where large structures or spacecraft

components are manufactured and/or assembled and checked out

5. A transportation node where payloads and vehicles are

collected, stationed, processed, and launched and where fuel

is manufactured

6. A safe habitat for space crews

7. A conmunications and/or relay station for manned or ummnned

spacecraft

8. An adaptation area (in variable "g") in preparation for long

space flights

9. A storage node for food, fuel, spare parts, etc.

I0. A variable "g" research facility

ii. A commercial manufacturing facility (drugs, crystals, etc.)

12. An energy collection and relay station

13. A diagnostic, medical, and convalescent facility

14. A tourism attraction

15. A horticultural research and food growth facility

16. A technology demonstration facility

17. A control center for manned and urmmnned spacecraft

Note that the above requirements are based on civilian use of the Space

Station. Military applications could well add other requirements.
The list of capability requirements are not, by any means, listed in

order of priority. The priority could be determined by the station

capability or by the national needs at that time. For example, co,m_rcial

needs for mlcrogravity processing could be a priority, or the use of the

station as a transportation node to colonize the Moon or Mars might be of

the highest priority.

The second part of the study involved a survey of the state-of-the-art

of spacecraft subsystems and the predicted technology trends in those

subsystems. The major subsystems have been identified in several reports as

the following:

I. Electrical Power

2. Guidance, Navigation, and Control

3. Structures and Mechanism

4. On-board Propulsion

5. Thermal Control

6. Environmental Control and Life Support

7. Extravehicular Activity

8. Communications and Tracking

9. Data Management

The Initial Operational Capebility Space Station uses state-of-the-art

subsystems. In some cases, these are essentially the existing subsystems in

the Shuttle Orbiter or those felt to be advanced to the point where they are

proposed for the Initial Operational Capability Space Station.

Trends in subsystem technology _re examined by literature search and

by use of the NASA Technology Model.

3
A conceptual 2025 Advanced Technology Space Station was outlined, based

on the trends in subsystem study and on a few guidelines devised as a result

of the work statement, the literature search, and group discussions. These

guidelines were:

I. Make use of synergy between subsystems.

2. Provide a semi-closed life-support system assuming adequate

on-board power availability and relatively low-cost resupply

capability.

3. Provide artificial gravity for normal crew activities as well

as a zero gravity capability for experiments. For this

preliminary study, a lunar "g" was selected for living

oonditions.

4. Provide shirtsleeve working envirorment and passageways to

reduce extravehicular activity.

Finally, a list of suggested topics for additional research for the

suggested or alternate Space Station configurations is presented.
A BRIEF STUDY LEADING TO A CONCEPTUAL DESIGN

OF AN ADVANCED TECRNCLOGY SPACE STATION

FOR ABOUT THE YEAR 2025

1.0 INTRfIX_TION

Space travel and exploration of the solar system has long been a dream

of mankind. Rocket research by Robert Goddard and many others, and

analytical studies by engineers and mathematicians such as Herman Gberth and

Konstantin Tsiolkovsky, began to open the possibility of manned space

flight. Finally in 1958 man took his first steps into space when Russian

Cosmonaut Yuri Gagarin was launched into low Earth orbit. This feat _ms

duplicated by American Astronaut John Glenn, and mankind _ms on its way to

the exploration of space. Since that time, progress has been fairly rapid,

and the United States has had a series of successful manned projects such as

Mercury, Gemini, Apollo, Spacelab, and Shuttle flights. Each project had a

specific purpose and, at the same time, provided logical steps to increase

the capability of advancing the exploration of space and utilizing the space

environment.

A current major space project involves the development of a manned

Space Station in low Earth orbit and operational within this century, a goal

which has been widely discussed for many years (Refezemce I). The Space

Station will provide an initial operational capability (IOC) for Earth

observation, scientific studies, and con_ercialization. It is anticipated

that the Space Station capability and size will increase as plans evolve and

technology advances, that the Space Station will assume additional

functions, and that subsystems will be changed or replaced to take advantage

of new technology.
At somepoint in time, a totally new advanced Space Station may evolve

because of new or additional requirements or subsystem changes. Some of the

possible requirements, or functions, of such an Advanced Technology Space

Station and associated technology advances will be reviewed in later

sections of this report. A time frame around the year 2025 is selected to

permit new technology development and to reach well defined capabilities

required of a future Space Station. The purpose of the present study is to

evaluate some of the possible characteristics of such an klvanced Technology

Space Station.

It should be noted that because of the limited n_er of manhours

devoted to the very broad study areas, the results will necessarily be

conceptual. However, results of this study led to a suggested configuration

and to subsystem choices and identified areas for more detailed analyses.
2.0 ADVANCED TI_q05OGY SPACE STATION CAPABILITY R_

In order to determine the capability that might be required of an

A_vanced Technology Space Station, it is necessary to consider what missions

it might perform or support. A reasonable place to start is a review of

proposed or envisioned missions that might require the use of a Space

Station. A current list of proposed sp_ce missions is given in Refe_ 2.

The missions fall into seven categories, as shown in Table 2.8-1.

Supplementary information on the missions is given in Refezenue 3. An

ac]ditional source to examine for future space missions is "The Report of the

National Commission on Space" (Beferemce 4). The Commission was appointed

by the President of the United States and was charged by Congress to

formulate a bold agenda to carry America's civilian space enterprise into

the twenty-first century. The goals proposed by the Commission were "To

lead the exploration and development of the space frontier, advancing

science, technology, and enterprise, and building institutions and systems

that make accessible vast new resources and support human settlements beyond

Earth orbit, from the highlands of the Moon to the plains of Mars"

(Refezemce 4). The recommendations of the Commission on space imply some

additional mission categories such as:

i. Htmmn colonization of the Moon and of Mars

2. Mining the planets and moons of some of the planets

3. Space tourism

Another source for possible space missions is the National Space

Transportation and Support Study (Rsferemue 5).

When all the above categories of missions are considered and projected

to about the 2020-2030 time frame, it is possible to generate a list of

7
Mission Category

Solar System Exploration

Astrophysics

Earth Sciences

Life Sciences

Communications

Space Transportation

Utilization of Space

Environment

No. of Missions in

Funding or Development Status

A P C 0

3 4 Ii 5

7 ii 7 15

7 6 8 0

4 0 0 1

4 0 3 1

9 4 2 3

4 1 0 3

Total

Missions

23

40

21

5

8

18

8

KEY

A = Approved, funded and on-going

B = Planned, start within 5 years

C = Candidate, start within I0 years

0 = Opportunity, start after i0 years

8
functions that could be performed on or by a manned Space Station. A

possible scenario to consider is a large Space Station in Earth's orbit,

possibly accompanied by tethered or free-orbiting platforms, while other

space probes or manned missions are being launched. The following list

suggests some roles that a Space Station might serve:

1. A permanent observatory to look down upon the earth and out

into the universe

2. An orbiting science, medical, materials, and new technologies

laboratory

3. A service and repair facility for payloads, spacecraft, and

platforms

4. An assembly facility where large structures or spacecraft

components are manufactured and/or assembled and checked out

5. A transportation node where payloads and vehicles are

collected, stationed, processed, and launched and where fuel

is manufactured

6. A safe habitat for space crews

7. A commmications and/or relay station for manned or _ed

spacecraft

8. An adaptation area (in variable "g") in preparation for long

space flights

9. A storage node for food, fuel, spare parts, etc.

i0. A variable "g" research facility

11. A commercial manufacturing facility (drugs, crystals, etc.)

12. An energy collection and relay station

13. A diagnostic, medical, and convalescent facility

9
14. A tourism attraction

15. A horticultural research and food growth facility

16. A technology demonstration facility

17. A control center for manned and _ed spacecraft

Note that the above requirements are based on civilian use of the Space

Station. Military applications could well add other requirements.

Note that the list of capability requirements are not, by any means,

listed in order of priority. The priority could be determined by the

station capability O r by national needs at that time. For example,

commercial needs for microgravity processing could be a priority, or the use

of the station as a transportation node to colonize the Moon or Mars might

be of the highest priority.

i0
3.0 SPAt8 STATION _3BSYST_WS, T_CRN[XEG_ ASSESSMENTS, AND SYNITY_IS

The _dvanced Technology Space Station and the IOC Space Station are

comprised of subsystems which require tailoring and integration into an

operational unit. The steps to configure an Advanced Technology Space

Station first identified the principal subsystems and then reviewed the IOC

configuration together with the assessments and projections of technologies

to provide the basis for synthesizing the _dvanced Technology Space Station.

3.1 Major Subsystems for Review

The major subsystems identified in the early Space Station studies

(Reference 6) have continued to appear in the later studies. The subsystems

are:

I.

2.

3.

4.

5.

6.

7.

8.

9.

Electrical Power

Guidance, Navigation, and Control

Structures and Mechanisms

On-board Propulsion

Thermal Control

Environmental Control and Life Support

Extravehicular _tivity

Communications and Tracking

Data Management

3.2 Reviews and Assesm_ents

The reviews and assessments of the technologies began with the

descriptions for the Dual Keel I0C Space Station (Rmferemue 6) as indicated

in F_jure 3.2-I as the point of departure for assessing the status and

ii
I_ poor QUALm'

Z

m

Q

14

II

III

oH

0

14

0

°rl

m

@

0

I0

i,,,,,I

@

!

I0

I:1

0

q,-I

,,p

I!

14

°1,4

0

¢J

@

-o-I

0

o'9

ml

i-I

I

¢q

-o4

12
r

projections for technology accomplishments which could be incorporated into

the Advanced Technology Space Station. In the assessments for the

individual subsystems which follow, the IOC configuration and its relation

to the current status of that technology are the first items addressed.

These subsystem definitions plus the forecasts for technology accomplishments of Rmfezemoe 3 became the general base for synthesizing the subsystems

incorporated into the P_vanced Technology Space Station.

3.3 Advanced Subsystems Synthesis and Synergies

The descriptions for the _ahranced Technology Space Station resulted

from assessment and iterations of configurations. In each of the

iterations, the concept of synergistic interactions between subsystems

received particular attention. The final configuration, shown as Figure

3.3-I and described more completely in Section 6, represents considerations

for accomplishing all seventeen of the functions identified within one Space

Station, while at the same time achieving significant synergies between

subsystems. In that context, the discussions of individual subsystems begin

with the IOC and lead to the Advanced Technology Concepts with potential

synergies identified. The synergies appear summarized as Section 5, and the

description of the proposed configuration for the Advanced Technology Space

Station appears as Section 6.

13
IQ

-H

14
4.0 _ s_r_'=OL_"-'JrLqE-,_n_'T,_ _r AND _IDA"rE

The Space Station design has already undergone several changes, and the

design will continue to evolve as it proceeds through the development steps.

In this study, the dual-keel configuration (Figure 3.2-I) is used as a

starting point because it is the one on which much study already has been

done in defining the overall station requirements and the various

subsystems. The subsystems themselves are described in Referem_es 6 area 7.

The following sections are brief stmmaries of the current subsystems,

technology trends, and candidate advanced subsystems.

4.1 Electrical Power Subsystem

4.1.1 State-of-the-Art Electrical Power Subsystem

The design of the Electrical Power Subsystem (EPS) is docmnented in the

Architectural Control Doctment (Referemoe 7). Copies of this doctment,

which are updated as revisions are received from the Johnson Space Center,

are maintained at the Langley Space Station Office (Building 1212) and at

The Bionetics Corporation (20 Research Drive, Hampton, Virginia).

The 75 kW power output IOC EPS, as given in the June 24, 1986, update

of the EPS, is a hybrid generation subsystem consisting of two solar dynamic

(SD) and two photovoltaic (PV) modules. There is a single SD generation

system in each SD module and two solar array wings in each PV module. The

SD supplies two thirds of the power, and the PV supplies one third of the

power (50 kW and 25 kW respectively). Energy storage for the SD system is

by phase-change salts, and for the PV system NiH 2 batteries are specified.

15
Power source oonverters provide 20 kHz, single phase, 400 volts AC to Power

Distribution Control Assemblies (PDCA's), which in turn supply the Space

Station loads. F_L-e 4.1.1-I shows the simplified features of the EPS

architecture (Refere_oe 7).

The IOC PV systems will use S_fA Lockheed planar, flexible substrate,

deployable and retractable silicon arrays. The SD system has not been

chosen, but the current choices being considered are limited to a closed

Brayton Cycle (CBC) or an Organic Rankine Cycle (ORC).

There are ten external load areas which are supplied by a modified

dual-ring bus system. The four IOC modules will each be supplied through

two penetrations utilizing transformers to provide isolation for a single

point ground. For life critical loads, the EPS is capable of providing twofault tolerant power. Table 4.1.1-i s_m,arizes the major features of the

current EPS.

The IOC EPS is loaded with control computers and microprooessors for

automatic operation, fault detection and switching, collector pointing,

status reporting, and interfacing with other subsystems. The Data

Management Subsystem (DMS) has a major interface with the EPS for reporting

EPS status and for interfacing the requirements of other subsystems back to

the EPS. The overall coordination of power management is performed by the

Power Management Controller.

4.1.2 EPS Technological Trends

Space Station growth may take the form of larger and larger structures

with ever-increasing power demands or moderate growth and multiple special

application individual stations, each with its own unique power

16
17
TWoI=VModules

o Lockheed silicon flex.

deploy.lretract, arrays

o 12.5 kWelm0dule

o 25 kWe total

o Energy storage, NLq2 batteries

Two SD Modules

o Type of SD units not specified

o Limited to ORC or Brayton

o 25 kWe/module

o 50 kWe total

o Energy storage,

salts

phase-change

o TWo failure tolerant power to life essential loads

o Modified dual-ring distribution bus

o Single point ground

o Main distribution bus 20 kHz, 440 volts AC, single _mse

o Power distribution and control assemblies (PDCA's)

20 kHz, 208 volts AC, single phase at modules and

attached payloads

o Hierarchy of control computers and microprocessors

throughout EPS

o Monitoring system commonality with DMS

*6/24/86 Version, Architecture Control Docmmmt (Reference 7)

18
requirements. Examples of individual stations with diverse power

requirements would be: low-power--geosynchronous communication and

monitoring units; medi_r--nRO earth science and payload servicing_

hlgh-power--manufacturing, interplanetary staging stations, or Lunar

Orbiting Space Stations to support lunar surface activities. Soma

activities such as in-situ fuel manufacturing would be especially demanding

in power requirements.

For future power subsystems, important considerations include the

/mpact of overall efficiency on size and weight and the need for high

reliability and graceful failure modes. Nuclear power systems impose an

additional requirement for assuring non-contamination from a launch failure

or re-entry and for obtaining the necessary approvals for launch of

radioactive material.

The overall efficiency of a power system is influenced not only by

source conversion efficiencles, but also by energy storage requirements and

on-board power processing requirements. For example, the original IOC

reference 75 kW PV system had a collection area of 1785 m2 to produce 75 kWe

continuous power for 37 percent earth shadow operation. At 1.35 kW/m 2 solar

flux, the overall conversion efficiency was less than 5 percent.

If a solar input EPS with an overall efficiency of 5 percent is

improved to 50 percent, the collective area is decreased by a factor of ten.

If it were possible to continue improvements and approach 100 percent

efficiency, the additional area reduction would only be by a factor of two.

This final area reduction is 5 percent of the original area. Thus, other

factors such as ease of maintenance and commonality with other systems might

be of greater importance than further efficiency improv_nent.

19
For any heat engine working between two temperature limits, the maximum

efficiency avail_le is that for a Camot cycle (_fezenues 8 and 9).

Efficiency = Tq-z_22 = _T = 1 - 22 *

T 1 T1 T1

The low temperature operational point of most terrestrial heat machines

is tied to the ambient temperature. For space applications, cold space can

provide low te_erature radiator operation--for a given A T, the lo_r the

high temperature operational point, the greater the efficiency and the

easier the maintenance requirements. The radiator area needed, however,

increases as the inverse fourth power of temperature. Improvements in the

design of low tm_perature radiators using new concepts such as liquid

droplet radiators will provide a bonus to heat engines for future Space

Station applications.

A power system will require a prime source of energy, an energy storage

method, and a means of transforming energy to a form suitable for a given

task. The prime energy sources currently being considered for Space Station

include solar, nuclear radioisotope or reactor, and stored chemical or

mechanical energy lofted into orbit. Recent advances in laser design and in

particle physics would indicate that fusion energy might be available in the

time period of this study (Referemee I0).

To first order, the energy source and energy conversion methods may be

considered separately. That is, a Stirling heat engine may be powered by a

solar source or by a nuclear source such as a fixed bed reactor, a space

*Footnote: _i = high temperature, T 2 = low temperature.

These are equivalent to T H and T C in

References 8 and 9. Letter subscripts are

avoided here because terms such as "cold" and

"collector" temperature and "radiator" and

"receiver" temperature can be confused.)

20
power advanced reactor, or a radioisotope heater. However, advantages may

accrue by synergistic relations between components, such as is indicated in

I_fezemcm 11, wherein a proposed 8.7 _ power system uses the same working

fluid for a Brayton heat engine and a nuclear fixed bed reactor.

The energy conversion methods can be subdivided into direct conversion

to electrical energy and heat engines to provide rotating or reciprocating

mechanical energy for conversion to electrical energy. For purposes of this

study, a few of the direct and heat engine methods of energy conversion will

be listed. Direct Conversion methods include solar cells (silicon, GaAs,

and multi-gap), thermomagnetic, thermoelectric, and thermionic schemes as

well as electrodynamlc using tethers. In the electrodynamic tether case,

energy is removed from the orbital velocity and restored by the propulsion

system. Heat engines include Pankine, Brayton, and Stirling. Perhaps new

approaches such as the use of shape memory alloy metal heat engines may

become practical for space power systems.

It is likely that high power output, on-board systems will be

superseded by the use of remote units transmitting power to a manned Space

Station or that multiple specialized Space Stations will reduce the need for

large power systems in conjunction with highly populated Stations. High

energy applications using relayed power might use tethered distribution

lines or laser or microwave power transmission (Bmferemce 12).

Very efficient use of waste heat and synergism between subsystems may

reduce per-unit operational power requirements. That is, in the evolution

of the Space Station, on-board power requirements would peak due to

operational and generation efficiencies exceeding growth needs. Another

factor that would reduce central power generation requirements would be the

21
increased use of self-contained modular packages which would provide local

power generation.

A 2025 Advanced Tec_mology Space Station may use power sources or

conversion methods very different from those covered here. The ability to

deliver mass and voltme to orbit may be less critical than at present, and

on-orbit production and refurbishment could make design lifetime less

critical than employing modular change-out and fail-safe redundancy

philosophies. The projections, however, of a few values as found in the

literature may be useful.

The NASA Technology Model (__ 2) contains useful projections for

some of the EPS components out to the year 2000. The technology forecast

table for electrical power is reproduced in Table 4.1.2-1.

For solar cell powered systems, the theoretical limits of efficiency

are already being approached and it will require a breakthrough in multiband material or direct conversion methods to significantly exceed 25

percent efficiency, which in turn fixes the power per unit area in the range

of 300 watts per square meter for unshaded operation without storage or

distribution allowances. Rmferemoe 13 describes a 75 kW GaAs concentrator

system with a 61 watt per square meter output.

For a comparison of solar and nuclear sources, the projections of

l_femmce 13 are reproduced in Figures 4.1.2-1 through 4.1.2-3. The power

per unit mass for a 300 kW solar Stirling system is about 25 watts per

kilogram. Figure 4.1.2-4, from Referemoe 13, includes a comparison of

masses when propulsion requirements are included, l_feremue 11 describes an

8.7 MH nuclear Brayton system utilizing a liquid droplet radiator with a 537

watt per kilogr_ design. Assumptions about shielding requirements, system

22
TA_LlllJ 4.1.2-1 POIrBR 8TB'I'IDW TL'CID_OLOGT I_JL_C'ABT8 (_CM 2)

o

Figure of Merit

Specific Power (W/kg)

Power Level (kW)

Design Lifetime (yr)

Cell Efficiency (%)

Speciflc Cost ($/W)_

Specific Power (W/m')

1985 1990 1995 2000

Solar Arrays

60 120 180 250

25 40 250 --

10 12 15 15

17.5 20 24 24

300 180 100 90

170 190 240 240

Energy Density (W-hr/k8)

NL/Cd, LE0/GEO

NL/H2, LEO/CEO

HEDRB_ CEO

Secondar 7 Cells

I0120 12125 15/30 20135

30/35 32/40 32/42 32/45

* 92 100 110

Design Lifetime (yr)

NI/Cd, LEO/GEO

NI/HT, LE0/GEO

HEDRB, CEO

7/12 8/13.5 9/14.5 -/15

4/7 7/11 9/13 -/14

* 5 6 10

Energy Density (W-hr/k8)

LIISOCI.

as/zn z

PrLmar 7 Cells

250 325

150 180

--m

m--

Design Llfe (yr)

Lt/S0CX2

._/Zn

8 10

2 2

--m

--o

Radioisotope Thermoelectric Cenerators

Efficiency (%)

Specific Power (W/kg)

Design Life (yr)

8.5 10.5 ....

4.5 6.5 8 --

6 7.5 8.5 10

Nuclear Reactors

Power Level (kW) *

Specific Power (W/ks) i0

Power Conversion Efficiency (%) 6

* -i00 I00

20 30 50

9 12 --

Fuel Cells

Specific Power (W/ks)

Power Level (kW)

Specific Cost (S/W)

110

50

25

115 122 130

100 175 --

20 15 10

Power Control S_stems

Specific Power Handling (W/ks) 72 80 85 85

*Technology not available.

23
O/

C)

I.F)

II

Z)

I:1.

,4-)

:3

C:)

L

0

Z

L

O .D

X

L QI C

C I, .p

U I C]

C L

O r C]D

(_J O

U L

(/) .,-, O

O ,.-, (-_

'_LI_ .,-, U- O

L._O: (.r) rv" Lr)

NPAN

_ O O O O O

cn

c c

o-i o .,_

•,-, _ L

L L ._

.,., rn Lr)

U_ L L

O O

L W 0J

•.-, U U

O ::)

U') Z "_

E

I

OU')

X

O

O

|

m

ill

• •

24
e

C3

C3

_J3

CL

_J

C3L0

CL

aJ

(D

Z

25
26
O)

C)

X

IU

f0

_J

0

r"

cO

a)

O_

L

O. C_

E

L

0 L

D"

Q

D

"- oV-rl

t'/l/I/llllill/////////l//_

,_

I'//////////////////1

... _ V///////////, L U_ U_

® + J _ [////h o

E | I "_

+ !

|_

.,.4,-i

QO

OQ

oQ

r4_.4

I

C) C_ C) C) C) (_)

_ C) C) C) C) 0

__ C) C) C_) C) C)

c_ c_ c_ c_ c_

u_

_J E

E

E

0.

0) 0

27
efflclencles, and energy storage requirements will greatly influence the

projections in this section.

4.1.3 Candidate Advanced Electrical Power Subsystem

As indicated in the previous section, high power generation and high

power applications may be isolated from the populated Space Station.

However, for this point design, an on-board system is chosen to support all

seventeen Space Station functions listed in Section 2.0. This choice, which

requires large collector and radiator areas as well as Sun inertial

orientation in conjunction with the rotational requirements for artificial

gravity, will serve to identify subsystem requirements for a large EPS.

For the point design, we assume a power subsystem of 2.5 M_e output.

The manufacturing and fuel generation requirements are the major energy

users. Table 4.1.3-I lists the power allocations which include a 36 percent

margin. The nm,ber of modules and spares to provide this output would vary

with operational requirements; for this case, we have chosen six units. Two

units supply power to the artificial gravity rotating section, and four

units supply power to the stationary section. These units would be selfmonitoring and self-repairing by automatic switching to redundant parts.

_equate intelligence would be built in to signal when redundant parts were

at a predetermined level and to alert to failure trends or generic failure

modes. The system would alert when manned or robotic maintenance was

required by either on-board Space Station action or return to Earth for

refurbishment.

The energy conversion method is a generic heat machine operating at 80

percent of Carnot efficiency. The actual type (stirling, shaped memory

28
4.1.3-I P_SOR_, _ 1_0l_ lIBJa _F_W8

Ftmction

Operations

Science

Transients

Medical

Manufacturing & Fuel

(bntingency

Total

Personnel

12

10

14

6

12

6

6O

PoweE

200 kW

I00 kW

20 kW

20 kW

1500 kW

660 kW

2500 kW

29
alloy metal, new approach, etc.) is not important in determining operating

limits. The 80 percent of Carnot is an arbitrary selection that sets the

heat engine A T.

The collector area is determined by the overall efficiency, the solar

flux, and the time spent in Earth shadow. The radiator area is selected to

be the same as the collector area, which is reasonable from drag and

structural considerations. The radiator temperature is determined by its

emlttance, view factor, and the power radiated to cold space. The inlet

temperature to the heat engine is determined by how close the heat engine

works to the maximum efficiency of a Camot cycle. These factors are

illustrated in Figure 4.1.3-1, which provides calculations for the point

design. The calculations are normalised and plotted as a function of

overall efficiency in Figure 4.1.3-2. The diminishing returns in reducing

the size of collector and radiator and the higher temperatures that must be

designed to as a function of efficiency increase are illustrated in this

Figure. Operating temperatures are shown on the radiation plot of Figure

4.1.3-3. A factor that is not included in these calculations but which

will definitely limit the maxLmmn efficiency obtainable is reradiation of

energy from the receiver as T 1 increases. A design factor that controls

this reradistlon is receiver aperture area, which in turn is controlled by

collector mirror surface smoothness and shape and by collector pointing

accuracy.

Another point should be noted relevant to Figuze 4.1.3-2. For a heat

engine operating at 80 percent Carnot efficiency and the overall efficiency

selected, the collector area and power output are directly related

independent of the power level chosen. For example, a 25 kW power module at

30
CRITERIA - . POgER OUTPUT 2500 kWe

. OVERALL EFFICIENCY 40%

• LOSSES BEFORE HEAT ENGINE 10%

• HEAT ENGINE 80% OF CARNOT

• RADIATOR: e = 0.85, vf = 0.9

v

]

f

,/,,f

Y

COLLECTOR

()in = --7_ ] HEAT ENGIRE

---,

0loss

RADIATOR

= 625 kW POHER OUT = 2500 kW

Qrad =

3125 kg

• SOLAR INPUT - 1.35 kW/m2 x 2/3 DAYLITE = 0.9 kg/m 2 AVERAGE

• COLLECTOR INPUT = Qin = 2500/0.40 = 6250 kW

• COLLECTOR AREA = 6250/0.9 - 6944 m2 (DIAHETER = 94m)

• RADIATOR AREA = COLLECTOR AREA = A1

• EFFECTIVE RADIATOR AREA A1 x e x vf = 6944 x 0.85 x 0.9 = 5312 H2 = A' • r

• RADIATED POWER PER m2 = 3125/5312 = 0.59 kg/m 2

RADIATED POWER IS ALSO • 5.6 x 10"11 T24 kg/m 2

• RADIATOR TEHP = T 2 = 320 K • 470 C

• HEAT ENGINE EFFICIENCY • 2500/(6250-625) = 44.4%

• EFFICIENCY HAXIHUH • EFFICIENCY H.E,/CARNOT FRACTION = 44.4%/80% = 55.5%

• INLET TENPERATURE = T 1 = T2/(1-EFFICIENCY HAXIHUN) = 720 ° K = 4470 C

Figure 4,1.3-i Point DesignCalculations

31
Qwork(kN):

EFFZCIENCY

QinputQloss=

I

o

1.00

(VARIABLE)

Qhe+Qloss

IOZ Qhe

Figure 4.1.3-2

T1

! I I I

_2 0.4

0_ B_CE_CY

+ _ o _1 +_)_

Qheateng: QNIEFFhe

Qradiate: Qhe-Qw

AREAinput=Qin/(1.35*213)

AREArad= AREAinput

Temperature and&rea

! !

0.6 0.8

i_r'=Ar*vf*e

QrlM^2=5.b.lO^-11.T2^4

TI=T2/( 1-EFFmax )

Versus Efficiency

IE - 5.6 x 10 "11 T 4 (kg/H 2)

UDIAT[D POiRIt,

(kW/_

101

1DO

10-1

10-2i

100 200 300

Figure 4. I. 3-3

OP£RATING

T(ItP - 7200K, 15 kU/M2

FLUX - l.$S kV/1N2

SOLAR INPUT • 0.5 kH/H2

RAOIATOR OUTPUT . 0.59 kW/fl 2

COLD SPAC£ o 115°K,10 "2 kg/_ 2

400 500 500 700 800

TEm,[_nm£. (°1()

Blaokbody Radiation

32
40 percent overall efficiency w0uld operate at T1 = 720° K and T 2 = 320 ° K

with a collector area of 69.5 _, whereas a 2.5 MWe system would operate

between the same temperatures with a collector area of 6,950 m 2. In actual

practice, there is an economy of scale, usually providing higher

efflciencies as the power level increases (Reference 14).

TWo methods of energy storage are employed. Phase-change materials

(not necessarily salts) are used as a buffer to regulate input to output

energy and as a direct source of high level heat for the thermal control

system. Water is electrolyzed, and hydrogen is stored in absorption

materials. The oxygen is stored cryogenically. In addition, energy stored

in angular momentum associated with the control subsystem and the artificial

gravity subsystem can be tapped.

4.2 Guidance, Navigation, and Control (_&C) Subsystem

This is a major Space Station subsystem and is critical to many

functions, including orbit maintenance, solar pointing, attitude control,

experiment stability, and proximity operation.

4.2.1 State-of-the-Art

The baseline control system (Figure 4.2.1-1) is a centralized system of

attitude and rate sensors, and control-mom_t gyros control effectors that

provide station stabilization. The navigation system uses inertial sensor

assemblies: rate gyros and accelerometers to provide attitude, attitude

rate, position, and velocity; star trackers to update attitudes; and the

Global Positioning System satellites to update position and velocity.

33
34
Primary attitude control will be through the use of control-moment

gyros. Supplsmentary control m0ments by use of magnetic torquers are

indicated in Referei_e 6, but not in _ 7. The reaction control

system will be used to perform any required translation maneuvers. In

addition, the reaction control system will be used to back up the controlmoment gyros and the magnetic torguers, if necessary. Reaction control will

be attained by placing thrusters in positions to produce torques around and

thrusts along three orthogonal body axes. The reaction control system will

be used to reboost the Space Station when required and may be used to avoid

collision with orbiting objects. All the control syst_ will be able to

operate individually or simultaneously.

The equipment list for the GN&C subsystem includes a pair of star

tracker triads, two 3-axis inertial sensor assemblies, six control-momant

gyros, and numerous components that provide electronic support to the GN&C

subsystem.

The GN&C subsystem has management responsibility for traffic control

and proximity operations of station traffic operating within the Space

Station's area of influence in order to insure that operational procedures

meet safety and mission constraints. The GN&C subsystem also will have the

capability of monitoring and controlling vehicles whose trajectories may, at

some time, cause them to enter the area of influence. The Station's area of

influence is, by definition, a 30° cone about the positive and negative

velocity vectors of the Station out to 2,000 kilometers and the entirety of

a sphere of radius 8 kilometers around the Station.

The Space Station GN&C equipment list is presented in Table 4.2.1-I

along with weight and volume parmnsters.

35
TABLE 4.2.1-I SPACE STATION GN£C EQUIPMENT LIST (REFERENCE 6)

Unit Description

n mini In il I --

DImenslons Vol. Vol. Welght Weight

Per Unit No. Unit Total Per Unit Total

(In.) Units (Cu.Ft.) (Cu.Ft.) (Lbs.) (Lbs,)

Actuators

CMG Assembly

Magnetic Bar

42 Sphere 6 22.45 134.70 420.00 2520.00

2.25 Dia.x98 6 0.23 1.35 109.00 654.00

Total Actuators 136.05 3174.00

Sensors

Star Tracker Triad

Hexad Strapdown

Spares for above

13x13x12 2 1.17 2.35 20.00 40.00

12x12x12 1 1.00 1.00 50.00 50.00

12x12x12 1 1.00 1.00 50.00 SO.O0

Total Sensors

Electronic Support

Magnetic Torquers

RCS Cont System

G&C Processors

NAV/Traffic Proc

Interface Devices

Solar Array Elec.

Spares for above

Radiator Elec.

Spares for above

Payload Elec.

4.35 140.00

9xgx9 6 0.56 3.38 20.00 120.00

12x18x8 3 1.00 3.00 35.00 105.00

7.SxlO.Sx15 3 0.68 2.05 20.00 60.00

7.5xlO.Sx15 3 0.68 2.05 20.00 60.00

7.SxlO.Sx15 18 0.68 12.30 5.00 90.00

8x9x6 6 0.25 1.50 12.00 72.00

8x9x6 3 0.25 0.75 12.00 36.00

8x9x6 2 0.25 0.50 12.00 24.00

8x9x6 1 0.25 0.25 12.00 12.00

2x5x6 2 0.04 0.08 3.00 6.00

Total Electronics 25.86 585.00

Grand Totals 166.26 3899.00

i

36
°

4.2.2 Technology Trends

Increased perfo_ in GN&C systems is driven by the apparent need

for large space structures, large optical assemblies and antennas, and high

precision orbit determination. The orbit determination requirements

(navigation) could be handled by on-board systems or could be handled by

existing highly developed ground tracking facilities, computer facilities,

and associated algorithms for data processing. Proximity navigation,

associated with collision avoidance, probably could be aided by on-board

radar tracking units.

Predicted technology trends for the primary components of the GN&C

hardware are given in Table 4.2.2-1 (from Rsfermlue 2). One emerging

technology that appears quite attractive is the use of laser or fiber optics

gyros for measuring angular rates. Fiber optic gyros are predicted to be

about 15 times as accurate as spun mass gyros by the year 2,000.

15 contains detailed discussions of inertial technology trends. It is

interesting to note that of the three types of gyros (spun mass, laser, and

fiber optics) only the fiber optics gyro is predicted to meet the

requirements for the Saturn 0rbiter/Titan Probe (needed by 1989) and that of

the Large Deployable Reflector (needed by 1993). The I_J drift accuracy

requirements are 0.03 mEad/hour and 0.01 mEad/hour, respectively (l_ferem_

2J.

Other advanced concepts for attitude control include the dual momentum

vector control concept (l_fexemue 16) and the Integrated Power and Attitude

Control System, IPACS (Referem_e 17). The latter integrates the angular

mumentum control and energy storage function. _ 17 indicates that

recent technology advances in composite rotors, magnetic bearings, and power

37
TJLBLI_ 4.2.2'1 NAVIGATIONs GUIDANCE, AND CONTROL TECHNOLOGY FORECABTB

(_mrD.maCB 2)

Figure of Merit

Position Accuracy (m)

1985 1990 1995 2000

Guidance and Navigation

5 2.5 1 --

Ceotd Residual (ca) 150 -- 10

Pointtngt Vtbratton_ and Figure Control

Viscoelastic Materials

Damping Coefficient (Z) I 2.5 4

Pointing Accuracy (arad)

Earth-Polnting Systems

Inertial Systems

Stellar Systems

1.5 x 10 -2 1 x 10 -2 7 x 10-_

5 10-_ 4 x 10 -3 . 3 x 10 -3 .

3 : 10 -4 1.5 x 10 -4 1.2 x 10 -4

3 x 10 -3

2.5 x 1_ -J

lxlO -1

Pointing Stability

(mlcrorad) 3 x 10 -2 1.2 x 10 -2 7 x 10 -3 3 x 10 -3

Pointing Stability Rate

(alcrorad/sec) 5 2 I 0.7

Maneuvering Settling

Time (sec) 0.6 0.25 0.2

Inertial Rate Unit Drift

(mrad/hr)

Spun Mass Gyros 0.2

Laser Gyros 0.07

Fiber Optics Gyros *

0.15

0.05

0.01

0.12

0.05

0.008

0.12

0.05

0.007

Active Earth Sensor

Accuracy (mrad)

Low Earth Orbit 5

Geosynchronous Orbit 1.2

Star Tracker Accuracy

(alcrorad)

8 ° FOV Charge Transfer

Devices 10

1 ° FOV Charge Transfer

Devices 1

3 2.5

_m

5

0.5

3

0.3

Momentum Wheel Angular

Momentum (N-m-s)

Reaction Wheels,

Medtum Speed 100

Reaction Wheels,

High-Speed 250

Control Moment Gyros,

I-DOF 3,000

Control Moment Gyros,

2-DOF %000

200

600

5,000

12,000

600

2000

9,000

20,000

2000

3000

11,000

State Variable

Estimation (no. states) 25 5O 250 1,000

*Technology not available.

38
control electronics have triggered new optimism regarding the feasibility

and merits of such a system.

4.2.3 Candidate _z]vanced Subsystem

An advanced candidate GR&C subsystem would contain all of the elements

of the current state-of-the-art subsystem (Table 4.2.1-1) but would use the

latest technology available. Selection of actual hardware components will

depend to a large extent on such factors as reliability, durability, ease of

maintenance and/or repair, availability, cost, size, and weight. Other

factors to be considered are redundancy, fault tolerance, and failure or

malfunction detection. Some of these factors are related to development of

control algorithms for the various functions.

"4

4.3 Structures and Mechanisms Subsystems

The Space Station (IOC) design and structural configuration will be

limited by the development of advanced materials, their manufacturing

concepts, and the volume/weight lift capabilities of the Space Shuttle to

low Earth orbit.

The assembly of the IOC Space Station may be accomplished using

astronauts performing EVA. An alternate assembly concept for the Space

Station is under development and could employ a free flying tele-rcbot with

a man-in-the-loop.

4.3 .I State-of-the-Art

The planned Space Station (IOC) will be assembled using as a foundation

a dual-keel truss structure to which pressurized modules and the mechanisms

39
of the operational Space Station will be attached. The construction will

cons4st of 5-cm-_iameter elements joined at spherical nodes to fabricate a

truss structure consisting of 5-m bays.

One preliminary design of a tubular element for erectable assembly in

space consists of a wall construction of 0.15-cm-thickness composite formed

of axially collimated carbon fiber reinforced epoxy resin. The composite is

clad on both the inside and outside diameters with 0.015-cm-thickness

aluminm, alloy. Threaded aluminm, alloy sleeves are adhesively bonded into

the tube ends, one end with a right-hand thread, the opposite end with a

left-hand thread to effect a turnbuckle concept. The strut attachment

fittings are assembled with the tubular element struts and adjusted by

rotation to achieve precise strut length.

Spherical alto,inurealloy nodes are numerically machined and threaded at

twenty-six locations to receive node attachment fittings. The attachment

fittings are prelaunch assembled to the nodes by cap screws (F_juze 4.3.1-

1).

The attac_mmmt of each strut end to its corresponding node attachment

fitting is accumplished in space by astronaut manual assembly of an erectable strut joint. A sidelatching of the joint occurs when the joint is hand

clasped transversely to the strut's axis. The astronaut secures the joint

by rotating a cam cover 45° to apply joint preload (Figure 4.3.1-2).

One concept of assembly of the Space Station dual keel will utilize a

Mobile Remote Manipulator System (M_4S). The MRMS is a platform which can

move one bay at a time being supported on the truss structure by MRMS guide

pins located at the structure's nodes.

40

i

L
Figure 4.3. $-I Erectable Strut lttachment Method

[LaRC Concepts Under Development)

Rotary preload and unlock cam

Detent spring and bell

_

_iei!!!! ee ring

Preload BallevlUa washers

Cam cover

n Latch bolt

Figure 4.3.1-2 Preloaded Ereotable Strut Joint--Bxploded View

(LaRC Concepts Under Development)

41
The NRMS provides manipulating capability for positioning Space Station

subsystems components. The MRMS provides two Mobile Foot Restraint (MFR)

platforms to position the astronauts during EVA erection and/or servicing of

the Space Station. The astronaut may control his MFR platform much like a

utillty serviceman operating a "Cherry Picker" bucket (Figure 4.3.1-3,

l_fere_ce 18).

The truss structure sl_herical nodes (Figure 4.3.1-4) will also provide

attachment locations to secure platforms, manned modules, and experiment

stations inside the truss structure to provide center of gravity location

control and not interfere with the operational clearances required for the

M_MS.

The astronaut assembly cono_pt has been demonstrated using neutral

buoyancy tests and in-space erection-disassembly studies with good

correlation of the performance times (IkM_ezemoe 19). The anticipated

assembly time for a 1,200 tubular element dual keel Space Station truss

structure is 20 hours based on one minute per element.

The habitation and laboratory modules will be fabricated from aluminum

alloy and welded to assure leak tightness.

4.3.2 Advanced Space Station Materials Technology Trends

The advanced Space Station will require improved structural materials

to assure fabricability of structures and enclosures that possess long life

during prelaunch, Earth storage, and on-orbit space environments. The

selection of materials suitable for each application will require

consideration of the following factors:

42
\

ORIGINAL PAGE IS

OF POOR QUALITy

Figure 4.3.1-3 Mobile Remote Manipulator 8ysta (Reference 18)

INSERTION DIRECTION

I _AL REMOVED

PLATFORM ATTACH FI)

PAYLOAD ADAPTER

F4gure 4.3.1-4

"TING

Plmttorm Attachment Method

(LaRC Concepts Under Development)

43
i° Definition of the requirements for a specific material

application

2. Determination of the effects of critical envirorm_ntal

conditions on the candidate material's properties and

possible interactions with other systems

3. _3iification of the material for its intended use

A summary of damage mechanisms of materials resulting from a space

environment is provided in Table 4.3.2-I (Reference 3).

Structural cumposites consisting of an organic resin matrix reinforced

with inorganic fibers offer high specific stiffness and specific strength

with adequate structural integrity for a variety of space applications. The

space durability of currently available resin matrix composites is projected

as approximately eleven years. Lifetimes of fifteen to thirty years in

space are forecast using improved materials and processing methods (Figure

4.3.2-1, Reference 3).

Polyimide resin matrix composite materials offer promise for increasing

the maximum service temperature of composite structures from the state-ofthe-art use temperature of 450 K to possibly 720 K by 1990 as predicted in

Figure 4.3.2-2 (Refe_ 3).

Metal matrix composites (MMC) composed of high modulus graphite fibers

embedded in an almuinum or magnesium matrix offer materials having high

specific stiffness, high thermal deformation resistance, low coefficient of

thermal expansion, and negligible outgassing. Large space structures and

space antenna systems require high dimensional stability to maintain the

system's alignment accuracy. F1c/_ze 4.3.2-3 compares the specific stiffness

44
_s.EEJ_ 4.3.2-1

(lq_mq_ 3)

Environmental Element

Ultraviolet Radiation

(Solar Electromagnetic)

fharged Particle Radiation

(Trapped radiation, solar

winds, solar event particles,

cosmic rays)

Pressure (Atmospheric)

Damage

Thermal (Solar Electromagnetic)

Micrometeoroid

Elect rodynamic Interactions

Atomic (_ygen BoMbardment

- Creation of lattice defects in

crystalline materials

- Chain scission of organic materials

Free radical formation

Color centers

- Crosslinking of organic materials

- Creation of lattice defects in

crystalline materials

Recombination centers

Absorption centers

- Chain scission Of organic materials

- Crosslinking of organic materials

- Secondary radiation damage

-Volatilization of low vapor pressure

fractions and materials

- Diffusion

- Vacuum welding

-Mechanical degradation, softening or

embr ittlement

- Chemical degradation

- Acceleration or deceleration of

above envirormental effects

Mechanical

Fracture or puncture

- Spacecraft charging

- Oxidized surface recession

45
80

40-

D

30 -

D

Z0 -

10

0

1970

SPACE

IDIVlRONIIENTLIFk'TIE (.yr)

j = PROBABLE

I I I I I

197S 1980 1985 1990 1995 2000

TECHNOLOGYREADINESSDATE

Figure 4.3.2-1 Space Environment Lifetime of Resin Matrix Composites

(Reference 3)

8OO

60O

TEMPERATURE(K)

4O0

GRAPHITE

POLYIMIDE

EXPLORER49

CONOENSATIOH-CURE

POLYIRIDE

HYBRIt

SHUTTLEOUITER

PROBABLE

I I I I

1980 1985 " 1990 1995

TECHNOLOGYREADINESSDATE

2OOO

Figure 4.3.2-2 Maximum Usable Non-Metalllc Service Temperature

(Reference 3)

46
versus thermal deformation resistance of several spacecraft structural

materials (Referenoe 3).

Space structures' weight reductions of 20 to 50 percent are projected

by the year 2000 using advanced composite materials with improved structural

analysis, design, and assembly concepts (Figure 4.3.2-4, Rafezence 3).

Earth manufactured structural elements will be designed for on-orbit

assembly for large structures using assembly aides such as those described

in Table 4.3.2-2. The free-flying teleoperators with man-in-the-loop remote

controls appear to be very versatile for non-repetitive assembly tasks and

to minimize extravehicular activities.

_SA LaRC is studying a space spider crane with four 60-foot legs and

two arms. The spider crane would be able to assemble various elements of

the Space Station and maneuver around complicated station elements such as

solar arrays or radiators (FJgEe 4.3.2-5). The crane would be battery

powered and controlled remotely by computer or astronaut. A rail system has

been considered at the Go_rd Space Flight Center for transport of a manned

extravehicular robot to install and service utilities and payloads from

within the geometry of the Space Station truss structure (Figure 4.3.2-6).

4.3.3 Candidate kAranced Structures and Materials Subsystem

A modified tetrahedral truss structure appears as a valid method of

construction for the Space Station using free-flying teleoperators.

Structural tubular elements of graphite fiber reinforced polyimide resin, or

graphite fiber embedded in magnesi_ or aluminum are candidate structural

composite materials.

47
£/p, INCHES

800 m

700--

600-

500--

400-

300-

200-

100 -

0

_.-0

.... Gr/Ng

6r/Hg • .....

43/v/o 48 v/o

• ....8

h O" 6r/A1

• OC-C

6r/Ep

6r/A1 48 v/o

43 v/o

• A1 UNCLASSIFIED

t _ I I I I J , I

0 20 40 60 80 100 120 140

KT/a L, 106 8tu/hr ft

l_Lcjul_ 4.3.2-3 Bpeotfic e_4._'_nu8, IVp , versus _ _tJ.an

li_ntLst_mm, ];F_;,, for _ _ Xaterfal8

(_ 3)

STRUCTURALVE]Gd4T/

STRUCTURALHEIGHT IN 1975

Ficjure 4.3.2-4

A. WHAT WILL BE

8 • WHAT IS POSSIBLE

I.Z

1.0 oOoooOQoooooo

o

• A

o

o

o

: D

TECHNOLOGY

i I

198S 1990

READINESS DATE

I I

1995 2000

8tructural Weight Reduction for Space 8ystems

(Reference 3)

°

48
49
Figure 4.3.2-5 Space Spider Crane (Under Study at LaRC)

i-#ETER STRUCTURE

ATTACHED

PAYLOADI'

\

U]'|L|TY

ATTACHED

PAYLOAD

UIL SYSTEM

5O
Inflatable structures can be carried deflated and folded into space.

The structure would be inflated in space on command and the resin impregnant

of the structure's reinforcements would harden in the space environment.

Large rotating seals for maintaining one earth atmosphere of pressure

within a rotating Space Station would provide a shirtsleeve work environment

for the astronauts.

Completely assembled laboratory and habitation modules weighing up to

150,000 pounds could be delivered to orbit by way of a heavy lift cargo

vehicle.

A helical-wrap large diameter tube could be stowed in a cargo vehicle

Hereby the helical band would be telescoped within itself to achieve a high

packing density. The tube would be telescopically extended in space with a

locking and sealing of the helical seam to form a gas tight cylindrical

structure.

A brief review of some factors to be considered in selecting materials

for Space Station application is given in the Appendix.

4.4 On-board Propulsion Subsystem

The on-board propulsion system provides the thrust capabilities to

accomplish reboost, maintain the attitude, and execute avoidance maneuvers.

A review of the future space missions shows the need for a major Space

Station in orbit at 500 km altitude with a 28° inclination. Orbital flight

requirements for higher (up to G_0) altitudes or at other orbit inclinations

would be served by auxiliary platforms. Therefore, for the Space Station,

reboost represents the principal continuing requirement for thrust. The

review of missions does imply a significant Space Station participation in

51
the assembly, checkout, and deployment operations for advanced missions. In

such operations, the thrust requirements to support docking, erection or

assembly, and undockings could each expand by more than an order of

magnitude. The considerations for Space Station evasive action or avoidance

maneuvers anticipate modest motions performed at low velocities. Any

hostile action or threat would be countered by other than evasive

maneuvering of the Space Station itself.

The descriptions and assessments of propulsion system technologies

which follow were generated at the same time the propulsion system for the

IOC Space Station received a major technical review which changed the

baseline from monopropellant hydrazine to 02-H 2. The discussions,

descriptions, and comparisons that pertain to hydrazine continue as valid.

The shift to an 02-H 2 based propellant system recognizes the cited

advantages and accepts the identified developmental achievements as part of

the planned effort to produce the I0C Space Station.

4.4.1 State-of-the-Art

The propulsion system initially selected for the I0C Space Station

utilized the established technology presented by catalytic combustion of

monopropellant hydrazine (N2H4) (Reference 6). The selection of hydrazine

is a workable compromise to achieve a straightforward operation of an

established system at a modest cost in fuel weight and res_ply

requirements. Table 4.4.1-I summarizes the pertinent performance

requirements and features for the system. The continuing design refinements

underway for the Station anticipate changes in the impulse (and propellant

weight) values listed, with the differences as fractions of the present

52
_a.BLE 4.4.1-1 _ _ _ P_EGLSI_

[l_I_ _ IHl1_Ilm, _ & S_CIFIC _ OF 220 SIC

A. THIU_ _IMHY_SEPE_I_CE_REMENTS

Parameter

I. Reboost, after 90 _ays

at 27O m (500 k_

2. Orbiter Dock; Undock

3. Attitude Control

Contingency

4. Co111slon Avo_l_e

5 ft/sec Oont_y

5. Altitude Transfer

20 NM Contingency

Total 0n-Board Propellant

Contingency On Board

Minimum _upply

Impulse Hydrazine

Ibf. Sec. Weight Lb.

483,000 2,200

26,000 120

147,000 670

61,500 280

831,000 3,780

7,050

1,039,500 4,730

509,000 2,320

B. TANKAGE AND OPERATIONS

i. No. of 40" Diameter Tanks

2. Tank Supply Pressure Range

3. Thrusters Arrangement

4. Thruster Force Range

9 Total, 6 Contingency, 3 Resupply

100-300 psi

4 Locations, 3 each direction,

36 total

25--75 pounds

53
values. The present system has no defined synergy with any of the other

systems on the Station and draws upon the Station's supply for both

operating electrical power and nitrogen to pressurize the fuel tanks. The

pertinent features of the system are described in terms of the thrusters,

propellant storage, and control.

4.4.1 .i Thrusters

The individual thruster design utilizes a catalytic bed unit which can

provide from 25 to 75 pounds of force in proportion to a blowdown fuel

supply pressure ranging from I00 to 300 psi. A triple redundancy results

from clusters having three independent units operating in three directions

to provide nine thrusters at each of four locations (36 thrusters total).

Reboost operations will establish the life requirements for both the

catalyst beds and nozzles and anticipate some configuration development for

the catalyst beds. The concept of on-board, in-place spare thrusters (e.g.,

extra units in each cluster) has been addressed and remains as a lifeextending option. Uprating by heated beds represents the near-term

development effort for an advanced system and could offer up to a 35 percent

improvement in specific impulse (Refe_ 20).

4.4.1.2 Propellant Storage and Distribution

The provisions for propellant storage utilize spherical tanks 40 inches

in diameter. The Space Station includes a fixed on-board tank cluster which

accommodates the contingency resource (nominally six tanks) (see Figure

4.4.1-1). Resupply will transport tanks sufficient to perform reboost and

docking (nominally three tanks) plus any additional tanks needed to

/

54
UPPER

LEFT

CLUSTER

LOWERLEFT

CLUSTER

|!

ON BOARDTANKAGECONTINGENCYSUPPLY

JLLJL

rp_rp

o0

_-]

OPTIONAL

SERVICING

INTERFACE

CLUSTER

SHOWINGLINE ROUTINE

(TYPICAl. 4 PLACES)

LOGISTIC HODULE

HYDRAZINESUPPLY

(REBOOSTAND

ATTITUDE CONTROL)

0

0

LOWERRIGHT

CLUSTER

ILg_re 4.4.1-I Referenoe Configuration Propulsion System

vith Ipproxlmate Equipment Locations

(Reference 6)

55
replenish the contingency. The concept presented envisions remote handling

of all fuel servicing activities (no EVA for connect, disconnect). The

tanks all include internal diaphragms for fluid control with the ullage

space filled and pressurized by (_2 from the central supply (LN2 on-board

storage or an auxiliary decomposition of N2H4). In operation, the ullage

pressure provides the distribution forces and the blowdown pressures at the

individual thrusters which determine the actual thrust levels achieved. The

distribution system provides triple redundant lines to each set of clusters

such that one line feeds three thrusters, one for each for thrust direction

at that location. The distribution includes solenoid valves that permit:

I. Isolation of each tank

2. Propellant feed to all thrusters from any tank

3. Propellant transfer between tanks

Ullage pressure provides the driving forces for all propellant movements including transfer between tanks. Therefore, the system is designed

to operate at tank pressures above 300 psi. Hydrazine has some limits for

temperatures during storage and transfer. The tanks and distribution lines,

therefore, include heaters for temperature control. These draw their power

from the on-board supply.

4.4.1.3 Thrust Control

Solenoid valves located at the inlet to each thruster provide the means

for initiating and terminating thrust forces. Impulse control becomes a

combination of ullage pressure and thrust duration. The concept of feeding

all thrusters from the same tank minimizes any effects due to changes in the

supply pressure; the distribution lines have been sized to balance or

56
minimize flow pressure losses. The operation of the system will permit

individual firing and timing for each thruster. The actual firing-sequence

control _ appear as the output from an on-board computer algorithm

contained within the navigation and control system. The anticipated

operations foresee only modest thruster usage for attitude control or

docking. Raboost presents the principal utilization with burn times of

several minutes. The contingency requirements for velocity change or orbit

altitude increase can involve longer times, but they are not scheduled

events.

4.4.1.4 Smaary of the State-of-the-Art

In summary, the propulsion system utilizes a straightforward

established technology which will require a minimum of configurationspecific development for the Space Station application. In exchange, the

Space Station accepts the extra increment of supply weight and the

particular safety-related handling constraints associated with hydrazine.

Reliability utilizes triple redundancies and accepts the additional nm_er

of valves and valve controls which result. The system shows no synergistic

participations beyond the potential use of N2H 4 d_sition for N2

replenishment. Actually, the system is dependent in that it makes demands

on electrical power. In total, the system appears to be an effective

compromise which reflects the diverging requirements presented by the Space

Station. In addition, the system offers a near-term potential Performance

improvement of as much as 30 Percent in specific impulse by utilizing heated

catalyst beds (Referem_e 20). Since thruster change-out is considered in

the IOC design, the potential uprating appears both attractive and

57
attainable to the point of justifying a continuing development for enhanced

performance monopropellant units in the 20- to 100-pound force range.

4.4.2 Review of Status and Technology for Propulsion Subsystem

Candidates

The candidates for Space Station on-board propulsion system

applications can be sum,mrized in terms of chemical, electrical, and other.

Table 4.4.2-I lists the present candidates in terms of the pertinent

performance and application considerations. For discussion purposes, each

of the candidates has been assigned a ranking relative to development

requirements and com[mtibility assessments. The ranking criteria appear in

Table 4.4.2-2. The organization of the table and assignments of technology

which follow reflect the energy source which effectively provides the

acceleration to the ejected matter. Chemical reaction obviously involves

the energies available through combustion. The electrical systems apply the

major portion of the thrust energy directly from an on-board electrical

power source. The "other" category involves interactions t_hat appear to

have an intermediary effect in the application of energy to the exhaust

effluent.

A survey of the candidates underscores the selection of monopropellant

hydrazine as the I0C propellant. The components of the hydrazine system all

exist and have proven capabilities from previous space applications which,

together with the straightforward handling and tankage logistics, more than

offset the performance penalty.

58
oPI

i,

N

U

r",- I'_ I_ it1 _

i

__ _ _i _' i _

59
!

I

i

!

ql I

_4J

il,

r_m

_ QJ

c_

.J-4

U_

i-I

M

Q_ .a-) ,.-I _,l

UJ.rl

6O
2_r2_ 4.4.2.-2 I_CILelI,SICm b'l,_m RNQ_]G _ _ 3)

A. _ __

General Tec .hnology Readiness Levels

Level 1 Basic Principles Observed

and Reported

Propulsion System Equivalent

Requires significant scientific

advance. No date for accomplishment.

Level 2 Conceptual Design Formulated Initial analysis shows potential

application. More than 10-year

development.

Level 3

Level 4

Conceptual Design Tested

Analytically or Experimentally

Critical Function/Characteristic Demonstrated

System has been operated; up to

one decade of development

required.

New system in development.

mate 5 years minimum time to

readiness.

EstiLevel 5 _nent/Breedboard Tested

in Relevant E_vironment

System available in less than 5

years.

Level 6 Prototype/Engineering Model

Tested in Relevant Environmerit

Technology is established. System ready for flight test evaluation. Adaptations available in

less than 5 years.

Level 7 Engineering Model Tested

in space

Technology is mature. System

operating on existing spacecraft.

Adaptations available in less

than 5 years.

Class 1

B. COMPATIBILITY AS_

Thrust and Exhaust products are both compatible. System offers

potential synergies with Life Support or others which could

simplify on-board operations and/or logistic support.

Class 2 Thrust and Exhaust products are compatible. Little or no synergy

opportunity.

Class 3 Thrust and Exhaust products are compatible. On-board operations

or logistics would become more complex.

Class 4 Concern or potential limitation due to one or more of the

following:

a. Thrust Level

b. Exhaust Product

c. Complexity in operations or supply

61
4.4.2.1 Chemical Propulsion Systems

The chemical propulsion candidates present the best developed systems

available at this time. The energies available are those of combustion as

oxidation or combustional breakdown and are well defined for all of the

reactions involved. Improvem_ts in specific thrust, therefore, relate to

Improvements in translating the energy of combustion into kinetic energy in

the effluent streams_ e.g., burn and exhaust at higher temperature. These

considerations lead to development of improved pump6 or pressurants for the

fuel supply, higher operating pressures in the combustion chamber, and

improved resistance to erosion or degradation of nozzles when subjected to

the flowing hot gasses. All of the chemical systems described have been

built and tested, most have flown in one form or another, and three

represent the major storable-propellant systems presently e_ployed in space

vehicles. Therefore, within the ch_ulcal-based propulsion system summarized

in Table 4.4.2-1, the lower value for Isp corresponds to present

capabilitles; the upper value represents a plateau of potential achievement

(R_'m'lmce 3).

_onopro__llant H_razine

The catalytic burning of N2H 4 represents the mature technology for

modest thrust range (25-250 Pounds) applications. The production storage

and control systems for this fuel have existed for nearly two decades. Onorbit storage times have exceeded several years. For Space Station, the

developments must address high purity production of the N2H4 and operating

life for the catalytic beds. The presently defined specific impulse of 220

seconds is well establislmd. Devel_t of heated bed combustors is

underway (Rife_m:e 20) and shows a potential for increasing Isp to 300

62
seconds. A continuation of this development should achieve the goal in

sufficient time for application to the IOC Space Station or at least an

early update. The completion of the thermal augmentation developments

appears as the last significant performance increase for the monopropellant

system. Small improvements may continue; however, an Isp of about 300

seconds appears as a near-limit for the technology. Monopropellant

hydrazlne will remain the base for comparison in evaluating any other

candidate propulsion system.

The opportunity for a synergistic interaction has been addressed by

Boeing (R_L_m 21) as use of the N2H4 to supply makeup N2 needed to

resupply the losses in the cabin atmosphere. The catalytic breakdown of the

hydrazine and recovery of the N2 has the potential to eliminate the need for

transport of LN2. A makeup rate of five pounds per day or 500 pounds over a

resupply period amounts to about half a tank of N2H 4. The synergy would

result in some extra }{2 (utilization not defined) plus the need for on-board

gas separators and compressors.

The mme olics, or

The hypergolics are well understood storable bipropellants; both have

been used in space flights. The NT0-UDMH option has an extensive

applications history, including storable boosters (Titan II), multiburn

long-life space probes (Viking Orbiter), and maneuver-attitude control

systems (Shuttle OMS-RCS). The hypergolics as utilized in the Shuttle RCS

demonstrate pulsed thrust operations in the range 25 pounds to 200 pounds.

The present hypergolics provide specific impulses up to 330 seconds. The

potential for increasing performance requires operation at higher chamber

pressures and an improvement in nozzle performance. The vernier engines on

63
the Shuttle (25-pound thrust) could be the starting point for a development

program which could achieve a 20-point increase in performance to an Isp of

350 seconds. It is estimated that a concentrated development program could

provide advanced propulsion capability in less than five years. On the

other hand, the performance margin over the competitive monopropellant would

not appear to justify the extra cumplexity in tankage or resupply unless the

N204 became a major element in the makeup of N2 and 02 for life support. A

chemical decomposition and separation that yielded 500 pounds of N2 would

also provide more than 1100 pounds of 02. A comparison of potential

synergies relative to the present system permits an assessment of value for

justifying the development of the application.

I. The present system (Table 4.4.1-I) shows a total of 7,050

pounds of N2H 4 on-board with 2,320 pounds minimum at each

resupply. The on-board propellant storage consists of nine

tanks, each 40 inches in diameter.

2. In the case of N2H 4 augmented to produce an Isp of 300

seconds, the reduction in propellant requirement would amount

to about 1,900 pounds and 600 pounds less at each resupply.

If the N2H 4 provided 500 pounds of replenishment N 2, the

total resupply requirement would still be 100 pounds less

than the present baseline value.

3. In the case of hypergolics at an Isp of 360 seconds, the

total propellant weights would become 4,200 pounds, and the

resupply would become 1,420 pounds. If N204 became the

source for N2 replenishment, 500 pounds of N2 require about

1,600 pounds of N204. The resupply requirement would then

64
increase to 3,020 pounds. However, the 1,100 pounds of 02

would be the 90-day requirement for a six-man crew.

The synergies offered are both attractive for an update of the present

Space Station. The use of N2H4 as an N2 source offers a means to eliminate

the need for LN2 resupply. The use of the hypergolics offers a means to

eliminate the need for LN2 and L02 and the need for recycling of fi)2 or the

electrolysis of water. In a power-limited Space Station, the use of N204

with its extra complexities in tankage could be offset by the simplified

life support system requirement (Referemue 21).

Fluorine Bipropellants F2-H__, CH4___

Within the options for chemical propellants, the fluorine reactions

provide the maximum energies and translate into the maximum values for

specific thrust. Values of Isp to 500 seconds reflect the fluorine-hydrogen

reaction. The improvements in performance bring attendant probless of

handling a toxic, corrosive liquid and the chemical effects of HF in the

exhaust stream. The nuclear industry has developed a comprehensive

capability for handling and controlling fluorine in all of its forms and

phases (UF 6 is the transport medium for gaseous diffusion enrichment

separation of uranium isotopes); however, these techniques may not be

directly applicable to flight support or on-orbit operations. The copropellants of H2 and CH4 have the potential for synergy with the life

support system as H2 from electrolysis of water or CH4 as an interim product

in the Sabatier process for reduction of 002 (Refezemce 22). In either

case, the presence of HF as an exhaust product appears to preclude fluorine

as a candidate for an on-board propulsion system. On the other hand,

fluorine-based bipropellants should be considered as candidates for the

65
propulsion systems used in spacecraft assembled aboard the Space Station.

Therefore, the Space Station should be capable of handling and storing

tanked fluorine as part of a spacecraft on-orbit erection and deployment

sec_erEe °

0xyg_en Bipropellants, 02-N_4. CR4-__2

The oxygen based bipropellants offer the combination of high specific

impulse and potential for synergy with the life support system. The

products of combustion for these combinations are compatible with long-term

exposures for the Space Station; the only concern may arise from 02-N2H 4

where the potential exists for some NOx residuals. All of the systems have

been operated with 02-H 2 performing as the main propulsion system for the

Shuttle.

These systems may be considered as available technology but not

available as equipment items with the particular thrust levels or control

systems which correspond to the Space Station applications. Thrusters in

the range from 25 to 75 pounds have not been built. The high pressure fuel

supply pumpa required for opti_l_n engine performance do not exist; the

valves and control systems for low-thrust, minutes-of-burn applications have

not been configured. No obvious barriers appear to inhibit developing such

equipment; a focused development of the O2-H2 or O2-CH 4 systems could

support an update of the IOC Space Station. Both offer the potential for

synergy with the Life Support System, in particular the utilization of onboard electrolysis of water for an 02 supply and the Sabatier process for

reducing 002 which yields CH4 as an interim product. As indicated in the

comparison for the hypergolics, 02 replenishment for a crew of six would

process about 14 pounds of water per day (12 pounds of 02). At an Isp of

66
440 seconds, the reboost requirement corresponds to about 13 pounds of water

per day. The reboost propellant for an O2-H2 system would utilize the

necessary redundant capacity incorporated into the main 02 regeneration

equipment.

A comparison of tankage shows a considerable weight savings in either a

cryogenic or water-supplied approach. For an Isp of 440 seconds, the

resupply is half the present requirement or 1,160 pounds, either as 1,030

pounds of I/)2 plus 130 pounds of LH2 or as water. As water, the resupply

amounts to one 40-inch diameter tank; as cryogenics, the resupply becomes

one tank for 02 and two tanks for H2. The on-board storage tanks would not

reflect the same advantage. The need to maintain the contingency fuel in

cryostorage amounts to five cryogenic tanks as c_mred to one temperaturecontrolled tank for the hydrazine.

In summary, the O2-H 2 or O2-CH4 propulsion systems rank high as

candidates for Space Station on-board applications, particularly if

configured for synergy with a life support system based upon electrolysis of

water and potentially the Sabatier process for reduction of CO2. These

propellants offer a means for simplifying the resupply requirements at the

expense of electrical power for use in orbit and the need for on-board high

pressure or cryogenic storage. At the present time, propulsion components

do not exist in the size ranges associated with the Space Station

application. No barrier has been identified which would delay such

developments; Space Station quality and compatible equipment could be

available within a half decade.

67
4.4.2.2 Electrical Propulsion Systems (Referemue 3)

The electrical propulsion systems are characterized by generation of

thrust as a direct interaction with the electrical power supplied. The

interactions can be heating as in a resistojet or arc jet, electromagnetic

induction as in the magnetoplasmadynamic (MPD), or accelerations of charged

particles byan electric field (ion). Each of the systems shows specific

thrust values which exceed those for any chemical system. On the other

hand, the working medium provides very little energy of its own; therefore,

electrical propulsion systems present real-time demands upon the power

supply within the spacecraft. In addition, these systems are inherently low

thrust configurations and have limitations on continuous operation. Table

4.4.2-3 presents a stmmary of the thrust levels and electrical performance

ranges available or achievable within a decade. As a comparison, the table

includes a value for the thrust required to accomplish reboost as a

continuous force. In addition, the table includes a value for relative

electrical performance if an 02-H 2 chemical system utilizing on-board

electrolysis of _ater is used as the source for the propellants. The upper

values for Isp as shown in Table 4.4.2-I, coupled with both the achievable

thrust levels and electrical performance, show that a successful development

program will provide system potentials for accomplishing reboost and docking

at power requirements which are less than 10 percent of the available power

in the I0C Space Station. Electrical propulsion applications for such

operations would be defined in terms of thrust levels and duty cycles. On

the other hand, these systems would not have the capability for responding

to contingencies. The Space Station contingency planning envisions thrust

levels in the range of I00 pounds sustained for minutes, and such

68
System .

Ion or

charged

particle

Resistojet

Arc Jet

MHD

CCMPARISC_:

Baseline

continuous

thrust for

reboost

02-H2 chemical from onboard electrolysls of

R2o

Thrust

Pounds

Electrical Perfo_

Thrust/kWe Sur_ied

Appl icat ions

Considerations

0 to2.5 N

(0 to 0.5 ibs)

0 to 50 N

(0 to 10 ibs)

0 to 50 N

(0 to I0 Ibs)

0 to 500 N

(0 to 100 lbs)

0.013 to 0.09 N

(0.003 to 0.020 lbs)

0.13 to 0.31 N

(0.03 to 0.07 ibs)

0.036 to 0.18 N

(0.008 to 0.04 ibs)

0.013 to 0.09 N

(0.003 to 0.02 Ibs)

Acceleration voltage

gradient limited by

arc breakdown condition

Temperature limit by

material and convection heat transfer

Materials limit in

arc plasma environment

Materials limit in

arc plasma environment

0.27 N

0.062 ibs

0.202 N

0.0452 ibs

(483,000 lb. sec. delivered as a continuous thrust for 90

days)

(Electrolysis at 70%

efficiency, thruster

Isp of 440 sec.)

69
performance capebilities are not possible with electrical systems. The

Space Station utilizatlon for electrical propulsion systems must be

addressed in terms of candidates for low-force applications used in

combination with a standby higher-force system for contingencies.

Electrostatic Field Acceleration _Systems (Rsferem_e 23)

Thrust generation by an electrostatic field includes the acceleration

of both ionized atoms (ion propulsion) or charged particles (colloids). The

NASA development program has focused upon applications that require thrust

levels in the millinewton (millipound) range and address precision attitude

control or station keeping at G_O. These types of thrusters have the

capability for relatively long-term sustained operations. The present goals

focus upon _@roved electrical performance with the intent to reach the 0.09

(0.02 pounds) newton-per-kilowatt levels within a decade. The thrusters

show attractive values for specific impulse but present inherent limitations

in attainable performance. They operate most effectively with heavy atoms

or colloids; therefore, not all materials can be used. In addition, three

operational features either impair the electrical efficiency or limit

thrust. First, all the thrust results frcm placing charged particles into

an electric field; consequently, the system must llve with limits on

available field strength (arc breakdown) and accept the energy losses

associated with charging the thrust media as particles or atoms. Secondly,

there is a limit to the nmNber of atoms that can be given an electric charge

within the thruster. Finally, once the atom or particle leaves the

accelerating field, it must be neutralized or a reverse acceleration will

occur (positive ions are attracted to the negative plate from both

directions). In assessing the long-term development for these types of

70
thrusters, they will remain attractive for applications requiring low

thrust. An order of magnitude increase in either thrust or electrical

efficiency does not appear realistic. A colloid thruster that could eject

trash generated from other on-board systems appears a possible Space Station

utilization for electric-field type accelerators.

Res_stojet _Systems (R_ere_ 2)

The reslstojet systems may be considered as a cold gas bl_own passing

through an electrically heated tube or through a grid. The heat transfer

surfaces provide the means for increasing the kinetic energy in the effluent

stream. Resistojets offer the best available electrical performance coupled

with the least sensitivity to the propellant constituents, while at the same

time providing specific impulse values which double those of chemical

systems. The system has i_erent limits imposed by the maximum operating

temperatures for the heat transfer surfaces and the attainable convective

heat transfer from the surfaces. A conservative long-term assessment of Isp

improvement suggests as much as 25 percent from temperature rises, and this

would translate to a 50 percent _crease in electrical efficiency. To the

Space Station, the resistojet provides an opportunity for attitude control

thrusters and possibly a two-t_s-per-orbit re_st that could utilize

trash gases or constituents in synergy with the life support system.

The arc jet augments a cold gas bl_uwn by passing the gas stream

through the plasma of a sustained arc. The stream temperature achieved in

passage through the plasma effectively determines the output Performance of

the system. The temperatures within the arc will dissociate compounds into

elements and produce an ionized elemental exhaust stream whose features

71
relate to the output limits of the system. Ionization and chemical

dissociations are unrecovered energy losses. Since arc plasmas exceed the

vaporization temperatures for any material, sustained arcs erode the

electrodes; consequently, most arc jets operate in a pulsed mode.

For a Space Station, an attractive application would utilize the

erosion as a source of material for ejection and become a trash burner for

overboard alL.ping of residuals from other systems. For instance, the Bosch

process for (D2 reduction results in elemental carbon, which could possibly

consumable electrodes in a gas blow, own arc jet.

Mauneto_Dlasm_mic (MPD) (Reference 24)

MPD propulsion utilizes the magnetic field generated by the current

flowing through an arc to accelerate the atun_ charged by the plasma of the

arc. MPD thrusters utilize concentric circular electrodes (e.g., a ring and

a cylinder) and create an arc at one end such that the gas flowing in the

circular annulus moves into the plasma and becomes ionized. The

circumferential magnetic field in the annulus as generated by the arc

current provides the electromagnetic interaction that accelerates the

charged atoms out of the annulus. Because of the electromagnetic

interaction, the gas stream does not require the pressurization needed for

resistojet or arc jet operations or the neutralizer of the ion system. On

the other hand, the system requires large currents to generate the

accelerating forces. The current and the plasma define the inherent

limitations for MPD thrusters. The electrical power source must provide for

arc current levels ranging from 5,000 to 40,000 amperes with 40,000 amperes

as the current density limit consideration. The operating voltages

associated with such arcs range from i0 to 50 volts DC. The instantaneous

72
power consumption then ranges from 50 kWe upward into the megawatt levels.

In the same manner as arc jets, the plasma temperatures exceed vaporization

for all materials and necessitate pulsed operation. However, the best

transfer of energy into the gas stream occurs at the time of initiation;

therefore, MPD thrusters benefit by pulsed capacitor-discharge operations

where arc times are in milliseconds at repetition rates of two to three per

second.

At the present time, MPD units have only operated with argon as the

flowing gas. While no limitations have been specifically identified

relative to the gas utilized, the noble gases have the advantages of being

both monatumic and inert. The performance projections for MPD thrusters do

not account for any compromises introduced by chemical dissociations in the

gas stream.

Summary of Evaluatio_q

The electric propulsion systems show attractive specific impulses.

However, thrust levels and operating limits make them marginal candidates

for application to the Space Station. The propulsion syste,_ all have other

potential applications which will continue to justify development

activities. The inherent limitations for these systems will keep them in

the low thrust range. However, a significant development which improves

electrical performance, coupled with an opportunity for synergy, may make

one of these systems an attractive choice for attitude control or reboost

application.

73
4.4.2.3 Propulsion by Other Techniques or Transfers of Energy

The four propulsion concepts listed in Table 4.4.2-1 represent

adaptations or applications of high energy sources to propulsion systems. A

significant development program addresses each of the technologies but none

specifically address the propulsion applications. In addition, each of

these technologies must achieve a major advance before a propulsion system

can be configured. The potential performance capabilities for each of these

systems make them candidates for application once the technology progresses

to the point where an application can be defined.

Electromaunetic" Mass Accelerators

The mass accelerator offers the advantages of a controlled thrust

generated by recoil from the ejection of almost any material. Launch

systems based upon mass accelerators using local material appear in the

descriptions of lunar base or planetary return systems. The present

national efforts to improve the overall performance in both electrical power

construction and eject velocities could bring the technology into maturity

within the next half decade. The concept has the potential for synergy with

life support by ejecting solid materials that need an overboard dump such as

the carbon from the Bosch reduction of 002 or other trash from crew

ac_tions (Referemces 3 area 23).

Ta_er R_ated Systems

A high temperature, high velocity gas stream generated by vaporizing

material with a laser presents an attractive option for a propulsion system.

The overall performance of the system relates to the energy efficiency of

the driving laser. The present national effort for development of lasers is

intended to provide systems which can produce beam energies and pulse

74
_p

durations compatible with a laser-heated ejection system and within the

power capabilities of the Space Station. A laser heated system would have

the potential for ejecting any form of trash generated on-board the Space

Station and ejecting the material in elemental forms (most probably as

charged ions).

Mass Conversion as Antimatter Annihilations (_ 25)

Antimatter generation and antimatter interactions represent major

investigations utilizing the high ene_jy particle accelerators on earth and

measurements from cOSmiC ray bombardments in high altitude experiments.

Antimatter is a well established concept, and the annihilation interactions

have been characterized for the products formed and energies released.

Among the researched candidates, the proton-antiproton interaction has been

identified as having the potential for controlled energy production and

application to propulsion. The interaction creates energetic subatomic

partlcles which in turn can energize a beem of protons into motion at

relativistic velocities. The concept for propulsion envisions the "burning"

of antiprotons in an excess of protons (hydrogen) with the mass ratio

balanced to optimize the kinetic energy delivered into the exit beam of

protons. The present source for antiprotons occurs with proton-proton

collisions at energies above 6 GeV. The antiprotons formed can be stored in

a magnetic ring in the same manner as regular protons can be stored and

accelerated in cyclotrons or similar units (Reference 26). In assessing the

availability of antimatter technology for Space Station application, the

developments appear to parallel those for controlled fusion. Applications

of antimatter will require major scientific advances in the technology for

75
generation, storage, and energy controls and none permit scheduling at this

time.

Free Radical Eneruv

The energy release, associated with free radicals (e.g., H+-OH -)

combining to form a stable molecule, has been defined and represents an

upper limit for purely chemical energy production. Free radicals have been

generated and controlled in laboratory environments. The combination of

particular interest involves the transformation of monatcmic hydrogen into

H 2 molecules and exhibits an energy about one order of magnitude higher than

combustion of hydrogen with oxygen (Reference 23). The developments

required before application to the Space Station include means for

generating free radicals in quantity plus the means for containment and

control of the combining reactions. The potential for application exists;

however, the enabling technical achievements cannot be scheduled at this

time.

4.4.3 Candidate Propulsion System for Advanced Space Station and

Recummendations

The requirements for propulsion remain essentially unchanged from the

those of the IOC Space Station. There will be a requirement for orbit

reboost and attitude control that is best served by thrusters providing from

I00 to 1,000 N (25 to 250 pounds). The requirement for contingency-based

velocity changes will continue. There will be no relief from the surface

contamination requirements which dictate exhaust product cleanliness.

Finally, resupply will need a continuing optimization. The recommendations

76
recognize these requirements and present a Space Station position relative

to all the candidate technologies for propulsion.

The propulsion system reccmsended utilizes 02-H2 to provide for orbital

reboost and attitude control supplemented by a cluster of small solid

rockets to provide the contingency boost requirements. The entire system

would operate in synergy with the life support system. _he reboost and

attitude control thrusters together with the Orbital Transfer Vehicles will

utilize 0 2 and H2 supplied from the electrolysis of water in a system which

supplies the makeup of atmospheric 0 2 and the reduction of CO2,

respectively. The generation of propellant becomes a major utilization of

on-board power. However, • the electrolysis cycle may be configured to

provide a means for achieving a oonstant load for the power system. _he net

effect is a near-optimum power generator operating with 02 and H2 produced

by electrolysis in proportion to the excess electrical power available at

any particular time. The storage of propellant requires maintaining some

on-board cryogenics with the H2 having some interim storage as recoverable

metallic hydrides.

4.4.3.1 Development Efforts Required and Supported for 02-H2

The utilization of O2-H2 for reboost and attitude control will require

the development of the following equipment items.

i. Thrusters producing I00 to 1,000 N (25 to 250 pounds) of

thrust for up to 20 seconds at an Isp of 440 seconds. The

units will need to show long-life capability for such

operation.

77
2. Control elements such as valves, throttles, actuators,

acctm_ators, etc., capable of operating with 02 or H2 under

high pressure, low flow conditions.

3. Pumps, valves and storage systems capable of handling H 2 and

02 generated by electrolytic cells through the steps of

compression, liquification, storage, and recovery.

4. Metal hydride storage and recovery technique for H2.

5. R_uote handling and transfer of cryo tanks between the Space

Station and the Orbital Transfer (or other spacecraft)

Vehicle.

The propulsion system could share development efforts with the life

support system for the water electrolysis cells and those elements of

equipment that reclaim _ter for electrolysis.

4.4.3.2 Contingency Propulsion

The needs for contingency propulsion can be met by clusters of small

solid propellant thrusters selected for minimal contaminants in the exhaust.

Anm_nium nitrate appears as the initial candidate fuel for the application.

The configuration for the rocket motor envisions a reusable case that can be

reloaded on-board, and resupply would consist of preformed grains and

igniters.

The use of ammonium nitrate as a fuel offers the potential for synergy

with the life support system as an on-board source fOE N2, H 2, and H20

through a chemical decomposition. The material shows only a modest Isp of

175 seconds; however it is produced in quantity as a constituent of

commercial fertilizers.

78
4.4.3.3 Recommendations Relative to Other Techniques and Technologies

The present activities and developments of propulsion systems or

technologies applicable to propulsion systems cannot be ignored in the

course of developing improvements to current or proposed new Space Stations.

Improvements will occur that may elevate a system or technique into a

preferred or advantageous category; consequently, attention is required to

utilize potential advantages from system technology advances.

Mnncg_ro_Dellant _H_razine

The efforts to uprate the present system should continue with an

augmented system ready for the first update of the IOC station.

The HvDercolics N_I4zL_W___

The flight experiences and operation with these thrusters and their

control systems should receive a continuous technical review to identify

effects which would benefit (or impair) their application to the Space

Station. The potential for synergy with the life support system should be

included in the study rexmmended for h_a_razine.

Fluorine Based Bipropellants

Monitor the development efforts toward defining specific requirements

for Space Station on-board handling and storage.

Oxyg__enBased Bi_Dr_opellants

The O2-H 2 system is the recogmsnded development (see Section 4.4.3

above). A study is re(xmmended to determine any potential benefit from

synergy with CR4 from the Sabatier (192 reduction process. Evidence of

significant benefit would then lead to a study toward defining a dual fuel

thruster that could operate O2-H2 or 02-CH 4.

79
F/ect ric _Systems

Monitor the results of developments for evidence of significant

improvements in thrust levels or electrical efficiencies that would offer an

advantage to the Space Station. An improvement in either resistojets or arc

jets would lead to a study for synergy with the ejection of materials

intended for overboard disposal (e.g., trash from life support).

Other Concepts

Monitor the results of research for technical advancements which would

bring the concept to the maturity level necessary before a propulsion

application could be configured.

4.5 Thermal Control Subsystem

Thermal management of the Advanced Space Station will be employed to

regulate the temperatures of various elements and subsystems plus their

environments to assure proper operating performance and reliability of

Station equipment and customer payloads in space. All power generated on

the Space Station must ultimately be rejected as waste heat.

The temperatures that may be encotmtered could range from 4 K to over

500 K with control bands ranging from 0.01 K to i00 K and heat loads or

leaks ranging from 0.i W to over i00,000 W (Referemoe 3).

4.5.1 State-of-the Art

The state-of-the-art thermal control systess currently available are

directed toward the slx control processes as follows:

1. Heat transport by way of pumped liquids and/or heat pipes

that operate on fluid phase change.

8O
2. Heat rejection by pmels that radiate to space the heat

transported to them.

3. Temperature control achieved with valves that govern the flow

of thermal fluids or variable conductance heat pipes.

4. Thermal storage achieved by transferring sensible heat plus

heat of fusion to appropriate substances contained within a

heat exchanger.

5. Refrigeration accomplished using thermoelectric cooling,

expendable coolant, or by mechanically implemented

thermodynamic cycles.

6. Thermal interfaces may require quick disconnects in fluid

lines and flexible hoses and rotary seals for joints that

rotate. Thermal switches and mechanical joining of structure

to exterior module walls coated with thermal control coatings

can provide controlled heat flow paths.

The state-of-the-art for thermal control is identified in Table 4.5.1-

1 by figure of merit, and the projected performance is forecast through the

year 1991 (Rmfezemoe 3).

The Space Station (IOC) will utilize both active and passive thermal

control systems. The active systems will include two-phase ammonia thermal

buses, capillary pumped loops, single-phase pt_ water, evaporator cold

plates, and two-phase heat pipe radiators. The passive systems will include

coatings and insulation to minimize supplemental heating requirements.

A centralized heat collection and transport function employs two

thermal buses to transfer heat by evaporation and condensation, thereby

maintaining a constant temperature over their entire lengths. One bus will

81
Figure of Merit

Spacecraft Thermal Power

Requirements

Spacecraft Thermal Control

Specific Power

Radiator Specific Mass

Radiative Cooler Focal Plane

Temperature

Heat Pipe Thermal Transport

Capacity

Spacecraft-Bome Cryogenic

Refrigerators

Cooling Capacity @ Temperature

Service Life

Connectable Thermal Interfaces

Spillage

Life

Rotating Thermal Joints

Flexible Hose

Thermal Slippings

Projection

SOA Forecasted Yea____r

<i00 kW 100kW 1991

9 Wlkg 18 W/kg 1991

18 kg/kW i0 kg/kW 1991

75 K 60 K 1991

105W-in 107 W-in 1991

1 W @ i0 K 0.i W @ 4 K 1991

1 yr 5 yr 1991

0.66 cm 3 Hone 1986

500cycles 500cycles 1986

104 cycles 180o 4x106 cy 180° 1986"

-- 2x106 rot'tns 1986"

*Space Station Technology

82
operate at 70° F and the other at 35o F. Single-phase pumped water circuits

maintain temperature control in the modules and interface with the thermal

buses outside the modules, thus avoiding the possibility of toxic ammonia

gas being released in the module. The thermal buses reject unwanted heat by

using an erectable centralized heat pipe radiator system (l_ferm_e 7).

4.5.2 Technology Trends

Thermal management technology trends are oriented to centralized

thermal control whereby Individual equipment can be integrated and

reposltioned with minimal effect on the central systems' ability to serve

the remain_ loads and allow for Space Station growth. Advances may

provide a time-shared programmable central controller to simplify the

control of heaters and thermostats and their associated control problems.

High performance cold plates (I-i0 W/cm 2) and distributed evaporators

(multlzone heat input) are needed to transport heat fluxes over greater

distances than today's 3,000-12,000 W design capabilities. Lighter heat

rejection radiators are envisioned as modular with on-orbit replaceable heat

pipe elements (Figure 4.5.2-I). Mechanical refrigerators are required

having a three- to first life with stable Performance at <75 K. Solid

state temperature controllers are needed to replace bi-metallic thermostats

because prelaunch functional verification of the large nmber of thermostats

required is very difficult.

Thermal control technology forecasts are provided in Table 4.5.2-1

(l_'ezcmce 2).

83
SPECIFIC IMSS

(kg/kW)

40--

30-

ZO -

10 -

SHUTTLE

SPACESTATION

n i i i i

1870 1975 1980 1985 l_JO 1995

TECHNOLOGYREADINESSDATE

- 20

- 25

- 33

- 50

-- 100

2O0O

SPECIFIC POWER

(_/ks)

Figure 4.5.2-1 Radiator 8pooifio Mass andSpecific Power

(Roforenoe 3)

84
TABLE 4.5.2-1 THERNIL CONTROL TECHNOLOGY FORECASTS

(UrZnNCS 3)

Figure of Merit 198.___55 199___O0 199__5 2000

Specific Power (W/ks)

Thermal Control S_stems

10 15 20 25

Radiators

Specific Power _W/ks)

l.l<,=P 'l;'luv [IrW/m"_

75 90 ....

0.5 0.7 !.0 2.0

Thermal Transport Capacity

(w-a)

Heat Pipes

1 x 10 4 2.5 x 105 --

Pumped Loop Heat Transport S_stems

Thermal Transport Capacity

(W-m)

HeaC Flux (Wcm 2)

5 x 103 5 x 104 I x 105 --

I0 15 30 75

Heat Flux (W/cm 2)

Cold Plates

0.5 2.0 I00

Passive Cryogenic Coolers

Lifetime (yr)

Radiators

Sollds

Liquid Tanks

4 7 10 12

4 5 6 7

2 3.5 5 6

Ac¢lve Cryogenic Coolers

Lifetime (yr)

65 K Stlrllng Cycle

II K Stirling Cycle

1 4 6 7

0.25 1 3 4

e.

85
4.5.3 Candidate Advanced Subsystem

The transport of A_vanced Space Station hardware from Earth to low

Earth orbit will be limited in size by the cargo bay of the heavy lift cargo

vehicle. The largest payload size anticipated is 33 feet in diameter by i00

feet in length.

The A_vanced SGace Station will employ both active and passive thermal

control subsystems to assure that proper operating environments are

provided.

The following six candidates are active thermal control systems.

i. Pumped flow two-I_msed thermal buses provide controlled heat

to selected Space Station locations and transport unwanted

heat to externally mounted space radiators. The thermal buses

will function similar to a two-phase heat pipe as shown in

Figure 4.5.3-1 but with a mechanical pump to return liguid

ammonia from the space radiator end to the heat input

locations (see _ferenue 27 also).

2. Capillary pumped loop technology provides a high heat flux

capability (up to 15 W/cm 2) to transport unwanted heat from

an experiment to a thermal bus (1_fereru_ 28) (Figure 4.5.3-

2).

3. TWo-phased ammonia heat pipe radiator elements from 50 feet

to 100 feet in length would be space-erectable to form largearea flat space radiators. Micrometsoroid or space debris

damage to the radiator assembly could be repaired on-orbit by

replacing the individually damaged radiator elements. Watt

densities of 3.1 W/cm 2 and heat transport levels of 14 kW86
HEAT INPUT HEAT OUTPUT

fill L .o.o.

HUT INPUT HUT OUTPUT

Figure 4.5.3-_L Banio Heat Pipe (Re£erenoe 3)

HEAT SOURCE /--VAPOR /--VAPOR r--HEAT SINK

._../._..=:=:===:..::;===::=====:=._.=:===_.===:===:.._.._:==...=..=.==..v==_. Fr.'._::.:';_'::.._ /RESERVOIR-_

<_ _.........................._

Figure 4.5.3-2 8ohematio of Capillary Punpe4 Loop Bngineer/ng Model

(Reference 28)

87
e

.

o

meters have been measured for a monogroove high-capacity heat

pipe. An Advanced Trapezoidal Axial Groove heat pipe having

a nominal outside diameter of 3.75 cm has a theoretical

transport capacity of up to 150 kW-meters (l_ferenoe 28)

(F_jure 4.5.3-3).

Variable conductance heat pipes used with experiments mounted

outside the Space Station modules can provide a variable

radiator feature reducing the heater power required during

cold-case operation (and non-operation) and yet allowing

full, unchoked heat rejection in the hot case (Refer.lee 28)

(r .re 4.5.3-4).

Electronic circuit board thermal control appears feasible

using a small ammonia two-phase heat pipe joined to the

perimeter of the board. The heat pipe can provide thermal

c_trol to chip-size piece parts and sometimes mechanical

parts (Refez_oe 28) (Figure 4.5.3-5).

Liquid droplet radiators can provide high heat rejection

rates using a reclrculating free stream of liquid droplets

(<i00 um diameter) to radiate heat in space. The basic idea

of the radiator concept is the large radiating surface area

provided by small droplets having a minimal mass. Low vapor

pressure silicone oil and molten low-fusing metal alloys are

candidate radiator droplet materials selected to minimize

offgassing and condensing on other Space Station surfaces.

The liquid droplet radiator can be constructed much lighter

88
56.9

I 1.55

r

f

(VAPOR

CHANNEL) ]5.01

(NONOGROOU O.Z3

SLOT)

(LIquID

CHANNEL) 9.98"__

• SPIRAL f

• 14ATL: 6061-¥6 ALUMINUM

• DRY tIT: 0.86 kg/a

• LIQUID VOLUME: 78.1 ml/m

• VAPOR VOLUME: 179.1 ml/m

• ALL DIMENSIONS IN am

Figure 4.5.3-3 Xonogroove Heat Pipe. Configuration (Reference 28 )

_,,_ L TOT

e

_ L EFFECTIVE

EVAPORATOR .J _ _ '_ _ '_ "_ _ /"-WICK

_e _ __ _lr "-] _I---CONDENSER---_ /

XORKING FLUID (NH3) _NON-CONDENS]BLE GAS (e_

6,) d

i I _I

I I

L' L. r I

....... CONTROLLER

TEHPERATURE SENSOR _ ...... ..;

I--STOOGE VOLUHE

ioo

t_2._Dt!

_J

(FOR ACTIVE CONTROL ONLY)

Figure 4.5.3-4 Typical Variable Conductance Heat Pipe

(Reference 28)

89
127 m

136 m

146 i

Figure 4.S.3-S circuit Board Heat Pipe (Reference 28)

DROPLET

COLLECTOR

POWER

DROPLET GENERATOR

Figure 4.5.3-6 Liquid Droplet Radiator (Reference 3)

9O
than heat pipe radiators, and the radiating area is immune to

micrometeoroid damage (1_fezes_e 3) (Flgure 4.5.3-6).

A liquid droplet radiator point design indicates a power to mass ratio

of 4.2 kW/kg which is two orders of magnitude improvement over heat pipe

radiators (l_fezemoe 11).

The following three candidates are passive controlled thermal systems.

1. Thermal control coatings and finishes applied to surfaces of

the Space Station will provide selective reflectances or

absorption of incident energy and selective emittance of

self-contained heat. Contamination-resistant coatings are

required to offset the effects of photolysis whereby

contaminants adhere more readily to sunlit surfaces. Atomic

oxygen, heat, electron fluency, and ultra-violet radiation

are other degrading environmental conditions affecting the

performance and life of thermal control coatings (R_erer_s

28 and 29) (F_jure 4.5.3-7).

2. Louvers used in combination with therml control coatings can

vary the Eadisting area of a surface, thereby modulating heat

transfer. The louvers resemble a venetian blind being opened

or closed by a thermally responsive bi-metallic spring

(Reference 28) (Figure 4.5.3-8).

3. Multi-layer insulation consisting of about 25 layers of

embossed an(] metallized plastic film interleaved with low

conductance separators can provide emittance values as low as

0.005. Seams, cutouts, and attachments increase the effective

emittance values (Reference 289 (Figure 4.5.3-9).

91
SOLAR ABSORPTANCE(aS)

/

/

/

o.+ - ALU,..tmzmeOLYI.IDE dill.-/+ "-'-- 7

- _Ikl_ llb-ww- ///

0.5- _ -- / ....

CONDUCTIVELY COATED IIm_"" "-"

0.4 TOLERANCE LIII|T

ON TRANSIT SAT

•/ SOL FOR /

m|l ) RADIATORS __..... --... -- -

TEFLON FEP (2 0.1 . •

o I ]+°YE"m l I "_

l 10 100 1,000

FLIGNT DURATION, DAYS

I

10,000

Ficj1]re 4.5.3-7 Extrapolated End-of-Life _bsorptance (Reference 3)

ACTUATOR ADJUSTMENT SCREM-_

BASE'-_

TRUCTURAL FRAME

LOWER BLADE (TYPICAL)

/

BONDEDFLANGE"-_- /

LONER HOT SECTiON_

UPPER HOT SECTION "-"

ACTUATOR HOUSINGADJUSTMENT

CYLINDER

ACTUATOR SPRING

Figure 4.5.3-8 lairchild and Northrop LouWor ]_•embly 8ohanatie

(Reference 3)

92
EFFECTIVEEIIITTNEE

.O6

.01

.005

.001

.000S

\

OAO--INSTALLED

:) \

0 _ 0/- _-II_TN.LED 0

_. c_-_--,m_,_

(_ LII- -BI.NKET TESTS

_*_ INCL SUPPOItTSVBIITS MiD SF.NIS

^0 \ 0 ,_ 0

I I I I I

20 40 60 80 100

IUSF.R OF ALUMINIZEDI'Ot.¥ESTF.tLAYERS

Figure 4.5.3-9 Multilayer Xnsulation PerformanceData

(Referenoo 3)

93
4.6 Environmental Control and Life Support System (ECLSS)

The functions of the ECLSS provide for:

IQ

2.

3.

4.

5.

6.

7.

Atmosphere pressure and composition control

Module temperature and humidity control

Atmosphere revitalization

Water management

Waste management

EVA equipment servicing

Fire detection and suppression

4.6 .i State-of-the-Art

The state-of-the-art in ECLSS is very similar to that of the Orbiter

and Skylab. Each of the seven functions is discussed briefly with proposed

changes from Orbiter and Skylab noted. In addition, Table 4.6.1-I

summarizes the status of pertinent equipment items or options considered for

application to the I0C Space Station (Rsferer_e 21).

I. Atmosphere Pressure and Composition Control (ACS) - The ACS

system provides partial pressure control within the

pressurized modules. It contains the oxygen and nitrogen

storage and resupply tanks and also the distribution

equipment throughout the pressurized capsules.

2. Module Temperature and Htmidity Control (THC) - The THC

controls cabin temperature, humidity, ventilation, and

cooling where required.

3. Atmosphere Revitalization &_PJ -The AR system revitalizes

the cabin atmosphere. It contains hardware used to close the

94
i

i

i

95
iu

96
!i

97
o

I I o o o

_ o _ _o_o_

| _o_ _o o_

_D

98
e

e

e

e

crew's oxygen loop and is a departure from Orbiter/Skylab.

Water is electrolyzed to provide metabolic oxygen. The byproduct hydrogen is used in a (D2 reduction process to yield

_ater and reaction by-products. The water is used to

replenish the crew's potable water supply, and the carbon byproduct is returned to Earth. The carbon dioxide produced by

the crew is routed to the (D2 reduction unit.

Water Recovery and Management - Water recovery and reuse is a

departure from Orbiter/Skylab. Water reclaimed from hygiene

and urine is recovered for use for hygiene purposes.

f_claimed water is electrolyzed in order to resupply the

metabolic oxygen, and the reclaimed condensate is the primary

source of the crew's potable water supply (along with the (D2

reduction water output).

Waste Management - Waste management consists of collecting

the urine waste water and collecting and processing the fecal

and trash waste. Use of the urine waste water has been

mentioned above. Fecal waste is vacuum dried, bagged, and

returned to the ground. Trash is collected and also returned

to the ground.

EVA Equipment Services - The ECLSS system provides fluid

services to the _4U and M_U.

Fire Detection and Suppression - This subsystem consists of

hardware to detect and suppress fires in pressurized volumes.

The technology is similar to that of the Orbiter/Spacelab.

99
F_ure 4.6.1-1 is a schematic drawing of the partially closed BCLSS.

The BCLSS is sized to accommodate the design loads of Tables 4.6.1-2 and

4.6.1-3.

Numerous studies have been made to examine methods for performing some

or all of the ECLSS functions, some experimental studies, and many

analytical studies (see Referemues 30 through 33, for example). It should

be noted that experimental studies have been made under Earth "g"

conditions. There appear to be no readily available results from tests at

zero or reduced "g" conditions, which could be important in some of the life

support functions.

4.6.2 Technology Trends

There are two major trends in the BCLSS. One is to improve the overall

recovery of metabolic oxygen and potable water by proper selection of the

processes for these purposes and also to examine relative costs (Referemue

3@, for example). The other trend is to make use of synergistic

relationships between the ECLSS and other subsystems. The two trends are

discussed in some detail in Sections 4.4.3 and 5.2

4.6.3 Candidate Advanced Subsyst_n

Life support functions on-board an Advanced Technology Space Station

will have to perform at levels that accommodate the particular crew

complement. The configuration definition studies that led to the concept

described below in Section 6 recognized that gravity would eliminate some of

the design complexities imposed upon the IOC equipment. The concepts for

synergy in the generation of 0 2 and H2 for both life support and fuel impact

100

4.6.1-2 BUaS 1ZSIGll ]r/]N_ _ 7)

General:

Number of Crew

Resupply Interval (days)

8.00

90.00

THC Subsystem:

i

Sweat and Respiration H20 (Ibm/man-day)

Metabolic and Sensible Heat (Btu/man-day)

Hygiene Latent H20 (ibm/man-day)

Food Prep. Latent H20 (Ibm/man-day)

Laundry Latent H20 (ibm/man-day)

Freezer Internal Vol. per Module (ft3)

Refrigerator Internal Vol. per Module (ft3)

Freezer Load per Module (Watts)

Refrigerator Load per Module (Watts)

Dish Wash Latent H20 (Ibm/man-day)

4.02

11,200.00

n_D)

nmD)

nmD)

_D)

nmD)

n_D)

0.06

_ES System:

Total IOC Station Atmosphere Leakage (ibm/day) 5.00

Air Lock Gas Loss (10% of Leakage gas/operation)

U.S. Module Volume (ft3) 6,166.00

AR Subsystem:

Metabolic 02 (Ibm/man-day)

Metabolic 002 (Ibm/man-day)

1.84

2.20

FDS Subsystem:

n_D)

W_4 Subsystem:

Potable H2O, Drinking (ibm/man-day)

Food Preparation H2O (ibm/man-day)

Metabolic H2O (ibm/man-day)

Clothes Wash H20 (ibm/man-day)

Dish Wash H20 (1bin/man-day)

Hand Wash H20 (ibm/man-day)

Shower H20 (1bin/man-day)

Urinal Flush H20 (Ibm/man-day)

Urine H20 (ibm/man-day)

Food H20 (ibm/man-day)

Hygiene Latent H20 (ibm/man-day)

4.09

1.58

0.76

27.50

12.00

4.00

8.00

1.09

3.31

1.10

0.96

102
Table 4.6.1-2 _X,SS ]]_I(_ _ (Coneluc]ec])

Subsystem:

Food Packing (ibm/man-day)

Urine Solids (ibm/man-day)

Fecal Solids (ibm/man-day)

Sweat Solids (ibm/man-day)

Wash H20 Solids (percent)

Shower/Hand Wash H20 Solids

Trash (ibm/man-day)

Trash Volume (ft3/man-day)

Fecal H20 (ibm/man-day)

_U Condensate (Ibm/man-hr)

ES Subsystem:

EVA 02 (Ibm/man-hr)

EVA 002 (ibm/man-hr)

EVA H20 (Ibm/man-hr)

_4U Condensate (Ibm/n_n-hr)

(percent)

1.00

0.13

0.07

0.04

0.44

0.12

1.00

0.i0

0.20

_mD)

0.18

0.22

(TED)

(TED)

103
Parameter

0O 2 Partial P

Temperature

Dew Point (2)

Potable Water

Hygiene Water

Wash Water

Ventilation

o2 Partial

Pressure (4)

Total Press (5)

Dilute Gas

Trace Contam (8)

Microorganisms

Units

deg F

deg F

]b/man-day

ib/man-day

ib/man-day

ft/min

psia

paia

m

rag/m3

C_3/m 3 (6)

Operational

3.0 max

65-80

40-60

6.26-11.35

12 (3)

28 (3)

15-40

2.83-3.35

14.5-14.9 (9)

N2

90-day

Degraded

(i)

7.6 max

60-85

35-70

6.26 (3)

12 (3)

0

10-100

2.4-3.45

14.5-14.9

N2

TED

1000 (7)

T-_D

1000 (7)

28-day

Emergency

12 max

60-90

35-70

6.26 (3)

0

0

5-200

2.3-3.45

14.5-14.9

N2

TED

i000 (7)

(i) Degraded levels meet "Fail Operational" criteria.

(2) Relative humidity shall be within the range of 25-76

percent.

(3) Minimum

(4) In no cases shall the 02 partial pressure be below

2.3 psia, or the 02 concentration exceed 23.8% of the

total pressure.

(5) All EK_SS subsystems shall be compatible with 14.7

psia total pressure.

(6) CFU - Colony Forming Units

(7) These values reflect a limited base. No widely

sanctioned standards are available.

(8) Based on N_3 8060.113 (J8400003).

(9) Hyperbaric chamber pressure shall be TBD psia

minimum.

104
the design of equipment. In such a context, four areas of technology appear

to offer particular benefits pertinent to life support functions. These are

stm_arized in the following four sections.

4.6.3.1 Semipermeable Membrane Development for Gas Separations

Semipermeable membranes are well established and provide the basis for

hyperfiltration and the molecular sieves used to separate 002 from the cabin

air. Developments in membrane technology are continuing for medical

applications (renal dialysis, oxygen permeable contact lenses, etc.) and gas

constituent separations. Microgravity processing appears as a potential

means for producing extra-thin films with controllable and uniform porosity.

These developments would lead to a straightforward separation system as a

series of membrane filters, each rejecting (or selectively passing) a

particular molecular constituent.

4.6.3.2 Wet Air Oxidation for Water Reclamation

Oxidation of carbonaceous materials in a water solution is a recognized

method for disposal of such wastes. The process has been developed, applied

commercially, and studied for application to flight equipment. In the

context of synergism, wet air oxidation is a recommended development effort.

The potential benefits together with a mass balance estimate appear below in

Section 5, System Synergies.

4.6.3.3 Electrolytic Cells for Generation of 0 2 and H2 in Quantity

The recommendations from the propulsion system indicated a shared

effort for the development of electrolysis cells. The IOC definition

105
identified three options under development; all were specifically configured

to operate in a microgravity envirorment (RafezGmcm 21). The recom_nded

development for the Advanced Technology Space Station becomes an

electrically efficient cell which can operate over the widest practical

range of current densities to provide 02 and H2 in response to fluctuations

in the on-board electrical power demand (e.g., the cells make the Space

Station appear as a constant current load for the generators).

4.6.3.4 Synergistic Utilization of the Carbon from fD2 Reduction

An increase in crew size coupled with wet air oxidation of organic

wastes will yield a significant amount of elemental carbon from the Bosch

process for (D2 reduction. In the process of reductions, amorphous carbon

is deposited on the iron catalyst in a form that presently has no identified

utilization. A utilization synergy into activated carbon, graphite, or

other appropriate forms would benefit the operation. High temperature

processing could be achieved by utilizing a portion of the solar beam

concentrations which drive the electrical power generators.

4.7 Extravehicular Activity (EVA) Subsystem

EVA is the term used to describe activities performed by the astronauts

outside of the pressurized spacecraft environment. The two general classes

of EVA are: _

i. Planned _EVA, which are tasks included in a normal mission

time line that perform such functions as erection of Space

Station components, assembly of spacecraft components, and

repair of satellites.

106
2. Unscheduled EVA, which are tasks which were not planned for

the normal mission time llne but might be required for repair

or replacement of failed components to enhance mission

SUCceSS •

In general, EVA is provided for assembly, maintenance, servicing, and

repair of the Space Station and any other large space structures; servicing

and repair of satellites; maintenance servicing and repair of the Orbiting

Maneuvering Vehicles (OMV) and Orbital Transfer Vehicles (OTV); and

..mm£ntenance, servicing, and repair of payloads and experiments.

4.7.1 State-of-the-Art

The current state-of-the-art in EVA is that existing for the Space

Shuttle flights. The EVA systems include the Extravehicular Mobility Units

(EMU) and the Manned Maneuvering Units (MMU). The I__J's are independent

anthropomorphic (body-shaped) systems that provide environmental protection,

mobility, life support, and co_unications for the crc_nnen (Fixjure 4.7.1-I).

The units have no propulsion and generally are tethered. The basic

components of the E_J consist of a space suit that includes the pressure

garment, a primary life support system (PLSS), and an ultra-high frequency

radio communications system. The EMU operates with a pure oxygen atmosphere

at 4.3 pounds per square inch, and the units weigh about _50 pounds. The

characteristics of the EMU's are shown in Table 4.7.1-I. A complete

description of the current (Shuttle) EMU is given in Refe_ 34.

The MMU's are self-contained propulsive backpacks (Figure 4.7.1-_

designed to increase mobility in EVA. An bRU is fastened to the E;4U and can

be put on or taken off by one unassisted crew member. F_jure 4.7.1-3 shows

107
IN-SUIT DRINK

BAG

COMMUNICATIONS

CARRIER

ASSEMBL_

LIQUID COOLING

GARMENT

t:

i.

i.

HELMET/VISOR

ASSEMBLY

LOWER

TORSO

PRIMARY

LIFE SUPPORT

SUBSYSTEM

UPPER

TORSO

CONTAMINANT CONTROL

CARTRIDGE

BATTERY

GLOVES SECONDARY

DISPLAYS AND OXYGEN

CONTROL MODULE PACK

Figure 4.7. I-I Bxtravehioular Mobility Unit Components

(Reference 34)

108
D- ,

4.7.1-I CgAm(_,gl'/CS CF YIW E_ _Bm:MWI

34)

Extravehicular Mobility Unit

Weight

A_l_ere

Cap_iiity

250pounds

02 , 4.3 psi minimum

8 hours per day at an

average of 1000

btu per hour

Manned Maneuvering Unit

Weight

Propellant

Total L_ulse

Max Range

Electrical Power

325 pounds

G_

1392 pound-seconds

3000 feet

752 watt-hours

(2 batteries)

109
Figure 4.7.1-2 Manned Maneuvering Unit [Reference 34)

Figure 4.7.1-3 Astronaut in EMU and MMU (Reference 34)

ii0
a crewman in an ERO with a MMU attached. The MM0 has complete six-degreeof-fr_ control authority with spacecraft-type piloting logic. It can

perform translational or rotational maneuvers about each axis individually

or in any combination. Control inputs are made through use of hand

controllers. The left-hand controller is used for translational _aneuvers,

and the right-hand controller is used for rotational maneuvers. The MMU

also can perform attitude-hold on coa_and. Some of the physical

characteristics of the _gCJ's are given in Table 4.7.1-1. klditional details

relative to the _gfJ's are given in Referel_e 34.

4.7.2 Technology Trends

Technology trends have been aimed at EMU improvements, such as

increased suit mobility, improved cooling, improved coBmmications, etc.

DevelopBent work has advanced the design of a space suit that would operate

at a higher internal pressure than the current 4.3 psi. The objective has

been a reduction in the amount of time required for pressurizing and

depressurlzing before and after EVA (Figure 4.7.2-I, taken from RBfeces_e

35), and has provided a more normal gas atmosl_here. The accmallative

physiological effects of many cycles of pressurizing and operating in a low

pressure oxygen atmosphere are not fully understood (RefetecEe 35).

_RU improvements have been concerned primarily with increasing the

amount of nitrogen propellant.

EVA has been very successful, as demonstrated in the Skylab and Shuttle

missions. However, there has been a growing concern relative to the

extended EVA time and effort projected for the assembly and maintenance of a

Space Station (References 36 _ 37, for example). This has given an

111
o

112
addltional impetus to the concepts of using mechanical devices in place of

EVA and using astronauts as a backup when necessary. The _ical devices

being examined have taken many forms, as shown in Fi@ures 4.7.2-2 and 4.7.2-

3 from l_f_ 38. The degree of automation has been from intelligent

robots to teleoperators and extends to a crewman in an enclosure

manipulating robotic arms. The area of robotics is receiving considerable

attention. (See Refet_m:e 39, for example.)

4.7.3 Candidate Advanced Subsystem

If technology continues to advance at its current rate, it would appear

that "intelligent" robots could be developed for the time period of interest

and that they would be used in place of EVA. The future use of robotics for

Space Station is likely to be through integration of efforts to design

tools, to tailor assembly procedures, and to develop repair procedures for

use of robots.

4.8 Commmications and Tracking Subsystem

The Space Station communications and tracking subsystem performs the

critical functions that provide the internal on-board audio and visual

channels, provides all of the external radio frequency links, and tracks the

positions of all the free flyers plus the rendezvous support for docking the

Shuttle. Each of these functions employs equipment and techniques that are

within the capabilities of established technology. Support for future

missions envisions an expansion of communications system capabilities to

levels which exceed the present limits of technology. Configurations for

future Space Stations must address con_unications support at carrier

113
WASTE- UQU1D&,_

NSTRLJMB_

PANEL

RCSFUEL(C_) UMT (,1EAC_

FOOD

Figuro 4.7.2-2 EVA Bubblo Concept (Re£oronoo 38)

H

Figure 4.7.2-3 BVA Hard Buit Conctpt (Roforonct 38)

114
frequencies which may not be compatible with transmission through the

atmosphere of the Earth; Fi_t_e 4.8-1 illustrates some of the concerns.

Within the presently used portion of the microwave spectrum, water

absorption, particularly from rainstorms, impacts Earth tranmnission within

the K band; multiple beams and extra power become necessary. Extension of

the operating frequency upward faces the oxygen absorption peak in the

vicinity of 60 GHz and an ever-increasing absorption by water. Throughout

the world, utilization of microwaves will continue to increase over that

portion of the spectrum from K band and below. Therefore, carrier frequency

assignments and transmission priorities will become more important

considerations as the open-to-Earth portion of the microwave spectrum

saturates from the total world demand. These considerations are addressed

in the descriptions of communications and tracking system technology.

4.8.1 State-of-the-Art

4.8.1.1 Space Station On-board Systems' Capabilities

The pertinent performance capabilities for the radio frequency links

installed on-board the IOC Space Station are smmarized in Table 4.8.1-1 for

the ccammication elements and in Table 4.8.1-2 for the tracking elements.

The principal data handling and communication links operate with the TDRSS

for transmissions to Earth. As presently configured, all ccammications to

and from the Earth pass through the TDRSS such that the Space Station's

capabilities reflect the system limits within the _DRSS. Uprating of the

Earth's data links for the Space Station envisions a new generation of relay

satellites with data rates in the Gigabit per second range. The other

115
Z

O

I-4

O

I'-4

Z

k4

10

1

0.1

.01

0-3

MICROWAVE BANDS

P L S C X

_T- _-7 -_ -TFT r"

/

0.225

0.39

_--K

K1

?

Q v w

!

! II

•! i!

a j

I g

! !

I

I I

( I

I

! I

! °

a

..... ! i

I

I '

I

i '

i ,

I

i ,

u I

t

' 1

I

I _'- ir

_i:i

I

OXYGEN / I

: _ n V

! / li

i ,

I

i ,

i i

, I

I o

_ W_ER/ ___.,'

1.55 3

r ,i

//

i00

15.35 33 56

i_.25 36

FREQUENCY, GHZ.

Figure 4.8-1 Atmospheric Absorption for Microwaves (Extracted

from Reference Data for Radio Engineers,

Howard W. 8ams and Co., Inc.,

Indianapolls, IN, 1979)

116
_°° °!!!! ! !!

117
,'." z -,:

_' I.iJ

,,_ ¢_ ¢..) .=J

_. Of-)

I.l.I ¢_ I_ ,'='

I I I .-i118
co_unications links address the on-orbit support functions. The links to

the Orbiter support the resupply and fueling. The links for EVA and MRMS

reflect on-board support activities. The Space Station support to

deployment of free fliers appears in the OMV and platform links. The

support to free flyer platforms envisions growth up to eight units with a

communications link extending line-of-sight to the horizon at the nominal

Space Station altitude.

The tracking capabilities support the operations within horizon lineof-sight to the Station. The capabilities listed reflect the particular

requirements associated with rendezvous and docking. The types of data and

measurements presented increase at close range to provide the necessary

monitor for berthing and engagements. The communications system within the

Space Station provides for voice and high resolution video in addition to

the data transmissions. The on-board communications utilize digital

techn_jues for transmission to permit recording and down-linking to Earth.

A limited number of audio channels include on-board encryption to permit

secure or private communications.

4.8.1.2 Space Station Cmamunications Technology Status

An assessment of the operations for the Space Station in LEO provides a

means for estimating the status of pertinent conmunications technologies.

Practical considerations dictate that earth communications links must pass

through a relay satellite in GEO. Operations as a communications relay

station must recognize the effects of Earth orbit and the changing solar

position throughout an orbit. In such a context, the principal limit

imposed by technology appears as the attainable transmitted bit rate and the

119
acceptable bit error rate. These features in turn become defined in terms

of operating frequency, transmitter powers, and antenna considerations. The

identified status of technology for data transmissions to earth appears in

smmnary as the capabilities planned or under development for the tracking

and data acquisition satellite _TDAS) which represents the next generation

replacement fOE the TDRSS (which is now moving into orbital operation). The

principal features of the TDAS include:

i. Data Pate - Data will transmit at an average rate of 250 Mbps

and employ up to 12 channels in each spacecraft. The carrier

will operate in the 33 to 36 GBz range (Ka band) and have

simultaneous transmissions to a number of ground stations (to

overcome atmospheric effects such as rain). The channels

will have burst rate capabilities up to 1,200 _ps.

2. Bit Error Pate - Nominal bit error rates will be one error in

105 to 106 bits. Coding techniques will extend the accuracy

for special cases.

3. Antenna - The down link will employ a 3-meter dish capable of

transmitting nine simultaneous beams (5 fixed and 4

adaptable, each scanning 1 by 2 degrees). The antenna will

be configured for a focal-length-to-diameter ratio of 0.45

and require a surface accuracy of 0.25 mm (0.01 inch).

4.8.2 Technology Trends

Continuing advancement for each of the system elements summrizes the

NASA developmental approach and trends for _ications technology. The

present development and research efforts focus upon specific improvements in

120
the performance of individual elements that lead to higher data rates,

improved data accuracy, extending the operating frequency spectrum, and

expanding the distances for transmission links.

4.8.2.1 Antenna Developments

The trends for antennas address multiple beams, size, and beam control

as geometric accuracy and phased-array techniques. _ne needs for multiple

frequency (multichannel) operation respond to the requirement for everincreasing data rate tranmnission to Earth; multichannel operation provides

the necessary option. The replacement for the _DRSS envisions nine channels

through one 3-Feter antenna (Referem_e 3). Other systems propose more than

100 channels into a large (up to 55-meter) antenna. Multibeam operation

shows no upper theoretical limit; however, practicalities in terms of

aperture blockage due to feed elements and related structure will set the

real upper limit and become characteristic for a particular unit or

installation. Developments in the elements which feed an antenna address

reduction in size of interfering materials as well as the electromagnetic

properties of those materials.

The geometric requirements for beam control relate the surface

aberrations to the wavelength transmitted; typically, the surface contour

must conform to the theoretical geometric shape (parabola, etc.) with

tolerances less than three percent of the wavelength. Antenna development

efforts address the structural implications in terms of stable materials,

fabrication techniques, and erection or deployment methods. Developments in

antenna geometric capabilities will continue in step with mission

requirements. Controls over beam direction can be accomplished within the

121
feed system either directly (as an off-axis feed) covering a few degrees or

by relative phase eontrol of the individual feeds into a mosaic of small

antenna elements (a phased-array). Phased-array antennas are included in

present advanced mission concepts. The technology has been well established

for radar (centimeter) microwave applications (e.g., the U. S. Navy Aegis

Class Cruisers). The extensions involve the development of power dividers

and phasing controls for operation in the millimeter wavelength range.

4.8.2.2 I_ Switching

Transmit/receive operations with multibeam antennas imply high speed

matrix-type switching in the RF portion of the system. The TDAS follow-on

to the TDRSS anticipates switching nine channels into 30 to 35 transmit/

receive options accomplished in nanoseconds. These requirements build on

existing equipment and address material selections or construction

techniques as used for monolithic microelectronics. The present

developments are building on the GaAs semiconductor system to provide the

electrical conductivity and the use of field-effect transistors to obtain

the speed and control. GaAs has shown the capability to operate in the

picosecond regime in some configurations. Ferrites appear as the alternate

material which can support RF switching applications. The development of

ferrite switches has proceeded to the point of demonstrating performance in

the 32 to 37 _z frequency range with low losses (0.25 db) and switching

accu_plished in less than 20 microseconds.

122
4.8.2.3 Transmitters and Amplifiers

Output power stands as the criterion for the elements which produce a

modulated c_launication channel as a feed to an antenna. The output power

provides the signal strength against the background noise which effectively

defines distance for transmission and data quality. Developmental efforts

pursue three separate approaches to these elements in terms of solid state

transmitters, electron beam units, and coupled cavities. Each shows

particular benefits against disadvantages. The solid state units focus upon

the application of the GaAs system and take advantage of its inherent speed,

power efficiency, stability, and reliability. The developments recognize

the inherent power limit of individual devices and address the combination

of outputs to attain the needed power level. Combinations which can deliver

up to 40 watts have been operated in the laboratory, and efficiencies

approaching 70 percent have been achieved. The continuing development of

solid state items will provide for communications systems where power

utilization is an important consideration.

The electron beam devices include the traveling wave tube amplifiers

and represent the established technology for high frequency, high power, and

long distance transmission. The units have had a continuing development in

terms of materials improvements and internal configurations. The present

levels of capability produce power levels up to 500 watts at about 50

percent efficiency; higher power levels have been attained. In operation,

these units require DC power supplies in the kilovolt range plus the

generation of magnetic fields for beam focusing. These requirements present

complexities to some space flight applications. The coupled cavity concept

may be considered an extension of the traveling wave concept. The most

123
recent development has seen a three-cavity unit produce power levels up to

one kilowatt and anticipates development up to the two-kilowatt range.

These units do require the same high voltages as the traveling wave tubes

and higher magnetic fields for beam oontainment. For some configurations,

the field requirements can only be met with superconducting niobium magnets.

4.8.2.4 Receiver Detectors

The performance of receivers relates to the intrinsic losses that occur

in the detector-amplifier combination. Receiver noise establishes the

signal level threshold needed for operation. The TDAS has identified

receiver noise limit effects in the range 4 to 6 db over the operating

frequencies of the system. The detector portion of the noise becomes 1.5 to

3 db and represents about a factor of two improvement over present its.

Developments in detectors focus upon the GaAs system generally configured as

field-effect transistors. Alternate configurations are under development,

particularly for frequencies at 100 GRz or above. Developments also address

cooled detectors. Cryo-cooled super conducting tunnel junction systems show

perfo_s approaching the fundamental quantum limits.

4.8.2.5 Optical Transmission Systems

Lasers with high speed off-on modulation provide the means for

obtaining serial data transmission at gigabit-per-second rates. The laser

beam operates with some form of pulse modulation (position, intervals,

amplitude, etc.) and utilizes an optical telescope as the signal receiver.

The detectors presently available operate in the tens of picosecond range;

the combination provides the basis for operating with fraction of gigaseeond

124
pulses. The laser-based optical systems impose stringent requirements for

precision pointing. Laser beams are narrow_ pointing accuracies must

achieve one microradian or less with line-of-sight jitter less than 0.2

mlcroradian. The lasers Ixesently identified for such applications are the

f requency-doubled, neodymium-yttrima aluminum garnet (_-Y_J;) at 0.532

micrometers (green), GaAs at 0.85 micrometers (red), the Nd-YAG at its

lasing frequency 1.06 micrometers (near IR), and 002 at 10.6 micrometers

(IR). The lasers operating in the visible are receiving the most attention

and their receivers would utilize a 40 an optical telescope focused upon an

avalanche diode detector. The principal applications envisioned for laserbased transmission become relay links between satellites in orbit with

particular emphasis for solar system probes. None of the wavelengths listed

have an assured tranHaission path through the atmosphere of the Earth.

4.8.3 _nded System

The system recommendations envision an expansion of the present

requirements plus the addition of relay communications from distant sources.

The expansion of data transmissions to Earth will face the complexity of a

saturated utilization for all of the transmission frequency bands that can

operate effectively through the atmosphere (including bands that have

significant attenuations). The relay data transmission links will operate

at data rates and carrier frequencies not generally compatible with

transadssions to the Earth (e.g., absorbed in the atmosphere, obscured by

cloud cover, etc.). The envisioned capabilities are su,_arized as follows.

125
4.8.3.1 Communication Links

The commmication functions identified in Table 4.8.1-1 will continue

to exist; the changes anticipated appear as data rates, numbers of

simultaneous links, and the operating frequencies. The two-way Earth

communications through a relay satellite will present the greatest

complexity. The co_m_ications links through a relay satellite need to be

reserve_ for priority transmissions (e.g., a requirement for near real

time). The Space Station will have the capability to fill all the relay

satellite's inputs at their maximum continuous data rates. For such

conditions, the relay satellite would operate with no processing (e.g., the

"bent-pipe" mode). All formatting, data compressions, etc., would be

performed aboard the Space Station. The present best estimate would be 12

channels minimma operating at 300 Mbps each. The carrier frequency will be

atmosphere compatible, such as Ku at 15-17 _ or Ka at 33-36 (_z, with

multiple ground receivers. Support data from processing or experimentation

will have a capability for dt_p tran_nissions to Earth at ground stations of

interest. These dumps would occur at compatible frequencies and data rates

during an orbital overfly. All 14 of the specific frequency channels

assigned to the United States must be accum_dated. Prudence dictates an

additional coverage for the principal assigned frequencies used by other

nations. The dump transmission system must be reconfigurable during flight

in terms of adjusting frequencies, data rates, and transmission codes. The

transmission links for platforms, EVA, OTV, and supply-docking support will

expand by a factor of five, with frequencies shifted upwards into regimes

attenuated or absorbed by the atmosphere of the Earth and will expand above

126
the frequencies presently assigned for other functions (e.g., 105 to 130 (_z

and up).

4.8.3.2 Tracking Links

The requirements for tracking links will expand with needs for multivehicle tracking for each mode identified. Except for the interrogation

link to a global position system, the other tracking links will need to

operate at frequencies which do not penetrate the atmosphere of the Earth.

optical links for precision docking and depl _oyment would expand by about

an order of magnitude to support on-board spacecraft erection, assembly, and

fueling operations.

4.8.3.3 Relay Link Operations

Laser-driven optical transmissions will become the established

technique for cummunlcatlons with remote satellites. The Space Station will

provide a relay capability particularly for lunar or libration point

spacecraft. Practicalities again dictate maximum bit rates frum remote

spacecraft. The need to commandeer the entire ground transmission

capability could relate to a short Period of real-tlme relay operation from

such a remote spacecraft. The Space Stations should be prepared to carry a

minimum of four high-bit-rate laser links with optical telescope receivers.

The system needs to have on-board reconfiguration to support differences

between data formats and code system. As an example, a lunar system may

need to use a different laser wavelength or coding technique from that used

for an outer planet explorer.

127
4.8.3.4 0n-board Commmications

The recording and transmission requirements for audio/video and

measurement data will _d in responding to increases in mission roles

and crew sizes. The requirements anticipate a hundredfold increase for

communication support in responding to a tenfold expansion of the crew

c(m_lement. Link transmission to Earth will not be available for the major

portion of on-board recorded data. Transfers of such data will employ a

high density semi-permanent record. The candidate considered would be a

form of laser optical disc record-erase technique now under development as

the Tera Bit Recording System (Joint DCD-km_A, performed by RCA). Such

records would be returned to Earth aboard Shuttle return flights. As an

example, the video data associated with the on-board assembly and deployment

of a spacecraft will generate a significant library of records with no need

for real time transfer to Earth. The written records associated with visual

inspections during Earth assembly bec(m_ video recordings for an on-orbit

assembly performed remotely by manipulators. Such data have established

needs for retention but not as real-time transmissions to the Earth.

4.9 Data Management SubB_tem

4.9.1 State-of-the-Art

The salient features of the Data Management Subsystem (DMS) as

presented in the A{D (Refem 7) and the IOC (Refezm_e 6) include: an

optical data transmission network, con_onality of both interface and

embedded hardware, multiple processors and software packages, large

128
information storage capability, and time and frequency references. The DMS

is largely conceptual at this time.

The functions of Space Station control and operation will be inherent

in the DMS if the role as described in the ACD is followed--"When

functioning in an integrated vehicle environment, the IRS provides the

connectivity between subsystems, payloads, crew/operators, and the ground

interface to accomplish the end-to-end function of control and monitor, and

information exchange."

Figures 4.9.1-1 and 4.9.1-2 from the ACD (Ref_ 7) show an overview

of the DMS configuration and function. The specific hardware items as

currently envisioned are listed below.

4.9.1.1 Standard Data Processor

The SDP provides a standard processing environment for host application

software. _here will be a family of compatible SDP's in terms of processing

and memory capacity. Each member of the family will be fully upward and/or

downward compatible with the other members, thereby providing a robust and

economical growth part in the processing architecture. An ADP may be

dedicated to a subsystem or may be shared across multiple subsystems,

depending on individual processing requirements. The largest of the SDP

family will have processing capability in the range of 4 mip's, with up to

32 nb of main memory.

4.9.1.2 EMbedded Data Processor (EDP)

The EDP provides a standard processing envirorment in a card level

implementation. The _)P will be incorporated into specialized subsystem

129
°

-il

14 _0,.I

• L kl

W 14.1--*

i "_ I_ L_J L__I 0 <l

! , o11

• • m

z c_

.J.J_

fm_

t_

4)

U

4)

4)

4)

v

0

-,-i

4.1

0

r_

,'4

I

14

,1111

_)

14

-,4

r,N

130
131
elements. It includes an interface to a standard backplane. The EDP also

includes an Operating System (OS) that is co_mtible with that running in

the SDP; as a common DMS element, the EDP supports the SSE concept. The EDP

does not include direct interfaces to DMS networks or buses; rather, these

interfaces would be sul0plied and configured by the subsystem Ol_J developer.

The EDP has processing capability in the range of 0.5 mips with 1 mb of

memory.

4.9.1.3 Mass Storage Unit (MSU)

The MSU is the mass storage unit provided by DMS for core station and

payload operations. The MSU has non-volatile memory characteristics.

Multiple MSU's are required to meet Space Station on-board storage needs.

4.9.1.4 Multi-Purpose Application Console (MPAC)

The MPAC incorporates display and control devices and provides the

operational crew interface to the Space Station core and payload systems.

THe MPAC is the electronic core resident within a crew workstation. The

workstation also includes the structural support and external configuration

of MPAC elements as well as equipment to support crew members in performing

tasks. The MPAC will use multifunctional displays and controls. There are

three types of MPAC's- fixed, portable, and helmet or head mounted. The

MPAC's will have sufficient processing capability and memory to support crew

workstation services. They will interface via an NIU to the DMS network.

132
_s

4.9.1.6 Network Interface Unit (NIU)

The NIU provides access to shared DMS local networks. It contains DMS

software but not non-DMS application software. The NIU is packaged in two

ways: (i) a full NIU capable of supporting the full range of DES end-toend communications services, and (2) a partial NIU capable of providing a

subset of data transport services. The partial NIU supports users who

require upper layer services, such as file transfer and address or

interactive sessions.

4.9.1.7 Multiplexer/De-Multiplexer (MDM)

The MDM operates as a remote I/O device for (i) acquiring,

conditioning, multiplexing, and transferring subs_tem sensor data to the

subsystem data processor; and (2) routing, conditioning, and distributing

commands and data from the subsystem data processor to subsystem effectors.

4.9.1.8 Bus Interface Adapter (BIA3

The BIA is a standard DMS-provided unit for connecting DMS and other

subsystem equipment to DMS-defined local data buses, implementing the lower

layer protocols.

4.9 .i.9 Bridge

A network bridge provides access from a DMS local network or a non-DMS

local area network to the global station network that runs between modules.

It contains DMS software but not non-DMS application software. Since the

DMS global and local networks are identical and use the same protocols, the

bridge does not need to perform protocol conversion.

133
4.9.1.10 Gateway

A network gateway provides access from a non-DMS local area network to

a DMS local network or to the global station network that rims bet_en

modules. It contains MS software but not non-DMS application software.

Since the DMS and non-DMS networks are assumed to be different and

implementing different low-level protocols, the gateway has to perform

protocol conversion to an agreed-upon standard interface at the data link

layer. This gateway is predicated on program-wide agreements on

communications protocols at the logical link layer and above; it will

greatly increase in cost and complexity if these agreenents are not

achieved.

4.9.1.11 Local Bus

The DMS shall provide one or two local bus standards as media for data

transport between subsystem sensors/effectors and BIA's. The BIA's in turn

provide connectivity to SDP's.

4.9.1.12 Local Network

AIRS local network is defined as any core or payload network that is

confined within a specific region of the Space Station, for example a

module. It is "owned" by the DMS in the sense that the DMS monitors and

maintains its health and controls its use through the NOS.

4.9.1.13 Global Network

A global network is defined as the core of payload network that is not

confined to a specific region of the Space Station and, in particular, runs

134
t"

between modules. It is "owned = by the _ in the sense that the

monitors and maintains its health and controls its use through the lqOS. The

global network is a high speed, 100 at_s, standard medium.

4.9.1.14 Time-Generated System

The time and frequency generation system comprises time and frequency

souroes,'_ators, processors, and oontrollers to generate accurate and

stable time and frequency information as required by Space Station systems.

4.9.2 Technol_Y Trends

On-board data processL_ is one of the most rapidly advancing areas of

spaoe systems _logy and as such can affect all of the Space Station's

subsystems and functions. Future systems will have higher processing rates,

will come in smaller pscMages, and will support networked architecture for

on-board signal processing (Referemce 2). Reliability and fault tolerance

will be required attributes.

Technology forecasts for (xm_onents of a data management system are

given in Referenoe 2 and are presented herein as Table 4.9.2-1.

4.9.3 Candidate Advanced :Subsystem

Because of rapid developments occurring in this area, no specific

subsystem components were selected for this study. It appears that such

selections may well depend on factors such as reliability, fault tolerance,

ease of replacement, and ease of repair.

135
Figure of Merit 1985 1990

On-board Processors

1995 2000

SImcebornecoa_uters

Performance (Mops) i0 40 I00 200

Signal Processors

Throughlm/t Rate (Mops) 20 100 1,000

Microprocessors

Performance (Mops) 1 15 30 60

1 x 105 2 X 105 5 X 105 1 X 106

Controller Chip

complexity (gates)

Magnetic Tape

Capacity (bits)

Transfer Rate _PS)

Data Storage Systems

4 x 1010 1 x 1011 3 x 1011 I x 1012

3 x 107 1 x 108 3 x 108 1 x 109

5 x 107 i x 109

<I x 106 4 x 106

Bubble Memory

Capacity C_its)

Transfer Rate CDPS)

Optical Disk

Capacity (bits) * 1 x 1013

Transfer Rate fl_ps) * 1 x 108

Artificial Intelligence/Expert Systems

Knowledge Base Size (rules) 1 x 103 1 x 104

Knowledge Base Throughput

(rules/sec) 400 4,000

Software Productivity

m

m

Productivity Rate

(codelines/man-year)

n

m

m

m

m

1.5 x 104 2 x 104

6,000 8,000

900 2,500 5,200 8,300

"t

*Technology not available

136
5.0 S[BSYSTEM SYNERSIES FOR SPACE STATION

Space Station synergies represent opportunities for subsystems to

beneficially share items of equipment, functions, or configurations that

enhance the total operation of the station. The assessment for synergies

reviewed each subsystem in conjunction with each of the others to identify

potential areas of beneficial interactions. These results are summarized

below; the synergy between life support, propulsion, and electrical power

has received an estimate of the overall mass balance effects from generating

O2--q2 as pr_-.g,_4n- fuel by e!ect_.rica!lyd_sing water.

5.1 Assessment of Synergies

The initial evaluation for subsystem synergies identified five

opportunities for beneficial interactions. These are summarized in a

sequence based upon the number of subsystems involved with a particular

synergy.

5.1.1 Life Support, Propulsion, Electrical Power, Guidance/Control, and

Experimentation in Relation to the On-board Generation of O2-H 2

Fuel •

The life support system provides the 02 for reconstitution of the cabin

atmosphere by the on-board electrolytic decomposition of water. The H 2

generated reduces CO 2 to elemental carbon and water. The system recovers

water from the air and recycles waste water to achieve a net balance that

actually shows a small degree of water surplus. The expansion of the

electrolysis capability up to the levels required for on-board propulsion as

fuels for attitude control, reboost, and auxiliaries together with a more

137
extensive recycling of wastes improves the water recovery balance and

provides a means for load leveling in the electrical power system. Fuel

generation at the rate of 13,500 kg (30,000 pounds) per month would utilize

less than i0 percent of the total power available and offers a substantial

opportunity for load leveling by absorbing any excess power. One of the

principal microgravity processes demonstrated during Shuttle flights

involves the electroI_mretic separation of organics in aqueous solutions.

The depleted stream from this experiment could pass through the life support

oxidizer and provide supplemental feed water for electrolysis. An initial

estimate of the life support interactions appears as Section 5.2 below.

°f

5.1.2 Thermal Control, Electric Power, Life Support, and Experimentation

As the Integration of Thermal Control Buses and Loops

The thernal control system provides the means for removing or supplying

heat at specific locations throughout the Station. The concept envisions a

series of loops such that one user's sink is another system's source.

Typically, the condensate loop for the power generation system would provide

sensible heat for experiments, cabin temperature control, etc. A low

temperature bus would extract heat from localized generators such as

experiments, electronic or electrical equipment, etc., and operate through

an exchanger to a space radiator.

5.1.3 Thermal Control and Structure As the Selection of the Pressure

Shell Materials to Provide a Heat Transfer Radiator

The construction of the pressure shells from materials such as graphite

fiber in an aluminum matrix can provide a direct thermal path through the

138
wall for radiation. In such an installation, the fluid loops would be on

the inside and in contact with the pressure shell; the external radiation

heat transfer could be enhanced by a locally applied coating that improved

the emissivity. As an alternate, a wound or wrapped pressure shell formed

from a tubular member would provide a means to circulate heat transfer

fluids locally within the walls of the pressure shell.

5.1.4 Electrical Power and Experiments By the Solar Collectors to

Provide a Concentrated Solar Flux for High Temperature Experiments

A Space Station power system based upon solar powered dynamic

generators (turbines or reciprocators) has large area solar concentrators

focused into high temperature receivers. The receiver and related structure

provide a mou_ting access for research experiments which can benefit from

the 6,000 K equivalent heat source.

5.1.5 Attitude Control Using Water Storage as an Inertial Balance in a

Rotating Spacecraft

The utilization of rotation to produce an artificial gravity in

habitated areas shows benefits for Space Station operation. The elimination

of gyroscope effects by counter rotations simplifies attitude control

requirements. A com_ter-rotatLng member constructed as the water storage

tank would provide a means for inertial matching by velocity adjustment as

well as providing the needed reservoir for water supplies.

139
5.2 Operating Synergy from 02-H 2 On-board Generation by Water

Electrolysis

The selection of 02-H2 as the general-purpose Space Station propellant

and fuel offers beneficial synergies between the life support, propulsion,

and electrical power systens. An expansion of the electrolysis capacity to

provide fuel for station attitude control, reboost, and the OMV appears to

benefit the life support system and electrical power as well as simplifying

resupply logistics. The description of the synergies will first stmmarize

the life support system used as the basis for comparison and then proceed to

show the effects on water utilization of on-board production of O2-H 2 in

quantities associated with fuel or propellant utilization.

5.2.1 S_mary of the Baseline Life Support System (Present IOC)

The presently defined IOC life support system provides the baseline for

constructing a oomparison; Figure 5.2.1-I (Referemoe 6) shows the principal

features and illustrates the daily mass flows required to support a sixperson crew. The pertinent features for later comparisons show that:

i. Electrolysis of water provides the 0 2 supply for

reconstitution of the cabin atmosphere.

2. The H2 from electrolysis performs the 002 reduction utilizing

the Bosch process.

3. Water from condensate and 002 reduction provides the potable

supply and balances the water O2-H2 requirements.

4. The water for electrolysis is drawn from the recycling of

water for hygiene which includes the processing and

decomposition of urine.

140
LZ

ou |

ulll ++- + |t ,,,

o.

,,,.

llpu

l:l

ill

i-4

Illwm

to

q:l Qm0

r-i

i-I

I-4

II

4-1k

IS

_-II

t_Q

141
5. Wash water has its own independent recycle.

5.2.2. Synergy System Parameters

The system envisioned for synergistic comparisons relates to a new,

larger Space Station that has a matured capability to perform all the

functions intended. The on-board population has been established at 60

persons. Life support requirements are adjusted upward as follows:

I. The requirements for atmosphere (including leakage) and food

together with waste outputs are direct multiples equal to I0

times the values shown in Figure 5.2.1-1.

2. The water used for flushing purposes is set at three times

the baseline usage and becomes a factor of 30 relative to the

baseline.

3. The water requirement for the hygiene loop and the water

requirement for wash are doubled to result in a factor of 20

relative to the baseline.

The system as defined envisions a significant increase in the utilization of

water, particularly for crew-comfort functions.

5.2.3. Synergized System Description

The system intended for O2-H 2 combined production supports an on-board

crew of 60 individuals and recycles all the w_ter from dish washing, clothes

washing, and hygiene-related functions. The system also oxidizes all human

wastes. The system flow and mass balance diagram appears as Figure 5.2.3-1.

The principal features are:

142
143
1. Electrolysis of water provides the 02 for reconstitution of

the cabin atmosphere, plus the oxidation of wastes and

propellant/fuel requirements.

2. The H2 from electrolysis performs 092 reduction from both

cabin air extractions and from the oxidation of carbonaceous

wastes.

3. The oxidation of wastes from humans, the galley, laundry, and

cleaning employs the wet air oxidation technique operating at

temperatures and pressures which provide both sterilization

and cumplete co, rotation.

4. The dissolved salt effluent (and residual solids) from the

wet air oxidizer are precipitated to dry crystals by a

distillation system which has the option to include reverse

osmosis as one of the extraction steps. The water output of

the wet air system eventually provides the input to the

electrolytic cells.

The values for mass balance involving wet air oxidation assumed the urine to

be 1.5 percent N2 as urea; waste from hand, clothing, and dish washing was

taken as 0.25 percent stearate soap. The oxidizable portion of feces

content %_s set as all carbon. The unoxidizable solids were set at 10 grams

per person per day. The auxiliary reduction of 002 is considered

conservative to provide an upper demand level for the electrolysis capacity.

5.2.4. Rationale for Wet Air Oxidation

The utilization of wet air oxidation for disposal of wastes and

accomplishing sterilization is based upon the present level of development

144
for the process together with an apparent potential for providing a general

purpose disposer of carbonaceous organics. A smmmry of the tec_mlque and

related technology presented in a recent technical journal (R_Eezemoe 40)

illustrated the applications of wet air oxidations to the disposal of sludge

and similar waste materials. The ability of water to dissolve oxygen at

elevated pressure provides a means to initiate exothermic oxidation of

otherwise incombustible organics and achieve a net power gain from the

process. The utilization of pure oxygen enhances the process. The

technology has proven effective in applications to the point where such

equipment is manufactured commercially for continued operation at conditions

just bel(_ the critical Point for water (critical Point at 21.5 MPa and 375o

C). The continuing evaluations by the LaRC have shown that operation above

the critical Point improves the oxidation of organics and reduces nitrogen

compounds to elemental forms (Referemoe 41). The reactor envisioned for

space would utilize a high strength reinforced material such as graphite

fiber in an aluminum matrix with the surfaces in contact with the reacting

liquids having additional oxidation protection (e,g., A1203, ZrO2,

chr(mlum). The system would operate in a manner to Perform a complete

oxidation. For conservatism, the diagram (F_/u_e 5.2.3-I) includes the post

oxidation steps of evaporation for solids removal while recognizing that

precipitation of inorganic solids (such as NaCl) would occur during

supercritical operations.

5.2.5 Assessment of System Operation

The pertinent considerations associated with providing 02-H 2

propellant/fuel in synergy with the life support requirements appear

145
summarized in Table 5.2.5-1. The requirements and effects are presented for

the cases of 237 kg (500 pounds) and 475 kg (1,000 pounds) O2-H 2 propellant

production each day. The lower rate would supply reboost and attitude

control with 4,750 kg (10,000 pounds) per month available for the OTV; the

upper rate would provide the additional capacity for a monthly fueling of

the equivalent to a Centaur booster.

In reviewing the operation, both rates of production would have to

first provide the life support requirement of 147 kg (325 pounds). The

total water processing for on-board life support functions is 1,938 kg

(4,273 pounds) per day. The reclamation of all water, including extractions

fram the _O-produced 002, provides a net 36 kg (79 pounds) daily gain in

actual water. At an Isp of 440 seconds, this weight of fuel provides more

than five tln_s the present impulse requirement for attitude control and

reboost. The generation of O2-H 2 for propulsion or fuel allows the use of

unprocessed fresh water for potable usage and reduces the quantity of

recycled water needed for life support. At a fuel generation rate of 237 kg

(500 pounds), only 13 kg (29 pounds) of recycled condensate needs finishing

to the point of potability. For a fuel generation of 475 kg (i,000 pounds)

per day, all potable water would be drawn from stores along with about 13

percent of the supplemental needs.

An assessment of the power impact compares the requirement for

electrolysis with the total available. An electrolytic cell operating at 70

percent efficiency requires 2.7 kilowatt hours of electric energy to produce

a pound of water. For a Space Station which provides 2.5 _e equivalent

continuous power, the electrolysis operation to produce 237 kg (500 pounds)

of fuel corresponds to 3.5 percent of the power available; 475 kg (i,000

146
Summary of Requirement

A. Electrolysis Requirements

o Electrolysis for Life Support

o Total Electrolysis Required

B. Water Needs or Utilization

o Potable Drinking. Quality

o Hygiene and Washing

o Life Support Electrolysis

o Total Requirement

C. Water Available from _._cle

o Cabin Air and Crew 002

o C02 Reduction from P_9

o Direct from P_O

o Total Recycled

- _dditional Water Created

by 7clmg

- Net Draw to Provide Fuel from

On-board Storage Supply

Fuel at 237 kg

(500 lb./day)

147 kg

(325 lb.)

386 kg

(825 lb.)

204 kg

(450 lb.)

1586 kg

(3498 lb.)

147 kg

(325 lb.)

1938 kg

(4273 lb.)

205 kg

(453 lb.)

40.3 kg

(89 lb.)

1728 kg

(3810 lb.)

1974 kg

(4352 lb.)

36 kg

(79 lb.)

191 kg

(421 lb.)

Fuel at 475 kg

(i000 lb./day)

147 kg

(3251b.)

601 kg

(1325 lb.)

204 kg

(450 lb.)

1586 kg

(3498lb.)

147 kg

(325 lb.)

1938 kg

(4273 lb.)

205 kg

(453 lb.)

40.3 kg

(89 lb.)

1728 kg

(3810 lb.)

1974 kg

(4352 lb.)

36 kg

(79 lb.)

417 kg

(921 lb.)

147
'JSMmg 5.2.5--1 A_lm_Ni_' (]E' ])AILYilZt'm_ l_[_'11qmJl_

(]q--EON_ _ (W O2-E 2 (Cancltx]ec])

Stmmary of Requirement

D. Water Utilization Sources

(a) Potable 204 kg (450 lb.)

o From Supply

o From Condensate

(b) Wash and Hygiene

1587 kg (3498 lb.)

o From Supply

o From Recovered Water

(Condensate and CO 2)

o From the _0 Output

E. Electrical Power Requirement

(70 percent efficiency)

Fuel at 237 kg

(500 lb./day)

191 kg

(421 lb.)

13 kg

(29 lb.)

Fuel at 475 kg

(i000 lb./day)

204 kg

(450 lb.)

0

231 kg

0

213 kg

(508 lb.)

1356 kg

(2990 lb.)

2220 kWh

(471 lb.)

243 kg

(537 lb.)

1129 kg

(2490 lb.)

3580 kWh

a,

148
pounds) of fuel equates to 6.7 percent of available power. Both of these

loads are within the anticipated cyclic variations in total station power

demand. The o_rd generation of 02-H 2 for propellants and the supply for

auxiliary fuel cells offers an attractive means toward load leveling of the

electric power generation system by operating the electrolytic cells with

the excess power available at any time.

149
6.0 ADVANCED T_ SPACE STATION CONFIGt_ATION CONCEPT

6.1 Approach

The subsystem candidates of Section 4.0 were selected to provide

maximum synergism for an integrated Space Station configuration. In

particular, the EPS, ECLSS, thermal, and propulsion subsystems were selected

to have an interplay of energy and material use related by a water,

hydrogen, and oxygen base to the extent practical. In addition, high-grade

heat energy of the EPS is usable directly by the other subsystems without

the double conversion from heat to electrical energy and back to heat output

where_r transport distances are reasonable. Lower grade heat energy from

the EPS heat engine outlet is available for thermal control, which

effectively reduces the thermal subsystem loads and reduces EPS radiator

requirements. Section 5.1 described the potential synergisms between the

selected subsystems on the integrated Space Station.

The integrated Space Station is selected to have portions that provide

artificial gravity and portions that are stabilized for Earth and space

viewing and for docking and servicing facilities. The rationale for

artificial gravity is covered in Section 6.3. In addition to the synergism

of subsystems and the requirement for artificial gravity, other factors

considered are the power output needed, the crew size to provide the

functions listed in Section 2.0, the Space Station orientation relative to

Earth and Sun, and the capability for additional growth.

150
6.2 Estimates of Power Requirements and Crew Size

_he seventeen functions identified for the Advanced Technology Space

Station received estimates for electrical power consumption and crew

requirements. _he estimates draw from the source references used throughout

the study and include a contingency margin. The estimates of power

requirements for each of the functions and the contingency appear in Table

6.2-I. For crew estimates, the functions have been grouped into six general

categories and summrized as follows:

NUMBER OF OC¢_PANTS

Operation 12

Science 10

Manufacturing 12

Transients 14

Medical 6

Conttugmcy 6

_TrAL 60

6.3 Considerations for Artificial Gravity and Space Station

Orientation

The major argt_ents for artificial gravity on a Space Station are the

need for the occupants to avoid physiological problems associated with

Coriolis acceleration, motion sickness, general disorientation, bone

demineralization, loss of muscle tone, and loss of traction. Some subjects

are more prone than others to various aspects of these possible problems.

151
1)

2)

3)

4)

5)

6)

7)

8)

9)

10)

11)

12)

Table 6.2-1 Jstimted Electrical Po_r Requtremmts

for the Simue Station Functions Listed in Section 2.0

Function

Observation Platform

In Orbit Science

Service and Repair

Manufacturing & Assembly

Broad

Category

Science

Science

Operations

Manufacturing

Electric Power

Requirements, kWe

i0

I0

5

500

Transportation Node

Crew Habitat

Co_mmications Center

klaptation Area

Storage Node

Variable "g" Facility

Operations

Habitation

Operations

Habitation

Operations

Science

Conmercial Manufacturing

Energy Collection & Relay

Station

13) Medical Facility

14) Tourism (Transients)

15) Horticultural Research

16) Technology Demonstration

Facility

17) Control Center for

Satellites

Manufacturing

Operations

Medical

Habitation

Science

Science

Operations

10

5

5O

5

10

5

1,000

25

20

i0

TOTAL EL_TRI(I_L POWER R__

TOTAL ELECTRICAL POWER fXEPUT

i i| i

10

5O

152"
Rafenmo_ 22, 36, amd 42 provide some background on effects of simulated

and actual gravity on various subjects. Although the minimum acceptable "g"

level and possible relation to exposure time is not defined (and is probably

dependent on the individual, training, etc.), some guidance is provided by

F/_suze 6.3-1 from l_fezemce 43.

Another argument for providing artificial gravity is the fact that some

physical processes, such as that of an ECLSS, work much better in a gravity

environment.

It appears, therefore, that if crew members are to remain in orbit for

extended periods of time, provisions should be made for some level of

artificial gravity. This could be accomplished by using a rotating section

of the spacecraft and using that section as the usual human habitat or for

periodic reconditioning as required.

It was decided that the Space Station should have a sun-oriented

element on which to mount solar collectors.

6.4 Sizing the Space Station

F/_ure 6.3-1 was used as a guide in selecting the size and rate of

rotation of the turning section of the Space Station. In order to fall

within the acceptable operating range shown in the figure while maintaining

a reasonable size (diameter) and rotation rate, a gravity level of 5.37 feet

per second per second was selected with a module or multi-module rotation

arm of 114.5 meters (375 feet}. This combination resulted in a rotation rate

of 1.14 rpm. The "g" level selected is one lunar "g" and could be used to

acclimate personnel for lunar base operation.

153
o

154
The selected diameter proy'zde4.:adeguate area for solar collectors,

radiators, hangers, and o_r _ce'4St_tlon adjuncts.

6.5 Space Station Configuration

The configuration selected, based on the considerations of the previous

sections, is shown in Figures 6.5-I 8rod 6.5-2. The large rotating torus

could as well be several modules on spokes or a regular polygon made up of

many cylindrical sections. Pertinent physical dimensions of the

configuration are given in Table 6.5-I.

The Space Station features a large cylindrical central tube and a truss

platform to which all other sections are attached. The box-like structure

at the _ttum is the repair, docking, and assembly area. The long cross arm

is rigidly attached to the central tube and is used for mounting Earthpointing devices. Four solar collectors are attached to the sun-oriented

platform and are the prime power source for any equipment in the cross arm.

Two solar collectors are attached to the rotating torus and are the primary

energy source for equipment in the torus.

The two disc-shaped structures, one on each side of the torus, are used

to reduce the net angular momentum of the Space Station about its axis to

zero. One of the disks rotates in the opposite direction to that of the

large torus. The second disk is fixed relative to the center core; however,

its contents (probably water) will rotate opposite to the torus. The rate

of rotation can be varied by using pumps. This could be used to trim out

the angular momentum. The axis of rotation of the torus is parallel to the

plane of the ecliptic and continuously points toward the sun. Therefore,

the Station must precess continuously at the rate of one revolution per

155
PAGE

156

II

Q

1.4

e_

o_

o_

_m

o

II@

i_r.I

O_

,..4-,'1

@.,t

o

,Om

?
q) e

0.,.4

¢I .,.I

_Q

0¢I

OA

0-,4

_m

q) q)

N

157
A. PRINCIPAL ELEMENTS A_D FUNCTIONS

Element Function

i. Rotating Torus Habitat and medical, life support and

fuel generation, communication and

control center, variable gravity

facility, component fabrication,

storage

2. Central Tube & Platform

3. Observation Tube

Structural core, materials transfer,

microgravity processing, solar

observatory and horticulture research,

orbital science and technology, air

lock, storage

Earth and space instruments,

communications antennas, energy relay

beams, storage

4. Docking & Erection Bay Rendezvous and docking for supply,

assemble fuel and deploy spacecraft,

pressure parts for crew transfer,

holding facility for OMV, OTV

5. Solar Dynamic Power

Collectors

On-board power generation by 6 units;

4 units on the platform and 2 units on

the rotating torus

6. Inertial Balance Rotators Angular mument_n of the torus nulled by

counter rotation, on-board water

storage

So

1. Torus Diameter to Ring

2. Ring Diameter

3. Spokes as a triangular Truss

around a tube diameter

4. Ring Volmne

RfTfATING TORUS, DESCRIPTION A_D

229 m

15.3 m

5.6 m

3.2 m

1.305 x 105 m3

750 ft.

50 ft.

18.3 ft.

10.5 ft.

4.61 x 106 ft.3

158
6.5-1 _ _ 8]_V_ 81_']_,

CIr _ _ ((_ontLnuecl)

5. Ring Weight, Total as sum of; 1.65 x 106 kg

Walls as 0.375" Graphite-Al 8.15 x 105 kg

Floor Load 30psf, 50% Vol. 8.02 x 105 kg

O2,H2 at atmospheric 3.00 x 104 kg

pressure, 50% of Ring Vol.

6. Moment of Inertia 2.16x 1010

7. Angular Momentum for 1.141rpm 2.58 x 109 kgm2

Provides 0.166g in Ring sec

3.64 x 106 ]bs.

1.80 x 106 Ibs.

1.77 x 106 Ibs.

7.24 x 104 ]bs.

1.59 x 1010 _2

1.90 x 109 ibm_2

sec

C. CENTRAL _0BE A_D _RM, DIMENSIONS A_D

1. Platform, as a Truss: Bays 5 m 16.4 ft.

Dia. ]58 m 520 ft.

2. Horticulture Domes as Dia. 25 m

Spherical Caps High 5 m

3. Central Tube Diameter

82 ft.

16.4 ft.

15 m 49 ft.

4. Central Tube Length

5. Total Volm,e

6. Microgravity Processing

Facility as a Levitated

Cylinder 8m dia, 25m long

7. Solar GbservatoryVolume,

15mdia, 10m long

100 m 329 ft.

1.762 x 104 m 3 6.24 x 105 ft3

1,260 m3 4.45 x 104 ft3

1,762 m3 6.24 x 104 ft3

8. Air Locks as Concentric

Doors

Full dia - 13 m 42.5 ft.

Half dia - 7 m 23 ft.

Smalldia- 3 m 9 ft.

D.

1. um h

2. Structure as a Square Truss

around Tube of Diameter

3. Volume of the Tube

OBSERVATION TJBE, DIMENSIONS A_D_

290 m

5m

3.2 m

2,210 m3

950 ft.

16.4 ft.

10.5 ft.

7,810 ft.3

159
6.5-1

E. DOCKING AND ERECTION BAY, DIMENSIONS AND

I. The Bay is a Cube 67 m 220 ft.

2. Structure as Square Truss

around Tubes of Diameter

5 m 16.4 ft.

3.2 m 10.5 ft.

3. Pressure Ports for Docking

that Includes Crew

Transfers

One at each of the bottom corners

of the bay

4. Manipulators, 10 Minimum

Coverage Internal to Bay

Coverage External to Bay

(All Faces)

All Maniuplators Have the

CapBbility for Interchanging _ Effectors

2 minimmn at every point

2 minimum at any point up to 15 m

(49 ft.), 1 minimum out to 30 m

(i00 ft.)

F. SOLAR DYNAMIC POWER G_ERATION AS 6 UNITS OF 425 KWE EACR

1. Collector Dieters (Overall

_ergy Conversion 0.4)

2. Radiator Area Total

(_ml to Collector Area)

39m 126 ft.

6,950 m 2 74,846 ft. 2

(Divides between torus and platfor,9

3. _ting & l_lication,

4 Units on Platform Power for solar observatory

mlcrogravity processing, Earth/space

experiments, communications, erections

Units on Torus Power for life support, fuel

generation, control systems, data

systems, on-board fabrications

160
I ,m

I. Reservoir Dimensions O.D. 92.4 m 300 ft.

I.D. 30.4 m 100 ft.

Thickness 15 m 49 ft.

2. Inertial Balance Estimate

Moment of Inertia Required

Weight of Water Required

Water Depth at Reservoir Rim

2.47 x 109 kcjm2

1.30 x 106 k9

0.3 m

1.82 x 1091bmft2

2.90 x 106 lbs

lft

3, __C_tionsfor Final Trim to R_II

Balance: Vary the Rotatlonal

of the Reservoir, Vary

the Water Level in the Reservoir,

Z_ovMe an Auxiliary Water

O_mnel and O_trol the Flow

Veloclty

161
year. It will take less energy to accomplish this if the net angular

mumentmn about the torus spin axis is zero.

In the subject configuration, the spokes of the torus and the large

cross arm are pressurized to provide a shirtsleeve environment. _he cross

arm can be used as a repair facility for Ct_V's, 01V's, etc.

162
7.0 _ TG_ICS FOR FUTURE RESEARO_

The present study has addressed the topics assigned by contract. It is

broad enough that the writers were required to review a considerable amount

of literature and to become more familiar with the Space Station and its

subsystems. As is usually the case, the study raised some questions and

suggested additional areas of research. These areas are as follows.

i° Continue to explore areas of synergism between subsystems°

2. Explore effects of artificial "g" on sutmyste_, subjects,

materials transfer, manufacturing processes°

3. Make sketches of the Space Station interior to see where

functions occur (Geodraw). Estimate weights and inertias of

Space Station and its elements. Determine CG locations.

4. Study rotational dynamics, estimating perturbations due to

docking, control inputs, and on-board materials movement.

5. Extend point design of EPS to include effects of radiation

losses from the heat engine receivers as a function of inlet

temperature and variation of system operating tm_erature

range as a function of Cannot efficiency.

6. Evaluate the manufacture of structural elements and pressure

vessels in space, including inflatable structures for

inflating and hardening in space.

7° Study robotics applications for fabrication, assembly, and

deployment.

8. Examine applications of free flying and tethered platfonm.

163
e

10.

Provide inputs to SAB for IDEAS programs to examine Space

Station drift, reboost requirements, thermal analysis, and

structural analysis.

Investigate Space Station surface electrostatic charging and

protective measures for damage protection.

164
REFerENCES

lo

o

3.

.

Longsdon, John M. : Space Station:

AIAA/_P_A Symposium on the Space Station.

18-20, 1983.

A Historical Perspective,

Arlington, Virginia, July

km_A Space Systems Technology Model. km.SA TM 88174, June 1985.

NASA Space Systems Technology Model. Vol. I, Data Base, Part A,

Technology Forecasts; Vol. I, Data Base, Part B, Technology Forecasts;

Vol. 2, Data Base, Technology Analysis; Vol. 3, Data Base, Future

Mission Payloads; km2A TM 88176, 1985.

Pioneering the Space Frontier. The Report of the National Commission

on Space, Bantam Books, May 1986.

5. National Space Transportation and Support Study. kronA, Civil Needs

Working Group, Volume 1 - Summary Report, March 14, 1986.

6. Space Station Reference Configuration Description.

19989, August 1984.

7. Architectural Control Document.

Station Office, June 1986.

k_SA Report JSCPreliminary Report, km_A JSC Space

8. Reynolds, William C.; and Perkins, Henry C.:

Thermodynamics. McGraw-Hill Book Company, 1970.

.

I0.

ii.

12.

13.

14.

Engineering

Roschke, E. John; and Wen, L.: Preliminary Systems Definition Study

for Solar Thermal Dynamic Space Power Systems. Jet Propulsion

Laboratory, 1984.

Croxton, R. Stephen; McCrory, Robert L.; and Soures, John M.: Progress

in Laser Fusion. ScientificAmerican, Vol. 255, No. 2, August 1986.

Beals, G.; et al.: A Lightweight Nuclear Powered O_V Utilizing a

Liquid Droplet Radiator. AIAA-83-1346, AIAA/SAE/ASME, 19th Joint

Propulsion Conference, June 27-29, 1983.

Garrett, L. Bernard; and Hook, W. Ray.: A Case for Space-to-Space

Power? km_A Conference Publication No. 2035, Vol. I, 1978.

Nainiger, Joseph J.; and McKessock, David B.: Mass Area, and

Propellant Requirement Comparisons of Several Potential Growth Space

Station Power Systems. Space Station System Directorate, Lewis

Research Center, PIR No. 71, October 1984.

Roschke, E. John: Solar Dynamic Systems for Power Applications. AIAA

24th Aerospace Sciences Meeting, January 6-9, 1986.

165
15. Inertial Technology for the Future. IEEE Transactions on Aerospace and

Electronics Systems, Vol. AES-20, No. 4, July 1984.

16. Lovelace, U. M.: A Microwave Radiometer Spacecraft. American

Astronomical Society 25th Anniversary Conference, Paper AAS 78-51,

October 30 - November 2, 1978.

17. Oglevie, R. E.; and Eisenhaure, D. B." Integrated Power and Attitude

System for Space Station. AIAA 23rd Aerospace Sciences Meeting, Reno,

Nevada, January 14-17, 1985.

18. Mikulas, M. M., Jr.; et al.: Deployable/Erectable Trade Study for

Space Station Truss Structures. km_W% TM 87573, July 1985.

19. Ease/Access Postmission Management Report. I_%SA Marshall Space Flight

Center, Brochure, From Shuttle Flight 61-B (1985).

20. Bruun, E. R.; and Kingsley: Monopropellant Thruster Performance

Augmentation Heater Development Program. Hamilton Standard Division of

United Aircraft corporation, Windsor Locks, connecticut, AFRPL-TR-802,

February 1981.

21. Space Station Program Configuration Description Document.

Environmental control and Life Support System, Boeing Aircraft company,

Report 483-50004-6, June 6, 1985.

22. Parker, James F., Jr.; and West, Vita R. : Bioastronautics Data Book.

SP-3006, Second Edition, 1973.

23. Hoard, R. M.: fRC Handbook on Space Technology Status and Projections.

CRC Press, 1985.

24. Randolph, L. K.; et al.: MPD Thruster Definition Study.

Marietta Corporation, AFPRIr4R-84-046, July 1984.

Martin

25. Vulpetti, A.: An Approach to the Modeling of Matter and Antimatter

Propulsion _stems. Journal of the British Interplanetary Society,

Vol. 37, September 1984.

26. Robinson, A. L. : Antiprotons Captured at CERN. Science, Vol. 233, No.

4771, September 26, 1986.

27. Baer, T.: An Interview with Robert Haslett, Space Station Thermal

Control. Mechanical Engineering, December 1984.

28. Wise, P. C.: Space Thermal Control Technology; Design Challenges into

the 1990's. RCA Astro-Electronics Division, Princeton, New Jersey,

IAF-85-373, 36th Congress of the International Astronautical

Federation, Stockholm, Sweden, October 7-12, 1985.

166
29. Lee, A. L._ and Rhoads, G.D.: Prediction of Thermal Control Surface

Degradation Due to Atomic Oxygen Interaction. Lockheed Missiles and

Space C0ml_y, Smmyvale, California, AIAA Paper No. 85-1065, AIAA,

20th _hermophysics Conference, Willi_sburg, Virginia, June 19-21,

1985.

30. Hall, John B., Jr._ Ferebee, Melvin J., Jr.; and Sage, Karen H.:

Environmental Control and Life Support Systems Tec_mology Options for

Space Station Application. Paper No. 851376, 15th Intersoclety

Conference on Environmental SlmtEms, San Francisco, California, July

15-17, 1985.

31. Reysa, Richard P.: Regenerative Life Support System Hardware Testing-A

Summary. Paper No. 860941, 16th ICES Conference, San Diego,

California, July 14-16, 1986.

32. Sumphrles, William R.; Reuter, James L.; and Schunk, Richard G.: Space

Station Envirom,ental Control and Life Support System Distribution and

Loop Closure Studles. Paper NO. 860942, 16th IOBS Conference, San

Diego, California, July 14-16, 1986.

33. Kovach, Licla S.7 and Miller, Craig W.: Environmental Control Life

Support for the Space Station. Paper NO. 860944, 16th IC_S Conference,

San Diego, California, July 14-16, 1986.

34. EVA Description and Design Criteria. _%SA Document JSC-10615, Revision

A, May 1983.

35. Moffatt, Miles; and Abeles, Fred: The Development of an EVA Universal

Work Station. Paper NO. 860952, 16th I_ Conference, San Diego,

California, July 14-16, 1986.

6. Furr, Paul A.T Mason, Conrad B.; Sears, William J.; and R0eles, Fred

J.: Physiological Aspects of EVA. Paper No. 860991, 16th ICES

Conference, San Diego, California, July 14-16, 1986.

37. Covault, Craig: Launch Capability, EVA Concerns Force Space Station

Redesign. Aviation Week and Space Technology, July 21, 1986.

38. Michaud, Roger B.; Miller, Ladonna J.; and Premeaux, Gary R.: Concept

for the Evolution of the Space Station Program. Paper NO. 860972, 16th

IfF_S Conference, San Diego, California, July 14-16, 1986.

39. Advancing Automation and Robotic Technology for the Space Station and

for the U. S. Economy. NASA TM 87566, April I, 1985.

40. Ohork, D. F.: The Wet Air Oxidation of Wastes.

_Igineering, Vol. I01, No. 2, December 1979.

Mechanical

1. Hall, J. B.; and Brewer, D. A.: Supercritical Water Oxidation:

Concept Analysis for Evaluating Space Station Application. 16th I_

Conference, San Diego, California, July 14-16, 1986.

167
42.

3o

DeCampli, William M.: The Limits of Manned Space Flight.

Sciences, September-October 1986.

The

Cra_er, D. Bryant: Physiological Considerations of Artificial Gravity,

Application of Tethers in Space. NASA Conference Publication 2364,

Volume 1, June 15-17, 1983.

168
_IX

SOME FACTORS TO CONSIDER IN S_BCTING MATERIALS

FOR SPACE STATION APPLICATION

Some factors are not relevant to the geometric configuration of a Space

Station but must be considered in the selection of materials or components

of the Station. These are reviewed briefly.

Thermal control coatings are currently in use to successfully control

spacecraft temperatures in space. Long-term exposure of the coatings to

the space envlrom_nt and possible contaminants (such as propellant exhaust)

require the evaluation of the coating's degradation over an increased

lifetime for the advanced Space Station.

Coatings that are electrically conductive with low resistivity and long

life are required to suppress Space Station charging (Referemce 3).

Lubricants and Bearinus

The development of improved bearings and lubricants having long life in

a space environment is essential for low maintenance of rotating or sliding

elements of the Space Station. Proposed solid film lubricants applied by

advanced methods have a projected lifetime of 20 years. The need for liquid

lubricants possessing self-restoring characteristics will drive the

development of highly stable, low vapor pressure, non-contaminating liquids

for space applications. Lubricant _arout rather than metal bearing fatigue

is the life limiting component of bearing assemblies; therefore, lubricant

replenishment is one approach to extending a beating's useful life.

Bearing modeling using computer analysis and design capability is

required to aid in the selection of bearing materials, design features, and

169
lubricant selection and to predict long-term lubricant and bearing

performance (l_cerenoe 3).

Supe_rconductinq_ Materials

Niobium titanium (NbTi) alloys are state-of-the-art superconductors for

operating temperatures below 10 K for steady state magnets with current

densities of 1-2 x 104 amps/cm2 at magnetic fields up to 8 tesla.

Applications that require fast changing magnetic fields will cause a

temperature rise in the superconducting composite and loss of superconductance with a temperature rise of one to several degrees Kelvin.

Niobium tin (Nb3Sn) is a brittle intermetallic which shows promise as a fast

pulsed and/or high current density superconductor operating at temperatures

up to 18.5 K. The manufacture of Nb3sn poses enormous problems to achieve

the required thermal-electromagnetic stability of the alloy. Reliable

manufacturing methods appear feasible by 1990 based on extensive

manufacturing, processing, and metallurgical investigations Performed to

date.

Dielectric insulation materials are in development for providing very

high thermally conductive electrical insulation for superconducting wire and

embedment of wire turns in superconducting devices. These insulation

materials will improve the thermal performance by limiting the temperature

rise of devices subjected to large loss terms of the superconductor due to

changing magnetic fields (Rmferem_e 3).

Optical Materials

Materials for fabrication of first surface mirrors and telescope

structures are listed in T_bles A-I, A-2, and A-3. The specific stiffness

of each structural material is one of the key parameters of determining the

170
TABLE A-I MIRRORS FOR SPACE OPTICS (REFERENCE 3)

Glasses Type Optical Comments

Utility

Pyrex Boro-sillcate Poor

cer-vit Boro-silicate Good

Fused SiO 2 Good

SlO 2

ULE Ti sillcate Very

Good

Zerodur Class-ceramic Very

Good

Heraues Fused quartz Very

Good

Castable; plentiful supply

No longer made; stockpile gone

High purity; frit bondable; CTE good

at cryogenic temperatures

Frlt bondable; weldable; 0 CTE at

room temperature

Ultra-low CTE; machinable; not

weldable; German source

From natural quartz; ultra-pure;

good CTE for cryogenic temperatures;

German source

TABLE A-2 MIRRORS AND SUBSTRUCTURES (REFERENCE 3)

_aterial Application Comments

Beryllium Faceplate Requires Ni coating to reduce scatter

Substructure Unproven for high quality appllcatlons

Mirror Difficult (expensive) to process

Aluminum Mirror

Graphite/

Epoxy

CarbonCarbon

Graphite/

Glass

Graphite/

Magnesium

Backing structures

Metering trusses

Faceplate

Substructure

(70 ca mirror)

Substructure

Faceplate

Substructure

Very high purity required

Good specific modulus

Possible CTE problems

Not for optical surfaces

Dimensional instability with water

absorption-desorptlon

0utgasslng potential

Good thermal conductivity and dlmensional stability

Acceptable CTE

Excellent fracture resistance; toughhess

Acceptable CTE

Low angle cross-ply gives needed

dimensional stability; low CTE

Very good thermal distortion resistance

No outgasslng

171
U

i

_o_

_ g

U- I-- S

>-g

..,.,I J..,G U

LIJ r,,_ E

,,.r ,.,_ U

i.. ¢..1 _

.J (._(J

u.l _ Ll. P,,.4 _

,.-I."-'

Q.V) U

LUE

..I

IX

_0_

eeeoe

.,..,

00000

00000

eeeee

O_ ON

eeoee

"__

I

_Z_

X

I

%

X

_B

o_

OOOOO

OOOO0

000_

_M

_Z_

..I

i..,q

I,nJ

I--

.i"

U_

_Z_

172

Z _,"* 0 I--C_

,_ _ _ h..._ >..,

l II N II N II

_ e
frequencies of fundamental vibrational modes of structures fabricated from

each material (R_eremoe 3).

MateriaIR for Propellant _Systems

A major thrust for liquid propellant systems is to reduce the weight of

tanks, pumps, and plumbing. Composite materials and high strength metals

made by rapid solidification rate technology or superplastic forming

processes are being used in a variety of systems. Carbon-carbon turbine

rotors, coMbustors, and engine afterburner flaps are being developed for the

cruise missile and aircraft turbine engines. These advanced materials

applications may provide technology spin-off for the advanced Space Station.

Long-life monopropellant space thrusters can be manufactured using a

hydrazin_tible AFE ethylene propylene rubber for valve seats, seals,

and expulsion devices and have become universally accepted for hydrazine

monopropellant applications. An equivalent elastumer is required for longlife applications of bi-propellant systems to serve in expulsion

applications of the oxidizer nitrogen tetroxide (N204) • Manufacturing

processes and demonstration of elastumer long-life compatibility with N204

remains to be established (R_zlm_e 3).

Materials for Solar Cells

Silicon is the primary solar cell material in use today with a

beginning of life efficiency of 15 percent. End-of-life efficiency is

considerably lower because of space radiation effects; therefore, solar

array design and subsystem efficiency must be sized on the end-of-life

efficiency.

GallimN arsenide (GaAs) may emerge as the next generation solar cell

material having a beginning of life efficiency of 15 percent. This

173
beginning efficiency is expected to increase with material development.

GaAs has a high resistance to space radiation and can perform at a higher

temperature than silicon solar cells. T_) promising concepts for use of the

CmAs solar cells are the flux concentrator and multiple bandgap arrays. The

flux concentrator will permit a reduction in size and weight of collector

panels and operate at higher efficiency than standard panels. Multiple

bandgap arrays will use existing technology to achieve higher conversion

efficiency by expanding the frequency bands at which light is absorbed

(Red.fence 3).

Materials for C_ryoqenic _Applications

Space Station storage of propulsion fuels and other fluids at cryogenic

temperatures requires materials that can function below 120 K without

degradation, cracking, or embrittlement. Cryogenic tankage problem areas

are primarily in conduction and radiation heat leakage paths through the

thermal insulation and structural supports. A secondary problem results

from thermal paths through fittings, ports, and attachments.

Fiberglass reinforced epoxy resin composites are presently used for

most cryogenic piping applications, but improvements in the matrix resin's

resistance to embrittlement and notch sensitivity are required to preclude

leakage and evaporation.

Mechanical refrigerators require low friction and non-%_aring materials

for long-service life. There is a need for matched expansion seals, which

permit the least pressure differential across the seals with extended life

capabilities.

The regenerators of mechanical refrigerators require materials with a

higher specific heat in the 4-12 K range than lead that is currently used.

174
Ceremics and metals that undergo magnetic phase transformations at very low

temperatures are being developed for this application.

Various felts and screens are used to transfer cryogenic fluids in the

zero "g" space enviro_snt. Improving and controlling the surface fluid

wetting characteristics of the felt or screens could increase the flow rate

of the fluld.

Low emissivity, durable coatings and coatings with variable or

controllable thermal conductance properties are being sought for use with

thermal switches (l_ference 3).

Q0ntaai uLQn

The advanced Sface Station will be assembled and serviced in space by

transporting structural elements, modules, subsystem components, and

resupply items from Earth to low Earth orbit by way of the available space

transportation systels, ttmanned cargo vehicles and new manned vehicles

will transport cargo from ground station to LEO. An orbital maneuvering

vehicle would transfer cargo from the urmanned cargo vehicle to the Space

Station. An orbital transfer vehicle would transfer people and cargo from

the Space Station to geosynchronous equatorial orbit and to other cislunar

orbital positions.

The functionality and mission effectiveness of various Space Station

subsystems may be impaired if contamination of telescope mirrors, infrared

detectors, and solar cell arrays (as examples) should result during

prelaunch, at launch site, ascent, on-orbit, or other space environments.

Surface contamination can occur due to Particulate, liquid, or vapors

deposited during the cargo transit envirom_nt or from materials outgassing

from the subsystem itself in space.

175
Possible sources of contamination for the Space Station sub6ystem or

associated spacecraft are identified in Flgure A-1.

Structural materials should not degrade in a space environment to yield

conta_ants. Ideally they should be precleanable prior to launch and not

degrade by outgassing, moisture desorptlon, ultraviolet exposure, electrons,

protons, mlcro_teorold impacts, or atomic oxygen when in orbit.

A comprehensive contamination control program must be implemented from

subsystem design through end use to be totally effective (Reference 3).

176
177

"o

m
1. Report No. 2. Government Accession No.

NASA CR-178208

4. Title and Subtitle

An Advanced Technology Space Station for the

Year 2025, Study and Concepts

7. Author(s)

M. J. Queijo, P. J. Butterfield, W. F. Cuddihy,

C. B. King and P. A. Garn .... ,,, , ,

9. Performing Organization Name and Address

The Bionetics Corporation

20 Research Drive

3. Recipient's Catalog No.

5. Report Date

March 1987

6. Performing Organization Code

8, Performing Organization Report No.

10. Work Unit No.

'11. Contract or Grant No.

Hampton, Virginia 23666 NAS1-18267

13. Type of Report and P.eriod Covered

'12. Sponsoring Agency Name and Address Contractor Keporl;

National Aeronautics and Space Administration

Washington, DC 20546

15. Supplementary Notes

Langley Technical Monitor - Melvin J. Ferebee, Jr.

May= Nov. 1986

14. Spon_ring A_ncy Code

506-49-31-01

16. A_tract

A survey has been made of potential Space Station missions that might exist

in the 2020-2030 time period. Also, a brief study of the current state-of-the-art

of the major subsystems was undertaken, and trends in technologies that could

impact the subsystems were reviewed.

then used to arrive at a conceptual

Factors addressed in the conceptual

The results of the survey and study were

design of a Space Station for the year 2025.

design included requirements for artificial

gravity, synergies between subsystems, and the use of robotics. Suggestions

are made relative to more in-depth studies concerning the conceptual design

and alternative configurations.

17. Key Words (Suggested by Authoris))

Space Station

Technology Trends

Advanced Systems

Space Systems

19, Security Cla=if. (of this report)

18. Distribution Statement

Unclassified - Unlimited

Unclassified

Subject Category 15

20. Security Classif. (of this page) 21. No. of Pages 22. Price

Unclassified 190 A09

N-305 Fo_sale by the NationalTechnicalInfo(mationService.Springfield,VirEinia 2216[