Space Sation Reference 

Configuration Description 

[blASA-TLI-E7 49 3) 5PACi i iA IIL ti BiPEREYCE b185-3 11 46 

CCYPXGUEATICN DISCfiiI" ,CN (EASA; 798 p 

HC AS3/tli E33 CSC- L~L L'nclas 

G3/18 21&26 

Systems Enginsering and Integralion 

Space Station i7rogram Office 

August 1984 

National Aeronautics and 

Space Adrn~n~stration 

Lyndon 6. Johnson Spaca Center 

tiouston. Texas 
SPACE STATION REFERENCE 

CUNFIGURATION DESCRIPTION 

SYSTEMS ENGINEERING AND INTEGRATION 

br'ACE STATION PROGRAM OFFICE 

AUGUST 1984 

NATIONAL AFRONAUT IC S AN3 

SPACE ADMINISTRATION 

LYNDOY B. JOHNSON SPACE CENTER 

HOUSTON, TEXAS 
ACA 

AC S 

ARF 

ASTP 

ATCS 

CER 

CMG 

C8T 

DDTAE 

DHS 

ECLSS 

EC S 

EMU 

ESS 

EVA 

FOV 

GNM: 

6PS 

GS E 

HDR 

HRF 

Hz 

If 

I DMS 

IN 

LAB 

Li st of Acronyms and Abbreviations 

Attitude Control Assembly 

Attitude Control System 

Animal Research Facility 

Apollo Soyuz Test Project 

Active Thermal Control Subsystem 

Cost Estimate Relationships 

Control Moment Gyro 

Communications arld Tracking 

Design, Cevelop, Test, and Evaluatioh 

Data Management System 

Environmental ControllLife Support System 

Energy Conversion Subsystem 

Extravehicular Mobil ity Unit 

Energy Storage Subsystem 

Extravehicular Activity 

Fiow Control Assembly 

Free Flyer 

Field of View 

Guidance, Navigation, and Control 

Gl obal Posi ti on Satell i te 

Ground Support Equipmer, t 

High Data Rate 

Habitation Vodule 

Health Maintenance Facility 

Human Research Faci 1 i ty 

Hertz (Cycles Per Second) 

Intermediate Frequency 

Information and Data Management Subsystem 

Initial Operational Capability 

Inertial Sensor Assemblies 

Laboratory Module 

Logistics Module 

Low Data Rate 
LVLH 

m 

MA 

MC 

ML I 

MMPF 

MMU 

MRMS 

MRW G 

nm 

Ocz 

ORU 

OiV 

OMV 

PA 

PAC 

PCU 

POP 

PF 

PID 

PMAD 

RC ti 

RFP 

RMS 

RC S 

SAA 

SE& i 

SS 

SSPE 

STS 

TBD 

TEA 

TDM 

TDRS 

WC S 

Local Vertical , Local Horizontal 

Meter 

Mu1 tip1 e Access 

Manned Core 

Mu1 tilayer Insulation 

Microgravi ty and Materials Processing Facility 

Manned Maneuveri ng Uni t 

Mobile R~mote Manipulator Systeln 

Mission Requirement Working Group 

Nautical Miles 

Operational Control Zone 

Orbital Rep1 acement Unit 

Orbi tal Transfer Vehicle 

Orbital Mdneuveri ng Vehicle 

Power Amp1 i f i cr 

Pci n ti ng and Control System 

Power Conversion Unit 

Perpendicular to Orbi t Plane 

Platform 

Proportional , Integral , and Differential Gain 

Power Management and Di stri bution 

Reference Concept Group 

Request for Proposal 

Remote Manipulator System 

Reaction Control S.Y stem 

Science and Applications 

System Engineering and Integration 

Space Station 

Space Station Program Elements 

Space Transportation System 

To Be Determined 

Torque Equil ibrium Attitude 

Technology Uevel opment Missions 

Tracking and Datd Relay Satellite 

Waste Control System 
CONTENTS 

Page 

1.0 INTROOUCTION .......................... 1 

2.0 BACKGROUND........................... 2 

2 3.0 SUMMARY............................. . 

3.1 CONFIGURATION DESCRIP'LON ................. 3 

3.2 SUBSYSTEM DESCRIPTION AND PERFORMANCE ........... 11 

3.3 HODULE ARRANGEMENT. .................... 14 

4.0 SPACE STATION MANNED CORE. ................... 20 

4.1 DESIGN REQUIREMENTS 

4.2 CONFIGURATION DESCRIPTION 

4.2.1 General Arrangement 

4.2.1.1 Subsystem Instal lation 

4.2.1.2 Customer Accommodation 

4.2.2 Fli ht Modes 

4.2.3 & 

4.2.4 -perties 

4.3 CONFIGURATION EVALUATION SUWRY. ............. 61 

4.3.1 Pavl oad Accommodations ............... 63 

4.3.1.1 Representative Pay1 oad Set 

4.3.1.2 Attached Payloads 

4.3.1.3 Pressurized Laboratory Modules 

4.3.1.4 Servicing Facility 

4.3.1.5 Growth 

4.3.2 Crew Accommodations. ................ 79 

4.3.2.1 Habitation Volume. ............ 79 

4.3.2.1.4 Personal hygfene areas 

4.3.2.1.5 Porkstations 

4.3.2.1.6 Crew stora 

4.3.2.1.7 dlty 
Page 

4.3.2.2 Nonhabi tati on Yo1 ume ........... 99 

4.3.2.2.1 Crew accomrodations in 

? aboratory vol ume 

4.3.2.2.2 Airlocks/hyperbaric 

chamber -.- -~ 

4.3.2.2.3 Safeven 

4.3.2.2.4 Crew accanmodations in 

loaistics module 

4.3.2.3 Crew Facility Equipment .......... 106 

Secondary structure 

Flatches 

iGaixuns 

IJETTties 

w 

RJXZance - faci 1 - i ti 

4.3.2.4 Volumetric Analysis ............ 115 

4.3.3 Dynamics and Controls. ............... 125 

4.3.3.1 Appl ied Forces and Torques 

4.3.3.2 Control System Definition 

4.3.3.3 Dynamic Model and Free Vibration 

Characteristics 

4.3.3.4 Mass Properties 

4.3.3.5 Performance Assessment 

4.3.3.6 Sumnary 

4.3.4 Power Capability .................. 265 

4.3.4.1 General Configuration 

4.3.4.2 Power Generation System 

3.3.4.3 Energy Storage System 

4.3.4.4 Power Management and Distribution 

4.3.5 Thermal Capability ................. 273 

4.3.5.1 Introduction 

4.3.5.2 Thermal Integration 

4.3.5.3 System Overview and Capabil i ty 
Page 

4.3.6 Colnraunications and Tracking Antenna Systems. .... 281 

4.3.6.1 Required Antenna Coverage and 

Operational Zones 

4.3.6.2 Antenna System Description 

4.3.6.3 Antenna Locations 

4.3.7 Launch and Assembly. ................ 305 

4.3.7.1 Introduction 

4.3.7.2 LaunchIAssembly Description 

4.3.7.3 Shuttle Payload Bay Packaging 

4.3.7.4 Component Instal lation Description 

4.3.7.5 Assembly OperationsICrew Time1 ines 

4.3.8 Operations ..................... 353 

4.3.8.1 Manipulator Operations 

4.3.8.2 Orbital Flight Operations 

4.3.8.3 Construction of Large Structures 

4.3.8.4 LogisticsIResupply Operations 

4.3.9 Safety ....................... 402 

4.3.9.1 General 

4.3.9.2 Habitable Module Egress Capability 

4.3.9.3 Protection/Control of Hi gh-Pressure or 

Hazardous Fl ui d Tanks 

4.3.9.4 Isolation of Modules After an Accident 

4.3.9.5 Reaction Time After the Occurance of 

a Leak 

4.3.9.6 Safe Haven Capabil ities 

4.3.9.7 EVA Operations 

4.3.9.8 Repair and Reactivation of Modules 

After an Accident 

4.3.9.9 Orbiter Docking Ports 

4.3.9.10 Other Safety Considerations 

4.4 SUBSYSTEM DESCRIPTION ................... 465 

4.4.1 Electrical Power .................. 405 

4.4.1.1 Introduction 

4.4.1.2 Photovol taic System and Generation 

and Storage 

4.4.1.3 Dynamic System Generation and 

Storage 

4.4.1.4 Power Managellent and Ui stribution 
Page 

4.4.2 Guidance, Navigation, and Control. ......... 415 

5.4.2.1 Functional Description 

4,4.2.2 Equipment 

4.4.3 Comnunications and Tracking System ......... 432 

Ccnmnunitations/Tracking System 

Definition 

Radio Frequency Communications 

Audio System 

Video System 

TrackingIDocking System 

Signal Processing 

Data Acquisition, TelemetryiCND 

Systems Idoni toring and Control 

4.4.4 Infomation and Data Management. .......... 451 

4.4.4.1 Introduction 

4.4.4.2 IDMS Architecture 

4.4.4.3 Platform IDMS 

4.4.5 Propulsion ..................... 459 

4.4.5.1 Introduction 

4.4.5.2 Subsystem Requirements 

4.4.5.3 Physicdl and Function Description 

4.4.5.4 Propellant Storage, Distribution, 

Managerne. t, and Resupply 

4.4.5.5 Concerns and Issues 

4.4.6 Environmental Control and Life Support System. ... 475 

4.4.6.1 Introduction 

4.4.6.2 System Functional Requirements 

4.4.6.3 Reference Configuration 

4.4.6.4 Conclusions 

4.4.7 Thermal Control. ................. .486 

4.4.7.1 Introduction 

4.4.7.2 Heat Acqui si tion and Transport 

4.4.7.3 Reference Configuration Heat 

Rejection System 

4.4.7.4 Passive Thermal Control IProtec tion 

4.4.7.5 Summary and Conclusions 
Page 

4.4.8 Structures and Mechanisms. ............. 542 

4.4.8.1 Introduction 

4.4.8.2 Truss Structures and Construction 

Considerations 

4.4.8.3 Prel iminary itructural Uesi gn and 

Analysis cf a Shuttl e-Launched Space 

Station Man~sd Habi tab1 e Module 

4.4.8.4 Module and Payload Attachment 

4.4.8.5 ModuleFIodule Interface and 

Docki ng/Berthing 

4.4.6.6 Solar Panel Deployment Mechanism 

4.4.2.7 Conceptional Rotary Joint Design for 

Space Station Appl i ca ti ons 

4.4.8.8 Deployment Verification 

4.4.8.9 Conceptional Design of a MRMS 

4.4.8.10 Sumry 

4.4.9 EVA 5ystezs. .................... 637 

4.4.9.1 i,?t~oduction 

4.4.9.2 Space Station Dedicated EVA Systems 

Configuration 

4.4.9.3 Government-Furni shed EVA Hardware 

5.0 UNMANNED PLATFORMS ....................... 652 

5.1 DESIGNREQUIREMENTS. .................. .653 

5.1.1 General Desi n Requirements 

5.1.2 d ay oa s an Missions 

5.1.2.1 Polar Pay1 oads and Missions 

5.1.2.2 Co-orbitincj Mission Paylcads 

5.2 CONFIGURATION DESCRIPTION ................. 659 

5.2.1 General Arrangement 

5.2.2 Ueiaht and Power 

5.2.3.1 Polar Platform 

5.2.3.2 Co-orbi ting Platform 

5.2.4 Launch and Servicing Operations 

5.2.4.1 Launch Operations 

5.2.4.2 Servicing Operations 
5.2.4.2.1 Servicing by STS/OMV 

5.2.4.2.2 Servicing by Space station 

5.2.5 Dynamics and Control ............... 680 

5.2.5.1 Disturbance Forces and Torques 

5.2.5.2 Reaction Control System 

5.2.5.3 Structural Dynamics 

5.2.5.4 Riai d Body Control Dynamics 

5.2.5.4.1 Simulation capabilities 

5.2.5.4.2 Rass ~ro~erties 

5.2.5.4.3 )lighk mbdesnulated 

5.2.5.4.4 ton trol reaui rements for 

disturbance environment 

5.3 SUBSYSTEMS. ........................ 726 

5.3.1 -- Structure ..................... 726 

5.3.1.1 Subsystem Functions 

5.3.1.2 Potential Hardware Implementation 

5.3.1.3 Space Station Refererice Hardware 

Appl icabil i ty 

5.3.1.4 Platform Reference Design 

5.3.1.5 Weight Table 

5.3.2 Mechanisms ..................... 731 

5.3.2.1 Subsystem Functions 

5.3.2.2 Potenti a1 Hardware Implementation 

5.3.2.3 Space Station Reference Hardware 

Applicability 

5.3.2.4 Platform Reference Design 

5.3.2.5 Weight Table 

5.3.3 -. Thermal Control .................. 732 

5.3.3.1 Overview 

5.3.3.2 Subsystem Functions 

5.3.3.3 Subsystem Capabilf ties and Sizing 

5.3.4 Electrical Power .................. 733 

5.3.4.1 Subsystem Requirements 

5.3.4.2 Potential Hardware Implementation 
5.3.4.2.1 Power generation 

5.3.4.2.2 

distribution 

5.3.4.3 Space Station Reference Hardware 

Appl icabil i ty 

5.3.4.3.1 Conversion, control, and 

di stribution eaui~nsent 

5.3.4.3.2 Solar arrav 

5.3.4.4 Platform Reference Des! yn 

5.3.4.4.1 Power netcation 

5.3.4.4.2 nergy stor3 --T- 5.3.4.4.3 ~,onversionanu distribution 

equipment 

5.3.4.5 Weight Table 

5.3.5 Colnmrnications and Tracking . . . . . . . . . . . . 736 

5.3.5.1 Subsystem Functions 

5.3.5.2 Potenti a1 Hardware Implementation 

5.3.5.3 3pace Station Reference Hardware 

Applicability 

5.3.5.4 Weight and Power Table 

5.3.6 Information and Data Management Subsystem . . . . . 739 

5.3.6.1 Subsystem Capabilities 

5.3.6.2 Subsystem Functions 

5.3.6.3 Potential Hardwcre Implementation 

5.3.6.4 Space Station Reference Hardware 

Applicability 

5.3.6.5 Platform Reference Design 

5.3.6.6 Weight and Power Table 

5.3.7 Guidance, Navigation, and Control . . . . . . . . . 743 

5.3.7.1 Subsystem Characteristics 

5.3.7.2 Subsystem Function!: 

5.3.7.3 Potential Hardware Implementation 

5.3.7.4 Space Station Reference Hardware 

Appl icabi 1 i ty 

5.3.7.5 Platform Reference Design 

5.3.7.6 Weight and Power Table 
5.3.8 Propulsion -..................... 747 

Characteristics 

Functions 

Potenti a1 Hardware Imp1 ementation 

Subsystem Performance 

Thruster Clusters 

Hydrazine Lines 

Redundancy 

Cold Gas System 

Weight and Power Tables 

5.4 SAFEN.. ......................... 752 

Introduction 

Protection/Control of Hi gh-Pressure or Hazardous 

kluid Tanks 

Frooul si on Systems - 

i mi ty Operations 

'tVA O~erations 

on Berthing Capabil i ties '&zz::y Consl deraCI ons 

REFERENCES ........................... 759 

APPENDIX A - ASSEMBLY OPERATIONS CREW TIMLINES EVALUATION ... 762 
I NTRODUCT ION 

This report documents the data generated by the Space Station Program 

"Skunk Works" over a period of 4 months which supports the definition of 

a Space Station reference configuration. The data were generated to meet 

these ob jecti ves: 

a. Provide a focal point for the definition and assessment of 

program requirements 

b. Establish a basis for estimating program cost 

c. Define a reference configuration in sufficfent detail to allow 

its inclusion in the definition phase Request for Proposal (RFP) 

Although the reference configuration has beet specifi~?d in the RFP, the 

specific data defining that configuration in this docr~ment should be 

viewed as preliminary, subject to change, arrd presented for information 

only. It was NASA's intent in providing this information that it be 

considered as a otential point of departure for the definition phase, 

not a set of approve % esign solutions. Basel ining of reference 

configuration characteristics will occur prior to definl tion phase 

Contract Start Date (CSD). 

This report addresses the IOC and growth of the Manned Station and 

Unmanrled Platforms. Section 3 presents a summar, of the configuration 

descr! pti on, subsysteas, and key eval uaticin results for the Space Station 

and Platforms. Sections 4 and 5 contain key design requirements and 

deccioiptions of the configurations of the Manned Core Station and 

tform, respectively. A functional evaluation is made of the 

tfi guration 's abil i ty to accomroda te payloads, suppolat crew an' 

~gi stics operations, accommodate ictegrated systems requirements, and to 

ae assembled and grown efficiently. Detdil ed subsystem descriptions with 

options considered and rationale used to select the specific subsystms, 

are a1 so presented. 

In the early phase of this effort, 14 tec; ical areas were identified 

which served as "targets" for systemlsubsystem alternative eva'luation and 

analysis. The data developed in these areas was considered too detailed 

for inclusiort in this document. The study results of these technical 

areas are published as "Sptze Station Subsystem White Papers" (JSC-20054) 

and the ti tles of each paper are identi f ied under the 11 sting of 

r~ference documents in the back of th!s report. 
The Space Station Program Office "Skunk YorksM activity was initiated to 

support RFP preparation, to develop a reference configuration, and to 

provide supporting data bases. The Reference Configuration Group (RCG), 

a subgroup of the "Skunk Worksu SE&I team, wa! tasked to develop 

configuration concepts :zd supporting data. As a point of departure a7d 

continuity from previous studies, five concepts were 3;!?cted for 

evaluation from those identified by the NASA Headquarters Cob-cept 

Dev~l opmnt Group (CDG). They included the 'UG Planar", the 

"Delta-Truss':, the 'Big T-lrussM, a cross-type structure referr?d to as 

the "Power Tower", and a configurrtion with a rotating solar array called 

the "Spinnerw. 

Early in the study, the five configuration concepts were narrowed to 

three based on the finding that the sal icnt features of a1 1 were 

incorporated in the remaining three. The three concepts selected for 

further study were the "COG Planar', the 'Delta-Truss', and the "Power 

Toweru. For these configurations, mission functional requirements were 

converted to design requirements and detailed Space Station and Platform 

configurations were developed. Systems and subsystems were sized 

accordingly. Cost sensi ti vi ty studies were performed and the integration 

of subsystems, crew, and customer requirements into the Station 

configurations were investigated. 

The "Power Tower" was finally selected as the reference configuration 

because it was seen as maximizing the accommodation of current user and 

growth requirements whi 1 e demonstrating acceptable design afid operations 

characteristics. It was also recognized that the "PlanarY and "Power 

Toweru configurations are members of the same family which differ 

basically in their placement of the manned modules and experiment bases 

with respect to the articulated solar collection devices. 

The RCG was comprised of approximately 80 technical specialists from 

various NASA centers, including L'SC, MSFC, GSFC, KSC, LaRC, LeRC, JPL, 

and Hqs. A listing of personnel involved in the study is contained in 

Section 6.0 of the report. 
3.1 CMFIGURATION DESCRIPTION 

A brief overvie~ of the Space Station reference configuration is provided 

in this section. More details are provided in subsequent paragraphs. 

The reference IOC Space Station configuration is shown in Figure 3.1-1. 

The Space Station operates in a local vertical-local horizontal (LVLH) 

orientation, with its keel along the local vertical direction and tk 

solar array boom perpendtcdl ar to the orbit plane (POP). The 

earth-pointed end of the Space Station contains earth-looking payloads. 

The zeni th-pointed end contains solar, stel 1 ar, and anti -earth viewing 

payloads and colwnications antennas. Won-viewing payloads are located 

at various places on the Space Station and the pressurized modules are 

located near the bottom of the keel. Servicing equiplnent is located 

along the keel on either side, with the front and back surfaces of the 

keel kept free for traverce of the Mobile Remte Manipulator System 

(URNS). The servicing anu refuel ing faci 1 i ties, OnV and OTV techno1 ogy 

demnstration equipment, and sate1 1 i te storage and equip~nt areas are 

located at vario~s places along the structure. 

GiRbaled solar array wings pmvade full power at any relative alignment 

of the Space Station and sunline. The solar voltaic wirer generation 

system was used in this studv and is shown on the configuration to 

demnstrate a design option rather than to advocate the ultimate system 

selection. Heat rejection is provided by a cc.ination of body-mounted 

radiators on the aodul~s, deployed non-rotatins radiators on the 

transverse boom, and deployed rotating radiators near the bottom of the 

keel. 

;'he long, thin shape of the Space Station gives it a natural proclivity 

tkmsrds gravi ty-gradient attitude control. Aerodynamic forces and any 

unbalanced side-to-side masses tend to cause s ti1 t of the keel off the 

nadir-zeni th line, but thz impact on the Attitude Control System can be 

minimized by maintaining reasonable side-to-side balance in mass 

distributions, and by allowing the keel to tilt slightly in the orbit 

plane, if required. Locations for attacknt of large transient masses, 

such as the Orbiter and large space construction eleaents, have been 

selected to minimize impacts on the Attitude Control System. 

One of the principal advantages of this configuration is the good viewing 

afforded to all payloads, both externally-lounted and 

internal ly-mounted. The configuration a1 so a1 1 ows good accorrodation of 

tether payloads and good accomodati on of carunication antennas. Good 

clearances are provided for Orbiter rendezvuus and berthing, and for 

construction, servicing, ad other operations activities. The deployable . truss-munted subsystems and distribution equfpment are mostly 

pre-i ntegrated (prior to 1 aunch) to mi niai ze on-orbi t time, conpl exi ty , risk, and especially EVA activities, during buildup and asse*ly. The 

mdul e-mounted subsystems and di stri bution equi plnent are 

pre-integrated also. Launch and assenbly of the IOC Space Station 

requires frol six to eight launches, including sole but not all 

pqyloads. Assembly is accmpl ished using the Orbiter WS and the Space 

Station MRHS after it is installed. Permiinently manned operation can 

begin after launch 5. 

The payload conpleklent sh~wn on Figure 3.1-1 is no doubt mre extensive 

than that which will exist on the early Station, but it was utilized here 

as a test of accamodation capability. Although physical locations are 

provided for this quantity of pqyloads, Space Station resources such as 

power, data and crew time would have to be time-shared amng the 

pay 1 oads . 

Some options for the truss structure on the Station are shown in Figure 

3.1-2. Some of these options are deployable, some are erectable, some 

are pre-i~tegrated with subsystems, and some have subsystems installed on 

orbit after deploynent of the structure. These are discussed in more 

detail in the structures section. 

A weight suarnary of the Station is provided in Table 3.1-1. It should be 

noted that these weight numbers do not include growth. The pqyload 

weight is probably the softest of the numbers shown, since the payload 

complement used in the study was optimistic. 

Five pressurized modules are utilized in the on-orbit configuration of 

the reference Station: two Habitation Modules, two Laboratory Modules, 

and a Logistics Module. In addition, a second Logistics Module must be 

provided in the program, for on-orbit exchange with the first one, to 

accomplish Station resupply every 90 days. Crew rotation or partial crew 

rotation will occur during resupply visits. 

One Laboratory Module is characterized as primarily a Life Sciences 

research module and the other as primarily a Materials Sciences research 

module, for purposes of this study. However, either module can be 

modified on-orbit to support other activities. 

The module arranqement is discussed in paragraph 3.3. Cost. analyses have 

shown that considerable cost avoidance can accrue from having cmmnalit. 

among the pressurized modules. Cursory systems assessments have 

indicated that a fairly high degree of commonality should be possible. 

More aetails are provided later in the report. 

A man-tended Station which is operated and maintained by a crew living in 

the Orbiter and which operates unmanned at other times may offer some 

cost deferral if pressur :zed mod~les on the Station are eliminated from 

the IOC version. Assuming that the Station wi '1 ultimately grow to a 

fully manned version, the overall configuration will not differ 

significantly from the reference concept described in thi s document (Fig. 

3.1-3). 

Platform concepts can be derived from the Station subsystems. Table 

3.1-2 identifies some implications of trying to achieve different levels 

TABLE 3.1-1 - IOC SPACE STATION WET WEIGHT (LBS) 

TRUSS ATTACH 

SUBSYSTEM HI41 HM2 LM LAB 1 LAB2 STRUC COMPNTS PAYLOA

ELECT PWR 

GN&C 

COMM & TRK 

> DATA HANDL 

5 u PROPULSION 

). ECLSS 

TC S 

STR & MECN 

CREW SYS 

PAYLOADS 

TOTAL 37,942 34,163 33,884 39,495 55,305 29,983 31,679 184,49
TABLE 3.1-2. - PLATFORM CONCEPT SPECTRUM 

DEGREE OF COMMONALITY WITH SS 

- HIGH - MDIUM - LOW 

Subsystenr -SS resource -Partof existing -New structure 

Module struct SS resource module -Non-SS subsystems 

-SS subsystems structure 

SELI -Most13 existing -Partially existing -Partially existing-to-mstly new 

Logistics/ -Conmiin 

Spares 

Maintainabil i tyl -Cmn 

Maintenance 

Growth -High degree; -High degree; -Probably 1 i mi ted; 

already scarred partially scarred scats must be 

added 

Sizeheight -Largest & -Medium size & -Probably small €st 

heaviest weight and lightest 

Cost -Low DLD but -Higher DAD, -Hi gher-to highest 

manned CER's; mixed CER's; DLD , but unmanned 

cost TBD cost TBD CER's; cost TBD 

0' TDRSS ANTENNA 

36' ANTENNA MAST 

FIXED RADIATOR/ 

HEAT EXCHANGER 

2' END OF FLAT 

PLATFORM EQUIPMENT 

MODULES BOTH SIDES 

OF FLAT BED 

- POWER 

- CdOH 

- R.F. 

- GN&C 

Figure 3.1-4.- CORE PLATFORM (WITHOUT PAYLOAD) 

. 
of co~onal ity. Figure 3.1-4 shows a generic platforn derived from the 

reference Station. Platforms that are sized to best met the user needs 

in a cost-effective manner should be studied. 

There is fro "safe haven" per se in the Space Station. Rather, the 

Station meets the basic safety requirements of 1) operdbility and safety 

after the loss of any one module, 2) survivability of the crew for 22 

days, and 3) rescuability by providing safe exit from and isolation of 

aqj one mndule from the others and sufficient life support, food, waste 

management, control /cownunications, and rescuabil i ty within the remaining 

three-modul e cluster . Assessments of berthing/docking appear in the approprfate areas of the 

report (dynamics, mec hani sms, operations, etc. ) . The growth versi on of 

the Space Station is discussed in later portions of the report. 

3.2 SUBSYSTEM DESCRIPTION AND PERFORMANCE 

A sumnary listing of some of the key features of the reference subsystems 

is provided in Table 3.2-1. Each of the subsystems is described in more 

detail in appropriate sections of the report. 

The Station reference subsystems consi st mostly of exi sting and 

state-of-the- art equipmert/designs. However, in some areas, it was 

deemed to be more cost-effective to provide some advance in the 

state-of-the-art. 

The Guidance, Navigation, and Control (GN&C) and Propulsion subsystems 

are probably the ones requiring fewest advances in the state-of-theart. The Crew, Information and Data Management, Communications and 

Tracking (C&T), and Structuresfiechani sms subsystems require a moderate 

degree of techno1 ogy/advanced devel opment. The Electrical Power, Thermal 

Control, and Environmental Control and Life Support (ECLS) subsystems 

probably require more advancement than the other subsystems. Further 

study mqy indicate that changes should be made in the subsystem 

concepts. One subsystem concept which should be given si gni ffcant 

attention in further studies is the solar dynamic concept of the 

Electrical Power subsystem. A1 1 presently identified technology 

advancement areas should be studied more thorou~4Sy, to verify their 

expected costs and performance. 

A top-level sumnary of the IOC SS performance in key subs.vstems areas is 

provided as T.Sle 3.2-2. The flexible body frequency is low, but 

analyses have indicated that there should be no interaction with the GN&C 

or other subsystems. The number of Control Moment Gyros (LMG's) is very 

moderate. Housekeeping power is fairly sizable, but possibly can be 

reduced. Some payloads should be able to utilize the pointing accuracy 

and stabil i ty of the Station directly, without requiring pointing mounts, 

. but others will require pointing mounts. 
TABLE 3.2-1 - KEY FEATURES OF SS REFERENCE SUBSYSTEMS 

ELECTRICAL POWER I NFORMT ION AND DATA MANAGEMSOT 

o PLANAR SOLAR ARRAY (DEPLOYABLE) o HOUSEKEEPING & PAYLOAD DATA BUSES 

o REGENERATIVE FUEL CELLS o HIERARCHICAL SYSTEM 

o GIMBALED ARRAY WINGS (2 AXES) o FIBER-OPTIC NETWORK 

o HIGH VOLTAGE AC DISTR. o FIXED AND PORTABLE CONTROL CONSOLES 

o SOME FLAT-SCREEN DISPLAYS 

- GN &C 

- COW & TRACKING 

o DOUBLE-GIMBALED CMG'S 

o MAGNETIC TORQUERS o KU-BAND & S-BAND (MULTI-FREQUENCY) 

o RATE GYROS, STAR-TRACKERS o DIGITAL COMMUNICATIONS 

ACCELEROMTERS o MULTIPLE-STEERED-BEAM ANTENNAS 

o CONTROL SIGNALS TO RCS o WIRELESS CREW COMM 

(REBOOST & BACKUP ACS) o RADAR TRACKING & OPTICAL BERTHING 

o ENCRYPTION/DECRYPTION 

PROPUI-SION 

CREW SYSTEMS 

o REBOOST & ACS (BACKUP TNRUSTERS 

o PRQPELLANT STORAGE & FEED EQUIPT. o PRIVATE CREW QUARTERS 

o BLOWDOWN SYSTEM o WARD ROOM & GALLEY (INCLUDING 

FOOD FREEZER & DISHWASHER) 

o HEALTH MAINTENANCE FACILITY 

- ECLS o AIRLOCK( S) WITH HYPERBARIC CHAMBER 

o EVA SUITS, WITH STOWAGE/CLEAN ING/ 

o REGENERABLE C02 CONCENTRATION MAINTENANCE 

REDUCTION o WASTE MANAGEMENT COLLECTION 

o N2 STORAGE &I RESUPPLY o SHOWERS & HAND WASHERS 

o CONDENSATE, WASH WATER, & o CLOTHES WASHER & DRYER 

URINE/FLUSH WATER RECOVERY o TRASH MANAGEMNT SYSTEM 

o WATER ELECTROLYSIS o EQUIPRNT MAINTENANCE FACILITY 

o WASTE MGT. PROCESSING & STORAGE o VEHICLE MANAGEMENT 

o SAFE HAVEN 

THERMAL CONTROL - 

PAYLOAD & SERVICING ACCOMMODATIONS 

o CENTRAL IZED/DECENTRALIZED HEAT 

REJECTiON o BASIC LAB VOLUME 

o FIA~D/ORIENTABLE/DEPLOYED/NON- o ATTACHED UNPRESSURIED PAYLOADS 

DEPLOYED RADIATORS (ALL HEAT o ATTACHED PRESSURIZED PAYLOADS 

PIPES) , SOME CONSTRUCTABLE Q OMV (IOC), HMU (IOC),& OTV (GROWH) 

o 2-PHASE HEAT TRANSPORT o SERVICING OF FREE-FLY ING SATELLITES 

o THERMAL CAPACITOR L PLATFORMS, OMV, MMU, & OTV 

o CONSTRUCT ION OF LARGE STRUCTURES 

STRUCTURES/MECHAN I SMS 

o 30' (CYL, PORTIOI;) PRESSURIZED 

MODULES 

o DEPLOYABLE TRUSSES 

o ERECTABLE STRUCTURE (MINIMIZED) 

o BERTHING 

o MOBILE RMS 

o ROTARY JOINTS (ELECTRICAL & FLUID) 
TABLE 3.2-2 - IOC SS PERFORMANCE SUWARY 

o FLEX BODY FLONDAKNTAL FREQUENCY: 0.123 Hz 

o NUMBER OF SKYLAB CMG'S REQUIRED FOR CONTROL: 

o WITHOUT ORBITER ATl&ID: 3 

o WITH ORBITER ATTACHED: 4 

o IOC POWER AVAILABLE 

o TO PAYLOADS: 50 kW 

o HOUSEKEEPING: 25 kW 

o IOC HEAT REJECTION AVAILABLE 

o FOR PAYLOADS: 50 kW 

o HOUSEKEEPING: 50 kY 

o POINTING 

G ACCURACY: + lo 

o STABILITY RATE: - + .02 DEG/SEC 

o THRUST LEVEL: 100-300 LB 

o REBOOST FREQUENCY: 90 DAYS 

o COMMUNICATIONS 

o DOWNLINK: 300 HBPS (LESS ENCODING) 

o UPLIW: 25 MBPS (LESS ENCODING) 

o DATA STORAGE: HOUSEKEEPING ONLY 

o CREW TIME 

o FOR PAYLOADS: 5 PEOPLE 

o FOR HOUSEKEEPIMG: 1 PERSON 
The thrust level during reboost varies from 300 pounds with full tanks to 

100 pounds later, since a blow-down system is uti 1 ized. Reboost occurs 

every 90 days. The comnunications capabil ity of the SS is that of the 

TDRSS link. Nominal crew time expected to be available for payloads is 

four equivalent people out of a six-person crew, with the fifth and sixth 

person usez: for SS housekeeping and miscellaneous activities. 

3.3 MODULE ARRAIJGEMENT 

Fi~ure 3.3-1 depicts the module arrangement used for the reference 

configuration. Thi s arrangement provides a "racetracku csnfiguration, 

i.e., each module (except the Logistics Module) has two exits. Although 

not a requirement, this probably enhances safety in the event of an 

emergency evacuation of any module. There is a high degree of modu'le 

comnonal ity, particularly among the four modules in the racetrack. This 

results in the fewest number of module types being required. This 

arrangement also pro-tides a minimum total number of elements and a 

minimum number cf in ierfaces between elements. Traffic through the 

Laboratory Modules is low, with the majority of traffic being in the two 

Habitation Modules. Traffic considerations and interface/integration 

considerations seem to make it preferable to have the Logistics Module 

and Orbiter berthed to the Habitation Modules, and to have the 

pressurized pay1 oad modules berthed to the Laboratory Modules. 

The above-described arrangement was selected after a fairly brief 

assessment early in the study to provide a reference module arrangement 

which could be utilized on all Space Station concepts being assessed. 

Some of t11e other module arrangements which were investigated during this 

assessment are shown in Figure 3.3-2. The first option utilizes "corner 

cubes" to join modules together, which increases the total number of 

elements, the number of element types, and the number of interfaces. 

Total module volume in this option is expected to be inadequate unless 

the modules are lengthened, and cost is expected to be higher. The 

second option has two d i f f e r e n t types of modules, which increases costs. 

This option also requires traffic to pass through the shorter 

(Laboratory) modules to get from one longer module (Habitation) to the 

other (Habitation 1. ihi s provides a very undesirable traffic flow. The 

third option shown in Figure 3.3-2 is the one selected as the reference, 

for the reasons a1 ready discussed. 

Figure 3.3-3 shows several other inodule pattern concepts arranged on a 

large truss, such as the reference Station keel. Some of these utilize 

interconnecting tunnels or other elements, which may provide some 

advantages but probably would increase costs. None of these has been 

assessed in this study, but these and other potential arrangements ;hould 

be studied further in order to arrive at the best nodule arrangement 

possible within the programtic and technical constrair,';~. 
LAB 

HAB 

KEY FEATURES 

RACETRACK 

HIGH MODULE COMMONAL

MINIMUM TOTAL NUMBER O

MINIMUM NUMBER OF INTE

TRAFFIC THROUGH LABS M

USABLE VOLUk,E IN MODUL

IMPACTED BY PORTS 

Figure 3.3-1 .- REFERENCE MODULE PATTERN 
4 SHORT MODULES (20 FT) 

4 INTERCONNECT MODULES1 

AIRLOCKS 

II. 

L 2 SHORT MODULES (20 FT) I 

01 2 LONG MODULES (40 FT) 

INTERNAL. Al RLOCKS I 

I 

4 MEDIUM MODULES (30 FT) 

INTERNAL AIRLOCKS 

CONFIGURATION 

COMPATl BLE 

X 

- 

X 

:RITERIA -. 

'RACETRACK INTERRUPTED BY AIRLOCK OPE

"STRONGEST DRIVERS 

RACETRACK 

Figure 3.3-2 - MODULEIMODULE PATTERN SELECTION 

"MIN 

MODULE 

DESIGNS 
(INCORPORATES 1 AIRLOCK & 

MOBILE MANIPULATOR FITTINGS) BERTHING M

OPTION 1 ORElTER INTERFACE 

RBITER INTERF

ORBITER INTERFACE 

Figure 3.3-3 - TYPICAL MODULE PATTERN OPTIONS (CONT) 
The number of berthing port util ired per mdule and the function of each 

port are defined in Table 3.3-1. Two radial and two axial ports per 

mdule would satisfy mst existing requirements and provide for slight 

growth. The mdule internal layouts provided later in the report allow 

for four radial ports per module, one of which is blocked off to increase 

usa5le internal volume. This allows for mre growth, but provides a 

slightly mre pessiai stic picture of available internal volume, since a 

good amount of volume has been all owed for each radial port. 
TO MAXIMIZE INTERNAL VOLUME AVAILABLE, AND TO MAXIMIZE MODULE COMMONALIT

AND 2 RADIAL PORTS PER MODULE 

BACKUP PORT FOR ORBITER AND/OR LOG MODULE 

Table 3.3-1 - NO. OF PORTS REQUIRED 
4.0 SPACE STATION MANNED CORE 

4.1 DESIGN REQUIREMENTS 

The primary design requirement for the Space Station is to provide a 

versatile, gr~wing, permanent, manned facility in space with the 

opabil i ty to enable the efficient acccwnpl ish.crerrt of sf grir'ficant bdvdnces 

.7 space science, techno1 ow, l;olrmerce, and transportation. 

Sgeci f ical ly , the Space Station is required to imp1 ement the fol 1 owing 

objectives. 

a. Establish the means for a permanent and productire presence of 

people in space. 

b. Establish routine, continuods. and efficient utilization of 

space for science, appl ications, techn2logy Aevel opment, and operations 

( incl uding servicing and refuel iy cr space assets). 

c. Develop further the caiiercial util itation of space. 

d. Develop and exploit the synergism effects of the manlmachine 

canbination in space. 

e. Provide essential system elements and operational practices for 

an integrated national space capabil ity. 

f. Stimulate the mutual benefits traditionally derived from 

cooperation in space with our allies and friends. 

g. Reduce the cost and complexity of living in and using space. 

h. Ensure leadership in space for the United States in the decade 

of the 1990's and beyond. 

Design rdqui rements have been derived from these objectives, both in 

terms of an Initial Operational Capabil ity (IOC) and a growth 

capabi 1 i ty. Prime considsration has been given to customer requi rernents 

and the operations requirements, and system requirements have been 

derived from these. At the basic level, the mission, or customer 

requirements, have been interpreted as necessitating three separate 

spacecraft: a permanently-manned Spdce Station in a 28-112' f ncl ination, 

270 nmi cietcular orbit; an unmanned co-orbiting platform, which is in a 

rendezvous-compatible orbit with the Spdce Station and a second platform 

in 3 98O, 430 mi, Sun synchronous (2 p.m.) orbit. The polar orbiting 

platform configuration characterized in this study contains primarily 

Earth obs~r-lt rdy ?ayloads, but is intended to be sufficiently versatile to 

support other payloads ar 5e operable in other near-polar orbits. The 

plat'orm which ir i7 the rendezvous-canpatible orbit with the Space . Stacion is rharacterized in this study as acconmodating payloads for 

celestfal and solar observations and demanding microgravity experiments 

with a mlnimum of mechanical disturbance and contamination. The platform 

is versatile enough to accommodate other types of payloads, if desired. 
The manned Space Station eleme~t of the program must be designed to meet 

a1 1 the remaining program objectives, as we1 1 as providing for the 

development of sensors and systems for Earth observations and a base for 

some celestial, solar, and Earth observations which require a manned 

presence. To meet these objectives, the configuration design must 

?rr-~i de for (1) mounting, pointing, servicing instruments which 

continuously view the Sun, the entire Earth's dis~, and the celestial 

sphere, and provide this capability simultaneausly except when prohibited 

by occul t~tion; (2) ci gni ficant continuous power for commercial an3 

scienti fic functions; (3) sign1 iicant pressurized volume for 1 aboratories 

dnd tnt c, ew required to service all users; (4) the capabil iij to act as 

a "Space Port" for the Shuttle, various free-flying satellites, the 

~o-r-4i ting platform, and space transportdtion vehicles such as the OTV 

and JMV; (5) the abil ity to store, maintain, assemble, and reconfigure 

vehicles and payloads of all types; (6) the abil ity to service and refuel 

free-flying spacecraft, platforms, and attached pqyloads; and (7) 

attachment and services to various unpressurized scientific and 

commercial pay1 oads. Specifically , at IOC, the manned Space Station 

element is require3 to provide at least 5OkW of continuous power a 

minimum of 3 fol:; crew persons to service customers, at least 45 m of 

laboratcry volume, an OMV, ports for attachment of two pressurized 

cust2mer payloads, attachments and service interfaces for at least seven 

un~ressurized instrument payloads, and provisions for large space 

coistruction and OTV development. 

The work set forth in this report is directed tov3rd a manned Space 

Station, a~d implicitly assumes that early manning is desirable. As a 

result, the possibil ity of a man-tended Station has not been explored in 

this study. "Man-tended" is taken to signify a mode in which the Station 

is visited at intervals by the Orbiter and operates unmanned the rest of 

the time. This mode offers some potential cost deferral if the 

pressurized modules are deleted from the IOC Station a1 together. This 

deletion would require the crew to 1 ive entirely out sf the Orbiter 

during their stay on orbit. If pressurized modules are included in the 

man-tended Station to enhance on-board operations and maintenance, there 

is little or no cost saving associated with it. 

Inherent in the program objectives is that the manned Space Station have 

the capability to grow in its ability to provide basic services to all 

types of customrs, -nd to accommodate additional users of foreseen and 

unforeseen classes. For design purposes, thi s requirement has been 

defined as approximately B~ce the poww level, crew size, and number of 

pressurized payloads of that expected at IOC; increases in the nuder of 

unpressurized attached or serviced payloads have not been specifically 

defined, but the major areas of growth in the types of attached pqyloads 

are expected to be in the size and complexity of sate1 li tes to be 

serviced, in the size of construction and assembly projects, and in the 

addition of increasingly capable OKV-and GTV-type vehicles. 

The requirement that the Space Station become a permanent facility 

imp1 ies that, in add: tion to growth capabil i ty , it must a1 so possess the 
QWWJAL PACE'SOF POOR QUALITY 

characteristic of being able to incorporate advances in technology as 

they occur. This compati bi 1 i ty wi th "modernization", or technol ogy 

transparency, is then considered a significant design driver. A1 though 

this requirement basically appl ies to all systems and subsystems of the 

manned core Space Station, its influence on configuration is reflected 

primarily in the power system area, and its influence on the module 

design primarily in the ECLSS, since it is anticipated that these areas 

wlll he particularly subject to post- IOC devel opment, and because 

provision must be made in the basic design for significant changes in the 

systems. 

To facil i tz5e the definition of speci fic configurations for the manned 

Space Station and platforms, a set of payloads for !OC has bettn 

identified to serve as a "pathfinder" in design. This set is shown in 

Table 4.1-1. This set is nut taken as an actlial expectcd manifest at 

IOC, so nruch as it is intended to erlcompass the types of near-term 

ayloads expected to appear on the SSPE's. It will be noted that these 

represent specific cases for a1 1 the generic user 

requirements. Tho Star1 ab (SAA0006) i s representative of a 

celestial -viewing instrument package, the pinhole facil i ty (SAA0009) of a 

solar observation instrument, and the plasma payload (SM0207) of a 

simultaneous viewing requirement, while the za~th observation technology 

experiment (TDM2260) demands a broad, continuous view of Earth. 

Microgravi ty requirements are represented by the ECG production uni t 

(COM1203), the material processi ng 1 abol-atory (COM1201) , and the EOS 

production unit (COM1202). The group of users for which the Space 

Station provides a base for maintenance, cmstruction, repair, and 

servicing is represented by the Depl oymentlAssmblylConstruction 

Techno1 ogy project (TUM2060) and the servicing technol ogy projects (TDM's 

2560 and 2570). 

In addition to these technol oav experiments, a definition of sate! 1 lte 

servicing requirements has been developed for use in preliminary 

definition of a reference configuration for the manned Space Station. 

This definition includes storage and servicing capabilities and is 

intended to provide for support of the specific satellites shown in Table 

4.1-2. These sate1 1 i tes are in tended to represent the range and types of 

syswms and vehicles to be accommodated at IOC; they are not considered 

as either exclusive or inclusive of the actual satellites wt-.ich wil; be 

serviced during the iOC period. 

Closely associated with the customer requirements are operational 

requirements, some of which also affect the overa'l configuration of the 

Space Station and the unmanned platforms. The primary requirement is 

that the SSPE's must all be launched, and supported by the Space 

Shuttle. This requirement has a significant effect on the assembly 

sequence and the modul arization of the structure and subsystems. Support 

by the Shuttle imp1 ies that the design of the SSYE's must provide 

convenient means for rendezvous, berth1 ng, and foadi ngiunl oading of the 

Orbiter. In addition to the Orbjter, the Space Station must also be 

configured to perform many of the same operations with free-flying 
satellites, the OMV, and the OTV (in the growth phase). The 

co~munications requirements are bui 1 t around the customer data 

requirements for near-continuous data transmission through TDRS and for 

comnunication and tracking of the Orbiter and free-flyi ng unmanned 

satel 1 i tes, vehicles, and EVA crewmen. 

The majority of the system requirements have more effect on the detail of 

subsystem design than on the overall configuration concept. However, the 

requirement for safe operation with the loss of one pressurized module 

affects the arrangement of the modules. The general requirements for 

commonality affect the modul arization of the structure and silbsystems, 

and the requirements fora maintainability affect the location of 

components and EVA and manipulator provisions. 
TABLE 4.1-1 - SPACE STATION MANNED CORE IN PAYLOADS 

SAAOOO5 Transition Radiation & Ion Calorimeter (TRIC) 

SAAO006 STARLAB 

SAAOO09 Pinhole Occt~l ter Faci! ity 

SAAO201 LIDAR Faci 1 i ty 

SAA3207 Space Piasrna Pay1 oad (SSP) 

COM1203 ECG Production Unit 

Materials Performance 

Depl oym~nt/AssemblylConstructi on Technol ogy 

Structurdl Dynamics Techco: ogy 

Earth Observation Instrument Technology 

Fluid Mdnag~ent Technology 

Atti tude Control Techno1 ogy 

Figure Control Technol ogy 

Environrnen1:al Effects 

Satel 1 i tc Zervici ng Technol ogy 

OTV Servicing Technology 

PRESSURIZED PAYLOADS 

SAAO307 Life Science Laboiiztory 

SA40401 Microgravi ty Kesedrcll Ldboreatory 

COM1201 MPS Lab #1 (Materials Processing) 

COM1202 E3S Production hit 

TDM2020 Materials Processing Technology 

TOM2520 Hiibi tat ion Techno1 ogy 

TDM2530 Medical Techno1 ogy 
CODE 

TABLE 4.1-2 REPRESENTATIVE IOC SATELLITES REQUIRING SERVICING 

SAAOO 12 

SAAOO! 3 

SA000 14 

SAAOO 16 

SAA0019 

SAAOQ 1 7 

MISSIONS 

Space Telescope (ST) 

Gam Ray Observatory (GRO) 

X-Ray Timing (XTE) 

Solar Maximum Mission (SPA) 

Far llltraviolet Spectroscopy Explorer (FUSE) 

Advanced Y-Ray Astrophys!cr Fac!!lty !!AXAF! 

Leasecraft 1 and Pavloads 

Leasecraft 2 and Payloads 

Leasecraft 3 and Payloads 

Leasecraft 4 and Payloads 

Leasecraft 5 and Payloads 
Taken as a whole, the design requirements dictate that the manned core 

station be an assembly of significant overall dimensions, with large 

areas for energy col lection and heat rejection, mu1 tiple pressurized 

modules, areas of unpressurized structure, and means for connecting 

elements fop orientation of the various elements and also providing 

access and clearance for viewing and operations. The requirements for 

the polar and co-orbiting platfoms demand small areas for energy 

collection and heat rejection and no pressurized volumes. But a high 

priority is placed on the viewing needs of observational payloads and 

must be compatible with support by the STS and/or the OHV. 

4.2 CONFIGURATIOK DESCRIPTION 

4.2.1 General Arrangement . 

The Space Station reference configuration is a set of deployed linear 

trusses to which pressurized modules, subsystems, and user equipment are 

attached. The principal structural cmponcnts are a keel and three booms 

at right ang:es to the keel. The IN configuration is illustrated in 

Figure 4.2.1-1. The coordinate system is Z para1 lel to the keel, 

positive toward nadir; X perpendicular to keel and boans, positive in the 

direction of fl ight; and Y parallel to the booms, positive to starboard. 

The origin is at the center of the intersection of the keel and 

transverse boom. 

Four of the pressurized common modules (two habitation and two 

laboratory) are arranged in a quadrangle to permit IVA crew mverwnt in 

case any one module becomes unusable. The keel is divided at the bottom 

to allow instal lation of the pressurized modules on the center1 ine of the 

keel. This maintains the principal axis in the orbit plane while the 

Orbi~er is berthed, avoiding excessive yaw and roll excursions. 

The materials processing mcdule (Lab 2) is located at the top of the 

quadrangle to keep it as close as possible ta the aass center of the 

station. The control station is at the front of HM2 for a direct view of 

berthing operations. The HM's are adjacent to each other to eliminate 

the need for excessive crew movement through the lab modules. The 

logistics module (LM) is berthed to HM2 to permit unloading with miniwusl 

disturbance to laboratory operations. 

Primary Orbiter structural interface is at the end of HM1 as shown in 

Figure 4.2.1-1. An alternate position at the end of Lab 1 is provided 

for emergency access in case the primzry port is inoperable. Both 

positions allow payload removal without interference. 

External airlocks are berthed to the port side hatches of )#I and W. 

This location leaves the entire forward face of the starboard keel 

extension open for manipulator travel . 

The conf iguratfon in Figure 4.2.1-1 employs a photovol taic power 

generation system. A solar dynamic system has also been considered as an 
option tor IOC. Such a system is illustrated in Figure 4.2.1-2. The 

arrangement is identical to the photovoltaic version except for the power 

system. Discussions in the following sections, with that exception, 

should be understood to apply to bbch IOC configurations. 

A sample growth configuration is shown in Figure 4.2.1-3. The principal 

differences are a solar dynamic power system providing 300 kW average 

power, two additional habitation modules and four additional 1 ab 

modules. The solar dynamic power system is shown to demonstrate a design 

option and is not intended to advocate a particular power system growth 

path. The lower keel is also expanded from one bay wide to three bays 

for improved structural redundancy and stiffness. Additional lower booms 

are placed at the ends of the keel extensions to preserve the field of 

view of the earth s~nsors. Propulsion system relocation to these booms 

may be necessary to maintain the Station's mass center bet wee.^ the 

thrusters. 

4.2.1.1 Subsystem Installation 

The electrical power system is instal led on the transverse boom (see Fig. 

4.2.1-1 ) outboard of the a1 pha gimbal joints, i ncl udi ng power generation, 

storage, and condi tioning . Thi s permits power transfer across the a1 pha 

joint to be 1 imi ted to ful ly-condi tioned AC power. Each array wing is 

gimbaled individually for beta adjustment to minimize principal axis 

shifts and to simp1 ify assembly and deployment. 

Primary heat rejection is provided by radiators on booms mounted on the 

lower keel. The radiators are rotated to maintain an edge toward the 

Sun. They are rewound during the dark portion of the orbit to avoid a 

continuously rotating fluid joint. 

TDRSS and GPS antennas are mounted at the ends of the upper boom for 

maximum upward view. Tracking and rendezvous antennas are mounted on the 

lower keel under the transverse barn for a clear view forward and aft 

along the flight path. 

The Attitude Control Assembly (ACA) is a rigid +foot cube at the 

intersectfon of the keel and the transverse boom. It c~ntains the 

control icifietrt gyros, star trackers, and other guidance, navigation and 

control components needed fc:- ccntrol of the Station bran the initial 

1 aunch . 

Identical sets of propulsion thrusters are mcunted at four lxations at 

the ends of the 1 ower boom and on the 1 ower keel, capable of f i ri ,,g 

forward, aft, and outboard. The lower thrusters (on the lower boom) 

normally provide a1 1 backup attitude control. The upper thrusters 

normally fi re only aft, for orbi t maintenance. Orbit maintenance 

maneuvers are performed at approximately 5 11.m. and 5 p.m. (orbit time) 

to minimize plume impingement from the upper thrusters on the solar 

arrays. 
A Mobile Remote Manipulator System (MRMS) is capable of mvi ng on the 

nodes of the truss to any location on the forward faces of the keel and 

booms. By moving outboard of the alpha joint and rotating the transverse 

boom, the MRMS can a1 so mve along the aft face of the structure. 

4.2.1.2 Payload Accumnodations 

Two primary locations are provided for observational users (see Fig. 

4.2.1-4). Solar and stellar sensors are mounted along the upper born to 

maximize their view of that part of the sky not OccLr tzd by the Earth and 

to minimize contamination fran activities at the inhabited end of the 

Station. Some mutual blockage of one stellar or solar instrument by 

another will occur in the region of the orbit poles. Any such lost area 

is re1 atively small and can be made up within a week or two as the 

Station orbi t regresses in 1 ongi tude. Horizon-to-horizon viewing of 

objects away from the orbit poles is poss'ble for all instruments wi+.hout 

blockage. Solar viewing is possible on every orbit because the Slrn is 

always within 52' of the orbit plane. Earth sensors are mounted on the 

lower boom; local vertical orientation allows continuous viewing 

limb-to-limb with no image rotation. 

Some solar sensors mw also be mounted on the transverse boom outboard of 

the alpha gimbal to simplify pointing if their mass characteristics do 

not adversely affect the Station's dynamic properties. 

Sate1 T iee servicing provisions are located along the keel. Two 

storage/servici ng bays are situated para1 1 el to the upper keel. TIti s 

1 ocation minimfzes contamination of optical surfaces and other sensitive 

components. Related ORU and tool storage is provided nearby on thtransverse boom. A refuel ing bay is 1 ocated on the 1 ower keel. 

Satellite propellants and cryogens are stored at the top of the keel 

extension near the ref~c! ing bay. 

The OMV and OMV kits are berthed alongside the port and starboard keel 

extensions. 

Large structure construction is carried out obi the aft face of the lower 

keel. Ample space for storage of materials and equipment f s available on 

the keel extefisions . 

Seuera'l berthing ports dre avai 1 able on the 1 aboratory and habi tat,ion 

modules for attachment dedicateti r~scr modules requi~ing IVA access. 

4.2.2 Flight Modes 

In normal operation, the Station is oriented with the keel approximately 

a1 igned with local vertical and the Y pr!ncipal axis (approxima.tely 

parallel to the transverse born) held perpendicular to the orbit place 

(POP) as showrt in Figure 4.2.2-1. A constant pitch attitude Is 

maintained ta balance gravity gradient and average aerodytramic torques. 

U4G's are used for momentum management to compensate for aerodynamic 

torque variations around the orbit. Long-t.em (more than a few orbits) 

OMWAL PP&' 1s 

OF POOH QULL~W 
SIDE VlEW 

UPPER 

37.5 KW DYNAMIC UPPER KEEL 

POWER SYSTEM 

MOBILE - 

MANIPULATOR 

LOWER KEf l 

RADIA

KEEL EXTENSl 

LOWER BCOM 

1- PA- .... 

FRONT VlEW 

Fl~ura 4.2.1-3 - GROWTH CONFIGURATION 
SAA 020iry I 

SIDE VIEW 

UPPER KEFL 

TCM 2510/ wi3 5'-,-54'- I- -- - 

. . - . . . . . -. . . .- -67 5 - 22 5 !c 36 - R

'NOTE SOLAR DYNAMICS SYSTEM A PRIME OPTION INCT SHOWN, 

Figure 4.2.1-4 - USER ACCOMMODATIONS 

variations in .ass properties or atmospheric density are accommdated by 

small adjustments in the flight attitude of the Station. 

The sular array is maintained normal to the Sun vector by a continuous 

rotation abwt the alpha gimbal at orbital rate and a slow adjsstment 

about the beta gimbal to track the Sun's motion relative to the orbit 

plane. The radiators are also rotated at orbital rate, but are "rewound" 

during the dark part of the orbit to avoid ccrn~inuously rotding fluid 

joints. 

During the early stages of buildup, a "streamlined" orientation is flown 

(see Fig. 4.2.2-2). The transverse boom is in the orbit plane and the 

gimbals are locked with the array masts POP and the array wings at local 

horizotttal . Thi s atti tude minimizes aerodynamic drag and enhances 

gravity gradient stabil i ty. The power system produces nearly ha1 f of 

full-rated power in this orientation, which is adequate until the Station 

i s permanently manned. 

The streamlined orientation can also be used in the unl ikely event of 

total control system loss. Should this occur, gravity gradient torque 

will force the transverse boom into the orbit plane. The power system 

can maintain housekeeping requirements under these conditions, but a 

reduction in user operations would be necessary until the control problem 

can be corrected. 

Elements 

The principal elements of the reference Space Station are: Habitation 

Modules, Laboratory Modules, Logistics Modules, and Utii i ties Elements. 

Figure 4.2.3-1 provides a 1 ist of the pressurized module functions by 

element, and each is briefly discussed below. As mentioned previously, 

the modules hatre a high degree of commonality, both at the overall 

configuration level and at the subsystems level. The basic pressure 

shell and berthing port arrangement is iuentical in all Habitation and 

Laboratory Modules. The Logistics Module is the same as the others, 

except that the radial port segment has been replaced by an 

externally-mounted tankage segment, and the pressurized portton of the 

Logistics Module is +.herefore only two-thirds as long as that of the 

other modules. All the Habitation and Laboratory Modules have the saw 

floor and ceiling arrangements except for the portion of the Habitation 

Module which contains the sleep quarters. Equipment racks have a high 

degree of commonality across the modules, as do the utilities. 

Each Habitation and Laboratory Module has a 22-day supply of food and 

medical supplies for two crewmen. In the reference Space Station 

concept, most of the ECLSS equipment is located in the two Habitation 

Modules. A more di stributed ECLSS system may have a greater weight and 

cost, but this may be offset by cost savings due to module commonality. 

. The EVA airlocks are external modules berthed to the Habitation Modules. 
OHJGWAL Fiik,~' is 

OF POOR QUALITY 
LAB $1 -.- 

PRESSURIZED MODULE FUNCTION ALLOCATION 

HAB #l -- O COWON FUNCTIONS* 0 

0 

0 

0 

0 

LAB $2 

O COMMON FUNCTIONS* 

COMMCN/FllNCTIONS* 

SHOWER/URI NAL 

UASTE WAAGEMENT 

PRiMARY ECLSS 

HEALTH MAINTENANCE 

WARDROOM 

GALLEY 

VEGETABLE CHAMBER 

COMMAND STATION 

HAB $2 

COMMON FUNCTIONS 

SHOWER/URIKAL 

WASTE MANAGEMENT 

PRIMARY ECLSS 

SLEEP QUARTERS 

WASHER/DRYER 

MAINTENANCE/RiPAIR 

STATION 

COWND STATION 

LOGISTICS 

O COMON FUNCTIO

EXCEPT: 

O REDUCED UT

NO HANDWAS

O EXPENDABLES 

SPARES 

DIRTY dATER ST

O TRASH STOWAGE 

"COFMON FUNCTIONS 

v 

O PRIMARY 8 SECO

STRUCTURE 

O UTILITIES 

O STORAGE 

O SAFE HAVEN SU

* PARTIAL CONTRO

O HANDWASHER 

FIGURE 4.2.3-1 
Each of the Habitation and Laboratory Modules incorporates the safe haven 

features mntiened previously . The Orbiter is structural ly compatible 

with any of the berthing ports on the modules for rescue, although the 

subsystems interfaces mqy be different than those between mdul es. 

One of the overall guide1 ines used in the internal layout activity was to 

try to keep the Laboratory Modules as free as possible of habitability 

and housekeeping functions. Control stations, util i ties, storage, and 

handwashers are included in the iaboratary Modu:es, and more equipment 

Ray have to be shifted from the Habitation Modules to the Laboratory 

Modules due to volume limitations in those modules. 

Another key guideline was to provide minimum disturbance in the sleep 

quarters area (ideally to have a "quiet" module and a living area 

module). This was not completely achieved due to other constraints. 

However, the wardroom and health maintenance facil i ty (kMF), which are 

noisy ai-ea;, arc l scated in a separate module from the sleep quarters. 

Packaging schemes which allow more dense packaging have been assessed 

hriefly and offer some improvement dn volume constraints. There is some 

equipment within each module to which access is not required frequently, 

and this equipment would lend itself well to more dense packaging. 

Access to walls is needed in all concepts for detection, isolation, and 

repair of pressure shell leaks. This requires that the equipment be 

movable or that some other equivalent approach must be taken in the 

design. 

A brief assesswnt was made of the on-orbit spares requirements. It is 

desired that these be kept to a minimum because of the premium placed on 

internal volume. It appears that the quantity of spares can be kept low, 

because of the modulari tylredundancy design concepts used for critical 

subsystems. 

The Logistics Module can be more densely packaged than the other modules, 

since it serves the "closet" function. Care must be taken not to exceed 

the STS down-wei ght limi t (32,000 1 b) . Thi s 1 imi t i s expected to be 

increased by the timeframe of the Space Station. The nonfluid experndable 

resupply items and spares are packaged inside the Logistics Module, and 

the fluids are stored in external ly-mounted tanks. 

As presented in the reference configuration, the habitation volume is 

comprised of two habitation modules which contain all those facil i ties 

and equipment required for crew sustenance and Station operation. The 

following areas and categories of equipment are contdined in the 

habitation vo'l urn: crew quarters, gal 1 eylwardroom, health maintenance 

faci 1 i tylexerci se area, personal hygiene area, workstations, stowage 

areas, waste management faci 1 i ties, and 1 aundry facii i ty. These areas 

contain equipment, the facilities which "hause" this equipment, and 

'access volume which facilitates its use and maintenance. The internal 

configuration of these modules is shown in Figure 4.2.3-2 (Hab 1) and 

Figure 4.2.3-3 (Hab 2). 
-POIIER EQUI PMEHT I.IARDR 

? 

J" 

'J! 

I' 

i, 

SIiOIIER/URII.4AL 

IiArU.? IIASIIER 

i 

WASTE - EQUI PE

MArl AGElclEr IT 

FACILITY 

Figure 4.2.3-2 - HA6 1 MODULE 

..- .. . 
-CAMERA 

\, -COMt4UFII CATIONS EQUIPIfENT RADI A1 

',, PONE R SL.EEP 

ECLSS OXYGEIl MASKS - r EQUIPMNT \ 

\ / \ 

DATA -l 

MANAGEMENT 

/ EQUIPMENT 

L---- MAINTENANCE/ 

REPAIR STATION 

Figure 4.2.3-3 - HA6 2 MODULE 
Crew accommdations in nonhabitation areas include the laboratory 

modules, the airlocks, the logistics module and safe haven provisions. 

Crew accosmodations in the laboratory volumes aid in crew health, safety 

and well-being and they support the crew in customer services. As 

presented in the reference conf iguratfon, the internal layout of these 

modules is shown in Figure 4.2.3-4 (Life Sciences Lab) and Figure 4.2.3-5 

(Materials Processing Lab). 
7HUMAN HESEARCtl RACKS 

POIdER EQUIPKNT 

,--SAFE HAVEN 

1,2.4 Mass Propfirties 

The weights and volumes for the ten subsystems were assembled for the 

reference configuration. 

ihe ten subsystevs were: 

a. electrical power 

b. guidance, navigation and control 

c. comnunications and tracking 

d. data hand1 ing 

e. propulsion 

f. enbironmental control and 1 i fe support systems 

g. active thermal control system 

h. structures and mechani sws 

1. crew systems 

j. payloads 

There w2re five iterations of welghts for the refew:~ce configurations. 

All five iterations showed a steady increase of wefghts. The increases 

were due mainly to components that were overlooked in a previous 

iteration. 

Because of the early decision to go with one module arrangement, the 

volume was only assembled the one module arrangement. There were 

three iterations of volume for the common rcce-track-module arrangement 

of the Space Station. 

The estfmated weights and volume of the Space Station are sumrized ;? 

the following tables. Although these estimates do not include an 

a! lowance for growth, growth can be expected to occur. 

The subsystem weights for each element of the Space Systems are presented 

in Table 4.2.4-1. These are the total wet and dry weights in pounds. 

These weip' - ~fter each iteration were djstrlbuted to all subsystems for 

their use, . well as for updating and correcting errors. The subsystem 

data were u~~lised forthe cost estimation of the Space Station. The 

weight data for the cost estimates were double checked fcr accuracy. 

Table 4.2.4-2 shows the Space Station volume in cubic feet for applicable 

subsystems thdt were in the pressurized wdules, as well as components 

that were attached tc the Space Station trdss structures. 

4 3 
Table 4.2.4-3 presents the total weight for the Space Station by elements 

and subsystems. 

Table 4.2.4-4 presents the total volume for the Space Station by elements 

and subsystems. 

Table 4.2.4-5 prescllts the 90-day resupply by subsystem for the Space 

Station. 

Table 4.2.4-6 presents the 10-year 1 i fetir,re replacement cycle for the 

Space Station 

Tables 4.2.4-7 through 4.2.4-15 list by modtile all the components of a 

slrbsystem that are cormion across the pressurized modulec. The units arc 

either number of components, number of men, or nunber of comparison 

quantity. Example: one unit might be 200 fzet of wire. 
TABLE 4.2.4-1 - SPACE STATION ELEMENT UEIGHTS - POUNDS 

TRUSS ATTCH 

SUBSYSTEM HM1 

- HM2 - LM 

- LAB 1 - LA82 -- STRUCT COMPTS P

ELECTRICAL 643 462 208 50 5 505 7,867 12,554 

POWER 

COMMUNICA- 834 2,105; 65 564 564 658 2,304 

TIONS AND 

TRACK I JG 

DATA 1,005 765 120 805 605 - 1,250 

HANDLING 

PROPULSION - - 2, '768 - - 6,951 

P 

ECLSS 6,312 6,312 929 929 069 - - 

ATC S 474 474 21 1 1,215 640 2,408 5,396 

STRUCTYRES 12,964 12,964 7,310 12,964 12,964 15,316 3,224 

AND MECH. 

CREWSYS. 15,710 11,080 9,125 3,883 3,883 - - 

PAY LOADS - - 13,148 18,630 32,275 - - 184

TOTAL S 37,942 34,163 33,884 39,495 55,305 29,983 31,679 184

HMl = HAB, MOD, 1 

HH2 = HAB, MOD, 2 

LM = LOGISTICS MOD. 

LAB1 = LAB, MOD, 1 

LAB2 LA8, MOO, 2 
SUBSYSTEM 

ELECTRICAL PONE R 

3 

TABLE 4.2.4-2 - SPACE STATION ELEENT ;IOLU)IT - FT 

CMUNICATIONS 

& TRACKING 

DATA HANDLING 

PROPULSIOW 

ECLS 

CREW SYSTtcfS 

PAYLOADS 

TOTALS 

- LAB 1 LAB 2 
3 

TABLE 4.2.4-2 - SPACE STATION ELEKNT YOLUWE - FT (CONTINUED) 

SUBSYSTEM 

TRUSS ATTC H 

P 

STRUC COnPNTS PAY LOADS 

ELECTRICAL POYER 392.0 - - 

6NK 161 .O - - 

COWUN ICATIONS 9.9 1055.3 2.5 

DATA HANDLING - 7.6 - 

PROPULSION - 130.0 - 

ECLSS - - - 

ATCS 310.3 427.1 44.6 

CREW SYSTEMS - - - 

PAVLOADS 

TOTALS 

(A) 8,898 

(8) 124 X 106 

783.2 1620 .O (A) 8945.1 

(0) 124 X 106 
ELEMNT 

HM1 

TABLE 4.2.4-3 - SPACE STATION WEIGHT - POUNDS 

WE I GHT -- SUBSYSTEM 

37,942 ELECTRICAL POWER 

LN 33,884 CIT 

LAB 1 ~9,495 DATA HANDLING 

LAB 2 55 305 PROPULSION 

TRUSS STRUCTURE 29,593 ECLSS 

ATTACHEC 

COMPONENTS 

PAYLOADS 

TOTALS 

ATCS 

STRUCTURES I 

MECHAN ICS 

C4EW SYSTEMS 

PAY LOADS 

TOTALS 
ELEMENT 

HW1 

HH2 

LM 

LAB 1 

LAB 2 

TRUSS STRUC 

ATTACHED COMP. 

PAYLMDS 

3 

TABLE 4.2.4-4 - SPACE STATIOt'l TOTAL VOLUME - FT 

YOLUME SUBSYSTEM VOLUME 

3,108.3 ELECTRIC& POWER 344.0 

3,704.4 DATA HANDLING 188.1 

3,599.1 PROPULSION 187 .O 

783.2 ECLSS 772.9 

1,620.0 ATCS 1,133.3 

(A) 8,945. CREW SYSTEMS 10,835.0 

(0) 124x10 

PAYLOADS (A) 18,246 

(B) 124x106 
--- SUBSYSTEM 

ELECTRICAL POWER 

GNK 

CMUNICATIONS 

TRACK1 NG 

TABLE 4.2.4-5 - SPACE STATION 90-DAY RESUPPLY 

DATA HANDLIMG 

ECLSS 

ATCS 

CREW SYSTEMS 

- WEIGHT (LBS) 

0 

0 

0 

VOLUME (FT~) 

0 

0 

0 

PAYLOADS 13,148 3,748 

TOTALS 23,143 4,323 
TABLE 4.2.4-6 - SPACE STATION TEN YEAR LIFETIW SPARE REPLACEMENT CYCLE 

SUBSYSTEM WE!GHT (LSS) VOLUME CFT~) - CYCLE (YRSI 

ELECTRICAL POWER 15,848 302 5-10 

COMMUNICATIONS 

& TRACKING 

DATA HANDLING 

PROPULSION 

ECLSS 

ATC S 

MECHAN 1 SMS 

CREW SYSTEMS 

PAYLOADS 

TOTALS 
TABLE 4.2.4-7 - SPACE STATION SUBSYSTEM CWN EQUIPMENT BY MODULE 

ELECTRIC.4 - POWER 

BW 

MBCU 

BDC 

UPC 

0 GA. WIRE 

8 Cd. WIRE 

12 GA. WIRE 

CIF 

SSS 

CYCLO (60 X 1) 

CYCLO (50 X 1) 

CYCLO (60 X 3) 

CYCLO (SO x 3) 

CYCLO (40 X 3) 

APMS 

OPS 

- HM! - !iM2 - LAB 1 - LAB2 - LM TRUSS STRUC 

24 

4 4 4 4 2 8 

24 

60 48 50 50 24 80 

1 

3/10 3/10 3/10 3/10 1/10 1 

1'3 2/ 10 1/3 1/3 1/12 1 

28 12 8 8 2 

4 4 4 4 2 

2 2 2 

2 2 2 

1 1 

1 1 

2 2 

1 1 1 1 

2 
TABLE 4.2.4-8 - SPACE STATION SUBSYSTEM COMMON EQUIPMENT BY MODULE 

CMG ASSEMBLY 

MAGNETIC BAR 

STAR TRACKER SENSOR 

HEXAD STRAPDOWN SENSOR 

MAGNETIC TORQUERS 

G&C PROCESSOR 

NAVITRAFFIC PROCESSOR 

RCS CONT. SYS. 

INTERFACE DEVICES 

TRUSS 

- HM2 - LAB1 - LAB2 - LM - STRUC 
TABLE 4.2c4-9 - SPACE STATIOCi SUBSYSTEM COMN EQUIPENT BY MODULE 

COmUN ICAT IONS i3 

TRACKING 

AUDIO TERMINAL 

SPEAKER MAKE 

MODULE CONTP,OLLER 

EMERGENCY COW. UNIT 

WIRELESS CW. 

WIRELESS COW. 

1. UNIT 

LASER DOCKING 

FIBER OPTIC TRANSCEIVER 

HDR 4-CHANNEL RCVR. 

FAR RANGE MA LDR 

MOD/DEMOD 

PROX. OPS MA LOR 

MDD/DEKID 

TDRS KU-BAND MOD/ 

DEMOD 

PROX. OPS MA LDR 

MOD/DEMOD 

TDRS KU-BAND MOD/ 

DEMO D 

TM PROCESSOR i3 

CONT 

P/L DATA INTERLEAVER 

FF SIG PROC. (LDR) 

FF SI5 PROC (HDR) 

TDRS SIG PROC. (KU! 

DIGITAL TV PROC. 

REND. RADAR SIG PROC. 

VIDEO LINE DRIVER 

RFI INTERFACE 

EXTERNAL INTERFACE 

UNIT 

CM &TRACK 

PROCESSOR COMPUTER 

DIGITAL STORAGE 
TABLE 4.2.4-10.- SPACE STATION SUBSYSTEM COWN EQUIPMENT BY MODULE 

CWUNICATIONS d 

TRACK I NG - HM1 

GPS LOW GAIN 

S-BAND LOW GAIN 

RADAR XPNDR LOW 6AIW 

MRMS S-BAND LOW GAIN 

S-BAC3 LOW GAIN MRMS 

TV 

S-BAND XPNDR (TDRS) 1 

S-BAND XPNDR (OREITER) 

S-BAND PAIREAMP (TI0 3 

XCVR (NRMS) 

RCVR (MRMS) 

XMTR ( ORBITER) 

LOW RATE MUX 3 

TDRS SIG PROC. 1 

ENCRYPTIDECRYPT 

ORBITER SIG PR3C 

GPS REC/PROC 

RADAR XPNDR 

CABLE 2 

MED RATE MUX 1 

TV CAMERA 2 

TV MONITOR 8 

STEREO DISPLAY 1 

VlOEO RECORDER 1 

LASER PLAYER 1 

PANITILT UNIT 1 

AUDIO INTERFACE 1 

CONTROL STATION 1 

LARGE DI SPLAY 1 

HIGH RATE HUX 

TRUSS 

- HM2 - LAB 1 - LAB2 - LM - STRUC 
TABLE 4.2.4-11 - SPACE STATIOW SUBSYSTEM CObl3DII EQUIMNT BY WDULE 

DATA HANDLING - - HMl - to42 -- LAB1 - LAB2 - LH 

DATA STORAGE 

TIME + FREQ. 

FACIMGT CON 

FIXED WORK STATIGNS 

PORTABLE WORK 

STAT IONS 

CBLNG/OPT. NET 

INTERIOR LIGHT 

SUPPOkT HARDWARE 

IDMS ID'S 

SUBSYSTEM ID' 5 

PAYLOAD ID'S 

!WS SDP'S 
TABLE 4.2 -4,-12 - SPACE STATION SUBSYSTEM CMQH EQUIPMENT RY MODULE 

- ECLS 

ORBITER P/C CONTROL 

FIRE DETECT ION/ 

SUPPRESSION 

DUMP + RELIEF 

FANS, SUCTING, 

MUFFLERS 

C02 REMOVAL SUB. 

CO REDUCTION 

T&E CONT~XI WANT 

02 GENERATION 

HUMID ITY/TEMPERATURE 

CONTROL 

MONITORING-ATMOSPHERIC 

REVITAL IZAT IOH 

PLUMBING, VALVES, FITTINGS 

INSTALL SUPPCiT 

POTABLE WATER rREATMENT 

HYGIENE WATER PROCESSING 

WASH WATER PROCESSING 

4;jiTORING-NA' -T MANAGEKNT 

URINALS 

0 GENERATION - EVA SERV. ~6 REDUCTION = EVA SERY. 

WA ? EF! MGT EVA SERVICING 

AIRLOCK SUPPORT - ULLAGE 

SAVE PUMPS ACCUL SvS 

SAFE HAVEN - POS + MASKS 

MODULE TANKAGE - POTABLE 

WATER 

MODULE TANKAGE - HYGIEKE 

WATER 
TABLE 4.2.4-13 - SPACE STATION SUBSYSTEM COmON EQUIPMENT BY MODULE 

- TC S 

COLD PLATES 

HEAT EXCHAHGERS 

PUMP/ACCUMULATORS 

PLUMB I NG 

VALVESIDI SCONNECTS 

CONTROLS/ INSTRUM 

BODY-MOUNT RADIATORS 

INTERFACE HX'S 

TtlERMAL STOZAGE 

- LAB 1 

32 

32 

7 

9 

7 

1 

1.4 

1.03 

TRUSS 

-- SiRUC 

3 1 

2 

1/4 

1 

1 

i1.2 

1 
TABLE 4.2.4-14 - SPACE STATION SUBSYSTEM CHON EQUIPhENT BY HODULE 

- ECLSS - MI - HHS - LAB1 - LAB2 - LM 

CABIN FANS, DEBRIS TRAP 

MUFFLEF 

DUCTING, FITTINGS 1 1 1 

PLUMBING, VALVES FITTINGS 1 1 1 

MICROB. CHECK VALVES 1 1 1 

PUMP PACKAGES 1 1 1 

WATER ACCUWULATION 1 1 1 

N2 CRY0 2 2 1 

WASTE STORAGE 1 1 4 

N2 CRY0 - MOD REPRESS TANK 2 2 1 

02 CRY0 - MOD REPRESS TANK 1 a 1 

COLLECTOR/COmODE 1 1 
TABLE 4.2.4-15 - SPACE STATION SUBSYSTEM COnrm EQUIMNT BY MODULE 

CREW SYSTEMS 

CLOTHING 

PERSONAL HYGIENE KIT 

EQPT CONTAINERS/RESTRAINTS 

HQUSEKEEPING SUPPLlES 

MCILLORY PROVISIONS 

WARD ROOM 

GALLERY 

WINDOWS 

FOOD CONTAINERS 

HANDWASHER 

CREW QUARTERS 

HMF, TOXICOLOGY , RADIATION 

NOISE 

CAMERA EQPT, ETC. 

STOW?.GE CONTAINERS 

SHOb,ER 

MA1 hTENANCE WORK STAT ION 

DRY AND WET WIPES 

SLEEP RESTRAINTS 

WASTE TRASH STOWAGE 

WASHER/DRYER FACILITIES 

FREEZER 

REFRIGERATOR 

EMU SERVICE STATION 

MU SERVICE STATION 

EMU 

HATCHES 

mu 

EVA TOOLS 81 MISC EVA EQFT 

A1 RLOCK 

AURLOCK EQUIP & 2YPERBARIC 

EQPT 
4,3 SPACE STATION REFERENCE CONFIGURATION EVALUATION SUWARY 

The reference configuration has been evaluated, with all the driving IOC 

requirements previously identified and the effects on each subsystem 

area, cnd met the requirements without significantly affecting the cost 

or threatening the feasibility of any of the major subsystems. The 

configuration provides excellent viewing opportunities for both Earth and 

celestial instruments, a1 though wide angle gimbal systems are required 

for the latter type of instruments. The local vertical, local horizontal 

fl ight mode permits a straight-forward approach to comnuoications and the 

use of establ i shed rendezvous and docki n!q/berthi ng techniques. The 

buildup and assembly sequence appears feasible using the expected Space 

Shuttle Orbiter capability. The configuration is assessed as being free 

of structure/control system interaction, requi res a moderate 1 eve1 of 

control authority, and tends to have a nontumbling attitude in an 

uncontrol led mode. The gimbaled solar arrays and radiators are 

efficient, since they operate continuously in the preferred solar 

orientation. Satisfactory locations have been identified for all the 

"straw man" external and internal IOC payloads and servicing/construction 

functions. The crew accmdations are judged to be adequate for a 

6-person crew, and for the facil i ties required for control , maintenance, 

and EVA support activities. 

In addjtion to the assessment against the IOC customer, operations, and 

systems requirements, the reference configuration has been evaluated 

against additional criteria. The more qual i tative criteria deal with the 

programnatic aspect of cost and risk, the ease of accomplishment in 

systems engineering , area of integration, and growth, as we1 1 as 

transparency to evol ving subsystems and p~l oad techno1 ogy . A cost assessment of the IOC Reference Configuration has revealed no 

significant cost drivers associated with configiiration unique features, 

nor have significant risk drivers been identified. 

The user accomnodations have been located on the con.7iguration based on 

viewing , contaminaticn, configuration, and functional requirements, and 

are considered to meet all requirements and to be near-optimum in meeting 

qual i tative criteria. 

The growth configuration has not been studied to the same depth as the 

IOC configurati ,n. However, at least one means of accommodating growth 

in any direction it might take, as sumarized in the requirements, has 

been identified. Ofie scenario discussed for growth in the power area 

encompasses techr~ological transparency in that qolar dynamic primary 

power sources would be substituted for the photovoltaic arrays envisioned 

zc iOC. Tt;is can be accompl i shed either by replacement of the entire 

power boom assembly, including mechanisms, or the initial structure and 

mechanicai systems can be "scarred" to Include the capability of 

accommodating the growth and system type change. The addition of modules 

for increased crew size and laboratory volume can be accommodated by the 

addition of structure to the keel extensions and ;he 
relocation or replacement of the lower boom assemblies with the 

propulsion and Earth viewing payload provisions. Growth in the 

construction and satellite and OMV servicing areas is accomodated by the 

erection or deployment of additional structure to the keel, and 

relocation of 1OC servicing equipment, and perhaps TCS radiators, as 

required. Growth in the viewing instrument attachment area is 

acconrpl ished by addition of structure and utilities to the upper and 

lower bcms, with relocation of propulsion and connrunications equipment 

i f necessary. 
4.3.1 Payload Accomnodations 

A large, re resentative set of payloads for the IOC Space Station manned 

core has een +a-T- 1 enti led. The purpose of this set of payloads is to 

bring out the wide range of customer requirements which will be imposed 

on the Station. It is not suggested that the entire representative 

payload set be present on the Station at IOC. Rather, the myriad of 

requirements brought forth by this representative set of payloads dre 

used to guide the design of the Space Station. The representative 

payload set provides a measure for assessing the capabil i ty and 

versatility of the Space Station configuration to accommodate its 

customers. 

4.3.1.1 Representative Payload Set 

The representative payload set is not an olficially accepted IOC payload 

set; it has no official status. The canponent payloads were chosen on 

the basis of the requirements which each !mposed on Space Station design 

and the proposed initial time period (1992) of their operation. 

Descriptions of the individual payloads are provided within the Mission 

Requirements Working Group (MRWG) Langl ey Ddta Base Documents and other 

MRUG reports. 

The representative set of payloads may be divided into the three broad 

categories of (1) science and appl ications (SAA) , (2) comercial (COM) , and (3) technology development missions (TDY) . These payloads and their 

MRWG identification codes are presented in Table 4.3.1-1. 

Servicing and refueling of free-flying spacecraft and platforms will a1 so 

be accommodated by the reference Space Station. A representative set of 

free-flying payloads which require servicing and/or refuel ing has been 

developed; this set of payloads is presented in Section 4.3.1.C. 

4.3.1.2 Attached Pay1 oads 

Seventeen payloads in the representative payload set are to be attached 

to the Space Station manned core. These payloads are identified in Table 

4.3.1-2. Requirements imposed by these payloads include: 

a. mass distribution and payload location 

b. simultaneity of viewing in inertial, solar, Earth, and 

anti-Earth directions 

c. precise pointing fields-of-view, accuracy, and stabil ity 

d. contami nation-free environment 

e. 10-5 micrograv i ty 
TABLE 4.3.1-1 - REFERENCE IOC PAYlOAD SET 

IDENTIFICATION CODE PAYLOAD NAME 

Transition Radiation 6 Ion Calorimeter (TRIC) 

STARLAB 

Pinhole Occulter Facility 

LIDAR Facility 

Space Plasma Payload (SPP) 

Life Science Laboratory 

Microgravi ty Research Laboratory 

MP5 Lab #I (Materials Processing) 

EOS Production Unit 

ECG Production Unit 

Material s Performance 

Material s Processing Technolsgy 

Depl oymentlAssmb1 ylconstruction Technol ogy 

Structural Dynamics Technol ogy 

Earth 03servation Instrument Technol ogy 

Fl uid Management Technol ogy 

Attitude Control Technol ogy 

Figure Control Technology 

Envi ronmenta' Effects 

Habitation Technology 

Medical Technol ogy 

Satel 1 i te Servicing Technology 

OTV Servici ng Technol ogy 
f . construction of 1 arge structures 

g. servicing 

h. attachment of tethers 

i . control consoles within pressurized module 

4.3.1.2.1 Mass distribution and payload location 

The 17 attached payloads represent a significant mass component of the 

Station. The weight of each attached payload is presented in Table 

4.3.1-2. The sum of the weights is approximately 110,000 pounds. 

Placement of these attached payloads will infl uence the mass distribution 

characteristics of the Space Station. The payloads are distributed on 

the reference ccnfiguration such that the inertial, solar, and anti-Earth 

pointing instruments are placed on the upper boom. Earth viewing 

instruments are placed on the lower boom. The technology payloads which 

have no pointing requirements are situated along the keel. The 

comnercial production units are attached to the Microgravity and 

Material s Processing Facil i ty module. 

4.3.1.2.2 Simultancity of viewing 

Attached payloads listed in Table 4.3.1-3 have specific pointing 

requirerents; dttacned payloads not 1 isted in this Table have no pointing 

requirements. 

The requirement for different payloads to simultaneously point in 

different directions is accommodated onboard the reference configuration 

by having the various instruments, or possibly groups of instruments, 

independently gimbdl ed. 

In general , instruments requiring similar pointing directions are grouped 

together. I~struments which require fnertial, sular, or anti-Earth 

viewing directions are situated on the upper boom. (The option, however, 

is held q?zn to place some solar pointing instruments on the transverse 

boom. 1 Instruments requiring pointing in the Earth direction are placed 

on the lower boom. 

4.3.1.2.3 Pointing fields-of-view, accuracy, and stability 

Payload requirements for pointie~g fields-of-view, accuracy, and stability 

are identified in Table 4.3.1-3. 

Placement of the pointing instruments on either the upper boom (for 

"outwardu-looking instruments) or the lower boom ( for "downward"-looking 

instruments) will maximize the possible field of view by reducing that 

area potentially subtended by the Space Station structure. 
TABLE 4.3.1-2 - ATTACHED PAYLOAD WEIGHTS 

CODE PAYLOAD WEIGHT 

Transition Radiation L Ion Calorimeter (TRIC) 

STARLAB 

Pinhole Occulter Facil i ty 

LIDAR Fa. il ity 

Space Pl , .ma Payload (SPP) 

EOS Production Unit 

ECG Production Unit 

Matzrial s Performance 

Depl oyment/Assembl y1Construction Techno1 ogy 

Structural Dynamics (mass requirements are 

covered by TDM2060 ) 

Earth Observation Instrument 

Fluid Management 

Attitude Control 

Figure Control 

Environmental Effects 

Satellite Servicing 

OTV Serv i c i ng 

12,675 lbs. 

7,055 

7,940 

4,190 

7,055 
TABLE 4.3.1-3 - POINTING REQUIREMENTS OF IOC PA'fLOAO SET 

POINTING IDENTIFICATION PAY LOAD NAME FIELD OF POINTING POINTING 

DIRECTION CODE VIEW ACCURACY Sf AS ILITY 

(DEGREES) (ARC SEC) (ARC SEC 

EARTH ShAO201 LIDAR Facil i ty 60 3600 3600 

EARTH SAAO207 Space Plasma 

Payload (SPP) 360 3600 3600 

EARTH TDM2260 Earth Observation 

Instrument Tech. - 360 .. 

EARTH TOM2510 Envirormental Effects 20 7 200 - 

ANTI - Transition Radiation 

EARTH SAA0005 and Ion Calorimeter 

(TRIC) 120 36000 36000 

SOLAR SAAOOQ9 Pi nhol e Oecul ter 

Facil i ty 3 10 1 

SOLAR SAAO207 Space Plasma 

Pay1 oad (SPP 1 360 3600 3601) 

SOLAR TDM2010 Material s Perfomance - 7 200 - 

SOLAR TOM2510 Environmental Effects 20 7 200 - 

INERTIAL SAAOO06 STARLAB 180 2 0.02 

INERTIAL TDM2410 Attitude Control 

Techno1 ogy - - - 

INERTIAL TDM2420 Figure Control 

Techno1 ogy - - - 

Payloads listed under more than one "POINTING DIRECTION" category 

have component parts with different point1 ng direction requirements. 
Pointing accuracy and stabil i ty requirements for payloads SAA0006 

(STARLAB), SAA0009 (Pinhole Occul ter Facil i ty) , and TDM2260 (Earth 

Observation Instrument Technology) cannot be ntet without some type of 

pointing augmentation system for the individual instruments. The nature 

of this pointing system has not yet been determined. 

4.3.1.2.4 Contamination-free environment 

Contamination is a serious concern for sensi tive-viewing instruments. 

Placement of those viewing instruments on the upper boom and the lower 

boom remove them from the vicinity of the sate1 1 ite refueling area and 

the Shuttle and OMV berthing areas on the lower keel . 4.3.1.2.5 - lo-s microgravity 

Payloads COM1202 (EOS Production Unit) and COM1203 (ECG Production Unit) 

have the requirement for a 10-5 g environment. These two pa:tloads are 

customer-suppl ied commercial production facil i ties. 

(Payload COM1202 requires attachment to the manned core modules through a 

pressurized port to permit IVA servicing. Payload COM1203 is 

unpressurized, serviced by EVA, and therefore does not neea to be 

attached at a pressurized port.) 

Since the Microgravity and Materials Processing Facility (MMPF) mociuie 

a1 so has a requirement for g, payload COM1202 is attached to that 

MPF rnodb'e. 

4.3.1.2.6 Construction of large structures 

Pay1 oad TDM2060 (Depl oymentIAssemb1 y1Construction Techno1 ogy) requires 

the construction of a 30-meter-diameter dish antenna. This antenna is 

positioned on the reference configuration lower keel for reasons of 

station mass properties and ease of EVA access. 

4.3.1.2.7 Servicing of attached payloads 

Servicing of the attdched payloads will be by EVA operation. (The 

exception to this EVA servicing is payload COM1202, which is pressurized 

and will ha~e IVA servicing. Attached payloads are situated at various 

locations on the structure: the upper boom, the lower boom, the lower 

keel, the mdules, and possibly a1 so the power boorn. These attsched 

payloads will require regular servicing. In addition, unanticipated 

repairs or trouble-shz2ting ma,y be necessary. Mobile RMS (MRMSj and EVA 

operations must be designed to work synergistically to allow servicing of 

these payloads. 

4.3.1.2.8 - kttncirment of tethers 

i'ayload SAA0207 l Space Plasma Pay1oad) will require both payload 

tethering "up" (up 30 krn in length from Space Stdtion along a cunducting 
tether) and tethering "down" (down 30 to 300 km from Space Station along 

a nonconducting tether). Tethering "up' can be accomplished from the 

upper boom; tethering "down" can be done from the lower bm. 

Payload SAA0207 further requires the deployment (and retraction) of a 

300-meter dipole antenna. This long dipole antenna needs to have the 

capability to be rotated plus-or-minus 60' from nadii- in order to fly 

either a1 igneh para1 lel or perpendicular to the Earth magnetic field . . r I r~tis. 

4.3.1.2.9 - Control consoles within pressurized mdul e 

Many of the external ly attached payl oads will r-equi re control consoles 

within one of the pressurized modules from which those payloads may be 

operated routinely. 

4.3.1.3 Pressurized Laboratory Modules 

Six of the payloads identified in the representative set of payloads 

(Tab1 e 4.3. I- 1 ) are conducted within pressurized 1 aboratory modules. 

Speci f ical ly , these payl oads are: SAA0307, Li fe Science Laboratory ; 

SAA0401, M-icrogravi ty Research Laboratory; COM1201, MPS Laboratory kl; 

TDM2023, Material s Processing Technol ogy; TDM2520, Habitation Technology; 

dna TgM2530, Medical Technol ogy. 

At IOC, there are two habitable Space Station modules to conduct. 

laboratory work and to provide olzrational support. For the purposes of 

this report, one module is identified to be a Microgravity and Materials 

Processing Facility module; it will require the maintenanc? of a 

mirrogravi ty level less than or equal to g. The other is 

identified to bo a Life Sciences Laboratory module; it is designed to 

conduct research into the problems and henomenological effects of 

long-term exposure of humans, animals, and plants to near 

weightlessness. The Life Sciences Laborator module Is required to 

maintain a microgravity level of at most 10-4 g. Additionally, the 

Life Sciences Laboratory will have equipme~t and work space in it for 

monitoring artd control 1 ing the various payloads attached externally to 

the St~tion. 

Both iaboratory modules are intended to have sufficient versatil i ty to 

support other science and appl ication missions. To that end, both 

1 aboratories will be capable of having equipment reconfigured on-orbi t. 

Science airlocks and optical-quality windows will be provided to 

faeil i tate viewing by internal ly-mounted payl oads. Parameters for 

viewing direction will constrain module orientations and internal 

arrangements of equi pme,lt. 

4.3.1.3.1 - Mr'c -ogravi ty and Materials Processing Facility module -

The Micrograv ty and Materials Processing Facil i'ty (WPF) module will 

support the development of unique materials and processes. An 
acceleration envilmonment of 10'5g, tr less, is required For periods of 

days to months. The WPF module will support the advancement of the 

knowledge base; the development of processes and process controls; the 

scal e-up to pi 1 ot yl ant operations; and the operation of pre-production 

snd commercial production foiil i ties. 

The module is requ iit.6 t~ be a sectional ized facfli ty composed of 

equipment stations that deliver the utilities of powcr, heat rejection, 

vacuum, control gases, data hand1 ing, comnunications, and canmand. The 

principal equipment within the module will include low and high 

temperature furnaces, crystal growth apparatus, containerless 

(levitation) furnaces, bi~logical materials separation facil i ties, a 

scientific air? ock station with equipment attachments exterior to the 

module, a sample preparation and characterization work statio,i, and 

instrunent and materials stowage. Attachment directly to the interior 

wdll of the module will be necessary for some equipment. 

The generic distribution of utility outlets and attachment provfsions 

wil l permit sel f-contained, integral equipment to be quickly "p'luggedu 

into an equipment station by tne crew. A high degree of equipment 

changeout and reconf igura tion will be essential to conduct required 

research, devei opment, and engineering. Speci fical ly, util i ty control to 

petmit ch,nge of the utility d,istribution to different equipment as a 

fufiction 01- tasks is needed. This utility controi function will 

accmoc', ..i. utility distribution to new processing equipment that is 

exchanged on-srbi t with obsol ete equipment. 

A schematic drawiag of the Microgravity and Materials Processing Facil i ty 

module is shown in Figure 4.3.1-1. 

4.3.1.3.2 Life sciences laboratory module 

The avail able voluxe within the Life Sciences Laboratory will be dirided 

between a human reseerch facility, an animal and plant research facility, 

and operations jcontrol facil i ties for payloads external ly attached to the 

Station. Few finn requirements exist at present for the volumes of these 

individual facilities within the Life Sciences Laboratory. For the 

purpcses of sizing the util ity requireinents of this module within the 

context of the overall Station cor;figuratfon, some assumptions of 

equipment mass and volume have been made. These assumptit)ns are intenaed 

only to support the viability of a mu1 tidiscipl ine Life Sciences 

Lahoratory and do not necessarily represent a consenstis within the life 

science comnuni ty. 

A schematic drawing of the Life Sciences Laboratory modu!e is shown in 

Figure 4.3.1-2. 

4.3.1.3.2.1 tiuman research facility 

The human research facil i ty will conduct Snvestigations of the 

physiological changes which human beings undergo due to exposure to 
'iMATERIALS PROCESS IHG RACKS 

DATA MAElAGElf llT EJUIPKFIT 

!HPHD WASllER 

I 

- -DATA MAN

OPEMTIONS '1 

7 .r FYiFFr ErjT \ CPM RA7 

/'- 

/" \ 

- ECLSS OXY GEti 

C:-I SS -' L . __ COMMUtI I CAT I OHS 

EQUi PMEMT 

EQUI PENT MAT

t-r~v~rr PRO

RAC

Flgun 4.3.1-1 - MICROGRAVITY AND MATERIALS PSOCESSING FACILITY 

microgravity for extended periods of time. (Much of the early research 

will concern NASA's need to develop countermeasure; to protect the health 

and safety of the Space Station crew.) 

Equipment for the humac research facil ity is assumed to require eight 

19-inch racks and two 38-inch racks. The contents of these racks are 

core 1 aboratory cqu-ipment, di scipl ine specific equipment, data management 

equipment, storage space, ond supporting hardware. A treadnil 1, 

ergometer, and body mass measurement devi~e will be situated in the 

center- aisle. In addition, there vill be floor and ceiling storage space 

and nounting provisions for experiment uniqwe hardware. 

Total volume for the equipment Ss approximately 280 ft3; tot61 

eqlripment mass is nearly 3175 lbm (1440 kg;. 

4.3.1.3.2.2 Animal/plant research facility 

The animal/pl ant research facil i ty will conduct long- term st~dies on 

microgravity-induced physiological changes in animals; on the mechanisms 

of gravity censing and gravity responses i~ animals and ,)lants; on the 

effects of microgravity on fundamental biological systems, including 

reproduction; and on biological and chemical system to close the 

ecologicai cycle and convert waste into food. 

Volume for equipment-within the anicii/&?ant resoarc* facility is 

approximately 160 ftJ. The equipment mass is nttdrly 4400 ?bm (2000 

kg). (A1 though t.his equipment volme is grxter than that for the human 

research facil i ty, the overal I volume of the ar aal/plant research 

facility is approximately the sane as thzt of the hman research 

facil i ty. ) Required racks will contain core 1 ~bor atory equipment, 

di scipl ine specific equipment, genera! 7urpcs.i woi kbeoch, advanced rodent 

holding facil i tizs, refri~er.?sl,r/freezer, plant growth irnits, closed 

ecology 1 i fe support equipment, and data mandgemerit eq!li pment . C e i l i n g 

stor$ge will also be provided. 

The animaliplant research facility is required to be encf osed in a 

sepdra te camgartinefit within the module. Thi:, cor,lpartment will have its 

own independent ECLSS system iproviaed by the user). The ECLSS will 

prevent cross-contaminatio~ between the rest of the Spat* Station and the 

animal dnd plant specimens. The ECLSS will exclude particles larger than 

0.5 microns on the air inlet and part!~les larger than 0.3 microns on the 

air outlet. Charcoal filters wi 11 el imii. .t sdor t-ransfe:. 

The animal and plant holding facilities should be as isolated as 

prictical (e.g., by movable harrie,~! from the ldboratory work area. 

This will provide better control of the 1 ight/dark cycle for specimens, 

more accurate tempereture and hu~idity control, and will isolate the 

animzls from disturbances caused by crew activity. 
Again, it must be stated that the particular choices of volumes and 

equipment for the human research facil ity and the animal/plant research 

faci 1 i ty were selected somewhat arbitrarily for reference conf igur-ation 

purposes. These volusles and equipment choices do not represent a 

concensus Qcisi~n of the life sciences cmunlty. 

4.3.1.3.2.3 Operations/Control Facilities 

Many of the externally attached payloads and servicing operatio?; will 

require consoles within the pressurized mdules from which those payloads 

and operdtions may be monitored and contrsl led. Equipment tnich fill s 

three 16-inch racks is provided inside the Life Sciences Laooratcry 

module fcr those col?trol functio~s (labeied as "Astrophysic. Racks" in 

Figure 3.3.1-2). 

4.3.1.4 Servicing Facility 

The Szr-vicing Facility onboard the IOC Space Station manned core will 

have the capabilities to service and refuel free-flying satellites (which 

have been brought to the Station), co-orbi ting platforms ( interpreted to 

be multi-payload spacecraft which can be berthed to the Station), 

paylaads attached to the Statiot?, the OMY, and the ObiV kits. The 

Servicing Facil ity will a1 so provide for the storage of satell ites, the 

WV, two OMY kits, ORU's, instruments, and tools. 

A list of representative frcee-fly-ing satellites which will require 

on-orbi t servicing or refuel ;ng at the Space Station is presented ia 

Tab1 e 4.3.1-4. Again, this 1 ist of representative free-f ly ing payloads 

has no official status. The set was developed to reflect the type and 

quantity of free-fl iers proposed for the 1991-1993 time period. 

Servici~~g and refuel ing requirements for this representative paylaad set, 

along with serious consideration for eventual expansion to growth 

canabil ities, were used to develop the Servicing Facility for IW,. 

Two dedicated work sites, or "bays" are required: one bay is nseded to 

perform servicing ogerations and the other to perform refueling 

operations. Several of the spacecraft which will be serviced or repaired 

contafn optical instruments which are highly sensitive to molecular 

and/or particulate contamination. Separate facilities for servicing and 

refueling operations are necessary to prevent possible ~ontai~iil~ation of 

optics which are exposed during servicing to unexpected leak~ge of 

propellants or other co~taminaats possible in the refueling area. 

This concern with the sensitivity of payload instruments to various 

ccntaninants dictates that the servicing bay be separated ar~dlor 

"upstream" from the refuel iy and fluid storage areas, from the Orbiter 

beything area, and from a,.: pressurized modules which may vent 

contaminants (e.g., laboratory or comnercih-i modules). 
TABLE 4.3.1-4 - REPRESENTATIVE SET Of !OC FREE-FLY ING SATELLITES 

REQUIRING SERVICING 

IDENTIFICATION FREE-FLY ING SATELLITE 

CODE SERVICING PAYLOADS 

Space Telescope (ST) 

Giua~a Rqy Observatory ! GRO 1 

X-Roy Tiair~g (XTE) 

Solar blaximula Mfssion (Sm) 

Advanced X-Rqy Astrophysics Facil i ty ( AXAF) 

Far Ultraviolet Spectroscopy Expl arer (FUSE) 

Leasecraft 1 and Payloads 

Leasecr aft 2 and Payloads 

Leasecraft 3 and Payloads 

Leasecraft 4 and Payloads 

Leasecraft 5 and Pqylodds 
The refueling bay and fluid storage area should be located so as to 

reduce any hazard potential to satel 1 ites being serviced, 

i nstrumentslpayl oads externally attached to the Station, or Station 

systems such as the solar arrays or radiators. 

Some satell ites may require solar protection while in the servicing and 

storage areas. If needed, thermal pr~tection may be provided by some 

type of shield or enclosure. 

An access corridor with sufficient clearance must be available for the 

OMV with attached payload to move close enough to the Statio~ so that the 

Mobile WS (MRMS) can grdpple dnd berth the OMV and yzyload. 

MRMS access to Servicing Facility elements is required so that payloads 

may be mved between the servicing, refuel ing , and storage areas. A1 so, 

ORU's must be moved between the Orbiter and the ORU storage area. 

A clear translation path is needed for the movement of EVA crew between 

the core modules and the Servicing Facility elements. 

The elements of the Servicing Facility will need to be provided with 

util i:ies including power, 1 ighting, CCTV, 1 iquid 1 ines, aad 

data/commdnica ti on. 

The elements which make up a Servicing Facil ity that acco~rmod~tes IOC 

miss iori servic iny dre iile ioi iowing: 

a. Servicing Bay :: a cylindrical vcl ume (not necessarily 

enclosed) which is 30 feet io diameter and 70 feet ia length. This 

vol ume a1 lows for the berthing of a 15-ft-diameter-by-60-ft-30ng 

satellite with clearances all around for moventent ~f EVA crew and the 

placement of workstations. The servicing area will have provisions for 

berthing pay1 oads either by a Fl ight Support Structure (FSS) , which has 

tilt and rotation capabilities, or by trunnior? latches. Moveable or 

reattachable berthing assembl ies wo~ld perni t the berthing of more than 

one pwload in this area. 

The servicing bay is sttached to, and para1 lei wi lh, the upper keel above 

the trdnsverse born. 

b. Refueling Bay: A cyl indric3l volunie w'th the same 

approximate dimensions ss the servicing are and simi! 2- berthing 

mechanisms. The reFueling hay is situated on the lower keel just above 

the radiators. 

c. Satellite Storay Area: A cylindrical volume with the 

same dimensions as the servicing area l i .e., 30-ft-diameter-by-70-ft 

length) aqd with Ihe same berthing mechanisms. (This volume i s i n excess 

of the approximate 15-ft-diameter-by-60-ft-1 ong vol ume which is actually 

required for storage purposps. However, a1 1 ocat ion of the additional 

volume would permjt this &rea to evolve into another servicing 
area for the growth Station.) The satell ite storage area is located 

across the upper keel from the servicing bay. 

d. Fluid Storage Area: An area which will provide 

facilities for storage of propellants, pressurants, and coolants for the 

payloads. It is located just beneath the refueling bay at the top of the 

keel extensions. 

e. OMV Storage Area: A cylindrical volume approximately 15 

feet in diameter and 4 feet in length. The OMV storage area is situated 

on the keel extension just beneath the radiators. 

f. OMV kits Storage .4rea: Two cylindrical volumes 

approximately 15 feet in diaeter and 4 feet in length. They are located 

on the keel extensions opposite to the OMV storage area. 

g. ORU Storage Lockers: Each enclosed rectangular locker 

is 3 x 5 x 5 feet. Ten lockers will be available for ORU storage. 

They are placed on the power boom inboard of the alpha joints for 

iocvenient access from the servicing bay. 

h. Pay1 oad Instrument Storage: An enclosed rectangular 

cmpartment which is 10 x 20 x 33 feet. It is situated on the lower keel 

opposite the refuel i ng bay. 

i . Tool Storage Lockers : Each enc! osed rectacgul ar 

compartment is 3 x 5 x 5 feet. Four lockers will be avdilaale for tool 

storage. They are located with the ORU storage lockers. 

j. Monitoring, Control, and Checkout Equipment: 

Monitoring, control, and checkout is provided for spacecraft and ORU's in 

storage as well as spxecraft undergoing servicing and refueling 

opl$rations. These tunctions will require volume for display consoles 

wi hin the pressurized modules. 

4.3.1.5 Growth 

Customer utilization ofi the growth Space Station will require an 

expansion of capabil 'ties present on the initial Station. However, it is 

not clear at this time just which capabilities will grow and to what 

degree, or how that growth will drive the Station evolution. 

An attribute of the reference Space Station configuration is thdt it can 

support growth in any or all of its initial capability areas: servicing 

and refoel i ng; constructim of 1 arge space structures; material s 

processing; 1 ife science research; i ;trophysics and solar physics; Earth 

remote sensing; or sensor development. GrorJth of some of these 

capabil 1 ties would require increased crew size (e.g., servicing, 
construction, life science research). Growth of other capabilities would 

require significantly increased power (e.g., materials processing). 

Whichever capabil i ties eventual ly come forward as growth requirements, 

the reference configuration can gracefully evolve to meet them. 
4.3.2 - Crew Accomnodations --- - 

Crew arcomnodations refers to al! areas which involve crew systems and 

crew support. The major portion of these accamnrodations i s i n the 

habitation vol me. The laboratory volume contains crew-related items if 

they aid in crew heal tin, safety and well-being, or if they support the 

crew in providing customer services. Crew accomnodations in all volumes 

are designed and arranged to the zero-g neutral body posture, traffic 

patterns, congestion avoidance, cleaning and ease of maintenance. 

Functional group interrelationships were a prime consideration in the 

basic arrangement. Other considerations in the design and arrangement of 

facil i ties and equipment incl uded the abil ity to support reconf iguration, 

grawth a,>d update; the ability to access facilities and equipment; 

standardization of crew interfaces and associated equipnient; and 

accomnodaticq of anthropometric strength and sire measurements. 

4.3.2.1 Habitation Volume 

In the reference configuration, the habitation vol unie consists of 'iwa 

habitation modules which contain all those facilities and equipment 

required for crew cllstenance and station operation. The fol lowing areas 

and categories of equipment are contained in the two modules: crew 

quarters, gdll ey/wa;-droom ( including housekeeping suppl ies) , health 

maintenance faci 1 i ty/exerc;se area, personal hygiene area, workstation 

areas, stowage areas, waste management faci 1 i ties, and 1 sundry facil i ty . These areas contain equipment, the facilities which "house" this 

equipment and access volume which facilitates its use. The internal 

configuration of these modules is shown in Figure 4.3.2.1-1, 4.3.2.1-la, 

and 4.3.2.1..ib (Hab 1) and Figures 4.3.2.1-2, 4.3.2.1-2a and 4.3.2.1-2b 

(Hab 2). 

4.3.2.1.1 Crew Ouarters 

Private quarters for each cremember are located in Hab Module 2. Each 

crewmember has been provided with 150 cu ft of volume which contains a 

sleepstation with bedding, a cm~unications unit, a desk, a CCTV/CRT, a 

portable workstation with certain comnand/sontrol functions, c+,owage 

vol ume (20 cu ft incl uding a "dresser" and area for personal i terns), a 

bulletin board, audio/video eotertaiment and a mirror. These 

accomnodations are shown in k igur-e 4.3.2.1-3 ($1 eep Cmipartmen ts) and 

Figure 4.3.2.1-4 (Sleep Compartment Furnishing in Working Position). The 

volume which has been a1 located is below desirable standards, but has 

been expanded beyond those minimum levels included in previo~ls 

investigatioas. Privite crew quarters util ize 88 inches cf /?e length of 

the module. 

The ~..iley/wardroom is locdted in Hah Module 1. The galle,y, as depicted 

'n Figure 4.3.2.1-5, cccupies 747 cu ft of whic'~ 225 cu ft is equipment 
r-POIIER EQUIPMENT UARDROOM7 

/ 

ECtSS 

WYAG

EQUI

FACILITY 

Flgura 4.3.2.1-1 - HAD MODULE 1. 


-COM4Uf4I CATIOriS EQUIPff t4T 

(X, 

W 

OPE

REPAIR STATION 

FI~UTO 4.3.21-2 - HA6 MODULE 2 



PULL-OUT LOCKERS -- -- PULL-O

MI RRO R --- TY 

LOCKER WITH 

FOLD DOIRI FRONT 

PAPIE L CRT (P

AUDIOCOM - - - -- -CRT K

'IRITIMG SURFACE -- --- PULL-O

(FRONT PANEL CF 

LOC (ER) 

----- LOCKER 

FOLD W

PANEL 

PULL-OUT LOCKERS 

- 

-I-- PULL-O

-- J 

Flgura 4.3.2.1-4 - SLEEP COMPARTMENT FURNISHINGS IN WORKING POSITION. 

volume, utilizes 163 inches of the length of the module and provides the 

following accomnodations: equipment and supplies necessary for the 

preparation and heating of food and drink; clean-up ar2 stowage for 14 

days is provided. Food and drink stowage is provided by the 

ref rigerator, freezer and ambient food stowage 1 ockers . Cooking 

accomnodations are provided by the oven and microwave oven. There is 

additional stowage for utensi 1 s/appl iances, housekeeping suppl i es, and 

trash. Galley trash collection and compaction for wet and bulk trash is 

provided. There is also a handwasher and a zero-g dishwasher. Galley 

equipment and food are GFE. 

The h~rdroom is an area, across from the galley, occupyi~: 576 cu ft, 137 

inches in length, designed to acccmnodate the entire crew simultaneously 

as a dining area as well as a meeting area, lounge, viewing and 

recreation area. This area contains tables (4 sq ft per person), 

audiolvideo entertainment equipmeat, game kits, a window, and IVA 

communications. 

The accomnodations provided were chosen to be above minimum levels in an 

effort to iacrease acceptability for long-duration flight and to 

contribute to the optimization of crew performance. 

4.3.2.1.3 Health maintenance facilitylexercise area 

The Health Maintenance FacilityIExercise Area is located in Hab Module 1, 

utilizing 72 inches of the length of the module, some 65 cu ft of 

equipment volume and occupying a total of 320 cu ft. The equipment 

contained in this vol ume provides for infl ight preventive, diagnostic and 

therapeutic meaical and dental capabil i ties. The health maintenance 

facil i ty (HMF) is principal ly for the appl ication of countermeasures 

inc1ud:ng exercise and the treltment of health problems. The WF and 

exercise area as depicted in Figure 4.3.2.1-6, share the same volume as 

well as some of the sane equipment. The approach to the design of the 

HMF is as follows. Its capabilities are tailored to the degree of 

mission complexity and defined "acceptable medical risks." The 

reqtii rements are derived from infl ight medical experience, mission 

re1 ated activities and hazards, and projected low probability medical 

scenarios which could be adequately treated with minimal equipment. 

The exercise area contains the equipment necessary to enable the crew to 

retain the requisite physical body tone and a1 so provides d means for 

recreation. Provisions a1 so exist in the exercise area for monitoring 

body mass and dimensional changes. 

4.3.2.1.4 Personal hygiene areas 

Personal hygiene areas which provide faci 1 i ties for body waste collection - and disposal, personal clean1 iness, and bathing are located in Hab 

Modules 1 and 2. Each area is 80 inches in length. There are two waste 

management systems (for both fecal and urine collection); two additional 

urine col lect ion systems; two personal hygiene stations for face and 

handwashing and oral hygiene; and one additional handwashing station in 

the galley and in each cf the remining hdbitable modules. 

8 9 
-- PATIENT 

(111 STOV

-. ----- . 4 

- Pity'. InLOGICAL . MOIIITOR1;IG --- - -- 

- ~..+ 

DE S I< - ----A- , . - - -- -~ 

-- l- . 

. . 

- -.. 

CLINICAL LA: --- -- - -- - -.- 

(MICROG IOLOGY ) .--. 

COpPUTER SUPPI.IE5 -- - - 4 ~~ 

COIITRnL DRUGS ----.. -- 6--- -- - 1.. 

EXERCISE E:]UIPI,lF:IT '---- . 

- -a 

"ICYCLE 3 - 

E RGOMETE I? 

TOXIC01 nC

MJriOR SUR

IV SYSTEI

If~1AGI~lS

-CL I Ii I CAL 

(CIiEr4ISTR

CLTllICAL 

(IfEI"TOLOG\

SUPPLIES 

TREADMILL 

Flgure 4.3.2.1-6 - HEALTH MAINTENANCE FAClt.lTY. 
The facil ities for collecting and disposing of vomitus, fecal matter and 

urine we sired to anthropometric specifications. Canntode compartments 

are iarge enough to pennit donning, doffing, and temporary stowage of 

clothias. 

The facilities for shaving, hair grooming, teeth clear'ng and 

expectoration are designed for easy cleaning and maintenance on orbit. 

Handwash facilities have hot, cold, and mixed water controls. 

A full body shower facility is provided in each of the Hab modules. Each 

has hot, cold, and mixed water controls, permits hair and scalp washing, 

and provider a temperature controlled (heated) private dressing area. 

There dre restraints to stabilize the crew while bathing. Means to 

faci 1 i tate drying 2nd cleanup. waste water coll ectioa, and transfer to a 

storage tank fgr processing dre provided. 

4.3.2.1.5 Workstations 

Workstations are 1pzuted throughout ,:he Space Station in all the 

habitable modules. They are conceived as generic stations capable of 

supporting Space Station oper<tions, customer services, crew health and 

equipment maintenance. Hab module 1 contains the following workstations: 

the work/control station for Space Station operations (Figures.2.1-7) and 

portions of the health maintenance facility (Figure 4.3.2.1-6). Hab 

mdul e 2 contains a 1 arger cmand/control station (Figure 4.3.2.1-8) and 

the maintenance workstation for equipment mdintenance and repair. Each 

of the Lab modules contain workstations. There is one operations station 

in Lab 1 (Figure 4.1.2.1-9, Life Sciences Lab) and two operation sta2ions 

in Lab 2 (Figures 4.3.2.1-10, 4.3.2.1-11; Materials Processing Lab). All 

workstatians are conceived to have the necessary and appropriate level of 

automation, tool s, equipment, and comnunications required for the 

specific tasks and dctivi ties to be accompl ished. 

4.3.2.1.6 Crew stowaae 

Twelve hundred cu ft of stowage has bee,? ,,rovided for a crew of six. 

Stowage vol~me has been a1 located to e~cn module as indicated in Table 

4.3.2.1-1. This allocation is the rr.sVl?t of the stowage requirements 

presented in Table 4.3.2.1-2. The ;op r stics module contains the greatest 

portion of the stowage volume including stowage for clothing for 14 days 

(this assumes a 1 aundry facil ity on..br~ard), for a 90-day supply of food, 

for bedding, and for food trash stlj:rage. The two Hab modules have 

approxi;nately equal amounts of stowage vol ume i ncl udi vg stowage in the 

gal 1 ey , crew quarters, personal hygiene areas, and at the various 

workstations. Lzbs 1 and 2 each has a small ?,w,unt of stowage volume 

provided. Ezch stowage container is 2 cu f? ar ! weighs 5 lbslcu ft (10 

Ibs) empty. 

These stowage requirements have been derived frrn !~revious infl ight 

experience as we1 l as extrapolation of consuma'~1e; for six people over a 

90- day period . 

KEY SO ARD 
MANAGEMENT AN I MAL 

EQUIPMENT LDATA ~ECLSS 

SAFE IAVEIIJ 

1 1 ANIMAL ECLSS 

TIiERMAL HUMAN OPERATIOllS SEALED 

EQUIPMEriT ECLSS STAT1 ON AN 11441. 

RESEARCH 

LA9 

Figure 4.3.2.1-9 - LIFE SCIENCES LAB OPERATIONS ''TATION. 

UTI
r 

MATERIALS PROCESS1:IG R4CKS 

DATA MANAGEMEtlT EQUIPIEHT 

I DAT

ECLSS - 

EQUI PMEFlT 

EQUIPMEIlT SAFE --\ 

I IAVEM 

Flgurr 5.3.21-10 - MATERIALS PROCESSING LAB. 

HAB 1 

TABLE 4.3.2.1-1 - STOWAGE ALLOCATION FOR CREW OF 6 (PER MISSION) 

GALLEY 

WARDROOP3 

MAINTENANCE WORKBENCH 

EQUIPMENT CONTAINERSIRESTRAINTS 

WORKICONTROL STATION 

PERSONAL HYGIENE AREA 

ADDITIONAL MISCELLANEOUS STOWAGE 

TOTAI. 

NO. OF STOWAGE 

CONTAINERS VOLUME 

(CU FT) 

HAB 2 TOTAL 106 .OO 212.00 

CREW QUARTERS 

COWICGNTROL STAT ION 

PERSONAL NYGICNE AREA 

EQUIPMENT CONTAINERSIRESTRAINTS 

ADDITIONAL MISCELLANEOUS STOWAGE 

LOGISTICS TOTAL 264 .OO 528 .OO 

CLOTHING (14 DAYS) 

FOOD (90 DAYS) 

BEDDIMG (90 DAYS) 

FOOD TRASH STOWAGE (EMPTY) 

AODITlONAL MISCELLANEOUS STOMAGE 

- LAB 1 TOTAL 56 .OO 112 .OO 

-- LAB 2 TOTAL 56 .OO 112.00 

STATION TOTAL 600.00 1200.00 
TABLE 4.3.2.1-2 - STOWAGE REQUIREMENTS FOR CREW OF 6.00 PER MISSION 

TOTAI. VOLUME 

SUBSYSTEM (CU FT) 

CREW QUARTERS 120.00 (20 CU FT/ 

( INCLUDES CLOTHING, PART OF PERSONAL CREWMEMBER) 

HYG!ENE PROVISIONS AND PART OF 

ANCILLARY PROVISIONS) 

GALLEY 74.30 

REFRIGERATOR (14 DAYS) 

FREEZER (14 DAYS) 

AMBIENT FOOD (14 DAYS) 

UTENSI L/APPLIANCE STOW 

TRASH STOW (EMPTY) 

HOUSEKEEPING SUPPLIES 

WORKSTATIONS 

ANC ILLARY PROVI SIONS 

(e.g., WRITING EQUIPMENT, FILM, 

CAMERA EQUIPf4ENT) 

SHOWER/HANDWASH/WMF (PERSONAL HYGIENE ) 28.50 

TISSUE DISPENSER 

PERSONAL WETWlPES 

TOWELS 

WASHCLOTHS 

LOGISTICS MODULE 528.00 

CLOTHlNG (14 DAYS) 

FOOD !90 DAYS) 

BEDDING (90 DAYS) 

FOOD TRASH STOWAGE (EMF fY) 

ADDITIOAAL STOMAGE 

STOWAGE CONTAINERS I 4DDITIONAL) 426.54 

EQU IPdENT CONTAI NERS/RESTRAINTS 11 .OU 

MAINTENANCE WORKSTATION 20.00 (PART OF WORXBENCH) 

WARDROOM 20.00 

ADDITIONAL HAB 1 90 .OO 

AODITIONAL HAB 2 60.00 

ADDITIONAL LAB 1 112.00 

ADDITIONAL LAB 2 112.00 

TOTAL 1200.00 

(2.222 CU FT/!'%i/'ilkY 

(5 LBS/CU FI -1 EMPTY 
4.3.2.1.7 Laundr) facility 

A laundry facil ity has been provided in Hab module 2. The reference 

configuration concept provides for a washer and dryer. The presence of a 

laundry facility onboard Space Station substantial 1 y reduces the amount 

of clothing necessary for a 90-dqy mission. With a laundry facility, 

only a 14-day supply of clothing is rlecessary. The clothing suppl) 

weighs 185.4 Ibs for a crew of 6 and occupies 63 cu ft of volume, while 

the washerldryer reighs TBD and occupies TBD ft3 of volume. Uitnout a 

laundry facility* a 90-day supply of clothing is necessary. This weighs 

1140 lbs for a crew of 6 and occupies 388.35 cu ft. 

4.3.2.2 Nonliabi tation Volume 

There is additional vol urne for crew accomnodations in the non-habi tai ton 

areas. This volume includes the two labor,' : ry modules, the airlocks, 

the logistics module, and safe haven. 

4.3.2.2.1 Crew accommodations in laboratory volume 

The crew accmodations in the laboratory volume aid in crew health, 

safety, and well-being (e.g., 1 ighting, windows, handwashers), and they 

support the crew in providing customer services (e.g., workstations, 

stDwage volume). Figure 4.3.2.2-1 depicts the internal configuration of 

Lab 1 (Life Science5 Lab), and Figure 4.3.2.2-2 depicts the internal 

configurdtion of !db 2 (Materials Processing Lab). Most of the internal 

volume of Lab 2 5s cmprised of racks for materials processing. The 

allacation of racks for specific types ~f processes is depicted in 

Figures 4.3.2.2-3 and 4.3.2.2-4. Crew accommodations in this Lab consist 

of a handwasher ias described in 4.3.2.1.4, Personal Hygiene Areas), 

cmunications equipment, and two operations stations (as described in 

4.3.2.1.5, Workstations!. 

Lab 1 contains faciiities and accomn~dations primarily for Life Sciences 

experiments. Powever, there are some racks allocated for Astrophysics 

experiments. Thdt portion of this module dedicated to Life Sciences 

experiments consists of the Human Research Facility (Pr.F) and the 

Animal/Plant Research Facility (APRF) . Cohabiting animals and humans in 

the same module requires special accommodations since each group has 

micr org~nisms that could adversely affect the other. Therefore, the 

ECLSS 'or the animals habitat must be separate from that provided for the 

crew. The environment i? which animal research is done should also be 

separate from the environment of the Space Station. Therefore, a 

"biological lock" isolating the APRF from the rest of the Space Station 

is required. A room for changing clothing and/or donning cap, gown, 

mask, and gloves is provided at the entrance of the APRF, as is a shower 

or handwasher (TBD). The ARPF occupies 670 cu ft. A floor plan is shown 

in Figure 4.3.2.2-5. Equipment includes racks, a surgical table, and 

biolock/shower. The nonhumar: Life Science comnunity has 1 isted a 1G 

centrifuge as part of their experimental equipment. The inclusion of 

such a centrifuge alters the internal configuration of this lab module as 

indicated in Figure 4.3.2.2-6. The ECLSS to support the ARPF is sized 
i 

HUMAN RESEADCH RACKS 

/ , /---ANIMAL ECLSS 

'iPOlJER EQUIPMENT 

SAFE IfAVErl 

\ ,-DATA MAIIAGEME

I / --- 

API I MAL - 

' / STAT I ON -ASTROPI IYSI CS RACKS AN I MC,L 

ECI-SS RESEARCH 

- TREADMILL LAB 

EOUIPPEr'T 

Figure 4.3.2.2-1 - LIFE SCIENCES LAB. 
\.NITER1 ALS PROCESSIllG RACKS 

DATA fANAGEMElIT CQUI PMErlT 

DATA 

. 

4 

& c. 

z 15 

,' 

. . 

ECLSS OXYGEII 

EqUI PMENT SAFE --" 

IiAVEII 

Flgura 4.3.2.2-2 - MATERIALS PROCESSING LAB. 
/ 

DATA t4ArlAGE!IEfIT EQU

Flgure 4.3.3.2-3 - MATERIALS PROCE881NQ LAB. 
- DATA MANAGEMENT EQUIPMENT 

r CAMERA 

I 

-CNEM 

/ /-- ECL

VRPGR CRY S'I'AL q!?C?'zrTH 

8 ISCTHERMAL Fd?pJACE 

HI GH GRAM EIiT 

CRY ST8. GROWTIIFICEF PULLiIiG - 

( ACOtJSTI C) & 

3 AX;T /'.COUSTIC 

L' 

ECLSS-J - COMMUFII CPTIONS 

ELECTROEPITAXIAL 

I 

SUPP9RT/TEST/EVl\LUATION 

CRYSTAL GROWT! i 

ELECTRO-IWGtIETIC & ELECTRO-STA

Flg~~re 4.3.2.2-4 - MATERIALS PROCESSINQ LAB. 
Flgun 4.3.2.2-5 - ANIMAUPLANT RESEARCH LAB. 
Y- HUMAN RESEARCH RACKS 

POUT R ;I PINT 

SAFE HAVEN 

1 ECLSS OXYGEN - THERMAL 

MASKS EQUIPMENT STATION 

-S

Flgun 4.3.2.24 - LIFE SCIENCES LAB WITH CENTRIFUGE. 
for four crew capabil ity; accomdating two mission specialists and TBD 

animals (rats, mice, small primates). It requires 375 cu ft of volume of 

which approximately 245 cu ft is equipment volume with the remainder 

being access volume to this equipment. The HRF consists of racks to 

support experimental equipment (Fig. 4.3.2.2-7) and a treacbnill for 

reselrch purposes. In total, it occupies 462 cu ft of volume. In 

addition, there is an operations station with racks to accomodate data 

management. 

4.3.2.2.2 -- Airloc!is/hyperbaric - chamber 

iwo EVA air locks, one with hyperbaric chamber capability, are planned for 

the Space Station. Details are presented in paragraph 4.4.9.2.3 EVA 

Airlock. 

4.3.2.2.3 Safe haven provisions 

Safe haven provisions dre located in each of the four habitable volumes 

(both Hab moduies and both Lab modules). It consists of a 22-day supply 

of food, dry and wet wipes, clothing, sleep restraints, waste, and trash 

stowage. Such a suppiy, sized for two cr.eumembers, is stowed in the 

module segment containing hatches and mechanisms and occupies 50 cu ft of 

each of these segments. Thus, if any one of the four modules must be 

evacuated and isolated, the remaining three modules contain safe haven 

suppl ies for six crewmembers. 

4.3.2.2.4 Crew accommodations in the logistics module 

The logistics module contain; a 20-cu ft refrigerator, a 60-cu ft freezer 

and 528 cu ft of stowage containers to support the crew for 90 days. 

There will be a 90-day supply of food and additional gitlley provisions, 

housekeeping supplies, clothing, bedding, personal hygiene items, 

equipment containers and restraints, ancillary provisions (batteries, 

writing supplies, paper, etc.) , and resupply items for the HMF and 

maintenance workstation. A1 1 of these accommodations occupy 1,675 cu ft 

of vol ume ( equipment, suppl ies and access vol ume incl uded) . A breakdown 

sf these accmc!. ,ions is presented in Table 4.3.2.2-1. 

4.3.2.3 Crew . acili .; Equipment 

The family of items required to outfit modules for habitation is defined 

herein. Such facil i ty equipment includes secondary structure, hatches, 

,mchanisms, util i ties (e.g., power, data/camn., thermal, and ECLSS! , 1 ighting, windows, and maintenance facil i ties. Crew Consumables and 

habitability provisl ns are not included in this paragraph. 

4.3.2.3.1 Secondary structure 

Secondary structure in the modules consists of walls, floors, crew 

transfer tl.~nn?l s, and framework for retaining and positioning stowage 

containers, habitability subsystems, utilities, crew workstations, and 
DISPLAY 

3ROP7RTII)IIAL 

C3U!lTEP. 

ELECTf?r)!lICS 

COMPUTER 

LSLE CE!iTRI FUGE 

, 

TV CAMERA C0Nr.l 

PNIEL 

VTR S!lIfCIl 

STORAGE 

LSLE ECI!O 

ECIK' SCMII;ER 

SCA'1'IER 

Cn!!VE!?TER 

EPSP ACCESS 

STORAGE 

IIORKBE I I CI 1 

RE FRI GERATOR 

FREEZER 

VIDEO RECORDER 

I 

VACUUM 

ELECTROFIICS 

I/F 

STORAGE DRAWER 

VIDEOMOFIITQR 

LSLE IIICRO 

DISC ZTORAGE 

VIDEO DISPLAY 

STORAGE 

HARD COPY UNIT 

CV/CP I/F PANEL 

CONTROL UNIT 

LSLE STRIP CliART 

GAS TAflK 

ASSEMBLY 

BAG- Ill-BOX 

LSLE GAMS 

GAHS ACCESS 

GAS CYLINDER 

EPSP ACCESS 

EPSP 

CARDIOPULMOIIARY PAY LO \D FL 

DATA F AALY S 

FILES 

DP,OP LEAF DESK/ 

TERMINAL 

SMIFiG OUT CIIAIR 

P/L FLIGllT DATA 

COMPUTER 

DISC SYZTEP! 

r PSP 

EPSP 

4 

I CRS I

INTERCOM 

BSTA 

TV CAMERA COr1P.I 

VIDEO ANALOG 

/ / I 

/ 

S t,IM I 

NECK CIIAMDER: 

ELECTRON I CS 

PRESSURE 

SYSTEf.1 

ALBrlP CONTROL 

COP 

STORAGE 

COMTAIMERS 

LON- C: 

I 

CENTRI FUGE 

EPSP 

S MUSCULOSKELETAL 

Flgun 4.5.2.2-7 - HUMAN RESEARCH AREA AND ASTROPHYSICS AREA. 
TABLE 4.3.2.2-1 - LOGISTICS MODULE 

MAN-SYSTEMS - CREW SUPPORT AND ACCOMMODATIONS EQUIPMENT 

EQUIP EQUIP 

SIZE TOTAL TOTAL 

NO. POWER LxMxH WEIGHT VOLUME 

XTEMS (WATTS) (FT) (LBS) (CU FT ) 

UTILITIES 

SEC. STRUCTURE 

HATCHES 1 .OO 

LIGHTING 

ECLSS 

FREEZER (EMPTY) 

REFRIGERATOR (EMPTY ) 

STOWAGE CQNTAINERS 264 

RESUPPLY FOR 90 DAYS TOTAL* 

* ITEMIZED RESUPPLY [ ALL GFE) 

CLOTHING 

PE 3SONAL HYGIENE 

EQPT. CONTAINERS/RESTRAINTS 

HOUSEKEEPING SUPPLIES 

ANCILLARY PROVISIONS 

GALLEY 

FOOD 

HMF 

MAINTENANCE WORKSTATION 

TOTAL LOGISTICS 9525.17 1120.19 

COSTING 

C F CFxWT WEIGHT 
payload facil i ties. Secondary structure material s include metal, 

fiberglass, fabric, and screen devices. 

Secondary structure within the launched module will be sufficient to 

withstand launch and crash loads. Once on orbit, such structure may be 

modified as required for support of new payloads and interior 

modification and/or growth. Should the module be returned to the ground, 

the secondary structural design must withstand 1 anding 1 oads. A1 1 

secondary structure is designed to permit access to equipment for 

maintenance considerSations. 

Secondary structure total weights for each module are 3854.3 lb (Hab 

Module I), 4756.0 lb (Hab Module 2), 2937.6 lb (Lab Module I), 2937.6 lb 

(Lab Module 2), and 1873.3 lb !Logistics Module). Sec~ndary structure 

volumes for each noaule are 74.22 cu ft (dab Module 1) and 126.42 cu ft 

(Hab Module 21, respectively. ather volumes are either considered as 

part of encompassing volumes or are volumes yet to be defined. 

Hatches 

Hatches are provided on a1 1 operational ports and crew airlocks. They 

are hinged and stowed such thdt crew and eqciprnent can traverse through 

the open ports as required. However, during such times, each may be 

closed t~ serve as a pressure bulkhead between msdules. 

When the radial hatches are opened, it is envisioned that they wjll move 

inward towdrd the center dpproximately 1 foot, then move back to the end 

cap and then dre restrdined. End ports or axial hatches rotate 90' down 

and are restrained due to operational volume constraints caused by 

utility line penetration. 

Hatch size is defined by mdintendnce and changeout requirements, 

irxluding the requirement to transfer a single !e.g., Spacelab-sized) 

rack from the Orbiter to its use location in the Space Station through a 

port, passageway, tunnel, etc. Hatches may be larger if design indicates 

the need. Sufficient volume around the hatch operational envelope is 

provided to prevent unusual body contortions or major reorientations of 

crewmembers when traversing through passages (other than turning the body 

long axis perpendicular to the plane of the hatch opening). Hatches are 

conceived to be capable of supporting either an open-or-closed hatch mode 

with manual contrc! requiring a maximum of 30 seconds. 

Hatcn weight has not been determined. However, the minimum diameter of a 

typiczl hatch is 50 inches, and volume (assuming 50 in. dia, 12 in. thick 

hatch) is 10 cu ft. There are six (four radial and two end hatches) each 

identified in all modules except the logistics module, which has only two 

end hatches. 

Crew mechanisms defined in this paragraph are limited to those which 

re1 ease/l ock the hd tch to the port. Weights are undefined. However, 
362.7 cu ft of volume is allocated, which incorporates the hatch and crew 

opera tisnal envel ope. 

Other crew-re1 ated mechanisms will be defined as Station definJ tion 

evol ves. 

All modules have selected utilities passing through theni (Figs. 

4.3.2.3-1, 4.3.2.3-2 and 4.3.2.3-3). Penetrations around a radial port 

and the opposite axial port permit entrylexit of the majority of the 

utilities. Additionally, twin penetration panels on the side bulkhead of 

each mdule suffice as other passages for utilities. 

Systems which require niidule penetrations include enviromental control 

(ECLSS), power, thermal,, housekeeping data, payloads data, and crew water 

(drinking, waste, wash). Line definition for the ECLSS includes two 

4- in. -di a l ines penetrating through the bulkheads, and expanding to 

6-in.-dia ducts. Air flow on one line provides supply to the mdule, 

while the other 1 ine is used for collecting exhaust air. The 

housekeeping bundle, an optical data network, ccnsists of six 1/2-in.-dia 

1 ines accommodating data acqui si tian/tel emetry , thennal data, control and 

navigation status, audio control and signal data bus, power distribution 

control bus (25Kw, 200 VAC), and a miscellaneous data bus. Internal 

util i ties entering/exi ti ng through the two bulkhead panels incl u& dual 

1-1/2 in.-dia coolant supply and return lines, dual 1-in.-dia lines for 

drinking water, for waste liquid water, condensate water, and wash water. 

A1 -a included are dual 3/8-in.-dia 02 supply and 1/2-in.-dia Y2 

supply lines. The thennal system has dual liquid water loops and a 

partial vapor loop passing throagh a heat exchange and a heat source. 

Volume calculations include 7.3 cu ft for ECLS, 35.3 cu ft for the 

optical data network, 17.6 cu ft for internal utilities, and various 

volumes for thermal (Hab 1-10.8 cu ft, Hab 2-10.8 cu ft, Lab 1-30.0 cu 

ft, Lab 2-19.1 cu ft, Logistics module -7.3 cu ft) insulation. Ueight 

calculations is a function of materials selected for lines. 

4.3.2.3.5 - Lighting 

The 1 ighting system provides both intericjr and exterior ill uination 

based upon appropriate htaaan factors specifications. Consideration has 

been given to luminaires location, brightness ratio and contrast with the 

surrounding area, glare, and shadows. Care has been given in locating 

switches to facilitate convenient operation. Night light route locators 

and switch illumination are placed in frequently darkened areas. 

Internal lighting as defined li given in Table 4.3.2.3-1. 

Y i ndows 

Uindows are located throughout the mdule cluster to facilitate work and 

recreational tasks. Each window is sized and located to permit viewing 

by two or more persons at a tie. Window workstations pemit, as 

required, shadowed viewing and restraint for crewmembers and equipment. 

Specific equipmerot which will be acconrrodated at each window workstation 
Flgun 4.3.2.3-1 - VIEW OF MODULES AND EXTERNAL UTILITY LINES. 
r DATA TRAtlSFEI: LIIlE 

ECS DIFFUSER 

\DATA TWISFER LINE 

-ECS CIFF

PO

ECS 

DATA TWISTER LIVE 

ECS RETRIEVAL 

FI~UTO 4.3.2.3-2 - CRDSS SECTION VIEW 01 OF MODULE AND INTERIOR UTILITIES. 
DATA T

EC

ECS RETRIEVAL- \---- POI-IER '--DATA TRANSFER 

Flwn 4.3.23-3 - CROSS SECTION VIEW U2 OF MODULE AND INTERIOR UTILITIES. 
I-LL 

Z 
include munted voice tape recorder, event timer, means to mount caneras, 

means to secure handheld cameras, mall 1 ight, method to secure paper and 

checklists, writing stotion, body restraints, CRT monitor and keyboard, 

naps, iaoving nap display with an optical driver to view the flight path, 

optical maps to identify future flight paths, nethod of measuring angles 

and the horizon if appropriate, control of adjacent 1 ighting, and easily 

deployed hood or curtain to block interior light. Baseline windows are 

not planned to be of optical quality or to be UV transmissive. However, 

some techniques will be used 50 verify that window transparency will bc 

maintained. The current configuration identifies Go 20-in.-dia and fc-tr 

i0-in.-dia windows in each of the two habitat modules. Four 10-in.-dia 

windows are located in the two laboratory modules. 

4.3.2.3.7 Mointenance Facilities 

The crew will serve CIS the main capability for accanpl ishing hardware and 

software maintenance. This capability will be enhanced via the 

maintenance workstation. Included in the workstation wii l be equipment 

permitting simple diagnosis of circuitrj, repair of hardware, and/or 

rep1 acment of el ements of a subsysten. A1 though the primary maintenance 

technique will be the changeout of Orbital Replaceable Unit (ORU's), the 

maintenance workstation wi 11 provide supplemental capabil i ty to diagnose 

and repair to the el mental level as appropriate. Additional items nay 

be fabricated frm components of failed items or from raw materials. 

Control led 1 ighting and magnification capabil ity will be provided. A 

dedicated microprocessor/screen/printer sill serve as a diagnostic tool, 

suppl mented by diagnostics with handheld instnments. Electrical test 

equipment, maintenance tool kit (powered and manual 1, and a sol derlng kit 

will be available. Suppqrt equipsent include crew equipment and small 

item restraints, a particle control system, measurement kit (tape, 

micrometer, thread Sages, etc. ) , fastener kit (nuts, bolts, washers, 

etc. 1, and an adhesive kit (tape, cement, etc.). 

A special feature of the workstation will be a portion that is removable 

and which converts into an EV compatible portable maintenance 

workstation. This de. ice can be transported, along with attached 

diagnostic and repair equipment and spares, to any location on the 

Station requjring maintenance, diagnosis, or repair. It will be capable 

of being positioned and rigidiy restrained at the worksite. Tools, 

diagnostic and repd i r equipment, and a computer-enhanced, audio-visual 

check1 is*, will be easily accessible to the cre-er. 

Estimated dimensions for the basi ,maintenances workbench are 19in. by 57 

in. by 84 in.) (92.3 cu ft); weight is estimated to be 238 lb. 

4.3.2.4 Volumetric Analysis 

4.3.2.4.1 Introduction 

A cmprehensive account of volumetric utilization was undertaken for the 

reference configuration. Assessments for dl1 systems and subsystems 
been included to the extent that current definition is available. Table 

4.3.2.4-1 shows volumes within the cylinder as they will be discussed in 

subsequent sections. 

4.3.2.4.2 Ailocations 

4.3.2.4.2.1 End cone volumes 

For the purposes of this analysis, er~d cone vollnnes are included only to 

shoe they are primarily occupied by utilities and hatches. The available 

end cone volume not occupied by this equipment is considered access 

vol unte for hatch operation and util i ty 1 f ne servicing and maintenance. 

4.3.2.4.2.2 Connecting end volumct 

A 9-ft long section ot the 30 ft cylindrical axis length is utilized for 

module interconnection where four radial ports are located SO degrees 

apart. Because gf the complexity of the connecting end volume, the 

1257-cu-ft vol urn contained therein is sumnarized separately (Table 

4.3.2.4-2) from the rema'ning cylindrical volume (21 ft of the 30 ft 

cyl indrical axis) . 

Figure 4.3.2.4-1 shows a cut-away vSew of the protrusion of radial port 

mechanisms into the connecting end volume. If a rectangular box 8.5 ft 

by 8.5 ft by 9 ft (650 cu ft, Fig. 4.3.2.4-2) is conceptualized in this 

end, space alsng one side (152 cu ft of volume) is available for 

equipment inst:llation where one radial port is blanked off. 

The volume mtsido the rectangular walls and inside the inner cylindrical 

surface i- dccupied by the radial port module connection mechanisms. 

Remaining volume inside this end "cubeu is utilized as hallway access for 

traverse between modules, and as access volume to the equipment 

(mentioned previously) installed agalt~s; the blanked off port. Some 

residual volume I128 cu ft) is available around the radial port 

mechanisms (inside the inner cylindrical wall and outside the conceptual 

cube wall ) but does not provide a form factor compatible witn 

conventional packaging. While it is certain this 128-cubic-f9ot volume 

wili be eventually utilized, for purposes of this analysis, it is treated 

as nonfunctional. 

4.3.2.4.2.3 Habitable volume 

The remaining 21 ft of the 30 ft cylindrical axls is considered habitable 

volume (2932 cu ft). A conventional layout (in terns of floors and 

ceilings) is util ized in the reference configuration, but the arrangemeci 

and 1 ayout methodology is not re1 evant to volumetric accounting, in that 

equipment vol ume requirements are basical ly constant (given that 

packaging form factors could a1 ter them somewhat). While installation 

methods and alternative layouts may provide more or less access. 

maintenance, and operational volume, the percentage gain or loss will not 

significantly change the fact that volume in HMl and HM2 is minimaily 

sufficient to inadequate. 

116 
TABLE 4.3.2.4-1 - VOLUMETRIC SUWRY BY LOCATION 

CYLINDRICAL AXIS 

(LENGTH IN FEET) 

(INNER DIAMETER = 13 ft 4 in.) 

VBLUME 

(CU FT) 

TOTAL VOLUME 30 4189 

HABITABLE VOI i!ME (21) 2932 

CONNECTING END (9) 1257 

END CUBE (VOLUME WITHIN CONNECTING END) ( 650 

CONNECTING END LESS EdD CUBE (9 ft LONG CYL. LESS CUBE) (607 ) 
TABLE 4.3.2.4-2 - CONNECTING END VOLW ALLOCATION SUWRY 

(ON TO BOTH IUl AND HQ) 

FUNCTION VOLUME (CU FT) 

EQUIPMENT (POWER, SAFE HAVEN, 

THERMAL, ECLSS (PARTIAL ) 

ACCESS 

HALLWAY TRAVERSE 

(OUTSIDE END CUBE) 

MODULE CONNECT MECHANISMS 

NOT UTILIZED 

TOTAL 


Table 4.3.2.4-3 dnd Table 4.3.2.4-4 chow equipment and access volume 

allocations by subsystem for the two hab modules. Access volumes are 

shown adjacent to equipment volu~nes and are in additior! to equipment 

vol umes . 

4.3.2.4.4 Access volume 

Access vol ume requirements Gary according to the equipment being 

util ized. For example, ? drawer , when pul led out and removed from its 

cabinet, requires sliq!~tly more than its own original volume plus enough 

room for someone to stand and perform the removal. A:! operable rule of 

thumb for this co:~di tion requires 2.7 times the equipnent volume for 

equipment pl us dccess vol ume. Howver, access vol ume can sometimes be 

shared by eq .ipment on opposite walls sharing a commcn disle provided 

simultaneous functions are not required in the same <;,ace. In other 

cases, a fartor greater than 2.7 is required if botn vertical and 

horizontal manipulation are required. In the case oY the reference 

configuration for the Space Station, insufficient detail exists to 

ascertain whether simultaneous operations ;,I 11 be performed in the 1 ab 

modules in opposing equipment racks. For the habitation and logistics 

modules it is generally assumed that this will not be the case, and the 

same assumption is general ized to the lab modules. Given these 

assumptions, a factor of 3.0 (or ! .5 for each side) times the equipment 

rack volume along one side of a module is considerea minimally adequate 

to provide space for an individual to operate or maintain the equipment 

and to allow passage during operations. This translates to half the 

depth of the equipment being available for a person to operate in front 

of the equipment and 1.5 times the equipment depth being availabie as 

aisle space for others to pass by with no room left to operate on the 

opposite wall. Alternatively, two people could op3:rate back to back with 

no room between them to allow passage. 

4.3.2.4.5 Conclusions 

The access volume for crew quarters requires that volume is needed for 

movement in and out of the area, and is satisfied by the 50 in.diameter 

tunnel common to the six entrances and the 150 cu ft per crzw quarters 

(which includes both equipment and access volume). 

The galley/wardroom share d common aisle and are conceptually simildr to 

the crew quarters in tens of volumetric req~irer~~errts. In addition to 

providing shared volume for meal preparation, the wardroom provides a 

common mezting location for recreation, station management meetings, 

dining, etc. If gal ley equipment must be removed, the tables can be 

removed to increase the maintenance access. 

Waste management faci 1 i ties, 1 ike the crew quart.ers, util ize access 

volume internally, and require only enough room externally for the door 

to operate. 

The respective external cccess factors of .52 and 1.6 for the oppc-ing 

volumes of hygiene and medical areas pennit resoval of the medical racks 

for maintenance, if required. 
TABLE 4.3.2.4-3 - HABITAT MODULE 1 SUBSYSTEM EQUIPKNT AND 

ACCESS VOLUlrE ALLOCATIONS 

(INSIDE CEILING AND FLOQR) 

WORKSTAT IONS 

EQU1P:EHT RACKS 

ST WAGE 

CREW 

SPARES 

HYGIENE 

GALLEY /WARDROOM 

MEDICAL 

SECONDARY STRUCTURE 

(OUTSIDE CEILING MD FLOOR) 

ECLSS 

LIGHTING 

UTILITIES 

STOWAGE ( SPARES ) 

(CONNECTING END) 

EQPT ACCESS 

(IN CU FT) 

39 

32 

192 

(236 

(48) 

228 

737 

88 

7 9 

SUBTOTAL 1395 

TOTAL 2310 

EQPT ACCESS 

162 ( COMMON ) 

19 ( CWN 

73 (COMMON ) 

331 3 7 

SUBTOTAL 585 3 7 

TOTAL 622 

G~d7 TOTAL 4189 

MIA 
TABLE 4.3.2.4-4- HABITAT MODULE 2 SUBSYSTEM EQUIPMENT AND 

ACCESS VOLUME ALLOCATIONS 

(INSIDE CEILING AND FLOOR) 

WASHER/DRYER 

MAINTENANCE 

DATA MGTICOMM 

CREW QI1lRTERS 

(INCL. 120 CU FT STOWAGE) 

HYGIENE 

WORK STATION 

SECONDARY STRUCTURE 

SUBTOTAL 

(OUTSIDE CEILING AND FLOOR) 

ECLSS 

LIGHTING 

UTILITIES 

STOWAGE 

(CONNECTING END) 

EQPT ACCESS 

(IN CU FT) 

TOTAL 231@ 

110 

203 

( COMMON 

837 

EQPT ACCESS 

162 ( COMMON ) 

29 ( COMMON ) 

7 3 ( COMMON ) 

321 3 7 

SUBTOTAL 585 37 

TOTAL 622 

GRAND TOTAL 4189 
It can be seen ffm Table 4.3.2.4-3, Table 4.3.2.4-4, and the internal 

configuration drawings that, while the desi rable equipment- to-access 

volume ratio of 3.0 is not achieved, there Is sufficient volume for 

operational activities to occur. In some cases, operations plus passage 

by another individual can be accmplished. In all cases, the total of 

access volume factors for the opposing pieces of equipment across the 

aisle wi 11 permit removal /rep1 acement and ma1 ntenance. 

Utilizing the packaging factors from HM1 and HM2, It should be possible 

to realize 70 cubic meters (1059 cu ft) to 45 cubic eters (1589 cu ft) 

of equipment volume per laboratory module, or 60 cubic neters (2119 cu 

ftS to 90 cubic meters (3178 cu ft) total. 

The volumetric allocations of Labs 1 and 2 in the reference configuration 

are not discussed in this section, as the functions and allocations are 

di scussed el sewhere. However, the Materials Process1 ng (Lab 2) module Is 

configured with approximately 18 cublc meters (640 cu ft) of equipment 

volume with the capability to add an approximately equal amount for a 

total of 35 cubic meters (1235 cu ft) to 40 cubic meters (1412 cu ft) of 

equipment volume. Equipment volumetric overhead (ECLSS, utilities, 

workstations, data management, power, thermal, safe haven, 

comnunications, secondary structure, 1 ighting, hatches, etc. i to the 

1 aboratory modules is approximately 21 percent. Access vol urn 

requi rments added to this equipment overhead increases the total 

overhead to approximately 42 percent. 
4.3.3 Space Station dynamics and control analysis 

Analyses of the rigid body control and orbital altitude maintenance 

requirements, f 1 exibl e body control , and flexible body response due to 

typical disturbances are documented in Sections 4.3.3.1 through 4.3.3.6. 

The results should be considered preliminary since details of many of the 

structural and control aspects of the configuration were not available 

during the analysis cycle. 

Sections 4.3.3.1 through 4.3.3.4 deal with input data for the analysis 

while 4.3.3.5 presents analyses results in an attempt to dernonstr3te the 

perfomance of the station in the natural and induced dyna+nic 

environments. A sumry of the results is presented in Section 4.3.3.6. 

4.3.3.1 Appl i ed Forces and Torques 

This subsection contains a description of mdels which have been used to 

generate loading events for preliminary analyses of the reference 

confi guration of the Space Station. 

Aerodynamic and gravity-gradient models generate forces and torques 

acting upon the Space Station as a fur~ction of orbital position and 

vehicle attitude which are used for the purposes of control system sizfng 

and configuration trade studies. Orbiter berthing, crew motion, and RCS 

firings models generate significant loading events which are used for 

control and response analyses and trade studies. 

Due to time limitations and availability of data, the following 

addi tional di sturbance sources were i denti fi ed kut not incl uded in thi s 

study. 

Manipul ator Dyr~amics 

Construction Dynamics 

Manufacturing Dynamics 

Rotating Machinery Dynamics 

Shuttle RCS Plume Impingenlent 

Crew Motions During Nominal Operation 

Momentum Management schemes1 

Solar Pressure 

These loading e.cnts are not considered to be an exhaustive list. 

However, taken together with the ones which were studied, this set is 

belJeved to be representdtive of the set of disturbances which are 

significant for dynamics and control systems studies. 

l~ee White ?apers "Space Station Fluidic Momentum Control ler, 'and 

"Space Station Large Area Magnetic Torquer.' 
4.3 3.1.1 Gravi ty-gradient to-s 

Gravity gradient torques act upon any satellite with a mass distribution 

which t-esul ts in unequal principal axis inertias. 

Two of the principal axis inertias for the reference configuration are 

approximately equal in magnitude. The minimm axis is half an order of 

magnitude smaller and the configuration will fly with the minimum axis 

$1 igned within a few degre~s of nadir. 

To compute the gravity-gradient torques, the mass moments of inertia in 

body axes ars calculated from the geometric definition of the reference 

configuration (see mass properties in Tables 4.3.3.4-1 through 

4.3.3.4-41 Assuming a uniform gravity-field, the orientation of the mass 

momen:s of inertia in the gravity field is used to calculate the 

resultant torque acting rjpon the Space Station. 

The detail for the model of gravity-gradient torques is contained in 

References 3 and 4. 

4.3.3.1.2 Aerodynamic draq and toraues 

The referencr configuration util izes large-area solar arrays to generate 

power and large-area radiators to dump excess heat energy. These arrjWs 

and radiators are control led with alpha and beta rotary joints to track 

the Sun. As a result of the solar-tracking, the drag area of the station 

varies as the orbit is traversed. 

The reference configuration which is not geometrically symnetrical about 

f ts center of mass and is flown wf th its x- axis in the orbit plane (see 

Fig. 4.2.1-1) with the power bocms perpendicular t.,~ the orbit plane. The 

aerodynamic torques about the x- and z-axes would be cyclic with respect 

to inertial space if the atnospheric density was constact and 

nonrotating. However, since the atmosphere rotates with the Earth and 3 

is not constant (Figure 4.3.3.5-ll), secclar torques occur about these 

axes. The aerodynamic torque abwt the y-axis is due to center clf mass 

to center of pressure offset and wi I1 result in secular torque if i, is 

not nulled with gravi ty-gradient torque. 

Drag area of the reference configuratio~ is caiculdted by a solid area 

projection which includes the effects of orbital position and shading. 

Aerodynamic forces are ca1cu;ated 3s a function of the area, the local 

atmospheric d2nsi ty and the center-of-mass velocity with respect to the 

atmosphere. Center of pressure offset with respect to the center of mass 

is then used to calculate the aeradynamic torques. 

the detail for the model of aerodynamic drag force and torques is 

contained in References 3 and 4. 
4.3.3.1.3 Atmosphere variation - mode: 

The nominal altitude range of the Space Station is from 250 to 270 mi. 

Control system sizing studi2s were performed at altitudes of 220, 250 and 

270 nai. At altitudes in the range of 100 to 300 nmi, the atmospheric 

denslty depends upon the degree of solar activity and the diurnal effect 

as well as the altitude. The atimspheric density mdel is a camplex 

function of Space Staticn altitude and Iccation relative to the incoming 

Sun-1 i ne. 

Tlre nominal altitdde range of the Space Station 9s from 250 to 27G nm:. 

Control sjstem sizing studies were performed at a1tit:ides of 220, 250 and 

2712 nmi, At altitudes in the range r)f 100 to 300 nmi, the atmospneric 

density depends upon the degree of solar activity and the diurnal effect 

as well as the a1 ti cuCe. The atmospheric density mdel is a complex 

function of Space Stati~n altitude and location relative to the incoming 

Sun-1 ine. 

Recent empirical models of the upper atmosphere density, based on drag 

data from various satel 1 i tes (reference 5) and direct mass-spectrometric 

measurements of upper atmosphere (reference 6) have been included in this 

anaiysi s. 

These atmospi1zr4c acdels reflcct the observational evidence for both lrjng 

and short-term variations In upper atmcsphere density, composition, and 

temperature. 1-ong-torn! variatims are dominated by the 11 year solar 

cycle (Figure 4.3.3.1-1) related variations in the extretne ultraviolet 

(EUV) radiation from the Sun and semianncal and seasonal variations. 

Short-te:m variations are dominated by the diurnal variation caus2d by 

the daylight side heating of the Sun and variations associated with 

energy inputs related to geomagnetic activity. The statistically 

predicted solar flux indicatcrs, Figure 4.3.3.1-1, as well as the 

geomagnetic index, Ap, were obtained fi-om MSFC Mission Analysis 

Division. The 2 sigma peak F10.7 values of 242 and correspcnding 2 signa 

Ap value of 22 was cnosen for the analysis to be used to size the CMG 

requirements. 

On rare occasions, extreme solzr magnetic actit ity results in short-term 

(less than 12 hours) atmospheric densities three times as !argc as the 2 

sigma case. Present plans calls for controlling the Station with the RCS 

during these occasions if necessary. 

The following assuwptions were used to derive a qodel for the disturbance 

to the Space Station contrcl and stabilization system due to berthing of 

the Shuttle Orbiter vehicle: 

- Significant forces along the x berthing axis only 

- Stdtion initially at zero rate stato 
- Orbiter closing at (3.1 ft/sec rate along berthing axis 

- Co~stant force berthing mechanism 

- 0.5% di splacenent of berthing mechani sw 

The beimthing disturbance is modeled as a W-ib force acting along the 

berthing axis for a period of 1 second. This is considered to te a 

conservati ve mdel for design purposes. Although dockiti? forces 

experienced an SKYLAB were an order of mgnitude higher, it is 

anticipated that the berthing operation ray actually produce forces which 

2re an order of magnitude be1 w the design mdel level . 

The time history for the berthing force mdel is given in Figure 

4.3.3.1-2. 

The z-axis moment arms for the reference configuration in body 

coordinates are 78 ft without pqyloads and 120 ft with pqyloads. 

4.3.3.1.5 - Shuttle launch -and landing load factors 

Load factors for Shuttle launch and lan&ng are given ia Table 

4.3.3.1-1. Load factors for intermediate flight events are given in 

Reference 3. These load factors wcrae swcified for preliminary 

structural design purposes and are included nere for completeness, 

although they were not required or used for dynamics and control 

analyses. 

4.3.3.1.6 -- Crew mtion (Xick-offistopping) 

The following assumptions were used to derive models for the df stuvbance 

to the Space jtatian control system due to on-board crew rovemnt: 

- Station initially at zero rate state 

- Free-flight distances ef 30 ft and 10 ft 

- 25-1 b peak force 

The first crew mtion model is 8n end to end traversal of a habitat 

module. The force generated by the crew member increases 1 f nearly to a 

maximum of 25 pounds over an interval of 1 second and tkn decreases to 

zero as wal i contact is 1 ost. At the opposing wall, the force increases 

instantly to 25 lbs as the wall is conticted and decreases lfneiirly to 

zero over a 1 second interval. The time between force pulses of 13 

seconds is calculated from a free-flight di stance of 30 T-zt. 

The tinrr- history for this crew motion model is giver! in Figure 4.3.3.1-3. 

The force is applied alo~g the x-axjs, a,.: the mmnt arms for the 

referenc configuration are 78 feet dithout pwloads ana 120 fWt with 

pqyloadr. Uhen the Orbiter is berthed to ttK statiorr, the z-axis nolllent 

arms become 25 feet without pwloads and 59 feet with pqyloads. 
The time i~istory for this crew motion model is given in Figure 4 3.3.1-3. 

The force is appl ied along the x-axis, and the nnlnent arms for .s 

reference configuration are 78 ft without pqyloads and 120 ft with 

pqyloads. When the Orbiter is berthed to the station, the z-axis mment 

arms become 25 ft without pqyloads and 59 ft with payloads. 

4.3.3.1.7 RCS re-boost firing ddsturbance 

The following assumptions were us~d to derive a mdel for the disturbance 

to the Space Station control and stabilization system during a reboost 

firing of the RCS thruster. 

The Swtion is initial Iy at zero at~glilar rate state. The RCS control 

 stern will off-modulate focr jets to hold the attitude of the Station 

within a + 1.0" deadband while applying the desired Delta-V along the 

flight paB. Eventually, the Station will reach a limit condition about 

a 1.0 degree bias or attitude error. The time history for the RCS firing 

disturbance model is given in Figure 4.3.3.1-4 for the initial 500 

seconds of reboost. 
TABLE 4.3.3.1-1 - SHUTTLE LIFT-OFF AND LANDING LOAD 

FACTORS FOR PREL IHINARY DESIGN 

UWD IN6 LANDING 

(9.6 FT/SEC SINK (6 FT/SEC SINK 

LIFT-OFF UTE 1 UTE ) 

1 

TIME, SEC 

Figure 4.3.3.1-2 - ORBITER BERTHING FORCE 

TIME. 13 

SEC 

Figure 4.3.3.1-3 - CREW MOTION DISTURBANCE 

132 

4.3.3.2 Control System Definition 

4.3.3.2.1 Control monlent gyro 

A proportio~al ,integral , and differential gain ('PID) type control method 

was chosen for the Space Station attitude control system. This method of 

control is relatively simple and is well within the state of the art. 

The integral gain is especially useful in nulling very small pointing 

errors. 

There are several methods for calculating PI3 gains. However, only one 

will be discussed here. Figure 4.3.3.2-1 represents the Attitude Control 

System (ACS) rigid and flexible body dynamics for a single body fixed 

control axis. The following definitions apply: 

TD = disturbance torque (N.m) 

Tco, = torque applied by control actuators (N.m) 

TR = resultant torque as seen by the vehicle (N.m) 

I = vehicle inertia about the body axis (~g.rn2) 

Kbi = model gain of the i 'th flexible body mode (unitless) 

Wi = frequency of the i 'th flexible body mode (Rad/s) 

'i = dampf ng ration of the i 'th flexible bo4y mode (unitless) 

Gp = proportional cgt~trol gain (N-m) 

GI = integral control gain (N-m/s) 

GD = differential control gain (N.m.s) 

S = La place operator (l/s) 

Tc = commanded torque (N-m) 

m = tota t number of flexible body modes (unitless) 

Some simplifying assumptions have been made to aid in analysis: 

a. The range of bending mode resonant frequencies was considered to 

be low enough to allow the sensor and control actuator dynamics to be 

neglected. 

b. The body axes are assumed to coincide with the principal axes. 

Thfs assumption decouples the body axes and justifies a planar model. 
c. The sample rate was high enough to a1 low the system tc be 

considered as continuous. 

d. The control sensor and control actuators have been colocated. 

This final assumption allows the model gain to be written as 

Where $i is the normalized slope of the i 'th bending mode at the sensor 

location and Mgi is the corresponding generalized mass. 

Temporarily ignoring the flexible body ddynamics the rigid body transfer 

function from disturbance to command can be written. After 

simplification the transfer function is 

Setting the characteristic equation to zero and dividing by the vehicle 

inertia gives 

Gain calculations are made by setting the above equation equal to the 

product of a second order equation dependent only upon the control 

parameters and a first order equation determined by the integral gain. 

Thus 

Where 

w c = control frequency ( rad/sec ! 

5, = damping ratio (unitless) 

W1 = integral 1 oop frequency (rad/sec ) 
After mu1 tiplying and equating coefficients of 1 ike powers, the fol lowing 

relationships can be derived: 

GD= I(ul f 25 w) C C 

G = I(W~ + ZS~W~W~) 

P 

- 2 GI - IucbJ1 

In practice, it is comnon to assume the control and integral frequences 

to be related by: 

where K is a constant 

If this substitution is made, the gain equations become 

Generally the control frequency of the rigid body should be kept 

sufficiently low so that there is a wid? separation with frequencies of 

other degrees of freedom. For the purpose of this analysis, the first 

mode elastic body frequency will establish the upper 1 imit on the control 

freque~oy of the rigid body. The control system parameters must be 

chosen such that the transient response to a disturbance does not impose 

excessive torques on the spacecraft (as would be the case for a very fast 

responding system) and yet not so "sluggish" that large attitude errors 

accumul ate. 

a?
The rigid body control frequency was selected to be one decade below the 

first elastic bo@ mode. This is sufficiently below the first mode so 

that strong coupling does not exist and unreasonable settling times do 

not result. Further, the values of K and cc for each configuration were 

selected to be 0.1 and 0.707 respectively and the elastic body damping 

ratio was chosen to be 0.005. 

4.3.3.2.2 React~on control system -

The RCS is used primarily to maintain attitude control during reboost 

maneuvers, but also serves as a backup to the CMG system. The RCS must 

counteract disturbance torques due to thruster mi sal i gnment, center of 

gravity offset and thrust mismatch between pairs of thrusters used to 

impart velocity changes during these maneuvers. Since the pitch 

thrusters are used in pairs for the reboost maneuver, they are 

off-modulated for attitude control. Figure 4.3.3.2-2 shows a block 

diagram of the RCS. The following nomenclature is used: 

E - - Attitude Error 

K D - - Compensator Gain 

D t3 - - Control System Deadtand 

h - - Hysteresis 

T L - - Control Torque 

TD - - Disturbance Torque 

J - - Space Station Inertia 
TD - + - 1 SAMP

/ - 1 

IS* t HO

n C 

W - I: Kbin 

&3 J 

i=l S' + 2viWi + wi2 

I 

Figure 4.3.3.2-1 ATTITUDE CONTROL SYSTEM 

4.3.3.3 Dynamir '!adel and Free Vibration Characteristics 

The purpose of this section is to present a preliminary analysis of the 

flexible body dynamics of ttie reference Space Station. Since the Station 

must accmmc~date growth, both 30C and growth configurations are 

addressed. Primary consideration i~ given to the I9C Statiotl with the 

results thereof p;*ovi di ng a reference for considering other stages of 

construction and growth. Furthermore, the IX Station is considered with 

and without payloads and/or attached orbiter and with the solar arrays 

oriented either normal to the flight path or normal to the Nadir. 

For both the IOC and grbw~h configurations, free-free vibration 

frequencies and modal shapes are presented. Transient responses due to 

nrhiter berthing, crew motion, RCS firing and certain MRMS functions are 

also presented in Section 4.3.3.5 . All analytical results were derived 

using the MSC NASTRAN structural analysis code of Reference 8. 

4.3.3.3.1 Andlytical model of :OC station -

The IOC Station is showr, in Figilce 4.3.3.3-1. Nine-foot bay square 

trusses as discussed in Reference 9 are employed for the lower, 

transverse arid upper booms, keel extension, and lower and upper keels. 

An analytical finite eleirent model using MSC MSTRAN was developed 

employing beam type members to produce a continuou~ beam repr?sentation 

of the discrete member box truss. Considerable literature pres~ntly 

exists for such continuous beam representations (e.9. References 10 and 

11). However, for the particular truss construction of the reference 

confi guration, equivalent continuous beam stiffness properties do psi 

exist in the literature. Nevertheless, some similar trusser have been 

examined in the literature. These stiffnes!; propertie; take the form, 

EI = (CEI) b2 (EA) 

where b is the dimension of the cubic repeating truss bay and EA is the 

axial stiffness of every truss member. Upper bounds for the coefficierits 

CEI and CGJ may be derived by assuming that the cross section battens are 

infinitely tiff. These yield values of CEI and CGS of 1.18 and 0.35, 

respectively. 

Recently, through a personal correspondence with the authors of Reference 

10, formulas fnr the stiffness properties of the box truss beam with 

flexible uattei,s were developed. This led to values of CEI and CG; of 

1.07 and 0, -especti vely . Unless otherwise stated, these 1 atter 

values were ,-d for the NASTRAN finite element beam modei . A sumtary of 

propertrier used for the truss heam representation is provided in Table 

4.3.3.3-1. 
The continuous beam representation tor this truss also revealed some 

torsional bending coupling 3s well as some higher crder beam effects. 

For the preliminary nature of this study it was considered appropriate to 

neglect these effects. They may be included in future studies. 

The solar arraj blankets of the IOC Station were not modeled in detail 

except for their mass which was distributed along the arrqy mast which 

supports each Erray. However, approximate frequencies for the b1snket.s 

are pi'ovided herein. The solar Lrrqys are assumed to be supported by 

30-inch diameter double laced continuous lnngeron coil aule masts whese 

properties are given in Table 4.3.3.3-1. The radiators are assumed to be 

supported b;! 2-foot diameter aluminum tubes of 0.1 inch thickness. 

A1 1 pay1 cads are modeled as ; umped masses with appropriate center-of - 

gravity offsets. The Station modules are treated as rigid beam mehers 

rigidly connected to one *nuther in tbe appropriate race track 

arrangement. Both payload and module weights are provided in Section 

4.3.3..:. Inertial properties of the analytica! mdel are givgn in Table 

4.3.3.3-2 and the undefornxd geometry of the model is shown in Figure 

4.3.3.3-2. 

It is believed that the rotary alpha and beta joints of the station as 

well as joint connections in the lattice trusses will involve some degree 

gf nonlinearity. lnis is an important subject for future investigation, 

but is beyond the scope of the presznt study. !t is also assumed that 

the alpha rotation rate of one cycle gsr 34 minutes is slow enough that 

the arrays may be considered statioliar~t. 

Since the Station is constructed in orbit, vibration modes of the Station 

during construction should also be addressed, however, this is beyond the 

c:.,pe of the present study. 

4.3.3.3.1.1 Solar arrays normal to flight path 

Tables 4.3.3.3-3 through 4.3.3.3-6 provide the free vibration frequencies 

and a brief verbal description of their corresponding mode shape when the 

arrays are normal to the flight path. In Tables 4.3.3.3-3 through 

4.3.3.3-6 the mode shape desc. iptions are provided for the four cases of 

IOC Station without attachments, with payloads, with Orbiter and with 

both payi,?ads and Orbiter. Each table contains the results for all four 

cases and a brief description of the node shape associated with one of 

the four cases whose values are given in the first column of L-esults in 

each table. Thus the order of the columns for each table is changed so 

that mode shape descriptions are provided for all four cases and a glance 

across the columns of any of these tdbles indicates changes i~ frequency 

as attachments ame added or subtracted. 
Selected mode shape deformations of the IOC Station are presented in 

Figures 4.3.3.3-3 thvcugh 4.3.3.3-7 for the Station wi thout attached 

Orbiter artd payloads; in Figures 4.3.3.3-8 through 4.3.3.3-12 for the 

Station with attached payloads only; in rigures 4.3.3.3-13 through 

4.3.3.3-16 for the Station with attached Orbiter only; and in Figures 

4.3.3.3-11 through 4.3.3.3-20 for the Station with both pzyloads and 

Orbiter attached. 

It is not always otrvic!:s how to characterize individual mdes. However, 

in Tables 4.3.3.3-3 thrwgh 4.3.3.3-6 an attempt is made to characterize 

the mode shapes using symbols defined in the table. Further, it is 

attempted to reflect the relative contribution of different parts of the 

Station to each mode shape by the order in which the symbols are given. 

In general , hqwever, mode shapes generally fa1 1 into the fol lawing throe 

categories: 

(a) Array mast dominated modes 

(5 j PrSiiiirj siruc~ure (keel deformation) dominated 

modes 

(c 1 Radiator mast dominated mcdes 

It is evident from Tables 3.3.3.3-3 through 4.3.3.3-6 that the station 

h2a inany closely spaced modes partidlly due to the repeti ti ye nature of 

its construction, !e.g.,eight solar arrqys). Modes which involve the 

array masts by themseives have frequencies near 0.16 Hz. (See Figures 

4.3.3.3-4, 4.3.3.3-9, 4.3.3.3-14 and 4.3.3.3-18 as well as Tables 

4.3.3.3-3 through 4.3.3.3-6. ) These modes are effectively the cantilever 

modes of tbn individual masts, (with the solar blanket mss distributed 

over their length), and as such are insensitike to the presence of 

pqyl oads and/or Orbi ter. 

Below the cantilever mast modes are the lowest modes of the Station. In 

general , the 1 owest s tat1 on modes i nvol ve ccupl ing of arrqy mast bending 

and keel tvist or bending, with the radiators and upper and lower booms 

moving as rigid t~dies. (See Figures 4.3.3.3-3, 4.3.3.3-8, 4.3.3.3-13 

and 4.3.3.3-17.) Depending on attachme~ts, these modes occur at 

frquencies of 0.096 to 0.15 Hz. A stiffer keel would raise these 

frequencies, but never higher than ths array mast cantiiever frequency of 

about 0.15 Hz. As shown in Tables 4.3.3.3.3-3 through 4.3.3.3-6, when 

payloads and/or Orbiter are attached to the Station, these frequencies 

decrease and keci deformation becomes more pronounced because the keel is 

dynamically softened due to the presence of added mass. Consequently, 

hen these mdes have frequencies close to 0.16 they are denoted as array 

mast dominated modes and otherwise as primary structure dominated modes. 

Above the calltilever array mast modes are additional coupled modes. In 

the absence of attachments there are two modes invclving cougling of 

arrqy mast bending and rigid body keel pitch. (See Figure 4.3.3.3-5.) 

These mdes resemble those occurring in a free-free beam having a heavy 

a;. 
mass ,t its center and as such are slightly higher than the cantilever 

modes. When payloads and/or Orbiter are attacned the frequency of these 

modes decreases slightly, (they can never go below the canti lever modes), 

and become indistinguishable in mode shape from the cantilever modes and 

are hence 1 isted as such in the Tables as array mast dominated modes. 

Above these array dominated modes are those dominated by deformation of 

primary struture; that is, keel twist, keel roll bending, and keel pitch 

beading. (See Figs. 4.3.3.3-6, 4.3.3.3-11, 4.3.3.3-15, and 1.3.3.3- 

19.) Primary structure dominated modes stsrt at a frequency of 0.2094 Hz 

in the case of no attachments and as lcw as .I97 Hz when both payloads 

and Orbiter are attached. Bending of the transverse boom occurs at a 

frequency of about 0.526 Hz in the absence of attachments, at 0.486 Hz 

whev payloads only are ,itached, at 0.509 Hz when Orbiter a'~ne is 

attached and at 0.475 Hz when both payloads and Orbiter are attached. 

Thus the transvese bending mode is nearly insensitive to attachments. 

Finally, the lowest radiator mast dc inated mode occurs at a freqbency of 

0.572 Hz in the absences of attachmen,,. (See Figures 4.3.3.3-7, 

4.3.3.3-12, 4.3.3.3-16 and 4.3.3.3-20.) The frequency of these modes 

change little when attachments are added to the Station. 

4.3.3.3.1.2 Solar arrays normal to the Nadir 

Vibration modes for the case of the solar arrays oriented normal to the 

Nadir are quite similar to those for solar arrays oriented normal to the 

flight ~ath and hence are not shown in this document. Nevertheless, 

transient response results for array oriented normal to the Nadir are 

given in section 4.3.3.5. 

4.3.3.3.1.3 Solar array modes 

As mentioned earl ier, the YASTRAN finite element model does nct allow for 

solar array blanke-des inasmuch as the arrays are onlv taken into 

account as nonstructural mcs; distributed a1 ong the array support mast. 

In order to produce a first approximation to the array blanket modes it 

is assumed that each blanket is attached to a mast near its root and at 

its end, and that the array behaves as a membrane grounded at these 

I ccations . As a consequence, the I owest bl anket frequencies are gi ven 

appw iimately by the formu! a, 

frequency in Hz = ( 1/2)[square root (P/L/M )I 

Where P is the blanket tension, L is the blanket length, and M is the 

blanket mass. For the IOC station, the blanket length is 80 feet and 

weighs 1200 pounds. Ter ion in the blanket must be reacted by 

compression in the array mast. It is assumed that P is 50 percent of the 

tiast axial strength. For the 30-inch diameter array mast, the axial 

compressive strength is about 2000 pounds. Then, a first approximation 
to the low?st blanket frequency would be C.29 Hz. In addition, the axial 

compressi~ rlwould reduce the arrqy mast frequency from about 0.16 to 0.11 

Hz. Presence of the axial compression will also reduce the bending 

strength of the mast. This is an important consideration since 'it is 

shown in Section 4.3.3.5, Tables 4.3.3.5-11 attd 4.3.3.5-16, that the 

array mast root has the smal ,est margia of safety on ths station. 

4.3.3.5.2 Growth confiauration 

The growth configuration is depicted in Figure 4.3.3.3-21 and the 

associated analytical model is depicted in Figure 4.3.3.3-22. The growth 

configurat.ion differs from the IOC Configuration in size, number of 

modules, and the use of a solar dynamic system for pwer generation 

rather than a solar array based system. A single bay upper keel is used 

in this study and the rildi6jtor mast is assumed to be an aluminum tube of 

10.5-inch diameter with 0.10-ixh wall thickness. 

Growth Station modes are similar in character to IOC modes and fall into 

the fol lowing general categories: 

a. keel twizt 

b. coupled keel roll bending and keel twist 

c. keel pi tch bending 

4. array mast modes 

e. radiator mast modes 

Selected mode shapes fcr each of these categories are displayed in 

Figures 4.3.3.3-23 through 4.3.3.3-27 for the Growth Station wi thwt 

attachments and for the dynamic solar collectors oriented normal to the 

flight path. Each figure shows a front and side view of the mode shape. 

The lowest mode involves keel twist and occurs at a frequency of 0.108 Hz 

which is about 8.8 percent lower than the IOC Station in a similar 

condition. 

4.3.3.3.3 Concludina remarks 

A finite element dynamics model and free vibration characteristics of the 

reference gravi ty-gradient Space Station have been presented. Vibration 

modes for both the IOC and the growth co~figurations have been presented 

for the the fol lowing focr cases: without attachments, with pqyl oads 

attached, with Orbiter attached, and with both payloads and Orbiter 

attached. 

For the IOC Configuration, modes were characterized into three basic 

types, modes dominated by arrqy mast deformation, modes dominated by 

primary structure (keel 1 deformation, and modes dominated by radiator 

&st deformations. 
Cantilevered array mast modes occur at about 0.16 Hz. Modes of the solar 

array blankets were not included iq the finite element analysis but the 

lowest frequency of these modes was approximated as being 0.29 Hz. Modes 

with somewhat lower frequencies involve cwpled arrqy mast and primary 

structur? (keel) deformat:on. They may be either array mast dominated or 

primary structure dominated. Stiffening the keel would increase the 

frequencies of these modes, but never above the frequency of the 

cantilever arrqy mast modes. Modes with frequencies sl i gthly higher than 

the array mast canti levered frequencies are a1 so array mast dominated, 

but involve some primary structural motion. Increasing the mass of the 

primary structure would tend to lower these frequencies, but never below 

the cantilevered I:-. ly mast frequency. Finally, in the absence of 

Station attachme.lts, modes haying a frequency of about 0.21 Hz and higher 

are dominated by prjmcry structural deformation. 

Growth confi guratioq fr ?quencies were a1 so provided, but detai 1 ed 

characterization was not ai ven. The 1 owest growth configuration 

frequency was 0.108 Hz in tt~ ,Le-tce of attachments. This is only 8.8 

percent lower than the lowest IOC :onfiguration frequency in the absence 

of attachments. Consequently, it appeared that the Growth Station would 

not respond radically different than the IN Station. 
TASLE 4.303.3-1 - PROPERTIES OF ANALYTICAL REFERENCE 

CONFIGURATION SPACE STAT ION MDDEL 

CWONEMT BENDING TORSIONAL MASS PENDING TORS I ONAl 

STIFFNESS STIFFNESS LENGTH STRENGTH STRENGTH 

(FT-LBS-FT) ( SLUGS/FT ) (FT-LBS ) -- 

BOOMS 81 KEELS 1.31 E+9 3.18 E+8 0+25 35,000 15,000 

30-INCH ASTRO 3.13 Et6 2.C8 E+5 2.3 3,480 208 

MAST 

TABLE 4.3.3.3-2 - INERTIAL PROPERTIES OF ANALYTICAL MODEL 

WEIGHT C. G. COORDINATES MOPENTS Of INERTIA 

- CASE ( LBS (FT) .- (LB-FT-SE~) 

WITHOUT L69.000 (1.1,0,84.2) (8.63 E+7, 

PAYLOADS 7.81 E+7, 

AND ORBITER 1.20 E+f) 

WITH 373,200 (-1-ls0,128-2) (2.06 E+8, 

PAYLOADS 1.98 E+8, 

ONLY 1.45 E+7) 

WITH 508,800 (5.7,0,34.4) (1.45 Et8, 

ORBITER 2.37 E+8, 

ONLY 1.35 E+7) 

WITH 608,600 (3.59,0,68.3) (3.21 E+8, 

PAYLOADS 3.14 E+8, 

AND ORBITER 1.62 Et7) 
TABLE 4.3.3 -3-3 10C REFERENCE CONFIGURATION FREQUENCIES IN HZ 

WITH MODE SHAPE DESCRIPTIONS FOR THE CASE WITHOUT 

PAYLOADS AND ORBITER ATTACHED 

MODE SHAPE DESCRIPTION* WITHOUT W I'TH WITH WITH 

PAYLOADS PAYLOADS ORBITER PAYLOADS 

AND ONLY OhLY MD 

3RBITER ORBITER 

W/k RP 

AM/KRP 

KT/KRB/M/RRR/TBRR/tdRR 

KT/KRB/AM/TBRR/RRR/LBRR 

KPB/AM/RM 

KRB/AM/Rb'/LBRR 

TB/W/RM 

RM/KPB 

RM/LORR 

RPl jTY 

KPB/RM/TB/AM 

KPB/AM 

KPB/AM 

AM 

KPB 

AM/KRB 

KEY: 

AM - ARR:.Y MAST BENDING RM - RADIATOP MAST BENDING 

LB - LOWER BOOM BENDING TB - TRANSVERSE BOOM BENDING 

KT - KEEL TWIST KPB - KEEL PITCH BLNDING 

KRB - KEEL ROLL BENDING KRP - KEEL RIGID BODY PITCH 

RR!? - RPDIATOR RIGID BODY ROLL MRR - MODULE RIGID BODY ROLL 

LBRR - LCWER BOOM RIGID BODY ROLL TBRH - TRANSVERSE BOOM RIGID BODY ROLL 

* MODE SHAPE DESCRIPTIONS CORRESPOND TO FIRST COLUMN OF VALUES ONLY 
TABLE 4.3.3.3-4 IOC REF ERENCE CONFIGURATION FREQUENCIES I& HZ 

WITH MOUE SHAPE DESCRIPTIONS FOR THE CASE OF 

PAYLOADS ATTACHED 

MODE SHAPE OESCPIPTIOM* LSITH WITHOUT WITH WITH 

PAYLOADS PAYLOADS ORBITER FAYLOADS 

ONLY AND ONLY AND 

ORBITER ORBITER 

KEY: 

AM - ARRAY MAST BENDING RM - RADIATOR &ST BENDING 

LB - LOWER BOOM BENDING TB - TRANSVERSE BOOM PENDING 

KT - KEEL TWIST FPi3 - KEEL PITCH BENDING 

KRB - KEEL ROLL BENDING KRP - KEEL RIGID BODY PITCH 

RRR - RADIATOR RIGID BODY ROLL MRR - MODULE RIGID BODY ROLL 

LBRR - LOWER BOOM RIGID BOD" ROLL TBRR - TRANSVERSE BOOM RIGID BODY ROLL 

* MODE SHAPE DESCRIPTIONS CORRESFOND TO FIRST COLUMh OF VALUES ONLY 
TABLE 4.3.3.2-5 IOC REFERENCE CONFIGURATION FREQUENCIES IN Hz 

WITH MODE SHAPE DESCRIPTIONS FOR THE CASE OF 

ORBITER AXACHED 

MODE SHAPE DESCR!PT1ONf WITH WITH WITHOUT WITH 

ORBITER PAYLOADS PAYLOADS PAYLOADS 

OhLY ONLY AND AND 

ORBITER ORBITER 

KEY: 

4M - ARRAY MAST BENDING R# - RADIATOR MAST BENDING 

LB - LOWER BOOM BENDING TB - TRANSVERSE BOOM BENDING 

KT - KEEL TWIST KPB - KEEL PITCH BENDING 

KRB - KEEL ROLL BENDING KRP - KEEL RIGID BODY PITCH 

. RRR - RADIATOR RIGID BODY ROLL MRR - MODULE RIGID BODY ROLL 

LBRR - LOWER BOW RIGID HODY ROLL TBRR - TRANSVERSE BOOM RIGID BoaY ROLL 

* MODE SHAPE DESCRIPTIO>S CORRESPOND TO FIRST COLUMN OF VALUES ONLY 
TABLE 4.3.3.3-5 IOG REFERENCE CONFIGURATION FREQUENCIES IN HZ 

WITH MDE SHAPE DESCRIPTIONS FOR 'THE CASE OF 

ORBITER ATTACHED 

MODE SHAPE DESCRIPTIOH*' KITH WITH WITH WITHOUT 

PAYLOADS ORBITER PAYLOADS PAY LOADS 

AND ONLY ONLY &Nb 

ORBITER ORBITER 

KPB/AM 0.0963 

KPB/AM 0.102 

KRB/KT/AM/RRR/LBRR 0.1139 

0.1515 

0.153 

0.1613 

0.1626 

0.1629 

0.1631 

0.16315 

0.1638 

3.1642 

0.1645 

0.1647 

0.16475 

0.1648 

AM/KRB 0.16485 

AMIKRB 0.1695 

KTIAMIRRRILBRR 0. I370 

KPB/AM 0.2627 

KT/AM 0.2983 

KRB /AM 0.3169 

KRB/AM/LBRR/RRk 0.3713 

M/TB 0.4280 

TB 0.4745 

RM 0.5572 

RM 0.5671 

RM 0 . 5874 

RM 0.5885 

TB/TPB/AM/RM 0.6889 

KEY : 

AM - ARRAY MAST BENDING RM - RADIATOR MAST BENDING 

LB - LOWER BOOM BENDING TB - TRANSVERSE BOOM BENDING 

KT - KEEL TWIST KPB - KEEL PITCH BENDING 

KRB - KEEL ROLL BENDING KRP - KEEL RA, :D BODY PITCH 

.RRR - RADIATOR RIGID BODY ROLL MRR - MODULE RIGID BODY ROLL 

LBRR - LONER BOOM RIGID BODY ROLL TBRR - TRANSVERSE BOOM RIGID BODY ROLL 

* MODE SHAPE DESCRIPTIONS CORRESPOND TO FIRST COLUMN OF VALUEE ONLY 


MODE 9 - .I385 HZ MODE 10 - .I

MODE 11 - .I 502 ti2 

Figure 4.3.3.3-3 - LOWEST ARRAY MAST DOMINATED MODES WITH SO

PRIMARY STRUCTURE DEFORMATION IN THE 

ABSENCE OF PAYLOADS AND ORBZT ER 
MODE 13- .I613 HZ 

- 

MODE 14 - .I630 

MODE 15 - .I636 HZ MODE 15 - .I642 H

Flgure 4.3.3.3-4 - ARRAY MAST CANT LEVER MODES IN THE ABSENCE 

OF PAYLOADS AND ORBITER 
\ 

\ 

c. 

m 

VI 

MODE 23 - .I752 HZ MODE 24 - .I7

Flgure 4.3.3.3-5 - ARRAY MAST DOYINAI ED MODES WITH SOME 

PRlMARY S'TRUCTLJRE DEFORMATION IN THE ABSEN

OF PAYLOADS AND ORBITER 
MODE 25 - .2094 HZ 

MODE 27 - .2640 HZ MODE 28 - .52

Flgure 4.3.3.3-6 - PRIMARY STRUCTURE MODES IN THE ABSENCE O

PAYLOADS AND ORBITER 
MODE 30 - .5721 HZ MODE 31 - .586

MODE 32 - .5920 IiZ 

Flgure 4.3.3.3-7 - RADIATOR MAST DOMlNATED MODES IN THE ABSENCE OF PAYLOA
MODE 10 - .I304 H

MODE 11 - .I317 HZ MODE 12 - ,1542 HZ 

Flgure 4.3.3.3-8 - LOWEST ARRAY MAST DOMINATED MODES WlTH SOME 

PRIMARY STRUCTURE DEFORMATION WlTH PAYLOADS 

ATTACHED 
MODE 14 - .I613 HZ MODE 15 - .l623 H

MODE 16 - .I631 HZ MODE 17 - .I631 H

Figure 4.3.3.3-9 - ARRAY MAST CANTILEVER MODES WlTH PAYLOADS 

ATTACHED 
MODE 25 - .I733 HZ MODE 26 - .177e H

Fi~dro 4.3.3.3-10. - ARRAY MAST OOMINATED MODES W!Tii SOME 

PRIMARY STRUCTURE DEFORMATION WITH PAYLOAD

ATTACHED 

MODE 34 - ,5762 HZ MODE 35 - .5785 H

MOUE 36 - .5897 HZ 

Figure 4.3.3.3-12 - RADIATOR MAST DOMINATED MODES WITH PAYLOAD

ATTACHED 
MODE 8 - .I200 HZ MODE 10 - .I345 HZ 

MODE 11 - .I376 HZ 

Figure 4.3.3.3-713 - LOWEST ARRAY MODES OOMINATED BY PRIMARY STRUCTURE AND A

DEFORMATION WHEN ORBITER IS ATTACHED 
MODE 14 - .I613 

MODE 95 - .I632 HZ MODE 16 - .I632 

Figure 4.3.3.3-14 - ARRAY MAST CANTILEVER MODES WHEN ORBITER IS AmA

..a---L--_il : .. . 1.- . . .T ' V.)i ,, . - .I . 

' . , c- l.t{,..:-,?z.u.-.-:.' %'.,-, -. 

MODE 30 - .5850 HZ MODE 31 - .5676 H

MODE 32 - .5881 HZ MOD L 33 - .5890 H

Figure 4.3.3.3-16 - RADIATOR MAST MODES WHEN ORBITER IS ATTACHED 
MODE 9 - .0965 HZ 

MODE 11 - .I139 HZ 

Flgure 4.3.3.3-17 - LOWEST MODES DOMINATED BY PRIMARY STRUCT

DEFORMATION WITH BOTH PAYLOADS AND ORBITE
MODE 12 - .I515 HZ MODE 13 - .I530 HZ 

MODE 14 - .1613 HZ MODE 15 - .I626 HZ 

*a t) - n- Figure 4.3.3.3-18 - ARRAY MAST CANTILEVER MODES WITH BOTH P.qk ~.+Aub 

AND ORBITER ATTACHED 
MGDE 28 - .2627 HZ 

MODE 29 - .2983 HZ MQDE 30 - .3168 HZ 

Figure 4.3.3.3-19 - PRIMARY STRUCTURE MODES WITH BOTH PAYLOADS 

AND ORBITER ATTnCHGD 
MODE 34 - 5572 HZ MODE 35 - .5671 H

MODE 36 - 5874 HZ MODE 37 - .5885 H

Flgure 4.3.3.3-20.- RADIATOR MAST MODES WITH BOTH PAYLOADS AND 

ORBITER ATTACHED 




Figure 4.3.3.3-25 - SELECTED GROWTH STATICIN COUPLED KEEL ROLL BENDIN

AND KEEL TWIST MODES IN THE ABSEENCE OF ATTACHEMENTS 
Figure 4.3.3.3-26 SELECTED GHOWTH STATION KEEL 

id 

PITCP BEINDING MODES IN THE ABSENCE 

4.3.3.4 Mass Properties 

Mass properties for the IOC Station used in the rigid body analyses are 

shown in Table 4.3.3.4-1. Six configurations are characterized in this 

table. The minimum mass configuration, 124,321 kg, represents the basic 

Station without the Orbiter, payloads, or servicing items attached. The 

maximum mass conf i guratian, 312,028 kg, hds all these elements attached. 

The Station element location, masses, and projected areas are shown in 

Table 4.3.3.4-2. Payload and servicing item locations and masses are 

shown in Tables 4.3.3.4-3 and 4.3.3.4-4, respectively. The mass 

properties us,d for the flexible response and controls analyses are noted 

in the appropriate paragraphs. As is usually the case, these differ 

slightly from the values listed in the tables. 

mmP.L PACtf L-'? 

oE mR QUALITY 
WEIGHT 

INERTIAS 

I XX 

IYY 

1 zz 

IXY 

IXZ 

IYZ 

UNITS 

TABLE 4.3.3.4-1 - STATION MASS PROPERTIES 

WITHOUT ORBITER WITH ORBITER 

PAY LOADS PAYLOADS 

BARE PAYLOADS SERVICING BARE PAYLOADS SERVICING 

124321 157964 20502? 231238 264900 306394 

98.72 193.93 256.25 177.57 317.35 417.00 

34.71 190.10 251.13 181.81 323.50 422.52 

7.549 10.10 11.398 18.70 22.99 25.10 

0 0.11 0.094725 SMALL 0.36 0.38 

-0.7073 -1 -09 -6.18063 -22.14 -30.73 37.14 

0 -1.47 -0.54212 9W.L -0.60 0.45 

C.G. LOCATION 

XCG METERS 9.1338 -.31979 -0.27027 5.48 4.43 3.32 

Y CG METERS 0 -0.32 -0.34699 0 0.19 0.23 

ZCG METERS 55 * 78 48.22 43.739 71.63 65.10 59.80 

NOTE: THE REFERENCE COORDINATE SYSTEM IS SHOWN ON FIGURE 4.3.3.5-2. 
J-I 8x1 ,x 83 82 @ xt 81 49 9 3 3 03 4 rq .y, - 4 :J - W! 4 tj? 0 0 AD a Ua 

LL! rrr ggj ~~~r~~~:,~~~;~...~-~rrm ms r.r.4~ 

I- .............................. X Id QQCCrtQCt dDi3$~U.i..'~C\l\~~ir1D9ilbf-t QU L0l;)i.cz .-A-.+.-.+~.+ pr, P: ~~7r-h~ r-r- I-w kt- - 1 I I I I I LnU~LlrP' 

'? 

LC 

,. d 

,=, -I--r-r t-+cno:so~r.~.-ttr~aac~~ 9.0 rr~varu 

~JLO r-:I r~r-mm~m~tpln~~a~m~ 64 *D\O(D-II it .............................. j - $= .3 p-0 I:fi Q )ib ~7 i)d F,. I= efi id id qz 0) a c. 8~ 5~ b-1 )n 8 Q. 

IJ; t 4 ;: +'4* :I 0'4 1-4 I -- Pi?, 

gw Ill1 L r 
I TEV 

PRYLLlROS 

TOM POL@ 

SHA 0005 

SHA @a86 

5AH 8089 

SHA G287 

TOM 2928 

TOM 24:Q 

TOM 2510 

rnn 256e 

TClM 2068 

TTIM 2078 

TDM z7,i u 

iOM 1283 

TLlM 2168 

SHA 32B1 

SHR 8287 

TOM 2578 

CENTER OF MRSS COURDINRTES 

METERS 

x Y 2 

-19.28 

0.88 

-18.97 

18.97 

4.57 

3.66 

-3.66 

8.88 

9.14 

0. BB 

0.08 

-3.05 

-6.96 

lb. 67 

0.68 

0. 

-4.57 

WEIGHT 

Table 4.3.3.4-3 - PAYLOAD LOCATIONS AND MASSES 

PUWEP TUUER PEFEFENIX CONF IGURRT ION 

CEP4TER OF 

I TEfl 

E 

SRTELL!; E SERVIC. ING 

SEYV Jr I:JC ATTPT~~I =SS-TYPE 1 e. so 

5TOF.AGE I~TTACHMEYT B.'Q 

REF1 .CL ING Hf TACHME-NT .A, 

OMV STORUGC HTTHCHrlEYS- .( 

3x3 b U6 

.IF ' -. i. rrsb361 BOXES 63.88 

~-r I. C.T~:-HCE tr(;xt~ '0. CIP 

i T130' t'I2 'F>* YFFUEL. RYkP 0. VFt 

INS" ?' .V+ ,;-: L'rJRHGE SHELTER -'d. 91 

Ft ':- L c 'CIFHGE TfiNES 0. 613 

CtJEL L'.'ORRCE TANK 4.013 

St3 ..LL I TE CIiECYOUT EOU1:- 12 P,~J 

'MU rHEtCOUT EOUIP 3, r:*b 

SET UkU'5 En. 

SET OF TOOLS a. BU 

CJMV b 1 TS b. BB 

',:T INST."'. ' ?;iF -8.91 

Ff.fEL 5, lib:, 0. 

3HTEI.L I TE 0. FJB 

C.,~TEL.L_ ITE Gc. C?+3 

MRSS CUOROIMTES 

METERS 

'I' 7 

L 

WC I GHT 

Table 4.3.3.4-4 - 3ERVICINO ITEM LOCATCONS AND MASSES 

181 
4.3.3.5 Performance Assessment 

4.3.3.5.1 Rigid body control dynamics 

The objective of this analysis was 03 determine the control requirements 

for Space Station operations including attitude control and orbit 

maintenance as a functior; of the natural on-orbit dynaaic envimnlmt. 

The dynamic environments simulated included grzvi Q gradient toques, 

aerodynamic drag, and aerodynamic toques. System requi t'enrents for the 

base1 i ne confi yuration were determined for parmetric variations of 

a1 titude and mss properties. 

- Siwlation capabilities 

Two indcyen~ently developed programs with similar capabilities were used 

to predict the on-orb1 t dynamics o'i' the Space Stat!oq, the Space Stiition 

Dynam,l'cs (SSDYNAMICS) and the Xi gid Body Control Dy,.~mics (RCD) program 

(References 12 through 16). 

The SSOYNAMICS program, developed at JSC, ir;:+fal izes with the Space 

Station on a user prescribed orbit and calculatrb the time hi stories of 

altitude and attitude as a function of the dynamic enviro~ments 

encountered. A simple, feed bacK control system is modeled. The 

associated control torque time histories and moa:ntun storage 

requirements for maintaining the Station at a prescribed orientatton or 

maneuver, e.g., solar inertial or earth ff xed (LVLH), are conputed. The 

control torques are computed ar 3 1 inear function of the attf tude and 

rate errors of the Space Station (Figure 4.3.3.2-1). CMG torque limits 

are imposed where app:icable. Motion of the solar arrqys and radiators 

relatf vo to the main body is included in the aoalysi s. #ass and 

aeradyr;imic properties for the Space Station are computed as a function 

of the Space Station's individual components. The atimphere &mi ty :'.. 

coraputed as a function of the time of year, altitude, longitude, 

latitude, F10.7 solar flux index, and Ap geomagnetic itraex. The 

atmosphere is simulated as rotzting with the Earth's surface. The 

equations of motion which are based on Newton's second law of mtion and 

Euler's mment equations are solved using a variabls step Runge-Kutta 

integratf on routine. 

The Rigid Body Control Dynamics (RCD) prcgram is a major software pPoGrrl,r 

within the Interactive Design and Evaluation of Ade!anced Spacecraft 

(IDEAS) computer-aided design and arialysis system developed at Langley 

Research Center (Reference 16). RCD uses equation of mtion essentially 

the same a: those employed jn SSDYNAMICS to compute on-~rbit 

environmtstal and mneuver forces ana toques at user-speci fjed circular 

orbital a1 ti tudes amd spacecraft orientation. Mass and aerodynamic 

properties are autom>tically computed and input by other IDEAS mdules +,I 

dC3. MawnCun storage dnd desi?turation req~il remen ts, contrr-1 system 

predesign control , and maneuvers are cal cul b ted for one orbi t 2nd for the 

spacecraft 1 ifetjm. Principal features of 6:CD are shown in Figure 

4.3.3.5-1. Results from both the SSDYNAMICS and the RrJ proyams are 

sumrized in Section 4.3.3.5.1.3. 
4.3.3.5.1.2 Flight mde selection 

The reference apace Station flies with its keel stabilized by CMG'S to a 

quasi local vertical local horizontal (LJLH) mo& with ;is zclar arrqys 

rotating at orbital rate using the alpha joint drives, to point at the 

Sun. The Sun beta condition is handled with beta joint drives. The 

power booms are maintained perpendicular to the n-h!t ;!?n~ to simplifjr 

the alpha and beta joint drive system. The co~figuration is pit: ::cd in 

the orbit i lane to an average torque equilibrium attitcC: i ihA) to 

preclude momentum buildup in the pitch plane. The pitch plane gravity 

graaient torque varies over an orbit due to rotation of the solar arrvs 

(Figure 4.3.3.5-8). The aerodynamic torque also varies over an orbit due 

tu solar array rotation and due to the diurnal density variation. A 

gravity gradient drift flight mode was analyzed and discarded after 

several simulations since the atmospheric variations and solar arrqy 

rc'ations caused a close to resonance pendulum notion of the keel bo* 

'h peak to peak oscillations of approximate 13' at an altitdde of 270 

and 19.5' at 250 nmi fw the IOC mass condition. 

4.3.3.5.1.3 Control requirements for the natural environment 

Simulations o, the six different mass properties conditions of the 

Station were conducted with parametric variations of the a1 titude to 

determine the minimum CMG control requirements for attitude 

maintainance. The simulation used a 2 sigma atmosphere condition and 

included short-term atmospheric density variations. Cyclic momentum, 

secular momentum, trim angle, number of CMG'S, and back-up RCS momentum 

storage were summarized per orbit in Table 4.- .3.5-1. 

Four skylab equivalent CMG'S wil sl!f 1 ice for the nominal altitude range 

oi 250 to 270 nmi . However, if the a1 t- tude i s reduced to 320 nmi , '.be 

IOC + Orbiter cqnditic~ will require slight1 j six skylab CMG'S. The 

secu1ir.r momentum bL';;dup per orbit shows the qeed to frequen'.ly 

dessturate the CMG'S. Ma~netic torquers can be csed in lieu of RCS to 

prevent structural respirnse arcel erations and contani nation. A1 so 

another method of nulling the secular torque buildup about the Station xand z-axes involves rotating the Sta:ion about an axes such that the 

solar ai-ray boom axis prescribes a ti1 ted cone with respect to the orb1 t 

plane. The resulting secular gravitj -.gradient torques can be made to 

null the secular aerodynamic torques. However this method was pot used 

since it violates the requirenrent that the solar arrqy booms remain 

perpendicular to the ormbit plane. Future trade studies mqy be conducted 

to determine the COSL :rade-off between the two before-mentioned nethods 

of secular momentum a1 leviari~n. 

Data for the preceding table was Lerived from detailed flight 

simultations. Time histories from one such simulation of the IOC 

confi guration including payloads at 250 nmi are presented in Fjgures 

4.3.3.5-3 through 4.3.3.5-11. The c?ordinate system used with Figures 

4.3.3.5-3 through 4.3.3.5-11 is shown in Figure 4.3.3.5-2. 
This case was chosen since its mass properties resulted in large out of 

plane toque requirements. The simulatdsn starts at the sub-solar pofnt 

in the orbit at an inclination of 28.5' and with an Earth longitude of 

0.0'. Thus, the local tine is 12:OO noon, and the latitude is 28.5' 

(Fig. 4.3.3.5- 3). The station is pitched in .:he orbit plane to an angle 

of - 0.59' such that the pitch axes momentum requirements at the end of 

an orbit are zero (Fig. 4.3.3.5-4). 

Figure 4.3.3.5-4 shows the inertial toque impulse vector with respect 

to an inertial coordinate system fixed in the plane of the orbit. It was 

concluded from this figure that 3.3 Skylab CMG's would be required for 

attitude hold. 

This nuher was arrived at by vectorally adding the peak-to-peak y-axes 

Prentum xquirement of 12000 ft-lb-sec to the concurrent x and z 

resultant morntun requiremnt of 9000 ft-lb-secs resulting in a total 

maerrturn storage of 15000 ft-lb-ser. By properly initilizing the M's, 

they need only to stor- half of the peak-to-peak cylic momentum. Thus 

half of 15000 ft-lb-secs require 3.3 equivalent Skylab CnG's. Note that, 

during this same orbit, a slight secular mmntum of L!OO ft-lb-scc 

accumulated about the orbit z-axes. As mentioned above, this secular 

aocaentura will requirc some form of desaturation. 

The resultant drag force over an orbit is shown in Figure 4.3.3.5-5. The 

peak drag fwce is seen to be slightly over 0.2 pounds and occurs at a 

tine of 0.16 orbits. krgency RCS propellent reqdirements for '..he case 

of full failure arc8 calculated from the integration of the aasol ute 

value of the control tchque vector with respect to the space station body 

axes. These torque impulse requirewnts are snown in Figure 4.3.3.5-6. 

It fs seeq th:t at the end of an orbit, the toque impulse vecto~ is 

3Qa00,24000,1C40 ft-1 b-secs. 

Figure 3.3. 1 7 shms the control toque vector requir-d to hold the 

station LVLH over the mit. The os ,ilations abwt the y-axes are caused 

by the solar arrqy articulations and varidtiolls in aerodrag. 

The gravity gradient toques are shown in Figure 4.3.3.5-8. The y-axes 

cmponent varies due b the solar array articulation. The aerodynamic 

torques are shcwn in Figure 4.3.3.5-9. The diurnal variation is apparent 

in the y-axes component of toque. The slight variation about the x-axis 

is caused by the aerodrag variation;. The aerodynamic drag impulse is 

shown in Figure 4.3.3.5-10. The peak impulse at thz end of an orbit is 

600 13-sec. The aerodensi ty occuring during the orhit is shown in Figure 

4.3.3.5-i1. The density varfes by approximately a factor of 3 over the 

orbit 

4.3.3.5.1.4 Control requirements during system fail urss 

A series of sinula!ions were conducted for the IX configuration to 

. understand the control require-ents and resulting dynamics in the event 

that a Space Station system eit let tai!ed or ras temporarily shut down 
for service. Three types of cases were investigated. The first type 

being a case in which the CMG system fails to provide av control 

torque. The second type being a case in which the CMG system fails to 

provide control toque and a back-up system provides attitude hold about 

the z-axes. The third type is a series of cases where the alpha 

joint/joints were failed or shut down with the solar arrqy either 

parallel or perpendicular to the main keel. 

The motion of the IOC Space Station with respect to a rotating orbit 

coordinate system for the first fa4lure mode case shown in Figure 

4.3.3.5-12 and Figure 4.3.3.5-13 for five orbits. The IOC is seen to 

rotate about its keel in Figure 4.3.3.5-12 and to swing in a pendulum 

manner in Figure 4.3.3.5-13. 

The motion of the IOC with respect to a rotating orbit coordinate system 

for the second failure mode case are shown in Figure 4.3.3.5-14 and 

Figure 4.3.3.5-15 for a 270 nmi and 250 nmi altitudes respectively. The 

aerodynamic orbital variation is seen to cause a near3 resonance pendulum 

motion with peak-to-peak ascil lations of 13 and 19.5" for the 270 and 250 

nrni altitudes, respectively. 

The control requirements related to the third type of failure are 

sununarized in Table 4.3.3.5-2. The CMG requirements for these cases 

include the angular impulse required to bring the solar arrqys up to the 

Station's angular rate and the angular impulse caused by the natural 

environments. A singular alpha failure requires 0.25 Skylab CMG's to 

accelerate the arrqys to station angular rates and 0.5 Skylab CMG's for a 

double alpha joint failure. These failure cases are only slightly more 

severe than for nominal operations. However the peak secular momentum 

(3000 ft- lb-sec) about the z-axes has increased greatly over the fiolainal 

case (1000 ft-lb-sec ). This wit 1 greatly impact the desaturation 

requirements, 

4.3.3.5.1.5 Control rquirements for the induced environment 

A few simple rigid body calculations were made in an attempt to develop 

insight into the effects of manipulator motion and crew ,mtion on the 

mmntum storage and toque requirements for the Station. For example, 

crew kick-off in the lower habitation module results in the following: 

Approximate distance to CM = 90 ft 

Peak torque = 25 lbs x 90 ft = 2,250 ft-lbs 

Angular momentum = 0.5 X 2250 = 1125 ft-1 bs 

Peak torque output of 4 CMG's = 4 x 160 = 640 ft-lbs. 

These results suggest that the CMG's will torque si).turate during the 

later part of the assumed crw kick imp~lse. This is not necessarily a 

problem but it does indicate how easily lrew scion can overpower the 

CMG' s. Also the corresponding momentum storage requirement i s equivalent 

t.3 about one-half the capacity of a reference CMG. Now consider the 
case where a 32,000 pound payload is being removed from the Orbiter bay 

by the Station manipulator while the Orbiter is docked to the Station. 

Assum? : 

Pqyload/manipulator tip velocity = 0.2 ft/sec 

am extended so that disQnce to CH = 100 ft 

angular momentum to start motion: 

H = mass x v x moment arm 

H = 32000/32.2 X .2 X 100 = 20C00 ft-lb-sec 

Peak torque - depends on start-up transient 

Equivalent CMG's - 200001 2300 = 8.6 

The later calculations show that if it is assumed that the Station is to 

be control led by CMG's during manipulator operations, then the nullber of 

CMG's requ4red car! be significant relative to those required of control 

due to the natural environment. Obviously additional work must be done 

in this area. 

Other similar calculation were made such as moving payloads on the upper 

boom with the manipulator. These results were essentially the same as 

those reported above. 

4.3.3.5.1.6 Controllability Essesslaent during assembly 

The on-orbi t configurations of the reference configuration after the 

first two launches are shown in Figwe 4.3.3.5.-16. The mass propertSes 

for each configuratio~ are shown in Table 4.3.3.5-3. Simulation of the 

mass properties conditions for each of the configurations were conducted 

in order to determine the flight orientation and minimum CMG requiremnts 

for attitude maintenance. The simulation was perfonwed at nmi, using 

a constant dencjt.3 representative of the MSFC 5-70 atmosphere, with +2 

:? 9 solilr flux predictions. The momentum storage requirements are 

shown in Table 4.3.3.5-4. This table shows that the muentun storage 

requirements for attitude control are well within momentum storage 

capabilities cf 1 Skylab-equi valent CMG. The corresponding force, toque 

and momentum time hi stories for each configuration are shown in Figure 

4.3.3.5-17. 
4.3.3.5.1.6 Controllability assessment during assembly 

The on-orbi t configurations of the reference configuration after the 

first two launches are shown in Figur2 4.3.3.5.-16. The mass properties 

for each configuration are shown in Table 4.3.3.5-3. Simulation of the 

mass properties conditions for each of the configurations were conducted 

in order to determine the flight orientation and miniwln CMG requiremnts 

for attitude maintenance. The simulation was performed at 270 nmi, using 

a constant density representative of the MSFC 5-70 atmosphere, with +2 

signa solar flux predictions. The momentum storage requirements are 

shown in Table 4.3.3.5-4. This table shows that the lnoraentuln storage 

requirements for attitude control are we1 1 within nruinentum storage 

capabil ities of 1 Skylab-equi val ent CMG. The corresponding force, torque 

and momentum time histories for each configuration are ~hown in Figure 

4.3.3.5-17. 

4.3.3.5.1.7 Control requirement sensitivities to payload placement 

The configuration of the station was designed to maintain symmetry with 

respect to the x-z plane causing the z principal axis to lie in the x-z 

plane. However, the addition of payloads to the configuration causes the 

z principle axis to shift out of the orbit plane. A study was conducted 

to determine the control sensi ti vi ty of the IOC+PAYLOADS configuration to 

variations in the numb~r of attached payloads. Mass properties were 

computed for each variation by removing a pqyload, computing the mass 

properties, replacing the payload, and then removing the next payload 

until 21 1 the payloads had been examined. The removal of the OTV 

Servicing Technology TDM 2570 caused the largest out-of-orbit plane shift 

of the z principal axis. The corresponding CMG control requirements were 

increased to 4.8 over the nominal case's 3.3 for an altitude of 250 nmi. 

These results indicate that a strategy should be developed for payload 

placement which minimizes the impact to the CMG control requirements. 

4.3.3.5.1.8 Orbit maintenance 

An investigation of the orbit a1 titude loss due to aerodynamic drag and 

of the impulse required to reboost was performed. Computations were made 

for predicted maximum and minimum atmospheric densities and station 

masses for the IOC timeframe. The following paramter values were used 

to characterize the system: 

Drag coefficient = 2.3 

Minimum mass 124,321 kg 

maximum mass 204,892 kg 

Reference are2 for minimum mass = 1475.8 m2 

Reference area for maximum mass = 1976 m2 

Atmospheric model - MSFC J-70 with + 2 sf gna solar 

flux predictions. From 

NASA SP 8021, 

March 1983 revision. 
Plots of decay time histories for initial altitudes ranging from 250 to 

270 mi are given in Figures 4.3.3.5-18 and 4.3.3.5-19. Detailed data 

from which the figure were derived is sunnarized in Table 4.3.3.5-5. The 

data on the table indicates that for an initial altitude of 270 mi, the 

minimum mass Station will decqy to an altitude of 250 mi in 90 days when 

the 2 sigma high density atrosphere is assumed. These values, which are 

intended to be representative of the most severe condition, satisfy the 

present system design requirements. 

Oecqy rates for the high mass Station (payloads and servicing items 

added) are lower than those for the low mass configuration. 

Estimates of the reboost requirements are also shown on Table 4.3.3.5-5. 

These calculations were based on a Hohman transfer from the lower to the 

higher circular orbit. Reboost from an orbit of 250 nmi to an orbit of 

270 mi requires 2600 pounds of propellant having a specific impulse of 

220 pound-second/pounds. The corresponding nuRlber for reboost following 

a 90-dqy decay from 270 nuti and maxiwa station mass is 3,564 pounds. 

4.3.3.5.2 Flexible body attitude control dynamics 

Both time and frequency response cases were run to study control 

system/fl ex body interaction. Bending mode data obtained from NASTRAN 

were run with a PID controller to obtain open and closed loop responses. 

The band width of the control system was set a decade below the lowest 

bending mode frequency. Sensor and actuator dynamics were not modeled. 

Configuration data are presented in Table 4.3.3.5-6. 

4.3.3.5.2.1 Pointiag accuracy 

a. Crew disturbance torques Disturbance torques due to 

crew motion were simulated in all three axes of the Space Station The 

shape of the disturbance which is shown in Section 4.3.3.1 simulates an 

astronaut pushing off a module wall and then contacting the opposite 

wall. 

Disturbance torque results are sumnarized in Table 4.3.3.5.7. Time 

histories of pointing error are shown in Figures 4.3.3.5-20, 4.3.3.5-21, 

and 4.3.3.5-22. 

b. Berthing torques A disturbance torque resulting from the 

Orbiter berthing with the station is shown in Section 4.3.3.1. 

The magnitude of the torque and the resulting pointing errors and 

settling times are sumnarized in Table 4.3.5-8. Time histories of 

pointing error are presented in Figures 4.3.3.5-23, 4.3.3.5-24, and 

4.3.3.5-25. 

c. Torque Limited Actuators The crew disturbance and berthing 

cases were repxed with the control actuator torque limited to 
640 foot-pounds. This simulates the condition of using four CMss where 

the maximum CMG output toque is 160 foot-pounds. Results for both crew 

and berthing disturbances are shown in Figure 4.3.3.5-26 to 4.3.3.5-31. 

Comparing these results with the nonlimited cases shows essentially no 

difference for crew disturbances. Berthing di sturbances, on the other 

hand, prodcce attitude error excursions well in excess of lo in the 

torque 1 imi ted case. 

d. + Time Res onse Results Analysis of the pointing error 

plots indicates a sta e contromm. Damping is good and the control 

system is able to hold the pointing error in each axis to within the one 

degree pointing requirement for crew di sturbance torques. Pointing 

errors greater than one degree are seen during berthing when the control 

torque is limited to 640 foot-pounds. The large settling times observed 

in all the runs are indicative of the low bandwidth controller being used 

to avoid control /fl ex body interaction. 

4.3.3.5.2.2 Control stability 

a. Frequency response Open 1 oop frequency response results 

in the form of Bode Plots and Nichols charts 3.e shown in Figure 

4.3.3.5-32 - 4.3.3.5-37. 

b. Frequency results Analysis of the frequency response 

plots shows acceptable rigid body gain mrgins and phase margins on the 

order of 60°. It is anticipated that the inclusion of actuator and 

sensor dynamics in the system model will 1 ower the phase margin 

somewhat. The important point, however, is that the results shw no 

control systemif 1 ex body interaction. Those bendi ng modes which do show 

greater than unity open loop gain are at frequencigs well above the 

control system bandwidth and are phase stable. Of course, it nqy ae 

desirable to adopt some type of filtering to insure attenuation of these 

higher frequency modes. 

4.3.3.5.2.3 RCS results 

A set of RCS cases were run to simulate a center of gravity offset during 

the reboost maneuver. Figure 4.3.3.5-38 illustrates the assumed 

condi tions. 

The figure shows that during a four jet reboost, center of gravity offset 

results in a 4500 ft-lb pitch disturbance toque. The control system 

will off modulate jets to hold the Station within a + 1.0' deadband. 

Eventually, the Station will reach a 1 imi t cycle conai tion about a l.OO 

bias or attitude error. The frequency of these periodic firings can 

obviously exci te structural bending modes. Using the configuration and 

autopil ot data given below in Table 4.3.3.5-9, the autopilot hysteresis 

(h) was parameterized to determine its effect on 1 inli t cycle frequency. 

Using a value of .05" for hysteresis, Figure 4.3.3.5-39 shows toque 

acting on the vehicle as a function of time. Figure 4.3.3.5-40 

illustrates 4n the phase plane the tendency of the station to limit 
cycle about a 1.0' error. The 1 imi t cycle frequency is approximately 

0.05 Hz which is below the first bending mode frequency. Toque 

histories for several different values of hysteresis were used to drive 

an open loop flexible model of the station. These results appear In 

Section 4.3.3.5.3 of this report. 

Transient response sf IOC Space Station 

The purpose of this section is to present and discuss preliminary results 

for the transient response of the reference IOC Station to the dynamic 

load cases of section 4.3.3.1. This includes responses to berthing, crew 

kick-off, RCS firing sequence for Station reboost and MMS operations. 

The analytical model used to derive these results is discussed in Section 

4.3.3.3. Section 4.3.3.3 a1 so suimnarizes the nodal charact. r of the IOC 

and growth Stations. All the results of this section are derived from 

the MSC NASTRAN finite element structural analysis program of Reference 

8. 

4.3.3.5.3.1 Orbiter berthing 

Berthing of the Orbi ter to the Station is simulated by a 500-pound pulse 

of one second duration as described in Section 4.3.3.1. The pulse acts 

in the negative x direction (opposite to the flight path). To simulate 

mi sat ignment of the Station berthing port to the Orbiter berthing port, 

the effect of an dddi tional torque pulse about the Nadir axis is also 

examined. The duration of this toque is also 1 second and has a 

magn~ tude of 2100 foot-pounds. Station responses of particular interest 

due to berthing are the pi'ak acceleration levels at the modules, center 

of the upper boom, tip of the transverse boom. and tip of an array mast. 

In addition, the response of a 5000-pound experimental p~vload on a soft 

spring support is also examined in order to assess the practicality of 

isolating pqyl oals which have 1 ow acceleration level requirements. 

Accelerations were obtained for both zero and one-ha1 f of 1 percent of 

critical damping. Peak accelerations were tabulated for the conservative 

assumption of no damping. 

It is observed that the longest period of vibration, of a mode whose 

shape invol ves motion of the berthing port, is, from Table 4.3.3.3-3, 7.2 

to 10.4 seconds, depending on attachments. Thus, it takes from 3.6 to 

5.2 seconds for locations furthest from the berthing port to experience 

the affect of the berthing load and 7.2 to 10.4 seconds for the berthing 

port to experience the wave reflected from these remote locations. This 

is cmsiderably longer than the puls duration itself, so that the 

berthing loading specified herein approximates an impulse. 

4.3.3.5.3.1.1 Solar arrays normal to the flight path 

Tim hi stories of acceleration response to berthing at the selec tea 

station locations when the solar arrays are oriented normal to the flight 

path are given in Figures 4.3.3.5-41 through 4.3.3.5-44. In these 

figures, one-half of one percent of critical damping is assumed in 
aac h mode. T ~ r.7 e 4.3.3.5-10 sumrizes the peak acceleration taagni tudes 

at these locatibns for the conservative case of zero damping. Results 

are expressed in multiples of g, (i.e., 32.2 ft./sec./sec.), and as 

ratios to the acceleratioc of each location were the entire Station 

rigid. This ratio, denoted as the amplification factor, "a', provides a 

measure of fl exibi 1 i ty effects. 

Also of interest are the peak bending mments and torques occurring at 

the intersection of the keel extension and lower keel, transverse boom 

root, alpha-rotary joint, top of upper keel and array mast root. Table 

4.3.3.4-11 contains peak moments and toques at these locations and their 

margins of safety (M.S.) using a factor of safety of 1.5. Strength 

values used in calculating the margins of safety are given in Table 

4.3.3-3-1. 

Peak acceleration magni tudes at the selected Station 1 ocations range from 

about 0.0012 to 0.016 g's. With the exception of the center of the upper 

boom, peak acceleration magnitudes change little wfth the attachment of 

payloads. Further, the amplification factor at the tip of the transverse 

boom is very large when payload^ are present. This occurs, not because 

the flexible boa zrceleration is very large, but because in an entirely 

rigid station with payloads attached, the instantaneous axis of rotation 

due to the berthing load 1 ies nearly along the transverse boom. Hence the 

rigid body rotation of this location is near zero. 

The effect of damping on Station response can be appreciated by conparing 

the peak values in Table 4.3"3.5-10 where no damping is assumed with the 

peaks occuring in Figures 4.3.3.5-42 through 4.3.3.5-48 where 0.5 percent 

damping is assumed. For example, the peak response at the center of the 

upper boom is about 39 percent lower In the presence of damping. The 

large damping effect is basically attributed to the fact that the highest 

accelerations in the absence of damping occur long after the applied load 

is removed. Thus damping s! gni ficantly reduces these 9 ater occurring 

peaks and earlier occurring peaks becow the maximums experienced by the 

station. It is be1 ieved that the conservatism introduced by using the 

peaks in the absence of dampiny will compensate for the uncertainties 

presently existing in load definition. In future studies, when load 

definition is better, it would be desirable to use peak values in the 

presence of damping . Inasmuch as no acceleration requirements during a berthing operation 

exist at the present time, 1 i ttle can be said about the acceleration 

1 eve1 s themselves. However, the internal structural force resultants 

require consideration re1 ati ve to allowable Station strength. 

Table 4.3.3.4-11 contains peak bending moment and torque magnitudes at 

selected locations on the station. The attachment of payloads has little 

effect on the root bending moment at the upper end of the keel extension, 

but more than doubles the root bending moment of the transverse boom. 

Moreover, since the addition of payloads to the Station makes the 
Station unsymnetric about the keel axis, torques exist in certain 

locations of the structure where they did not exist in the absence of the 

pay1 oads . 

Margins of safety assuming a factor of safety of 1.5 are also shwn in 

the Table. All margfius of safety exceed a value of 4.7 wfth the lowest 

values occurring at the root of the array mast. Thus, this location 

should be given careful scrutiny in future studies when the loading 

envf ronment is better defined. Finally, the results indicate that the 

flexible deformation of the arrqy mast, frtm root to tip, does not exceed 

4 inches. 

As discussed previously in this sectf on, the Orbiter berthing load is 

essentially impulsive for this low frequency station. As shown in 

Reference 17, but for a different Station confi guration, internal force 

resultants can be considerably higher when load duration is longer. To 

more accurately account for 1 oad durati on as we1 1 as berthing 1 oad level, 

future studies should consider simulation of the berthing operation. 

This could be accomplished by using a simplf fied berthing mechanism model 

between the Station berthing port and an Orbiter wfth an initial berthing 

velocf ty. Presently, initial berthing velocities are believed to be 

about 0.1 to 0.3 feet per second. 

4.3.3.5.3.1.2 Solar arrays normal to the Nclir axis 

Response results when the solar arrays are oriented normal to the Nadir 

are sunnnarized in Table 4.3.3.5-12. Results are similar to those of 

Table 4.3.3.3-10 for arrays oriented normal to the flfght path. The 

primary difference between results for the two array orientations occurs 

at the tip ~f the array mast. The peak acceleration for arrays oriented 

normal to the flight path is up to 5 times larger than that for arrqys 

oriented normal to the Nadir. Thf s f n turn indicates larger bending 

moments at the arrqy mast root when arrays are oriented normal to the 

fl i ght path. 

4.3.3.5.3.2 Response to crew-kick-off motions 

As described in the Section 4.3.3.1 of this document, two differ-cnt 

crew-motion cases are examined, one based on crew kick-off across the 

lower habitation module, denoted as transverse crew motion and the other 

based on crew kick-off along the lsngth of the habi tation module, ¬ed 

as axial crew motion. 

4.3.3.5.3.2.i Solar arrays oriente~ normal to the flfght path 

Figures 4.3.3.5-45 thru 4.3.3.i-48 illustrate the time history of the 

acceleration response due to ixial crew mtfon at the saw seiected 

Station locations as that used for herthi ng when the solar arrays are 

normal to the flight path. For the r-esults presented in the figures, 

one-half of one percent critical damping is assumed in each mode. 
Responses of the modules, as shown in Flgure 4.3.3.5-45 clearly indicates 

the occurrence of the crew-kick-off. Table 4.3.3.5-13 :**mrizes the 

peak acceleration magni tucks of the responses for the conservative 

assumption of no damping. Acceleration values are given in multiples of 

g and amplification factors, are also provided. Internal loads are not 

provided for crew motion since they are not expected to be significant 

from a strength point of view. 

Table 4.3.3.5-13 indicates that attack !ng payloads to the Station 

increases some accelerations and decreases others. In general , attdc hi nc, 

the Orbiter to the Station decreases acceleratfon levels. Again, as 

explained under the berthing discussion, amp1 ification factors when only 

payloads are attached to the Station are high along the transverse bmi 

since it lies along the instantaneous axis for rigid body response to 

crew motion. Yhen both payloads and Orbiter are attached to the Station, 

the instantaneous axis of rigid body rotation for these crew motions lies 

along the upper end of the vertically oriented solar array masts, so that 

amp!i ffcation factors are very large for this location. Moreover, when 

payloads and/or orbiter are attached to the Station, there is generally 

1 f ttle difference between the peak acceleration responses for transverse 

and axial crew motions. 

4.3.3.5.3.2.2 Solar arrays oriented normal to the Nadir Axis 

Response results when the solar arrays are oriented normal to the Nadir 

are summarized in Table 4.3.3.5-14. A comparison with the results shown 

in Table 4.3.3.5-13 indicates that, in general, the peak accelerations 

are lower when the arrqys are normal to the Nadir for a1 9 configurations 

except when the payloads and Orbiter are attached to the Station. 

Certain payloads require accelerations on the order of 10-5 g's. Table 

4.3.3.4-13 indicates that this requirement is only met when the Orbiter 

is attached. Hence it will be necessary for payloads to have an 

isolation system which allows them to function at very low acceleration 

levels. To ascertain the possibility of doing this, the design of a soft 

spring isolator and its analytical verification is discussed f +he next 

subsection. 

4.3.3.5.3.3 Isolation of experimental payloads 

In this section isolation from axial crew-kick-off is examined for a 

5000-pound materi a1 science payload (COM 1203), 1 ocated or, the 1 aboratory 

module. Isolation Ss provided by the use of a soft ;!)ring support for 

the experiment. Design of the isolation spring stf frness was determined 

by considering a simple Go-mass problem. Response of a Go-mss system 

to a step load applied to one-mass is well known. On the basis of this 

system and the requirement to have accelerations on the order of 10-5 

g's, the spring stiffness is approximately 5 pounds per foot. Verifying 

this isolation design through NASTRAN analysis of the IOC Station with 

the i sol ated experiment included, but no attached pqirl oads and Orbiter, 

reduced the acceleration level before l solation of 0.00015 g's to an 

acceptable level of 0.000012 g's . 
4.3.3.5.3.4 Response to RCS reboos t firing sequence 

The RCS reboost firing sequence used for the transient aciiysi s )presented 

herein is presented in Section 4.3.3.1. Each RCS is assumed to de:iver 

75 pounds of thrust. Two thrusters are located on the lower boola and two 

are located 175.5 feet up the keel. The RCS thrusters are assumed to 

fire in pairs in a staggered sequence thus rocking the Station in the 

pitch plane. 

Acceleration response time histories at selected Station locations are 

provided in Figures 4.3.3.5-49 through 4.3.3.5-52. Peak accelerations in 

g's are sunmarized in Table 4.3.3.5-15. These results ?ssume one-half of 

one percent (.5X) of critical damping in each mode. Tht RCS f-irizsequence for reboost can occur over one or more orbits. However, '3 the 

present study only 500 seconds of this sequence is simulated. The 

results indicate that acceleration levels are not increasing in an 

unbounded fashion during the first 500 seconds of response history. 

After about 500 seconds, the thrusters are turned off and the responses 

decay to zero. Furthermore, at Station locations far from the Station 

center of gravity, acceleration levels are generally higher due to 

reboost than due to other load cases. The peak acceleration at the upper 

boom, which accommodates several payloads, is nearly 0.025 g's. However, 

1 ow-g requirements for experimental payloads are relaxed during rebocst 

operations. 

Table 4.3.3.5-16 displays peak bending moments and torques at selected 

locations on the Station. This table also displqys margins of safety 

assuming a sdfety factor of 1.5 and strength values ieu to perform these 

calculatims are given in Table 4.3.3.3-1. All marg; 's of safety are 

pos:tive. As in the case of berthing, the lowest margins occur at the 

roots of the arrqy masts. They are somewhat lower for reboost than for 

berthing. (Compare Tables 4.3.3.5-11 with 4.3.3.5-16. ) Since there is 

considerable freedom in designing the RCS firing sequence and in 

adjusting the magnitude of the RCS t!mst, it is important in future 

studies to give special attention to margins of safety at the roots of 

the array masts during reboost as these mqy become dangerously low. 

4.3.3.5.3.5 Response to MRMS operations 

The Mobile Remote !4anipuiator System (MRMS) can produce loads during its 

operations when accelerating and decelerating payloads. In order to 

assess the Station Qnamic response to MRMS operations, it is assumed 

that a worst case occurs when the MRMS is located at one end of the upper 

boom and the arm of the MRMS is decelerating a payload weighing 10,000 

pounds. Deceleration loads for the Shuttle RMS am as gf ven In Reference 

18 are used after being scaled for a 10,000-pound pqyload. This results 

in applied loads on the Station of 17.2 pounds In the flight path 

direction and 877 foot-pounds of twlst torque about the negative Nadir 

axis. These loads are assumed to occur as step loadt with a 1 second 

duration. 
Peak acceleration mgni tudes in g ' s and amp1 i fication factors 

(acceleration of flexible station at speciffed location ratioed to 

acceleration of rigid station at the same location) due to the MRS 

operation, are summarlzed in Table 4.3.3.5-17. Pcceleration responses at 

a habitation module, center of the upper boom, tip of transverse boom, 

and tip of an array mast are provided. All of these are quite high. 

Peak acceleration response at the end of the upper ooom where the MRMS is 

located is found to be 0.0484 q's. This leads to an amplification factor 

of 33.9. The implication of this on the station perfomwcc is a 

potentitl control problem in performf ng MHMS activities on Flexible 

portions of the Station when payloads are not attached to the Station. 

It is probable that the presence of payloads in the vicinity of the MRMS 

operation is what plays the major role in reducing acceleration responses 

at the MRMS. 

4.3.3.5.3.6 Effect of bqy size on peak accelerations 

It is of interest to examine the sensiiivi ty of the station responses to 

the choice of various bay dimensions. In Figure 4.3.3.5-53 this is 

accomplished by considering peak acceleration responses at four station 

locations, namely, application point of applied load, center of upper 

boom, tip of transverse boom, and tip of an arrqy mast. Peak 

acceleration responses are presented for cubic bay dimensions from 1 to 

15 feet as amplification factors ratioing the actual acceleration at sew 

location to the corresponding acceleration at the same location were the 

entire Station fully rigid. These results were compiled using eatlliep 

valuer. for CEI and CGJ ~hich are now known to be too high. Nevertheless, 

these results provide trends and insight into the effect of variations of 

stiffness about the nominal 9-foot bay IOC reference configuration on 

Station response 1 eve1 s. 

As the bay dimension is varied the Station truss structural stiffnesses 

change in proportion to the square of the bay dimension. Only the array 

roast and radiator mast sti ffnesses remain unchanged. 

In the limit as the bay dimensio~ goes to zera, a11 parts of the Station 

become completely isolated from the rigid modules and are thus unaffected 

.y any loads applied ct the modules such as berthing and crew mtfon. 

Indeed the curves of Figule 4.3.3.5-53, for a1 l but the berthing port, 

confirm this trend down tc the bqy dimension of 1 foot. The dashed 

continuation of the curves to the origin represents the consequence of 

this postulate. The curve for the berthinq port, which is on one of the 

habitatiun modules, limits to a finite value as the bay dimension goes to 

zero. This value i 5 the response of the rigid modules due to the applied 

load ratioed to the response of an entirely rigid Station due '9 the same 

appl i ed 1 oad. 

As the bay dimension is increased from iero, the response of the berthing 

port decreases because nore of the mass of the Station is allowed to 

participate in the response. The decrease in this response is not 

dramatic, yoing from a ratio of about 1.55 at a 1 foot bqy 
dimension to about 1.2 at a 15-foot b~y dimensjon. Thus acceleratios 

levels at the modules are not very sensitive to truss stiffness. On the 

other hand, the response of all other locations rernoved from the modules 

increases rapidly as the bqy dimension is increased. 

As the bay dimerision is increased the m sdal frequencies of the Station 

undergo continuous change, while the frequency content of the applied 

pulse remains unchanged. For certair. bay dimnsions the Sacation modal 

frequencies will better 1 ineup with the frequency content of the applied 

loads and hence larger responses will r2sult. This helps to explain the 

rapid variations in response peaks as the-bay dimension is varied. 

For very large bay dimensions, the truss stiffness goes to infinity, and 

the curves should asymptote to limiting values. These limiting values 

will not be un;iv as though the Station were entirely rigid, since the 

array and radiator masts are always flex- 9le. Interestingly, a 15- foot 

bay diwnsion does nct in this sense closely approximate the rigid 

lintiting case. In order to ascertain the limiting risid str~c;ure 

values, results are shown for a 1000-foot bqy dimension with the 

understanding that this is entirely an academic situation with no 

practical application. Figure 4.3.3.3-53 indicat;.; that ampli fication 

ratios limit to values between 1 and 1.2 depending on Station location. 

The fact that very small bay dimensions leads to essential isolation of 

parts of the Station has benefits and drawbacks. It is beneficial that 

the acceleration levels and hence also internal l~ads are kept low in 

certain parts of the Station removed frm the mdule location. On the 

other hand, too small a bqy dimension can have the drawback of rendering 

the same parts of the Station uncontollable. At pcesent, no require~lent 

to precisely control portions of the Station far removed from the modules 

has been identified. However, if in the future this became a 

requirement, control of remote locations using actuators at the modules 

would be essentially impossible when the bay dimension is too low and 

consequently arr expensive distributed control system would probably be 

required. Thus it appears wise to avoid too ' ur a bqy dimension. Figure 

4.3.3.4-53 indicates that the minimum bqy dimnsion desirable in orde- GO 

avoid possible future control problems on remote Station locations 

depends on the location itself. The worst location is the upper boom. 

Responses at this lwntior! fall off rapidly when the bay diwnsion is 

less Van 8 feet. Consequently, it is wise to ksep the bay dimension 

above 8 feet to avoid possibie furr re control problems ' !stiitions 

rmte from the modules. 

4.3.3.5.3.7 Concluding remarks 

Preliminary analytical results for the dynamic response of the reference 

Space Station configuration have been presented and di scussed. 

Disturbances due to Orbiter berthing, crew motion, RCS reboost firing 

sequence and MRMS operations have been examined. The Station with and 
without pay1 oads and/or Orbiter was considered and arrqys were oriented 

either normal to the flight path or now1 to the Nadir. Time histories 

and sunrrary tab'les of peak accelerations, amp1 if ication factors, peak 

bending ~nomnts and torques and mrgins of safety using a safety factor 

of 1.5 were presented and discussed. 

Considering a1 1 four load cases, maximum peak accelerations occur at the 

tip of the array mast due to RCS reboost firing dith a magnitude of 0.028 

g's. This led to relatively high internal bending moments at the root of 

an arrqy mast with a margin of safety of 4.13. Peak accelerations also 

at an array mast root due to berthing were not much lower and 'Se 

corresponding margin of safety was 4.71. Since neither tne RCS firing 

sequence for reboost nor the bertning I aaas are very we1 ; dciined at 

present, future studies with better defined applied loads should give 

special attention to margins of safety at the root of the array masts. 

Furthermore, arrqy mast flexible deformation from root to tip did not 

exceed about 4 inches. 

It was found that crew motion leads to exceeding the 10-5 g order of 

magni tt~de accaleration requirement for the laboratory module. However, a 

soft spring -- lator was designed for a 5000-pound experimental payload 

and its performance was verified in the full iinite element analysf s. 

This led to an acceptable acceleration level of 0.000012 g's for the 

isolated experiment. 

Operations of the MRMS at one end of the upper horn produced relatively 

large dccelcrations a, tile MPYS location itself. This needs future 

attention as it could lead to problems in controlling the MWS 

operations. 

An examination ot the effect of bay dimensiun on Statior! response to 

Orbiter berthing indicated that the bw dimension would have to be 

extremely large to provide dynamic responses approxir~~ting the 1 imf tfng 

case of a rigid keel. Even the use of erectable :5-f00t bays would fall 

far shorr of achieving this. The reference config1,ration with its 9-foot 

bays is therefore a very flexible st~.~~ctr?e. It appears to be able to 

met dynamic pe:formarlce requirements as Uley are presently defined. 

However, it i a notea herein, that if additional performance requireaents, 

such as on payloads located some distance from the station modules are 

later imposed on he Station, an add! ti,rial control system would become 

necessary. Dynamic r.osults indicate that a bay dimensi-? much less than 

9 feet would probably require a distributed control 5ystem. 

Rigid body control requirements for si operational configurations of the 

IOC Station were determined. The conf iguratians di ffered accordf ng to 

;he number and type of i tems attached to the basic Station. These items 

were iile Orbiter, payloads and servicing squipme.lt. In establ f shing the 

requirements, the orbital altitude range was varied between 220 and 
270 mi and the cui-rasponding atmspheres were characterized by 2 sf* 

high densities which varied over ara >?hit and rotated with the Earth. 

The analysi s results indicate that four eq~ivalent Skylab CMG's ri 11 

provide sufficient lloRntum storage and toque to control the starf;.: in 

the natural envi ronment (grabi ty-grz5ient and aerodrag) for tk 

prescribed altitude range of 25u to 270 mi. SORE secular noslentua 

buildup occurs on the rol l and yaw axes which must be nulled. Three 

different methods for doing this were discussed. Secular mentun 

buildup is precluded an the pitch axis by flying with a small pitch angle 

(less than 10') such that the angular momentum accumulated over an orbit 

due to gravity-gradient is the negative of that due to aerodynamic 

torque. 

A limited number of system failure cases we- investigated. Under the 

rather severe assumption that all control s: ,tetns failed, the Station 

rotated about the keel and started to swing in a pep~~lium manner. A 

subset of this case assumed that a1 1 coqtrol was 1 ,,t except for ;ire 

z-axis (yaw). In this case, 'he Station oscillated in the orbit (pitch) 

plane. The case of a failed alpha joint was also investigated. The 

nuder of CMG's required to provide angular impulse to bring the solar 

arrays to the Station (keel 1 rotational rate varied from two to three 

depending on the assumed fai 1 ure conditions. Th. associated secular 

momentum buildup was about three times the nominal value. 

A cursory investigatio~ of the control requirements due to induced 

environmer.ts was also perfor~ned. Torque generated by crew lwtfon exceed 

the torque capability of four CMG's ?nd requires about one-hal; +he 

rnolsentun storage capability. Torques and momentum due ts 

dockingberthing and moving of payloads with the MFNs were estimated to 

be two to three times the capability of four CMG's. This wwld seem to 

indicate that control of the Station doring bertning/docki ng and 

subsequent MRMS activity should be accomplished witn the RCS or possibly 

the control requirements should be reduced. 

Orbit decay and reboost c3lculations for the maximn; and minimum mass IOC 

configuration were performd. For the highest atmosphere densities and 

most severe drag-to-mass ratios, the IOC Station deccvs from a 270 nmi 

orbit to 250 nmi in 90 days which mets the system rquirement. The 

reboost propzllant required for a 90-day decay period and highest mass 

configuration is about 3600 lbs. for a propeilant having an ISP of 220 

seconds. Flexible body attitude control dynamics studies were made for 

the IOC configuration without the Orbiter, pay1 oads, and servi cing i tens 

added. 

The control frequency was set otle decade below the lowest bending 

frequency. Vibration modes jn the analyses were included to 

approximatsly one decade above the lowest mode. The peak pcinting error 

due to crew-kick-of' was less than 0.02' at the navigation base. The 

maximun settli ng time was 150 seconds. Similar Ci sturbance responses for 

decking/berthing were lcss thaa 2' and 400 seconds, respectively. 
Associated frequency response investigations show adequate gain and phase 

margins dnd no significant control systemlflex body interaction was 

observed. A simple RCS control systea was designed and torque histories 

which simulated a four RCS jet orbital reboost were also generated. 

These were used In an uncontrolled ?esponse analysis to investigate 

structural 1 oadf ng. 

Structural enamic models of both the IOC and Growth Station 

configurations were developed and utilized for the flexible control and 

transient response analyses. Continuous beam representations were 

employed for the lowr, transverse and upper beams, keel extension, and 

lower and up?er keels. Pqyioads were modeled as lunped masses and the 

station modules were treated as rigr'd beam faeders rigidly connected to 

one anolher in the acetrack arrangeent. No atterpt was made t~ del 

nonlinear effects such as those potential ly assori ated with the large 

rotating joints. The frequency of the fundamental vibration node of both 

the IOC and Growth Station is approximately 0.1 Hz. The lower frequency 

modes are associated with coupled motion of the solar arrqys and keel for 

the IOC configuratior: and with keel twist for the growth configuration. 

Transient response results for disturbances from Orbiter berthing, 

crew-kick-off , RCS reboost and MM operations were developed. bxfu 

bending moments at the base of the solar arrqy =st were well below the 

allowable. Arraj tip maximum deflections were on the order of 4 inches. 

Acceleration responses due to crew-kick-off exceeded the 1.0 E-5 g 

requirement in the laboratory module. Consequently, a soft spring 

isolator was designed for a 5000-pand experimnt located in the 

laboratory module. This led to at1 acceptabla acceleration level of 1.2 

E-5 g for the isolated experiment. 

An examination of the effect of bay dimension on Station rspo ,° to 

Orbiter berthing indicates that the keel i s far frm ri gid. Increasing 

the bay width to 15 feet did not significantly change the response 

level s. 
TABLE 4.3.3.5-1 - SUMURIES FOR IOC 

SPACE STATION PER ORBIT 

ALT CYCLIC SECULAR TRIM CUG' S RCS 

WNTW MOKNTUI ANGLE HNTW 

VECTOR VECTOR VECTOR 

(HI) (ft-lb-s) (ft-lb-s) degree (ft-lb-s) 

xlooo xlooo xlooo 

IOC: 

270 0,9.5,.9 OS0,-9 -2.2 2.3 1,21 ,O 

250 .2,13,1.3 -2,0,1.3 -3.2 2 -8 2,30,0 

220 1,23,2.7 1,0,2.7 -6.1 5.1 3,57,0 

IOC +Orbiter: 

270 019.2,1 l,O,l 5.66 2.3 1.8,23,0 

250 4,12.?,16 .2,0,1.6 4.88 3.4 2,35,0 

220 1,22.5,3 0,0,3 2.53 6.3 4,69,0 

IOC + Payloads {Service +Orbiter: 

270 2,8,2,3.8 O,O,l,O 4.97 2.3 10,20,2.5 

250 2, 11,4.0 0,0,1.6 4.74 3.0 10,31,2.5 

220 2.4,17.5,5 0,0,3.0 4.05 5.3 11,60,2.5 
TABLE 4.3.3.5-2 - JOINT' FAILURE SWV RESUiTS 

ALT CYCLIE SECULAR TRIM CMG'S RC S 

MOMENTUM MOMENTUM WGLE MM NTUH 

(MI) VECTOR VECTOR VECTOR 

( FT-LB-S) (R-LB-S) DEGREE (FT-LB-S) 

xlOO0 xlOO0 xlOO0 

IOC - SINGLE ALPHA JOINT FAILURE PERPENDICULAR TO KEEL: 

2 70 5s5,5 2.5,0,1.7 - 1.7 1,20,1 

IOC - SINGLE ALPHA iOINT FAILURE PARALLEL TO KEEL: 

2 70 lS9.5,3 OS0,3 - 2.45 2,2C,7 

IOC - DOUBLE ALPHA JOINT FAILURE PARALLEL TO KEEL: 

273 .4,9.2,1 .lsOsl - 2.5 1.5,24,0 

TilSLE 4.3.3.5-3 - MASS PROPERTIES SUWRY 

FLIGHT 1 FLIGHT 2 INERTIAS (ku-d yL6) 

I XX 

IYY 

I zz 

I XY 

I XZ 

IYZ 
TABLE 4.3.3.5-4 - WOMENTW STORAGE REQUIREKNTS FOR REFERENCE 

COWFIGURATION AFTER FIRST TUO LAUNCHES 

CYCLIC SECULAR 

MDWNTUM HOENTUM TRIM 

VECTOR VECTOR ANGLE 

( nmi -sec ) ( nuti -sec ) - (DEG) - CFIG'S 

FLIGHT 1 252,10,155 45,0,10 0 1 

FLIGHT 2 29,0,34 4,0,6 5.01 1 

TABLE 4.3.3.5-5 - ESTIMATED ORBIT DECAY AND REBOOST CHARACTERISTICS 

Ini tal Fi :la! Reboost Decay 

A1 ti tude A1 ti tude Propel 1 ant Tine 

( nmi 1 ( nmi 1 (Kg) ( Dws 1 

Min/Max Mass Min/Max Mass Min/Max Mass 

Maxilwm density - 8/1991 launch 

300 291/292 540/780 90/90 

270 250/253.5 1190/1620 90/90 

258 210/211 2420/2880 90/90 

300 223/220 4840/79 70 386/478 

270 220/220 3000/4950 154/ 190 

250 220/220 1790/2960 78/91 

Minimum density - 4/1994 launch 

300 296/2%.3 250/340 9o/m 

270 260/262 610/790 90/90 

250 232/236 1070/ 1380 90/90 

300 220/220 4840/7970 1965/2169 

270 220/220 3000/4950 360/507 

250 220/?20 1790l2960 151/188 

Reboost ISp = 220 lb-sec/lb 
TABLE 4.3.3.5-6 - CONFIGURATION DATA 

Ixx - 47510000 ~g m2 

IYY - 38570000 Kg m2 

Izz - 13160000 ~g m2 

No. of Modes: 

X Axis 13 

Y Axis 21 

ZAxis 21 

Frequency Range: 0 .09 - 1.07 Hz 

Control Frequency: 0 .0094 Hz 

TABLE 4.3.3.5-7 - DISTURBANCE TORQUE DUE TO CREW MOTION 

DISTURBANCE TORQUE X Y Z 

To (NT - M) 281 3 2813 508 

POINTING ERROR 

(OEGREES) 

SETTLING TIME 

(SECONDS) 
TABLE 4.3-3.5-8 - DISTUBRANCE TORQUES DUE TO BERTHING 

DISTURBANCE TORQUE X Y 

TD 49506 49506 

POIkTING ERROR X Y 

( DEGREE) .42 .5 

SETTLING TIME 100 400 

( SECONDS ) 

TABLE 4.3.3.5-9 - AUTOPILOT AND CONFIGURATION DATA 

Kg = l.Odeg/(deg/s) 

DB = 1.Odeg 

h = Ot00-ldeg 

TL = 10,575 ft-lbs 

Tg = 4,500 ft-lbs 

I = 70,837,100 slug ft2 
TABLE 4.3.3.5-10 - PEAK ACCELERATIOW RESPWE IN G's AT SELECTED 

STATION LOCATIONS DUE TO ORBITER BERTHING FOR 

SOLAR ARRAYS ORIENTED EOW TO THE FLIGHT PATH 

LOCATION BERTHING CENTER TIP OF TIP OF 

PORT OF TRANSVERSE ARRAY 

UPPER BOOM MAST 

CASE BOW 

Y I THOUT 0.00476 0.0162 0.00460 0.00123 

PAY LOADS ag1.70 as5.21 az4.48 a-5.05 

Y ITH 0.60432 0.00280 0.005% 0.00126 

PAY LOWS as1.77 as2.44 ax92.3 a=17.5 

KEY: a = anplification factor=(accel . of flexible station )/(accel . of 

rigid station) 
TABLE 1.3.3.5-11 - PEAK BEHDI016 mMWS AND TORQUES IN FOOT-LBS 

AT SELECTED STATION LOCATIONS DUE TO ORBITER 

BERTHING FOR SOLAR ARRAYS ORIENTED NOW& TO 

THE FLIGHT PATH 

LOCATION KEEL EXT. TOP OF TRANSVERSE ALPHA ARRAY 

ROOT UPPER BOOM JOINT MAST 

KEEL ROOT ROOT 

CASE 

Monent: Moment: Moment: bnt: hnt: 

6930, 0 .OQ4 2890. 1172. 406. 

UITHOUT M.S.=7.15 M.S.=* M.S.=18.5 M.S.=# M.S.=4.71 

PAY LOADS Torque: Torque : Toque: Tarque: Torque: 

350 . 0.0 i009. 1008. 0.16 

M.S.=67.6 M,S.=* M.S.=22.8 M.S.=# M.S .=865 

Monent: Moment: Mown t : Momn t : Momn t : 

6990. 248. 712" - 3170. 381. 

WITH M.S.=7.08 H.S.=227. M.S.16.93 M.S.=O M.S.=5.09 

PAYLOADS Torque: Torque: Toque : Toque: Torque: 

510. 728. 2500. 2400. 0.0 

M.S.=46.1 M.S.=32. M.S.=8.6 I4.S .=# M.S.=* 

KEY: M.S.=Margln of Safety=(strength)/(l.5 x mmnt or torque) - 1 

*: Margin of Safety exceeds one mill ion 

#: All owable strength for alpha jot nt presently unavailable 
TABLE 4.3.3.5-12 - PEAK ACCELERATIOO RESPONSE IN g's AT SELECTED 

STATION LOCAT IONS DUE TO ORBITER OERTHlNC 

FOR SOLAR ARRAYS ORIENTED NOW TO TllE NADIR 

LOCAT ION BERTHING CENTER TIP OF TIP OF 

PORT OF TRANSVERSE AHRAY 

UPPER BOOM MAST 

CASE W)W 

W I THOUT 0.0044 0.0125 0.06275 0.00275 

PAY LOADS a=1.52 a=4.46 d=3.09 6=3.09 

W lTH 0.0041 0 .OW7 0.0056 0.0056 

PAY LOADS a=1.71 a=1.14 a=18.1 a=lH.l 

KEY: a=amplificdtion factor=(flexible reoponse)/(rigld station 

response 1 
TABLE 4.3.3.5-13 PEAK ACCELERATIOW RESPONSE IN go s AT SELECTED STATIW 

LOCATIONS DUE TO AXIAL AN0 TRANSVERSE CRUJ KICK-OFF 

FOR SOLAR ARRAYS ORIENTED NDM'ci TO THE FLIGHT PATH 

LOCAT ION HABITATION 

WDULE 

-4 

Y ITHOUT AXIAL KICK 1.49 X 10 

PAYLOADS as1.23 

AND -4 

ORBITER TRAMS. KICK 2.35 x 10 

a-1.54 

CENTER OF 

UPPER 

POW 

-4 

5.37 X 10 

aa1.0 

-4 

8.07 x 10 

a4.2 

TIP OF TIP OF 

TRANSVtRSE ARRAY 

BOOM HAST 

-4 -4 

1.5X10 5.47X10 

?=?. 5: a4.46 

-4 -4 

1.51 x 1' 5.16 x 10 

a=i .94 as4.2; 

-4 -4 -a 

WITH AXIPL KICK 1.10 x 10 1.25 x 10 2.68 1; 1 i.18 x 10 

PAYLOADS a-1.39 as2.18 a42.7 ,,1f7.? ONLY 

-4 -4 -4 -4 

TRANS. KICK 1.86 x 10 1.42 x 10 3.10 x 10 6.89 x 10 

as1.52 at2.47 a=95.4 as19.1 

- 5 -4 - 5 -4 

Y ITH AXIAL KICK 5.68 x 10 1.88 x 1C 6.77 x 10 2.03 x 10 

ORBITER as1.08 a363.1 as4.27 as55.9 ONLY 

-5 -4 -5 -4 

TRANS. KICK 5.34 x 10 1.92 x 10 7.42 x 10 2.54 x 10 

a=1.01 ~64.3 az4.69 a49.8 

-5 - 5 - 5 -4 

Y ITH AXIAL KICK 4.91 x 10 2.27 x 10 9.78 x 10 2.01 x 10 

PAYLOADS a=1.0 a=3 26 ar8.08 az94.6 

RID - 5 -5 -5 -5 

ORBITER TRANS. KICK 4.56 x 10 1.16 x 10 5.50 x 10 1.62 x 10 

az0.94 af1.67 a=B .5 ax76.2 

KEY: atamp1 i ficatlon Pactor=( accel . of flex! ble station)/( accel . of 

rigid station) 
TABLE 4.3.3.5-14 PEAK ACCELERATIOH RESPONSE IN 9's AT SELECTED 

STATION LOCATIONS DUE TO AXIAL CREW KICK-OFF 

FOR SOLAR ARRAYS ORIENTED NORMAL TO THE NADIR 

LOCAT ION HABITATION CENTER OF TIP OF TIP OF 

MODULE UPPER TRANSVERSE ARMY 

BOOM BOOM WST 

-4 -4 -4 -4 

WITHOUT AXIAL KICK 1.45 X 10 2.6 X 10 4.0 X 10 3.5 X 18 

PAY LOADS az1.58 a02.89 az4.44 a~3.~ 

AND 

ORBITER 

-4 -4 -4 -4 

WITH AXIAL KICK 1.35 x 10 1.17 x 10 2.1 x 10 2.5 x 10 

PAY' ZIADS ar1.96 a=1.77 am3.09 a=3.42 ONLY 

- 5 -4 - 5 -4 

WITH AXIAL KICK 5.3 x 10 1.55 x i0 2.1 i. 10 1.6 x I0 

ORBITER a.1.06 ax2.91 az4.17 az2.60 ONLY 

- 5 - 5 -4 -4 

W ITH AXIAL KICK 5.4 x 10 8.8 x 13 1.3 x10 1.4 x10 

PAY LOADS ap1.32 a14.1 aa4.1 a=3. le 

AND 

ORBITER 

KEY : a=ampli fication factor=( accel . of flexible station)/( accel . of 

rigid station) 
TNLE 4.3.3.5-15 - PEAK ACCELERATIOY RESPOYSE IW 3's 

AT SELECTL3 STAYiON LOCATIONS DUE TO 

RCS REBOOST FIRING SEQUEMCE 

LOCATION dABIT4T ION CENT€R TIP OF TIP OF 

WDULE OF TRANSVERSE ARRAY 

UPPER BOOn MAST 

CASE BOOn 

W ITHOUT 0.0034 0.024 0.012 0.028 

PAY LCADS 

WITH 0.0024 0.0062 0.003C 0.0037 

PAYLOADS 
TABLE 4.3.3.5-16 - PEAK 6ENDING mhiS AND TORQUES IN FOOT-LBS 

AT SELECTED STATION LOCATIONS DUE TO 

RCS REBOOST FIRING SEQUENCE 

L.OCATION LOWER KEEL ROOT OF 

ROOT LO'dER 

ROOM 

CASE 

Moment: Moment: 

9,134 12.6 

WITHOUT M.E.=8.62 M.S.=* 

PAY LOADS Toque: Toque: 

0 0 

M.S .=* M.S .=* 

TRANSVERSE 

BOOM 

,ROOT 

Torque; 

942 .O 

H-S.=24.7 

ALPHA ARRAY 

JOINT %ST 

ROOT 

Torque : Toqlre : 

856 .O 0 

M,S.=# M.5 .=* 

Moment : Moment : Momerl t : Nomn t : Moment : 

15,820 20.8 5,260 '1,750 357.0 

WITH M.S.=8.36 M.S.=2700. M.S.=9.70 M.S.=# M.S.=3.65 

PAYLOADS Toque : Toque : Torque: Toque: Toque: 

1,925 0 1,660 1,660 0 

M.S.=9.26 M.S.=O M.S .=13.6 M.S.=# M.S.=* 

KEY: M.S .=Margin of fafety=(strength)lll.5 x moment or torque) - 1 

+: Margin of Safety exceed: one mi: 1 ion 

#: Allowable strength for alpha joitit presently unavailable 
TAME 4.3.5-17 - PEAK ACCELERATIOY RE- I4 6's AT SELECTED 

STATION LOCATIONS DE TO MRHS OPERATIONS ON 

THE UPPER BBOn 

LOCAT ION HABITAT ION AT TIP OF TIP OF 

WDULE WS WSVERSE ATTAY 

80811 MAST 

CASE 

WITHOUT 0.000550 0.0248 0.0202 0.00070 

PAYLOMS a-5.7 ar33.9 ax22.0 a213.8 

YITH 0.00015 0.00030 0.0026 0.0050 

PAYLOADS ax1.15 a21.25 az17.6 itr24.L 

KEY: a=WLIFIChTIOY FACTOR=(,KCEL. RESPONSE OF FLEXIBLE STATION)/ 

(ACCEL. RESPONSE OF RIGID STATION) 
1 Eg$F:E I-= FREDESIGN 

CYCLIC MOMENTUM I 1-J 

CALCULATE 

ENVIRONMENT 

-TORQUE AND F3RCE 

ATMOS DRAG 

GRAV GFlAD 

CTHEF? 

-- 

MOMENTUM 

- DESATURATION 

ORBIT ORBIT/STATION KEEPI

REACTION THRUSTERS 

ONLY ACS 

Flgure 4.3.3.5-1 - RIGID-BODY CONTROL DYNAMICS MODULE (RCD
EARTH 

DIRECTION 

OF 

FLIGHT 

\ 

POP 

Flgure 43.3.5-2 - COORDINATE SYSTEM FOR RIGID BODY Sllf lLATlON 
H=25QNM, I NCL=28.5,THP=-1.59,ARRAYB~O.O,KR=l.JOdES,KA=5l72 H=ZSONM, I NCL~28.5,THP=-1.59,AR

SS POWER TOWER/IOC+PAY- JUNE 25TH SS POWER TOWER/IOC

Flgure 4.3.3.5-3 - SPACE STATION 

LONGITUDE, LATITUDE, 

AND LOCAL SUN TIME 

HISTORY 

ORBIT POSITION IN REV. C44 SI*rrcluru Jul K U4 

Ch?BIT POSITIO

Flgure 4.3.3.5-4 - TOTAL I

IMPULS

AN INER

ORBIT C

SYSTEM 
H-2SOW1. 1 NCL =~~.~,TH~=-I.$~.ARRAYB=O.O,KR~~.~~~~,KA-~~~~ H=2SONM, I ~L-28.5,~HP=-1.59,ARRAY8

ss POWER TDWER/IOC+PAY- JUNE 25~t-1 SS POWER TOWER/IOC+PA

ORBIT POSITION IN REV. 

fS4 Sbwlurn &I 1\ W 

Figure 4.3.3.5-5 - RESULTANT 

AERODYNAMIC DRAG 

FORCE HISTORY WRT 

BODY COORDINATE 

SYSTEM 

ORBIT POSITION 

Flgure 4.3.3.5-6 - ABSOLUTE 

EXTERNAL 

BQOV COO

SYSTEM 

H-SSONM. I NCL-28.5.THP=-1.5O,ARRAYB=O.OOKR=1.306ESJc~ -5172 H=25ONM. I NCL=~~.~.THP=-~.SQ,AR

SS POWER TOWER/IOC+PAY-JUNE 25TH SS POWER TOWER/IOC+PAY-J

ORBIT POSITION IN REV. tS4 Struclurm Jul K TMM 

OU~IT POSITION IN R

CS4 

Figurm 4.3.3.5-9 - AERODYNAMIC TORQUE 

HlSTORiES WRT BODY 

COORDINATE $YSTEM 

Figure 4.3.3.5-10 - TOTAL AERO

DRAG IMPUL

WRT' BODY 

COORDlNAT
H=250NM. INCLs28 5.THP=-1.59,ARRAYB=O.O,KR=1.306E5,KA=5172 ~=~~ONM.INCL=~~.~.THP=-~.O~~,ARR

SS POWER TOWER/IOC+PAY- JUNE 25TH SS POWER TOWER/IOC- J W 25TH. F

@ Y.0 

U 0 

t2 0 

.- 

k 

u; 11.0 

z 

W 0.0 0 

0 

I r.0 

4 

5 6.0 

h, 0 

s; 2 7 0 W 

a 

6 0 

5.0 

4 0 

0 0 0.1 0.2 0 3 0 4 0 5 0.6 0.7 06 0.9 LO Ll 0.0 t.3 1.0 1.5 2.0 2 3.J 3

ORBIT POSITION IN REV. ORBIT POS!TION IN 

LS4 Slrrlurom Jut It DH EY 

Flgura 4.3.3.5-1 1 - AERODYNAMIC DENSITY Flgure 4.3.3.5-12 - YAW An!; 

AS A FUNCTION OF HISTORY OF 

ORBIT POSITION WRT A ROTA

COOROMAT
H-250NM. I NCL=28.5,TW~-J.O75,ARRAYB=O.CI.ARRAYS FEATHERED 

SS POWER TOWER/IOC-JUNE. 25TH. FREE DRIFT ALL AXES 

Figure 4.3.3.5-13 -. ROLL AND PiTCCC 

HISTQRY OF KEEL WRT A 

ROTATING ORBIT 

COORDINATE SYSTEM 

I NCL=28.5.THP=-2.2,A~AYB

SS POWER TOWER/IOC-JUYE 2

Flgun 4.3.3.5-14.- ROL

ATTI

A RO

COO
H-250hM, I NCL=28.5.THP=-5.075.ARRAYB-O.O.KR-I.306ES,KA=5172 

SS POWER TOWER/IOC-JUNE ZSTH, 2 AXES CONTROL ONLY 

Figure, 4.3.3.5-15 - ROLL, PITCH, AND YAW 

AWITURE Hl8TORY WRT 

A ROTATINQ OR8lT 

COORDINATE SYSTEM 



250 - 

MIN WEIQHT 

(IOC STAT ION WITHOUT 

r' 

z 

w0 I50 - UUN

3 

E 

(" 

100 

60 

- 

- 

1 1 1 1 0 100 m -

3m roo #K, 80

TIME, DAYS 


. . 

LAUNCH 4/1

- 

- 

- MIN WEIGH

(IOC STAT

WITHOUT 

AND SATE

SERVICING

- 

I I 1 I 1 1 

0 ,400 800 1200 1600 2000 2400 

TIME, DAYS 

Figure 4.3.3.5-19 - ALTITUDE DECAY FOR MINIMUM DENSITY ATMOSPHE
i2 8 F 8 .- 8 0 

INN '30fl1111V 

Flgure 4.3.9.620.- X AX18 CREW Dl8TURBANCE 






Figure 4.3.3.5-27 - Y AXIS CREW DISTURBANCE TORQUE LIMITE
OMGtP(AL P:.%-;. .3 

OF POOR QLIALi'iV 
0 20 40 60 80 lam 120 140 160 1

TIME, Sec 

Flgun 4.3.3,s-29.- X AXIS BERTHINO DISTURBANCE TORQU
. . 

. . 

0 M MI W 120 t 190 210 2W 270 JDO 

TIME ,Sec 

Figuro 4.3.3.5-30 - Y AX18 BERTHING DISTURBANCE TORQUE LIMIT





ORIGINAL FiX.; ,% 

OF POOR GLIALlW 



POWER TOWER REORB IT BURN TiME HISTORY 

HYSTERS I S-0.05 DEGREES 

6.2 6.4 6.6 oa io 

ATTi TUOE ERROR, DECREES 

rbun 1.9.9.5-40 - PHASE PUW 

3E 

0 

0 

m 

60 

az 0- =E 8E 

am 

Wa I-w 

It 

oL so :P 

g: 

a0 

(ID 

w= a2 

5li 

YE 

(IDz 

sx 

I- Z 



E-2 WITHOUT PAYLOADS AND ORBITER 

1.5 

WlTH PAYLOADS 

WlTH ORBITER 

WITH PAY LOADS AN0 CJR

-4 % 

0 20 40 80 80 100 0 20 10 60 

TIME (SEC) TIME (SEC) 

FIQUW 4.3.9.8-46 TRANSIENT RESPONSE AT CENTER OF UPPER BOOM 

IN X DIRECTION DUE TO AXIAL CREW-KICK-OFF 


0.08 I LOWEP RCS I, I, 1 1 1 

-0.08 I I 1 I I 

0 200 400 600 800 lo00 

TIM. SECONDS 

Figure 4.3.3.5-49 ACCELERATION TIME HISTORIES IN FEET/SEC~ AT THE 

AND LOWER RCS LOCATIONS DUE TO RZS REBOOST FI

FOR loe STATION WITHOUT ATTACHMENTS 


0.X 0.8 

h) 

Q, 

W 0.15 0.6 

A 

A 

C 0.10 

C 

0.4 C 

E 0.05 E 

L L 0.2 

0 E 

F? R 

A -0.05 

A 0 

T T 

I ' -0.2 

0 -0.10 0 

N hr 

-0.15 4.4 

-0.20 -0.6 

0 200 40G 600 800 to00 0 200 400 8

TIME, SECONDS TIME, SECON

Flgure 4.3.3.5-52 ACCELERATION TIME HISTORIES IN FEETISEC~ AT THE SELECTED 

LOCATIONS DUE TO RCS REBOOST FIRING FOR IOC STATION 

WITH ATTACHED PAYLOADS 

-0.20 L. - 1 1 e 

200 400 600 800 r 000 

TIME, SECONDS 

TIME, S

I 4.3.5-52 ACCELERATION TIME HISTORIES IN FEET/SEC~ AT THE SELE

LOCATIONS DUE 10 RCS REBOOST FIRING FOR IOC STATION 

WI'.'H ATTACHED PAYLOADS 

4.3.4 .. Power Capabil i ty and Util iuatic~n 

4.3.4.1 Power capability 

The electrical power system consists of a generatiun subsystem, storage 

subsystem, and management and distribution subsystem. The subs~stems are 

of mdular desi gn to facil i tate repai r/replacent and growth. The 

jnitial statior! power system may be either solar dynamic or photovoltaic 

based. The growth station will be solar dyna~ic. 

Tbe power system provides c?ettrical power to support station activities 

and is ca9abl.c of accepting new technology as it becomes available, or as 

it is needed. The power syste~ must accomplish these objectives while 

simul taneously mi?imizing adverse effects on station altitude and 

attitude mintenance arld interference with Space Station customers. 

The pGwer system integration with other users of the Space Station is 

primarily through electricai interfaces. The identification of these 

interfaces helps determine the requirements and responsi bil i ties of both 

th~ power system and the Space Station users. Electrical power produced 

by the solar 3rrays or the solar dynamic power system is transferred 

across beta joints and distributed to the solar photovoltaic energy 

stcrage subsystem and management and distribution elements outboard gf 

the alpha joints. The DC power from the generation and storage 

subsystems (or AC from the solar dynamic power system) is converted to 

high frequency AC power and transferred dcross the rotary a1 pha joints. 

Power is then Ci stribdted viz power 1 ines a1 ong the upper and i ower keel 

to varfcii; S;;;zce Station users. This power is distributed to attached 

pay1 oads and comnunication antenn;: aim2 the upper boom. Likewise, 

power is distributed to the reaction rsr~ii'?l system c~tiuies, mobile 

manipulator system, them i :~nerol system, Space Station modules, and 

other .itt.arkA ",;j';uads on the lower keel. The managemeat and 

di stri bdtion subs.vstem controls the power and p:*ovi des i ndf vidual 

protection for each user interface, power source, and subsystea hardware 

and wiring. A properly designed management and distribution subsystem 

can functisn with several different generation subsystems, th~s 

accommnddti ng Space Station growth. 

The arzs requirenwnts and configuration of the power generation subsystem 

have a majoi* effect on power system integration with the station. The 

dray force experienced by the station is directly related to thc area 

rquiremetits, and the drag force affects the orbit altitude of the 

station and the amount of propellarfts required for altitude riintcnance. 

As tRe power system srcws in size and power requirement, the Swpact of 

drag forces wiii become more severe. Also, in the Space Zbati0.1 

reference configuration most of the station w~ss is near the bcttm uf 

the keel and most of the d~ag force is near che top. S~al ler area 

requirements will thus ailw the station to experience snaller 

aersdynamic torque and to fly mre near1 y vert4i ca! . The genrraticn 

subsystem conf i ~uratim was selected to mi nimi.re grsri ty gradient roll 

torques at high beta angles. The spacir~g of the generation subsystem 
elements frola the main boon and from each other was based primarily on 

shadowing and viewing considerations. This spacing also affects the 

Space Station keel length, because interference by the volume swept by 

the generation subsystem elements must be avoided. 

4.3.4.2 Power utilization 

4-3.4.2.1 Introduction 

The issue of power system size and capacity is fundamental to size and 

operational capability of the Space Station. There is a natural desire 

to be 'power-rich' .~t the inception; however, cost considerations dictate 

a compromise which will require close modeling of power use. The pmr 

utilization anr. ,ysis is an attempt to provide realistic time hf stories of 

power consuaq,t~on for each sbbsystem. 

The technique rquires an electrical equipnent list which defines the 

power draws for various operating mdes for each ~iece of electrical 

equiplaent. In addition to a crew tineline, a definition of eqcipllcnt 

oper8tion is required in fhe form OF activity blocks. These blocks 

correspond to a time period for which a particular activity is 

performed. Another set of input data &fines the 1 ink bebeen tiis pmer 

levels that exist fcr eqiri?~r?* ?eration and activity 5locks and 

timelines. The resuiting ana'ij+ls GTdui~i =. variety of speciF3c. 

information which descriues the powr corlfigur,;t:dn of the Space SbSion 

at any point in time and the associat~d statistics for the syste~. 

This technique has been developed and utilired on the Shuttle. The 

technique has been applied previously to a Space Station design stuw and 

a study for an eight-win, 120kW station which preceded the current studye .- 

An initial output was a set of f~~i- guidelines (Table 4.3.4.2-1) for the 

use in the developnent of tke subsystems. 

4,2.C.Z.2 ~eference crew activfty -

The two 12-hour shifts provide a nominal operation of 24 hwzr a day for 

the crew of six with a 6-day workweek for each crerrlllenber. The crev 

complement is two station specialists and fwr mission specialists. The 

Station specialists work 9 hwrs on iqeratians, 8 hwrs on overhead 

activities such as eating, exercise, planning, and personal time. Iht 

mission specialist spiit of activities is 9 hours and 7 hwrs. All 

crewembers hav~ r3 hours for sleep. 
TABLE 4.3.4.2-1 WIDELIMES FOMUUTED BY THE POYR UTILIUTIOU GROUP 

1. DON'T LEAVE REDUNDANT SYSTEMS POWERXI, ROTATE THE USE OF Enr_.Ymm 

STRINGS YHERE POSSIBLE 

2. USE LOU WATTAGE AREA L16HTING AND APPROPRIATE TASK LIMING FOR UOW( 

AREAS 

3. DON'T USE ELECIRICITY TO HAKE HEAT IF THERE IS A COST EFFECTIVE YAY 

TO USE EXISTING HEAT SOURCES 

4. THE STATION WILL SUPPLY CONDITIONED LirIlITY GRADE POYR. DO WOT 

SPEND POWER AND EIGHT TO RE-CONDITION THE POER AS ON ORBIER 

(12% OF POYER) 
4.3.4.2.3 Analysis considerations 

The power analysis has certain assumptions which affect the results: 

a. The power is provided to the subsystems and p~iyloads at 

the bus, which means that the power distribution and control subsystem 

accounts for the inefficiencies associated with the delivery end 

conditioning of the power. 

b. Interior low-level area lf ghttng is provided for normal 

transit through nodules. Dedicated task lighting is pmv$ded for worir 

places at 125 watts per person per task. 

:. One active work station is provided for each crewcan, 

d. A significant portion of pvload activities would be 

onboard computer or ground controlled. 

4.3.4.2.4 Prel iainary Results 

The pel im!nary power profile for the total station is fl iustrated in 

figure 4.3.4.2-1. The power level exceeds the 75 kW sys- capability 

primarily due to the size of the reference payload coaplent (Table 

4.3.1-1) and a c~rrent lack of understanding of payload operation 

phasing. The profile for the facility averages is shown in figure 

4.3.4-2-2. These results should be taken only as Indicative of the 

geiiersl operdiSng range cf the facii lty due to the preliainary nature of 

the subsystems definition. The di strl bution of power utilization amng 

the subsystems is given in table 4.3.4.2-2. The profile for the pqyloads 

only is shown in figure 4.3.4.2-3. 

The qua1 ity of this analysis is consistent with the mturi ty of the 

subsystem a;;? pqyload data. The housekoep!ng power requirements should 

be interpreted as an indication thr: ;he Station can potentially operate 

below the 30 kW housekeepins 1 eve1 . prior experience and analysis performed to support the current study tend 

to i~dicate thax station power consunption is not strongly affected by 

the number of crewmembers onboard or the level of crew activity. The 

power consumption is =re strongly a function of the design sire of the 

subsystems. 
0 

LIJ 

b? 
-1 1.: 1 

f " i5i i3 

rY: [I- 

. . . , 

TABLE 4.3.4.2-2 - PRELIMINARY DImIBLCTIU OF PWES :a I l LSZATIOY 

TOTAL 

WYER 

ALL 

COt3WNENTS 

SUBSYSTW OII-KW 

POWER GLNERATION 

POYER DISIRIBUTICIY 

WI DANCE COWTROL 

NAVIGATION 

TRAFFIC CONTROL 

IF SYSTEMS 

ANTENNAS 

AUDIO 

VIDEO 

SIGNAL PROCESSING 

-DATA ACQUISITION/TA/CDS 

1RACK I N G 

sysm PION ITORI~~~/CONTFUIL 

DATA WEEK 

INTEGRATED D C 

FACILTIES MANAGEKNT 

OPEMTIOWS/PUtrNI#G/S!JPPORT 

ORBIT PROPULSION 

FLUIDS CIANAGEENT 

PROPULSIOY STAGE MAGE)IEWT 

ECLS 

EVA 

ACTIVE THE= CONTROL 

STRUCTURES 

CECHMICS 

PASSIVE THERMAL CONTROL 

CREW ACC-DAT IONS 

HEALTH MI NTENAKE 

PAYL 0.W SUPPORT 

'M 

SUBTOTAL 

AVERAGE 

POER KW 

9YER 48- 

HOUR 

MISSIOH ; 

0 .oo 

2.65 

5.99 

0.00 

0.08 

1.59 

0.24 

0.52 

2.53 

0.55 

0.39 

0.12 

0.07 

1.50 

0.85 

0.00 

0 .oo 

0.00 

0.00 

0.04 

7.07 

0.35 

0.26 

0.00 

0 .QO 

0.00 

0.85 

0.53 

PERCENT 

u/o 

FAY LOmS 

29. PAYLOAD OPERATIONS 118.04 63.16 ?Q::.. , 6 - -- - - . - . - - - . 

TOTAL 183.25 90.05 4321 -7 - 
4.3.5 'Them1 capabll ity 

4.3.5.1 Introduction 

The thermal design of a Space Station is highly dependent on the 

configuration and is therefore sensitive to seemingly slight changes In 

the configuration. The fol lowing thermal eval uation criterl a wcrt 

considered for this study: 

a. Integrated station enviroment effects on thermal design 

requi re~ents 

b. Radiator area repuf r-nts 

c. K,ximm use of waste heat 

d. Minirize sensitivity of themal design to surface coatings 

cr ntami nation and degradation 

e. Desl gn cmpl exi ty and mi ntai nab1 1 i t,y 

h. Safe haven requirements 

I. Technology status 

4.3.5.2 Thermal integration 

A totally integrated Them1 Control System which efficiently uses all 

heat si nks' and sources i s ideal . Hodever, thl s ideal i sa mst be 

moderated by practicality. Yith this in mind, the Space Station 

configuration was revieued to identify functions and systems rquiring 

thermal control. Considering the evaluation criteria and the 

requirements specified in the RFP, levels of fntegraticn and ratfonrle 

were fomul ated. 

Table 4.3.5-1 presents the major identifiable functions/syskrs and the 

design approach and rationale selected. Where sped fic thermal 

requirements were not defined, scch as OWV, ON, and satellite se?n+Jclng, 

assumptions were made as to the thermal design approach. 
FUNCTIOM/SYSTEM DESIGN APPROACH RATIONA

0 LOGISTICS MODULE 0 AUTONOMOUS ATCS 0 SIMPLIFIES M

( BODY MOUNTED RADIATORS ) BREAK INTERF

0 LOW HEAT REJ

0 THERMOS BOTTLE PASSIVE 0 MINIMIZES TH

OESlGN INSENSITIVE TO LOAD EXCURSI

COATING DEGRADATION AND LIMITS E

OF UNPRESSUR

0 HABITAT AND 

LAB MODULES 

0 POWER SYSTEM 

0 INTEGRATEC ATCS FOR 0 EFFICIENT US

MODULE HEAT REJECTION SOURCES AND 

THROUGH CENTRAL 

BUS RADIATORS 

0 BODY MOUNTED RADIATORS 0 SAFE HAVEN A

PUILDUP CONS

0 PASSIVE DESIGN - THERMOS 0 MINIMAL THER

BOTTLE INSULATION AND EXCURSION ON 

NON-RADIATOR SURFACE SURFACE PROP

THERMAL OPT!CAL PROP- TENANCF 

ERTIES a/E = 1. .8/ .8 

0 AUTONOMOUS ATCS O STATIOH/BLATF

COMMONALITY 

0 MINIMIZE MAK

INTERFACES 

O PASSIVE TCS INSULATION 0 LIMIT COMPON

AND COATINGS TO LIMIT TURES AND THE

COMPONENT ENVIRQNMENTS EXCURSIONS O

TABLE 4.3.5-1 TCS LEVELS OF INTEGRATION 
FUNCTION SYSTEH - DES IGM APPROACH RATIONALE 

0 OW STORAGE 0 NO THERW HARDWARE 0 OMV PROVIDES OWN THER

INTERFACE REQUIRED CONTROL USING STATION 

O ANALYTICAL DEFINITION 

OF STORAGE AREA IMTEGRATED THERMAL ENVIROMMENT 

O OTV HANGAR 0 BASIC DESIGN IS PASSIVE 0 REQMTS. NOT IDENTIFIE

WITH INSlILAT~3W AND COATINGS TAILORED TO MEET OTV 0 OfV DES!GN SHOULD CON

LIMITS SSP THERMAL ACCOmQDA

CAPABILITY 

0 ATCS CAN BE ADDED AS 

UNIQUE REQUIREMENTS 

ARE IDfN?,FXEO 

C PRESSURIZED 0 HEAT SINK/SOURCE 0 ACTIVE THERMAL CONTRO

PAYLOADS PROVIDED BY ATCS INTE- REQUIRED UNLESS PAYLO

GRATED WITH CENTRAL BUS PASSIVELY DESIGNED WI

AND RADIATORS LOW POUER CONSUMPTION 

0 ATTACHED PAYLOADS 0 SAM AS UOVE 0 SAME AS ABOVE EXCEPT 

LOAD PASSIVE THEM 

CAN BE ACCOMPLISHED F

LARGER POER LOADS 

TABLE 4.3.5-1 - TCS LEVELS OF INTEGRATION (CONTINUED) 
0 PROPULSION 

9 SATE:: ',, i, 0 DEPLOYABLE THERMAL PROSERV I C ING/STORAGE/ TECTION SHROUD TO LIMIT 

REFl!cLING BAYS EIIV I ROMMENT FOR SOME 

SATELLITES 

0 PlALYTICAL DEFINITION 

3F INTEGRATED THERMAL 

ENV IRONi!4ENT 

0 PASSIVE TCS DESIGN; 

MAKE UP HEAT FOR TANKS 

AND FLUID DISTRIBtiT!ON 

SYSTEM 

0 REQMT. FOR 

TROLLED ENC

INFERRED BY 

POSSIBLE TH

OF VARIOUS 

L TCS DESIGN 

PiISHED FOR 

WXE UP HEA

D PASSIVE TCS XSIGN FOR C MINIMIZE MA

ENGINES AND HEAT SOAK INTERFACES 

BACK PROTECT IcK DEZTGh CCIW

O I'flSSIVE TCS DtSIGN ;3 GIH8.Q ML7T

UkiQVC TO LACii TYPG QUIRES SiMP

APPROACH (I

3 A'TCS INCREA

GREA!! IHTER

ANTENHAS DE

P!EXITY UD 

RUNS 

0 COYTIIJllCUS 

OR St.MDBY 

PLANNED. T

TIONING REQ

FOR CONTIHG

TABLE 4.3.5-1 TCS LEVELS OF I'iTEGRATION (CONTINUE
4.3.5.3 System 0ver.vlew and ?~paki?.lty 

TC.e candidate Act1 ve 'Ther~nal Control Subsystem ( ATiS) functtonal ly I1 lustrated by Figure 4.3 5-1 i~ a hybrid design concept that maximizes the 

use of local thetma! control for individual Station modules and satisfies 

the remai njng thermal controi requirements with a central ized system. 

Tl,e heat acquisition and transport function is accmpl ished with c 

thermal bus approach using a two-phase ,l..riring fluid ti, transfer heat by 

evaporation and condensation rather than by sensible heat changes of a 

single phase coo'lant. As a result, the t+erm3l bus essentially operates 

at a constant temperature cver the entire length of the flow circuit wlth 

pumping power requirements that are very sn~all compared to a single phase 

fluid system. Each manned Station module cantair~s mu1 tiple two-phase 

water heat transport circuits to accommodate Station iscthermal i ty and 

heat load capaci ty requirements., Hi th tC2 exception of the Logistics 

Modul e, the manned modul e water ci rcui ts i nierf ace wi th appropriate 

central two-phase ammonia thermal bus iirc~its outside the pre:.slrrized 

cabin areas operating at about 350F, 700F, ai;d 90T. Thew c;rr -"A: 

flow circuits in turn interface with an erectable. cer,,ra'i'., -; 1;;t pipe 

radiator systen. 1 his radiator assembly includes a girn~.t-a,,~a& to 

orient the radiators tcwdrd a minimum thermal environrr.errt .& s 

ml ni~ni zi ng radi at~r size. 

In additioli, each mann2d Staticn module has heat pipe radia-rs 

(operating at aboct 3S°F) integrated with the module meteroid pratection shield. Altho~g~~ hardwar,: ccommonz'ity iis a goal, the size of 

these body-wun+,ed radiators may varj froni module to modale depending on 

surface area avdi 1 abil i ty coi~sideri ny berth-in.; ports, windoh=, them1 

blockage, etc. To a 1 imi ted degree, these rnotiul e radi aiorSs augment the 

central radiators d~ri ng normal operations. However, wi tl~ the exctption 

of the Logistics Module. the principal function r' these radiators Is to 

provide an autonomous safe haven r-apability and to possibly assist durlng 

the Station buildup an0 assembly. To av9;d "make and break' fluid lin2 

operations fcr the 1 ogi stics module, body- mounted radi atcrs are usea 

exclusively to reject wasto heat. 

A sepdsate two-phase amonia t.hennai (operating at about 1159) Is 

proviasd for electrical power systm s~i:,ce heat rejection from th~ 

regenerztive fuel cell s, electro1y;is uni tc, nnd power. condi timing 

equipment. This flow ~ircuit also incl,dbs a thermal stdrage unit to 

minimize the radiator 2izc since tae nightiday orbital power s\.stern waste 

heat load varie~ by a fY*.tcr of about 2.5. tleat pipe radiators are 

provided for rejection cf the power Ejs\im raite heat. 

The IOC raeiator areas for th? reference 3TCS cor.c?pt irre sumarized 

be: ow: 9 L 

Radiator Area ( ft I* 

Body-Mt 96C -* 

Central 462' 

Power rystem :226 

I. og Modl~l e122 

* Ts2tal radidting ore3 

** Excludes Logi sticc Module 

277 
TYPICAL HAB OR 

LAB MODULE 

PRESSURIZED 

CABIN 

r"-----"""""-- ----------, 

, VAPOR (TYPICAL) 

I 115OF I 

I COND. 

I ! 

r __,t I 

THERMAL i_ 1 

I I LIQUID (TYPICAL) 

\POWER SYSTEM 

Figure 4.3.5-1 - ATCS DESIGN APPROACH 
Y i rtual ly a1 1 e i events of the ref t?r~ii~~ c~Si giirati oii i-ecei ve direct 

solar radiation i7 the local vertical flight mode. Ideally the passive 

thermal desi yn of habitable mod,. ,es for these conditions would use 

radiator-type coatings (i .e., low solar absorptivity to mi ttance, 

a!/€, ratios) in conjunction with an insulation design to minimize heat 

leak into the modules for those surfaces receiving sun while maximizing 

heat leak out for other surfaces. However. considering the desire to 

minimize coating refurbishment and the fact that the mdule total passive 

heat loss or gain is small in comparison to the total internal heat load 

i approximately 1 to 6 percent of tire Habitat Module 1 oad) , a thems 

bottle insulation design war selected. This desensitizes the design to 

enviromental var! ation: and coating degradation and a1 lows the selection 

of surface treatments such as anodized alwninum with an aC/b approachins 

1.0 fi>r the habitable aidules and external airlocks. 

The Om servizing and storage areas do cot require themal protection 

since the CW must be designed for the extremes of thermal environaent to 

accomplish i:s ~eission. Also, no active thermal control .,iterface is 

provided since the them1 design of the OPCV will probably be coapletely 

passi ve. 

Sztell i te storage, service, and refueling area requirements are varied in 

order to accommodate a Targe spectrum of ,arellite thermal design 

environments from stationary attitudes which represent the extreaes of 

temperatti~e capability to spinning modes which tend to a moderate range 

of temperatures. This imp1 ies a cdpabi 1 i ty to provide themal protection 

for sane payloads it? o,*der to 1 imi t the range of enviromts to which 

+%e sate1 l i te is subjected. Temperature extremes for unprotected 

satellites with various oc lE themzl properties are prese~ted in Section 

4.4.7.4.3 as we1 1 as interanal temperatures for a deployable 15-by-60- 

feet-1 ong cyl i ndrical thermal protection shyoud as a function of internal 

power generation. k diameter gre&ter than 15 feet will probably be 

required to meet access requirements. For tho.: satellites requiring the 

themal protection shroud the all owable satei l ite and eq?iipent heat 

generation/power will be 1 imi ted by the external w/E ratio of the 

shroud and the allowable rnaximtlm temperature environment of the payload. 

Therefore, the ability to dcsiqn the shroud to eliminate surface property 

refurbf sknt may be liai ted. The shroud themal design must be based on 

a review of payl oad and support i nj equi pmrit powei. and temperature 

requirewnts and servicing/refuel ing operations. 

The 3TV hangar, 33 by 33 by 66 feet, is considered to have a canpletely 

passive thermal design composed of wul ti1 ayer high performance insulation 

and appropriate hangar external surface thermal optical properties. 

Actual gerfommce requirements must be s?lected based on detnf led OTV 

design requi remnts and servicing operations. Active themal control 

interfaces with the central bus can be added as these requirements are 

identified. 

For an e/E of 1 .O, an internal heat load of approximately 2kW can be 

acccllaeodated without exceeding an average Internal hangar temperature of 

1000F. 
Propellants for sate1 1 i te, OW, and OTV refuel ing and the station 

contingency propellant are located within the truss work just below the 

lower keel fn the keel extension. Insulation closeouts are provided to 

fom an enclosure around the tanks. Since the desfgn approach for the 

cryogenf c fluids (hydrogen, helium, nitrogen) has not been selected the 

thermal desfgn fs left as an open issue. Horrver. it is expected that m 

approach sfmf lar to that of the mno- and bi -propellants can be applied. 

Applying anaC/€ = 1.0 (0.010.8) external surface treatment in 

conjunction with high performance mu1 tf layer f nsulatf on results in 

minim makeup heat to maintain the mono- and bi-propel lants between a 

400F and lOOOF operating range. For each 9-foot truss cube 

contalning the tanks, 54 watts of nakeup heat are required. The makeup 

heat is provi ded by elec trf cal heaters. Where practf cal , propel 1 ant 

lines will be run para1 1el to the central tnemal bus with a crn 

insulation cocoon, thereby obtaining required makeup heat frou the 

them1 bus transfer lines. Those 1 fnes not wrapped wf th the thermal bus 

lines require approximately 0.27 watts of makeup heat per foot. Makeup 

heat for operation of the catalytic beds for the station RCS engines was 

not establ f shed since this is highly dependent on the detaf 1 ed mgine 

design and structural integration approach. 

The 90-day stati 3n propel 1 ants are 1 ocated in the unpressurized section 

of the logistics rudule. The design of the logf stics module f s sfmflar 

to that cf the habitation and laboratory modules. In addition, the bodymounted radf ators are mounted to the greatest extent possf ble on the 

unpressurized portion of the mdule to provide an acceptable boundary 

temperature for hardware located in the unpressurized cmartment. A 

thermal curtain (multf -layer insulation) closes out the open end of the 

coapartnrent . The nurber and type of antennas for the reference configuration are 

varf ed and thef r partf cul ar thermal requf relaents vary dependf ng on 

lccatfon, internal power generation and duty cycle. Sf nce the antennas 

and associated avionics are expected to be either in an operational or 

standby node at all times, a passive design similar to that of the Space 

Shuttle Orbiter Ku-band is envf sioned rather than integration into the 

Active Thermal Control Subsystem. 

Passive thermal control of the solar arrays can be obtained to a uall 

degree by optimf zing the back face thermal opt1 cal propertfes. For an 

oc/E of 0.5 and 0.7, the maxi nun and lrinfur temperatures during an 

orbf t are 1660F and -1350F, and ior an oC/€ of 0.2 and 0.9 they 

become 1380F and -14FF, respectively. More complete results are 

presented in Sectf on 4.4.7.4.7. Ff nal selection of surface properties 

and materials treatment mist consf der the solar array detailed design and 

evaluation of the contamination cnvironnent to which the array will be 

subjected. 
4.3.6 Cwnications and Trackf ng Antenna Systems 

4.3.6.1 Required Antenna Coverage (Fie1 d-of-V i ew) and Operati onal Zones; 

4.3.6.1.1 Operational zones 

The operational zones specified for the Space Station Program are 

depicted in Figure 4.3.6-1. These consist of nine zonal areas: (1) 

proximity operations, (2) control, (3) departure, (4) rendezvous, (5 6 6) 

co-orbital , (7 8 8) non-co-orbital , (9) parking (see Section 4.3.8.2.1). 

Only zones 1 through 6 will be the responsibility of ae Space Statiorr 

for cmunications and tracking. 

figure 4.3.6-2 depicts cmunications and tracking systea coverage of 

these operational zones. The approach taken was to cover zone 1 ana the 

control zone with an antenna system which can achieve spherical coverage 

out to 20 mi since Space Station control of operations is liaited to 

this area. Zones 5 and 6 are covered by antenna systems which produce a 

conical beam of - +8.S0 out to 1080 mi (line of sight grazing Me 

atsosphere). 

These antenna systems will be described in Section 4.3.6.2. 

4.3 -6.1.2 Coverage required for an earth-oriented Space Station 

a. Antecna coordinate systea: Figure 4.3.6-3 shows the 

coordinate system used to define the cone and clock angle direction 

(theta and phi respectively) fm the Space Station antennas to any athtr 

vehicle. For an earth-fixed orientation, the z-axis is pointed Ward 

zenith (opposite the nadir direction), the x-axis is pointed along the 

Space Station velocity vector, and the y-axis is perpendicular to the 

orbit plane (POP). The boresight or peak gain direction of the antenna 

is usually asswd pointed along the z-axis, so that the thetalphi plats 

will show the angular variation relative to the antenna boresight. 

b. TDRS: Two Tracking and Data Relay Satellites (IDRS) will 

be located at westlongitudes of 41" and 1710, with a spare at 

830. Since rhe TDRS i s i n synchronous orbit, the look angle from the 

Space Station to TDRS will vary very slowly over 2 short period of ti=. 

However, in a 1-day period the cone and clock angles vary over a wide 

range. This can be seen in f igum 4.3.5-4 where the theta and phi angle 

variation is plotted for one 24-hour period. One orbit pass corresponds 

to one horseshoe shape set of points around the y-axis. Each point is 

separabsd in time by two minutes. For the whole year the cone and clock 

angles will fall within the shaded area shown it1 the figure. Note that 

the theta angle is never greater than 1000. The earth's horizon Os 

assumed to be at 800 corresponding to about the top of the ionosphere. - Note the large angular area over which the TDRS antenna does not have to 

look. This allows greater flexibility in the location of the antenna on 

the Space Station since such a large area of obscuration is available 

without interfering with the TDRS 1 ook angle di rcctions. 
L 

v 

ZONE 8* 

ZONE 33 

ZONE 52 

4 

100bMI 

(185 km) 

1 & 

m 

- 

ZONE ASSIGNMENTS: 

1. PROXIMITY OPERATIONS ZONE 

2. CONTROL ZONE 

3. DEPARTURE ACTIVITY ZONE 

4. RENDEZVOUS ACTIVITY ZONE 

5. LEADING COORBlTlNG SATELLITE ZONE 

8. TRAILINQ COORBBTING SATELLITE ZONE 

7. LOWER NON-COORBITINQ SATELLITE ZONE 

8. UPPER NON-COORBITING SATELLITE ZONE 

0. PARKING ORBIT ZONE 

NOTES: 

1. THlS DRAWING IS CURVILLINE

DRAWN TO SCALE. 

2. THlS ZONE IS CONTINUOUS AB

3. THlS ZONE BEGINS AT THE SP

4. THE MINIMUM OPERATING ALT

ZONE IS -- 100 NMI (-105 km) 

1 R 

4.84-1- OPERATIONAL CONTROL ZONE@.' 

ttt 

LIKC 

NXt 
PLUS = IN SUN SWAE = IN SMIM OIA)1BNO = INITIAL WBITS 

0.0 

PHI 

FIGURE 4; 3,6 -4 ANGULAR COVERAGE TO TDRS FOR AN EARTH ORIENTED SPACE STATION. 

ANTENNA Z-AXIS POINTED 180' f ROM NADIR, X-AXIS POINTED ALONG FLIGHT PATH. 
c. GPS: The GPS look angle r~quirements are similar to those 

of the TDRS. ever, since the GPS satellites are &t a highly inclined 

orbit (5501, there are periods when the cone and clock angle directions 

are all near the positive and negative y-axes. Thus for GPS, unobscured 

v4ew of the upper hemisphere (theta = 00 to 900, phi = 00 to 

36Co) will be required so that at least four GPS are always in view. 

d. Co-orbitin S/C: With the Space Station in an earth-fixed 

nadir pointed or 4 enta cn 6-"+8.50 cone coverage required around the 

forward and aft tracks stays fixed throughout the ortiit. However, small 

roll, pitch, and yaw excursions of about 50 can occur. Thus, with one 

antenna pointed forward and one pointed aft an dng~rlar cone of at least 

- +14O about both the positive and negative x-axes must remain 

unobscured. If larger angular rotations of t+e Space Station about the 

roll , pitch and yaw axes are expected, then larger unobscured fie1 ds of 

views will be necessary. 

4.3.6.2 Antenna System Description 

4.3.6.2.1 Communication 1 inks 

4.3.6.2.1.1 TDRS 

Two %foot diameter parabolic antennas will be used to carmunicace to the 

ground via the TDRS. Two antennas are necessary in order to minimize 

1 ink outages during handover between the two TDRS sate1 1 i tes. For a 10- 

tc 15-minute period each orbit both TDRS sate1 1 i tes are in simul tsneous 

veiew. During this period the cmunication 1 ink must be transferred from 

the TORS satellite that is setting to the one that has just risen above 

the horizon. If one antenna is used to track both sate1 1 ites, 

considerable data outage will occur during the period required to slew 

the antenna 1300 from one TDRS to the other and to re-acquire the TDRS 

signal. To minimize this outage period high slew rates would be 

necessary, requiring much higher power to drive the gimbals. Low power 

is required to drive the gimbals during the normal tracking periods since 

the lnnk angles change very slokly. Each antenna will contain a dual 

feed system, one for S-band and one for Ku-band. A tracking feed will 

a1 so be incorporated to maintain antenna poir~ting initial acquisition. 

The RF transmit and receiving electronics will be locateb near the feed 

to minimize cable losses. 

During the initial build up of the Space Station, two hemispherical 

coverage low gain (0-3db) S-band antennas will be used to provide a low 

data rate S-band 1 ink to both the TDRS and the Orbiter. Small microstrip 

antennas can be used for this application. These antennas are used for 

the TORS 1 ink when the Orbiter is not present. When the Orbiter is in 

the vicinity of the Space Station these antennas are switched to the 

Orbiter 1 ink. The TDRS 1 ink from the Space Station is turned off. 

Camnunication to the ground is now accomplished via the Orbiter TDRS 

1 ink. When the Space Station Assembly is completed, these antennas, 

after relocation, will be used to provide a low data rate S-band link to 
TDRS as a backup in the event of loss of tracking by the %foot high gain 

antennas. Separate RF transmit and receive electronics are located near 

each of these antennas. 

4.3.6.2.1.2 Orbiter 

As stated in the previous paragraph, low gain near hemispherical coverage 

S-band antennas will be used to carmunicate with the Orbiter in the near 

proximity of the Space Station during the assembly phase. After the 

reference configuration is completed, two additional 1 ow gain S-band 

antennas will be added to provide coverage to the Orbiter during 

proximity operations. 

These antennas can proviae coverage out to a range of about 5 n. mi. 

Beyond this range a low gain (10 db) horn antenna will be csed to provide 

comnunication to the maximum range of 20 n. ni. specified in the 

rzquirements. Two steerable horn antenaas will be required to provide 

coverage along the forward and aft zones. Since these antennas have a 

bermwidth of about 50 degrees, a single axis gimbal will be required to 

provide hemispherical (+goo) - coverage in the plane of the orbit. 

4.3.6.2.1.3 EVA, OMV, FF, Platform 

The communications link to these various users will operate at K-band (25 

to 27 GHz). Section 4.3.2.1 and Figure 4.3.6-2 discuss the required 

operational coverage zones. As pointed out in that section, two besic 

coverage zones will be provided. These are the near range contrcl zone 

and the far range co-orbital zone. Two antennas systems will be used to 

cover these zones. In the near range control zone (20 nmi sphere) a 

mu1 tiple access .link will be used to comnunicate with the EVA'S and OMV. 

This zone will be crvered by a set of swi tch~ble antennas with 

over1 appi ng beac.., to provide herni spherics; coverage. Figure 4.3.6-5 

shows the mounting configuration of seven horn antennas, and their 

transnit/receive electronics, which can provide hemispherical ccverage. 

The peak gain of each antenna is approximately 9 db with a crossover gain 

between antennas of 5 db. 

A separate transmitter and receiver are connected to each antenna so that 

mu1 tiple users can be served. The signal received from each antenna is 

down converted to a different IF frequency, correspondf ng to a different 

beam. Each IF frequency is mu1 tiplexed and transmitted by cable to the 

control center, where the signals are again separated and converted to a 

comron IF frequency. A switch unit connects tha mixer output to the 

dem+~dul ator corresponding to a particular beam. Several user signal s can 

be present in the same beam. In this situation more than one demodulator 

would be connected to the same mixer. The demodulator selects only ,the 

signal that matches the identifying code fed to it. COMA has been 

assumed for this design. TDFA could be used with a simpler design. The 

switching unit would be located at the antenna, and the mu1 tiplex unf ts 

eliminated using a single TC frequency. Using TDMA requires an increase 
f 1 

I 7 HORNS USED. 1 POLAR DI

OTHER 6 ARE POINTED UP A

AND EQUISPACED IN LON(3

---- 5 dB1 MINIMUM. 

RADIATING HORNS 

TRANSMIT AND RECEIV

AND SUPPLY POWER ' 

MIXER DRIVERS 

MULTIPLEXERS 

SYNTHESIZERS 

TO CONTROL CENTER 
in the trmsmitter power in a direct ratio to the number of users being 

served. Higher power microwave amp1 if iers may not be available at 25 

GHz. For this reason and other considerations the use of TDMA may be 

precluded, A more detailed trade-off analysis will be necessary before a 

mu1 tiple access system design is chosen. In any event this antenna 

design will work with any MA system. 

An antenna switching arrangement, similar to that used for the received 

anal, is fncorporated for the transmitted signal. The transmit and 

rt ive signals are connected to the same antenna using d deylexer, and 

RF 1 ters to provide isolation. Additional single element broad beam 

anter,nas with 0-3 db gain will be used to provide extra coverage to EVA 

activiLy in and around the inmediate Space Station structure. 

Far the initial Space Station configuration a single access corrrnunication 

link will be used to comnunicate with the platform and FF in the far 

range co-orbital zone (100-1080 mi). To provide this coverdye a 28-inch 

diameter K-band steerable dish is used. The peak gain of this antenna Is 

43 db. This is the gain required for the maximus range of 1080 mJ. At 

this range a FF located at the -10 mi edge of the coverage zone wil? be 

at an angle of -8.90 from the Space Station velocity vector. At +10 

nmi this angle is 7.80, for a total cone angle :overage of 16.70 

(t8.50). To cover this angular zone a two axis gimbal system is 

rzquired for pointing the antenna toward the cmunicating user. The 

antenna must be real i gned each tim the 1 ink is swi tched to a new user. 

Two of these antennas are required; one pointed forward anQ one pointed 

aft. A separate set of RF transmit and receive electronics will be 

mounted near the feed of each of these antennas. 

In the Space Station growth conf iguratfon a mu1 tiple access capabil ity 

will be incorporated for the far range co-orbital zone. To provide this 

capability the 23 inch diameter dish will be replaced with a multiple 

beam offset reflector antenna. Figure 4.3.6-6 shows the design of a 

mu1 ti beam offset reflector antenna which can provide the required 

coverage. This antenna contains about iOQ switchable feeds, which cas 

provide beams anywhere in the - +8.5O core. The peak gain of this 

antenna is a1 so 43 db. 

The antenna boresight must be pointed about -50 below the orbit track 

so thdt the +8.50 cone coverage of the antenna can reach the lower scan 

angle of 8.90. A two ad is gimbal system is required for proper 

alignment of the antenrbc iystem borecight. This gimbal f ng system will 

allow for realignment of the boresight which may be needed to compensate 

for any small pitch, roll or yaw excursions. Two antennas are required, 

one pointed forward of the Space Station, the other pointed aft. 

Each feed of the new multibeam far range antenna system contains both a 

transmit and a receive module, each consisting of a filter, mixer, buffer . amplifier, andeither a power amplifjer (PA) or low noise amplifier 

(LNA) . The transmit and receive rnodul es are di pl exed to the feed. The 

signals to and fran the transmit and receive modules and the coctrol 

center are at IF frequencies. A mixer frequency is also provided for 
I 1 FEED ASSEMBLY 

-4 d) 

FEED APERTURES 

a) PROFILE VIEW 

PEAK GAIN 

SCAN ANG

flgurr 4.3.0 4 - ANTENNA SYSTEM FOR FAR RANOE. 
the up and down heterodyne conversion. A beam switching matrix is 

incorporated at the antenna to switch the IF signal to fhe correct beam. 

Up to four beams ~dn beoperated siaultaneously. A gate signal is 

provided to enable the LM and PA of the operating bems, and to disable 

the corresponding aupl d f i ers on a1 1 the non-operati ng bems. 

In the growth configuration a higher gain (50 db) antenna will also be 

incorporated to provide high data rate (25 dps) capability eut t~ the 

maximum range (See Table ? .4.3-1) . This antenna is a !%foot diameter 

steerable dish with a single beaa. Only one user can be serviced at a 

time. Two of these antennas are also needed to cover the forward and aft 

directions. Since the beamidth of this antenna is about .50, a tuo 

axis gimbal will be required to realign the antenna to a different user 

in the - t8.50 (+14tfor - snall atti tuee angle excursions) coverage zone. 

4.3.6.2.2 - Tracking and navigation -

Two antenna systems are used to provide the Space Statian tracking and 

navigation functions. One is a rendezvous radar antenna using a 

steerable 3-f00t diameter dish operating at Ku-band. Two of these 

antennas are required to provide tracking of the Orbiter or W anywhere 

in the required 20 mi sphere. The other is a low gain (0-3 db), dual 

switched, hemi spherical coverage antenna, which can receive position 

locatian informdtion from the GPS satellites. Small microstrip antennas 

can be used for this application, TW of these antennas will provide the 

hemi sphericil coverage required. 

An additional pair cif low gain (0-5 dh) Ku-band microsttwip antennas will 

alsc be included for use wSth a rendezvous radar transponder which 

operates with the Orbiter rendezvous radar. The radar transponder is 

included on the Space Station in order to improve the tracking capability 

of the Orbiter rendezvous radar, especia1:y during the asserdbly phase. 

Jt also aids in distinguishing returns from the various parts of the 

Space Station structure, and thus avoids para1 1 ax errors when approaching 

the docking pcrt. 

4.3.6.2.3 Mobile Remote Manipulator System (MRMS) 

To transmit comnands and receive N and TM from the MIWS, an S-band ljnk 

will be used from the Space Station to the MRMS, One low gain near 

hemispherical caverage microstrip antenna will be located on each end of 

the MRMS. Two ormidirectional antennas with a dipole type pattern will 

be mounted on the Space Station. A biconical antenna or an array af 

micr~trip antennas mounted around a cylinder could provide the beam 

pattern coverage required. The antenna will have to be mounted away from 

the Space Station structure and the pattern shaped to minimize the 

effects of reflections from the structure. One of these antznnas will be 

mounted on the transverse born for use durjng the assembly phase. As the ' Space Station is cmpleted these antennas are moved to their pemanent 

1 ocations., 
4.3.6.3 Antenna Locations for the Reference Configuration 

Table 4.3.6-1 lists the total antennas to be used on the reference 

configuration. Included in the table are the antenna type, frequency of 

operation, and the elerents served by each antenna. In the initial 

configuration antennas Nos. 2 and 3 will not be incorporated. As pointed 

out in Section 4.3.6.2.1.3, a 28-inch dfareter K-band steei-able disk 

antenna will be used in the initial configuration instead of the 

multibeam antenna 1 isted for antenna No. 2. The antenna location 

discussion, however, will include both antennas Nos. 2 and 3. In general 

two antennas will be needed for each 1 ink. For proximity operations four 

additional low gain antemas (No. 5 in the table), with near 

hemispherical coverage, are incldded to provide extra coverage to EVA 

activity in and around the imnediate Space Station structure. Four low 

gain S-band antennas (No. 7) are shown. Two of these antennas and the 

GPS (No. 8) antennas and their RF electronics will be located on the 

transverse bocm for use during the early Space Station assembly phase. 

When the Space Station is camp ?te these antennas will have significant 

cbscuration frorn t;le solar panels and upper and lower keel structure. To 

a1 lcviate this protl2m these antennas and the RF electronics associated 

with them must be movtd to a less obscured location. However, if the 

moving of these antennas is not feasible, Go additional antennas and 

their electronics will be required. The two systems left on the 

transverse boom could be used for 1 imi ted redundancy capabil i ty. 

Assuming these antennas can be mvea, the two S-band antennas will be 

located on the upper boom to provide a 1 ow data rate backup 1 ink to 

TDRS. The two GPS antennas will also be located on the upper boom. The 

two additional low gaf n S-band antennas (No. 7) will be located on the 

lower boom to provide additioaal coverage to the orbiter in the proximity 

region. The S-band horn antennas (No. 6), which provide coverage to the 

orbiter out to 20 nmi , will be located on the upper boom. mese antxnnas 

provide the primary coveraye for the S-band 1 ink to the orbiter. 

Figure 4.3.6-7 is a plan vie:/ of the rc'erence configuration showing 

antenna locations. The reference conf ig~ration is flown earth-orientea 

with the bottom end of the Space Station always pointed toward nadir as 

indicated in the figure. For this situation the look angle coverage 

requirements are as discussed in Section 4.3.6.1.2. Both the TORS snd 

GPS antennas must have unobscured views in the upper hemisphere looking 

toward space. To provide this coverage these antemas are located on the 

upper boom, as shown in figure 4.3.6-7. The GPS antennas are located on 

each end of the boom with the boresight looking along the z-axis toward 

space. A1 though one hemi spherical antenna could provide the coverage, 

two are included since the placement of payload experiments on the boom 

may cause obscuration. Figure 4.3.6-8 shows an ot,scuration plot for the 

TDRS antenna located on the starboard side of the boom. The antenna 

+boresight reference is aligned 3s described in Section 4.3.6.1.2. The 

shaded areas in the plot ccrreipond to the angular volume where 

obscuration occurs. (Note: This obscuration plot and the ones to follow 

were made for the configuration which is ;;entical to that shown in 

Figure 4.3.6-7 except that it included three lower keel sections, and did 

r>t include the large antenna experiment TDM 2060. The results, however, 
ANTENNAS FOR REF. CONFIGURATION. 

CW. AND TRACK. ANTENNA FREQ. ELEMENTS 

TRACK TYPE GHz NO. SERVED 

1. TDRS SKU BAND 9 ft dish 2/14 2 TDRS 

2. ED. GAIN 2.5 ft multi- ATCO-ORBIT K.A. feed offset 2 5 2 

Ref. 

3. HIGH GAIN (50db) - 

CO-ORB IT S .A. 5 ft dish 25 2 :kl4?gR 

4. - GAIN I element 

PROX. OPS H.A. switched array, 

hemisph. cov. 25 2 

5. 0- Gain Sing1 e element, 

PROX. OPS M.A. near hmi sph. cov. 25 4 

6. S-BAND LOU 

GAIN (10db) Horn 2 2 ORBITER 

7. S-BAND LOW Dual swi tched , ORBITER, 

GAIN (0-3db) hemi sphere cov. 2 4 1 DRS 

GAIN - 3. GPS LOU Dual swi tched, 

S 

(0-3db) hemi sphere cov. 1.5 2 GP S 

ORB ITm, 

ED. GAIN (40db) 3 ft dish 14 2 OMV 

IG. REND. RADAR XPNDR Near hemisphere 

LOU GAIN (0-3db) coverage 14 2 ORBITER 

11, MRMS LOU GAIN See go- 

(0-3db) 2 4 Manip. Sys. -4 

1 

-- -- -~- --- 

TCTAL 28 

TABLE 4.3.6 -1 

293 
NOTES: I. Antenna #5 required to provide extra coverage to EVA activity 

in and around the faraediate Space Station structure. 

2. Two antennas of #ll are mounted on the 13R)11S. They are single 

element ant. with near henispherical coverage. The other two 

are mounted on each side of the keel structure. These ant. 

wi 11 produce ani-di rectional dipole type patterns. 

3. The Radar XPNDR. (#lo) is included on the Space Station to 

enhance the capability of the Orbiter's rend. radar to see the 

Space Station, especially during assembly. 

4. The S-band low gain ant. (#7) are used to provide coverage 

to the Orbiter and to TORS during the assembly phase, and to 

provide a low data rate S-band link to TDRS as a backup to 

the 9-ft dish (#I). Two of these antennas will yovide extra 

coverage to the Orbiter when it is stationkeeping near the 

Space Station. 

5. In the in:tial configuration antenna ti? will be a single beam 

28-in dish. Antenna #3 will be added in the growth 

configuration. 

would not be much different for a single keel configuration.) Figure 

4.3.6-8 shows that there is a clear field of view (FOV) over the required 

coverage angles indicated by the hatched area. Figure 4.3.6-4 indicates 

the required angular coverage. A similar clear upper hemisphere FOV will 

also be available for each of the GPS antennas except for slight 

obscurations that occur from pqyloads. Howe-.or, the two GPS antenna 

coverages complement each other so that payload obscurations are avoided. 

The low gain TDRS S-band antennas located near the GPS antennas will have 

the same coverage capability as the GPS antennas. 

The medium and high gain antennas (No. 2 and No. 3) used for the far range 

co-orbit cormunications need only a +I50 FOV around the forward and aft 

orbit track. This is easily o~tainex by placing these antennas on the 

upper portion of the lm;- keel, as seen in Figure 4.3.6-7. One each of 

these antennas is pointed along each track direction. Figures 4.3.6-9 li 

-10 show respectively the obscuration plots for the forwarC and aft lookfng 

5-ft diameter K-band antennas (No. 31, located cn the starhard side of the 

lower keel . For this plot the solar arrays were rotated from their 

position shown in Figure 4.3.6-7 by 900 in alpha and 520 in beta, 

placing the arrays in the position that can cause the greatest 

obscuration. The hatched in areas on each plot correspond to the +150 

FOV requirement, showing that a sufficiently clear FOV exists. A Ximilar 

set of plots would result for the medium gain antenna, and for the 

rendezvous radar antenna (No. 9) located on the port side of the lower 

keel. For the radar, however, a full heni spherical (+goo) FOV is needed 

around the posftive and negative x-axes. For the fornard looking direction 

shown in Figure 4.3.6-9, this FOV (theta=O-180, & phi+-90 d 270-360) is 

mostly clear except for some obscuration caused by the solar arrays. In 

the aft looking direction, where the required FOV is theta=O-180 & 

phi=90-270 degrees, a similar situation exists. This is shown in figure 

4.3.6-10. These areas of obscuration will reduce when the solar arrays 

rotate away from this orientation. The rsdar transpower antenna, which has 

the same FOV requirement as the rendezvous radat- antenna, will have 

obscuration similar to that shown in figurs 4.3.6-9 & -10. 

The S-band horn antennas (No. 6), which are located on the upper born 

pointing along the fomard and aft direction, will have a clear FOV over 

the fomard and aft hemisphere. The obscuration plots for the fomard 

looking antenna is shown in Figure 4.3.6-11. In this particular plot the 

z-axis of the antenna has been rotated so that it points along the x-axis. 

In this orientation the required FOV is the angular volume of thetad-980 

8 phi=O-3680 as indicated by the hatched area shown on the figure. Two 

of the low gain S-band antennas (No. 7) that are used to provide additional 

coverage to the Osbi ter will be located out on either end of the lower boom 

structure, pointed along the positive and negative Y-axis respectively (not 

shown on the figure) . 

The 7-element prox. ops. antennas (No. 4) are located near the top of the 

keel extension. They are placed away from the structure in both the x and 

y direction and pointed along the positive and negative x-axis, in order to i ' 
FIGURE 4,3.6 -8 UESCURATION PLOT FOR TDRS ANTENNA LOCATED UN UPPER BOOM. ANTENNA 

Z-AXIS PUINTED 180' FROM NADIR, X-AXIS POINTED ALONG FLIGHT PATH. 
n.. " 

PHI 

FiGURE 4.3.6 -9 CaSCURATION UT FOR FEDIUM AND H icy GAIN FAR RANGE ANTENNAS LOCATED 

ON FORWARD SIDE OF UPPER SEl ION G* I OWER KEEL., SOLAR ARRAYS ROTATED YO' IN AND 

52' IN P . ANTENNA Z-AXIS POIN iCD ,0° FROM NADIR, X-4x13 POINTED ALONG FLIGHT PATH. 
SOLAR - 4 

REQUIRED 

FOV 

PHI 

FIGURE 4.3.6 -10 OBSCURATION PLOT FOR fiEDIUM AND HlGH GAIN FAR RAHGE ANTENNAS 

LOCATED ON AFT SIDE OF UPPER SECTION OF LOWER KEEL. SOLAR ARRAYS ROTATED 99' 

IN I AND 52' IN . ANTENNA Z-AXIS POINTED 180' FROM NADIR, X-AXIS POINTED ALONG 

VELOi I TY VECTOR. 
PHI 

F!GURE 4,3.6. -11 OBSCURATION PLOT FOR ORB1 TER HORN ANTENNP LOCATED ON FORWARD 

SIDE OF ?HE UPPER BO(W1. ANTENNA Z-AXIS POINTED ALONG FLlGHT PATH. X-AXIS POIWTUl 

TOWRRD NAl;!!l. 
provide unobscured hmi spherical coverage in the forward and aft 

direction. Figures 4.3.6-12 L -13 show obscuration plots for these Go 

antennas, for the z-axis of each antenna rotated so they point along the 

positive and negative x-axis, respectively, as in Figure 4.3.6-11. 

Two of the low gain prox. ops antennas (No. 5) wi 11 be located on each end 

of the upper born pointed along the positive and negative y-axis, 

respectively. Figure 4.3.6-14 shows an obscuration plot for the antenna on 

the starboard end of the upper boom, with the z-axis of the antenna pointed 

along the negative y-axis. The solar panels for this figure were rotated 

900 in a1 pha and 520 in beta. The required hemispherical FOV is the 

angular volume of thsta=O-900 & phi=O-3600. This area is essentially 

clear except for the small area of blockage caused by the solar panels 

around phi=270°. The prox. op. antenna on the other end of the upper 

boom will have a similar obscuration plot, except the shawdows from the 

solar arrays are located around phf=90 degrees. The hemispherical FOV will 

lie in the positive y-axis direction. Tht other two No. 5 antennas are 

located on tne keel extension section with their axes pointed along the 

positive and negative y-axfs similar to the other antennas on the upper 

boom. 

Two of the MR.MS low gain antennas (No. 11) will be located on the top of 

keel extension, one on the forward side acd one on the aft side. The 

omnidirectional patterns of these antennas are aligned to provide coverage 

to the MRMS up and down each side of the keel. The other two antennas as 

stated earlier will be mounted on either end of the MRMS. 

Locations of the low gain broad beam antennas (Nos. 5, 7, 8, 10 6 11). are 

only suggested to provide good FOY coverage. The obscuration plots shown 

for the various antenna locations give line of sight blockage information 

only, and do not necessary represent the RF blockage which may occur. The 

actual antenna patterns from these antennas or dny of the broad beam low 

gain antennas will be affected by reflections and refractions from nearby 

str tures. A more detailed volumetric pattern computer analysis, of these 

antennas mounted on the structure, will be needed before true RF blockage 

car, be determined and suitable locations chosen. This also applies to the 

high gain pencil beam antennas, in which case the problem is from mu1 tipath 

ref1 ections from structures enteri ng the antenna side lobes. 
PHI 

FIGURE 4.3.6 -12 OBSCURATIOrJ PLUT FOP. 7 ELEMENT PRCX. OPS. ANTENNA LOCATED ON THE 

T,:I\RBOARD SIDE OF THE KEEL EXTENS IUN (NEAR TOP 1, ANTENNA Z-AX1 S POINTED ALONG 

FLIGHT PATH, X-AXIS POlNTED TOWARD NADIR. 
0.0 90.0 180.0 2M. 0 360.0 

PHI 

FIGURE 4,3.6 -1.3 OBSCURATION PLOT FOR 7 ELEMENT PROX. OPS, ANTENNA LOCATED ON THE 

PORT SIDE OF THE KEEL EXiEHSlON (NEAR TOP), ANTENNA Z-AXIS POINTED 180' FROM THE x 4 

FLIGHT -PATH, X-1, .IS POINTED 180' FROM NADIR, 9 
PHI 

FIGUPI: 4,3.6 -14 OBSCURATlON PLOT OF 1.01 GAIN PROX, 9PS. AhTENNA LOCATED ON THE 

STAk8ilARD END OF THE UPPt I! BOOM. SOLAR ARRAYS ROTATED 90' IN d AND 52' IN 4 . ANTENNA Z-AXIS POINTED IN STARBGARD Dl REI! ION, X-AX1 S POI HTED ALONG FLIGHT PATH, 
4.3.7 Launch and Asselably 

4.3.7.1 Introduction 

At present, no spacecra+t has ever been assembled on-orbit. The 

construction of a Space Station therefore presents a brcad spectrum of 

design and developlnent challenges that must be addressed early in a 

configuration analysi s. 

A Space Station reference configuration launch and assembly scenario was 

developed for use in determining Space Station design imp1 ications and 

requirements on the STS. It was recognized that there were many 

variations and options applicable to the launch and assembly of such a 

system as the Space Station; however, the following scenario is 

considered to be an achievable one based on the available inputs. The 

effort described herein includes the objectives that were established and 

all e:.planaticn of the assumptions that were made. Also inclu&d are 

descriptions of each pqyload package and th assewbly operations leading 

to IOi, Finally, estimates of the required .:m time1 ines for the 

assembly operations are included. 

The scope of this scenario is iimi ted to the launch and assembly/ 

installation of the major Station elements lsadi~g to IOC and does not 

address station payload related items. The intent is to convey a 

starting point for more extensive studies leading to a well defined 

1 aunch and assembly scenario. 

Only the assembly of the IOC Station is described. It is felt that the 

larrnch and assembly operations associated with the growth of the station 

would be quite similar to those presented here. 

4.3.7 .T Launch/F.ssembly Description 

Early in the study three general otjectives were selected and were 

incorporated into the reference scenario. The three ob jecti ves are: 

1. Minimized STS-supported EVA. EVA that is conducted from and 

supported by the Orbiter is a iimi ted resource. On a typical seven day 

mlssion, there is currently a limit of two, two-man EVA'S of 

approximately six hours duration each. In addition, should an emergency 

arise, the Orbiter may be required to return to Earth before a1 1 mission 

objectives have been met. For this reason an objective was to mjnimize 

the amount of EVA required in the early assembly phases of the Space 

Station. Some EVA wiil of necessity be required during the early phases 

of the buildup, however, the objective was to keep the durations and 

complexities to a minimum. After the station csn be manned and the 

iiecessary consumables are on board, EVA could be conducted without the 

presence of an orbi ter . Therefore, many operations required to achieve 

the final station conf,-jdration but not required to reach habitability 

were reser-ed for sta tion-supported EVA missions. 
2. Self-powered dnd controllable spacecraft fmn Flight 1. 

Another imortant objective for launch and assembly planning was ~Teave 

on-orbit f'm the fikt flight a spacecraft element capable-of generating 

some amount of power for attitude control, thermal control, and 

collmrnications functions. This goal dictates the order in which mjor 

station elements are launched through the first two flights. 

3. Early habitability. A final important objective was to man 

the station as early in the buildup phase as is practical without making 

rnajor subsystem changes. Having the- station manned allows the assemblyoperations such as systems checkout, interface verification, and 

utilities instal lation to continue behreen STS flights. This objective 

dictates the order in which station elements are launched beginning with 

the third flight. 

In establishing the launch and assembly scenari~, certsin assumptions 

concerning the STS were made- These include the following: 

1. Only one Orbiter is required for each launch/asswbly phase. 

2. 011ly one Logi st'cs Yodule is required to reach IOC (no 

changeout during buildup) . 

3. Flights 1 through 5 are capable of supporting EVA. 

4. Shuttle Remote Manipulator Syst~m i SRMS) is available on 

Flichts 1 through 7. 

5. The Orbiter would not be used in a hovering mode from point 

to point mode. 

It was also assumd that the GMV would not be used to assist in the 

assedly operations. 

The launch/assembly scenario is composed of seven STS flights and 

associated operations. Figure 4.3.7.2-1 illustrates the major components 

of the Space Station and indicates the flights on which they are 

manifested. 

The basic structure of the Soace Station is composed of 1 inear, 

sequentially aeployed, single fold, box truss type beam segments as 

described in the Yhite Paper entitled, "Space Station Truss Structures 

and Construction Considerations." The reference structure has strut 

members that are two inches in outer diameter. A single unit cf this 

beam is a 9- by 9-ft cube formed by 18 struts and eight connecting nodal 

joints. Utilities are pre-integrated into the structure, thereby 

reducing the task of routing cables, hoses, etc., to one of making 

relatively simple connections at the major structural interfaces. 

Certain seg~ents of the station structure, it is felt, could be 

efficiently assembled in a strut-by-st.ut sinner by the crew. This is 

especially true for short connecting segments which could become quite 

difficult to integrate, package, and deploy with 1 onger segments. 

The 1 lsting that follows is a sumry of the pqyload components making up 

each of the seven flights. 

FL I GHT MAJOR SPACE STATION ELEMENTS 

I - inboard solar array wing pairs 

- rotating power joints 

- power conditioning radiator arrays 

- inboard transverse boom structure 

- power conditioning equipment 

- control equipment 

- comnunication equipment 

- berthing structure 

- MRMS 

I I - lower keel st-ucture 

- port keel extension structure 

- starboard keel extension structure 

- lower boom sttucture 

- main radiator panels 

- closeout structure 

- main radiatcr booms 

- RCS 

- berthi straucture 

I11 - HM1 (habitation module 1) 

- A11 (airlock 1) 

- AL2 (airlock 2) 

I V - HM2 (habitatfon mcdule 2) 

- upper keel structure 

- upper boom structure 

- antennas 

V - LOG1 (logistic module 1) 

- port salar array wing pair 

- starboard solar array wing pair 

- port outboard transverse borm structure 

- starboard outboard transverse boon structure 

V I - LAB1 (laboratory module 1) 

- equipment spares 

- external experiments 

VI I - LAB2 (laboratory 2 j 

- equipment spares 

- external experiments 
Such iterus as the OMV, OWV hangar, and satellite servicing equipment 

have not been included in this scenario. Further study is needed f n this 

area to determine whether these items could be delivered on Flights VI or 

VIX, on subsequent flights, or possibly by resupply flights. 

*.3.7.3 Shuctle Payload Bay Packaginy 

1 major design constraint on the princi pal Space Station elements is 

cuupatibility with the STS for launch. This constraint includes such 

c -itical payload parameters as weight, volume, and center of gravity. 

For each of the seven STS flights in the reference scenario, an analysis 

was carried out to define t~ payload package that was ~i thia the 1 iaits 

set by geometry and latrnch capability sf the Orbiter. Figures 4.3.7. --I 

through 4.3.7.3-7 illustrate the laur~ch package for each flight. 

4.3.7.4 Component Installation D~scr~ption 

Following is a description of :hc: 3ssembly procedures using illustrations 

as appropriate. The approach taken was to develop only the details of 

the operations necessary to understand the total assembly task. 

FLIGHT I 

TISK/OPERATION FIGURE 

1. Install radiator paoels 1-1 

2. Remve payload package from launch 1-2 

restraints; rotate 90° and insert 

f nto deployment restraints 

3. Deploy port half of deployable 1-3 

structure; deploy stzrboard half 

of deployable structure 

4. Deploy solar array blanket boxes; 

deploy solar array blankets 

5. Erect berthing structure support 

bay; attach berthing port 

6. Install MRMS (Mobile Remote 

Manipulator System] 

7. Perform systems checkout; depart 

Mechanism 

EVA 

Orbiter 

systems 
I Radiator oanels / , Deployable strut'-rre 

Antennas (ntowed) Rotary joint-PYP 

Radiator heat 

r 

--- 

G) 

c..l 

0 

T - - --*- - 

Payload bay 

dynamic envelope 

Figure 4.3.7.3-1 Launch package 1. 
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FLIGHT I1 

TA!X/OPERATION FIGURE MTHOD 

Remove lower keel package from payload 11-1 M/EVA 

bay; unfold deployment rails; attach 

package to transverse boom 

Deploy 1 ower keel deployable structure; 11-2 Mechanism 

un fol d radiator support boans 

Unfold radiator arms; erect Go keel 11-3 MW/EVA 

extension bays on the port and starboard 

sides of the lcwer keel boom 

Install radiator panels in the port 11-3 MRUS/EVA 

and starboard heat exchanger boans 

Remove the port keel extension boom I 1-4 MRMS1EV.Q 

package fran the cargo bay; transport 

to attach~nent site; unfol d deployment 

rails; attach package to erected keel 

extension bay on port side 

Deploy port keel extension structure 

Remove the starboard keel extension 

boom package from the cargo bay; 

transport to attachment site; unfold 

depl oyment rai 1 s; attach package to 

erected keel extension bay on starboard 

si de 

Deploy starboard keel extension structure 11-6 

Erect keel extensi on interconnect bays 

Perform systems checkout; depart 11-6 

Mechanism 

MRPu/EVA 

Mechani sm 

Orbiter 

Systems 






FLIGHT I11 

TASK/OP€RATION' FIGURE ETHOD 

1. Remove HMl from payload bay; attach 111-1 WS/EJA 

to keel extensfon structure 

2. Remove AL2 fran payload bay; attach 111-2 MRMS/EVh 

t.~ HMl 

3. Remove K1 from payload bay; attach 

to HMl 

4. Pcrform system checkout; depart 111-3 Orb1 ter 

systems 


1. Remove HM2 from payload bay; attach 

HM1 

2. Attach A12 to HM2 

3. Remove upper keel structure package 

fr.~ payload bay; transport to 

attachment site; unfold deployment 

rails; attach package to transverse 

beam 

4, Deploy upper keel and uppe. boom 

structure 

5, Perfonn systems checkout; depart 

FIGURE 

IV-1 

METHOD 

MRHS/EVA 

Mechani sn 

Orbiter 

sy s terns 


C 

(I' 
FLIGHT V 

TASK/OPERATION FIGURE 

1. Hernove LOG1 from payload bay; attach V-1 

to HM2 

2. Remove port solar array addition 

package from payload bay; transport 

to attachment site; attach to 

transvarse beam 

3. Deploy port octboard transverse 

beam structure 

4. Remove starboard solar array addition 

package from payload bay; transport to 

attachment site; attach to transverse 

beam 

5. Deploy starboard outboard transverse 

beam structure 

6. Deploy salar array blanket boxes; 

deploy solar array blankets 

7. Perform systems checkout; depart 

Mechani va 

V-4 Orbiter 

Sy s teas 

It was assumed for scenario development purposes that the Station was 

permanently manned at this point in the assembly sequence. Mny 

operations could be conducted at this point including perfonaing EVA 

assembly tasks such as veri ficarion of attachments, installation of 

additional radiator panel s, and f nstal lation of permanent hard lines as 

necessary. Tbis will greatly reduce the amount of time that an orbiter 

has to be on-crbit to support the assembly operations. 
Port and starboard outboard solar array and /' 

deployable structure packages in cargo bay 

\ 

Figure V-1 Logistics modulo attached. 

, i . at- 1-A 

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FLIGHT VI 

TBSK/OPERATION FIGURE MET HOD 

I. R.move LAP2 from pay 1 oad bay; at tach 

to HM2 and to keel extension structure 

2. Perfom systems checkout, depart 

SRMS/EVA 

MRMS 

VI-1 Orbiter 

Systms 

No major Station element other than LAB2 has been designated for launcn 

on Flight VI. Such items as Station systems spares and external payloads 

co~rld be d~l ivctred on this fl ight as suggested in the payload sumsry 

given it? 4.3 -7.2. It is felt that as further studies are conducted many 

items will be identified as being required early in the build-up 

sequence. Hence, the available volume on this flight and on Flight VII 

was intentionally ill ustrated in the launch packaging figures irr 4.3.7.3. 

FLIGHT VII 

FIGURE 

1. Remove ! 431 frm payload bay; attach 

to HM1, and LAB2 

2. Perform systems checkout; depart 

SRMSIEVA 

MRMS 

Vi 1-1 Orb1 ter 

Systms 

As on the previous fl ight, there exlsts a volume for miscell aneous I terns 

or pay1 oads in the launch package. 

Figure <.3.7.4-1 illustrates the final configuration of the 1OC Reference 

Space Station. Assembly techniques described in the reference scenario 

would be appl icable to the envisioned growth configuration of the 

Station. 

4.3.7.5 Assembly Operations Crew Time1 ines 

P- a part of the study estimates were made of the crew timlines required 

to perform the on-orbit assembly operations These estimates were based 

on NASA Contrartor Report 2751, "Analysis of Large Space Structures 

Assembly", and on SRMS, EVA, and neutral buoyancy testing experience. 

Table 4.3.7.5 1 1 ists the estimated number of hours of two-man EVA'S 

reauired to perform the assembly tasks associated with each flight in the 

reference scenario. 

The results of this analysis, including ii breakdown of the EVA, SRMS, and 

YRMS operations required to perform the assembly tasks described in 

4.3.7.4, are detailed in Appendix A* 


4.3.8.1 Manipulator Operations 

Assembly of the Space Station, as well as the placement and servicing of 

external ly mounted payloads and maintenance of station ORU's a1 1 require 

a Mobile Remote Manipulator System (MRMS). Some of the more crit.ica1 

requirements of the MRPIS will be station assembly, module removal, 

OMV/OTV berthins in the hangar area, deployment of the OMV/OTV from the 

hangar area and as an aid to OMV, OTV and satellite servicing. The 

analysis conducted emphasizes the use of the current Shuttle RMS to the 

n~aximum extent possible for assembly. The MRMS was considered only for 

those operations which exceeded the reach capability of the Shuttle RMS. 

The MRMS analysis was performed using the RMS Desk Top Planning 

Simulation (RPS) developed for RMS mission planning activities and used 

to define RMS payload hand1 ing capabil ities and procedures for STS 

missions. The program was updated and modified to include the reference 

configuration, as well as a general ized manipulator in the sense that the 

length of the manipulator booms can be varied to accomnodate larger reach 

envelopes than the current RMS. The number of active joints can be 

reduced and the booms shortened so that a Hand1 ing and Position Aid (HPA) 

type of mechanism can also be accurately simulated. 

The manipulator analysis included herein is based on a kinematic model of 

the RMS in that no rigid or flexible body dynamics are included. This 

limitation, however, does not invalidate the feasibility of using the RMS 

for station assembly since all modules handled are within the weight and 

inertia 1 irni ts verified for standard RMS operatiors. The study results 

are based on the current RMS control algorithms and software and verify 

the reach capability as well as the maneuver path for both the RMS and 

the MRMS- The simulation also checks for singularities and joint reach 

limits. In summary, all maneuvers studied for the reference 

configuration assembly sequence should be valid with the exception of 

possible crew visibility constraints. RMS operator eye-point and CCTV 

viers can also be generated using the RPS simulation, and these result< 

wiil be cansidered in future analysis. The study focuses on assembly ~f 

the major structural components and modules of the Space Station only and 

was performed using the fa1 lowing assumptions and guide1 ines: 

a. Orbiter RMS use was restricted to the current port RMS. 

b. All construction was perf~med with the Orbiter firmly 

attached to the Space Station structure. 

c. The MRMS dimensions and control algorithms were identical to 

, those of the current RMS. 

d. The MRMS could move on either the front or back side of the 

upper and lower keels, the keel extension, and the transverse boom. 
e. Crew vis!bil i ty of maneuvers was not considered. 

f. Grapple fixture locations and orientations considered only 

RMS/MRMS reach capabi 1 i ty and did not consf der module structural design 

at the attach point. 

The RMS/MRMS maneuvers are not necessarily the most optimum maneuvers, 

but do verify reach capability, adequate clearances along the entire 

maneuver path, and freedom from encountering reach limits or 

si ngulari ties. 

4.3.8.1.1 Station assembly 

a. Flight 1 - The Orbiter port fMS is required to remove the 

initial package of truss structure components from the Orbiter payload 

bay and place it on the guide rails which permit the transverse boom to 

be deployed (Figure 4i3.8.1-1). The Orbiter RMS :rill also be required to 

assist in 2ositioning the MRHS onto the truss structure once it is fully 

deployed. 

b. Flight 2 - Once the Orbiter is docked to the structure 

following Flight 1, the MRMS can be used to remove the new structure 

packages from the Orbiter and position then for attachment to the 

existing structure. Figure 4.3.8.1-2a shows the MRMS removing the 

package that makes up the lower keel (step 1) and placing it in position 

for deployment (step 2). Figure 4.3.8.1-2b shows the sane maneuver using 

the Orbiter port RMS. 

c. Flight 3 - Flignt 3 requires the Orbiter to dock to thr 

auxiliary berthing ring at the module end of the lower keel as showr, in 

Figure 4.3.8.1-3. The MRMS is then used to remove the first Habitation 

Module (HM1) from the Orbiter payload bay (step 1) and position it for 

attachment to the keel extension (step 2). 

d. Fl ight 4 . The Orbiter is docked to the HM1 as shown in 

Figure 4.3.8.1-4. Tie rationale for docking the Orbiter to the HM1 as 

shown a1 1 ows the module to be entered directly from the Orbiter without 

an EVA. This berthing port is not the primary or auxiliary, but it is 

required to facilitate transfer of the second Habitation Module (HM2) and 

still allow access to the interior of the HM1. As on Flight 3, the MRMS 

is used to remove the HM2 from the Orbiter paylnad bqy (step 1) and 

position it as shown (step 2). 
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e. Fl ight 5 - bJi th the Orb1 &r docked to the prlrary berthing 

port on the HM1 in the orientation shown in Figure 4.3.8.1-5, the Orbiter 

port RMS is used to remove the logistics nodule (LW) from the Orbiter 

payload bay (step 1) and attach it to the auxlliary berthing ring located 

011 the truss structure (step 2). The MRHS is then used to remove the LM 

from the auxiliary berthing ring (step 3) and attach it to the tM2 (step 

4). Care must be taken when maneuvering the MRMS around the external 

airlock attached to the HI. For this and all remaining assembly 

flights, the Orbiter 1s attached to the primary berthing port on the H1. 

f. Flight 6 - A scenario similar to that of Flight 5 is used 

to transfer the LAB module (LABlj from the Orbiter to the MRMS for 

dttachment to the existing modules: however, the MRMS is required to 

operate on the opposite side of the lower keel from a1 1 previous flights 

(Figure 4.3.8.1-6). 

g. Flight 7 - The transfer of the second laboratory mdule 

(LAB21 from the Orbiter bay to the MRMS (steps 1, 2 and 3) of Figure 

4.3.8.1-7 is identical to that of Flight 6. The MRMS then positions iAB2 

(step 41, completing the racetrack configuration (Figure 4.3.8.1-7). 
L.. 

EXTERNAL 

AIRLOCK 

Firgure 4.3.8.1-5 - FLIGHT 5 PLACEMENT OF LOGISTICS MODULE REQUIRING, BOTH THE 

CRiG(NAt VA'c;,' 

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1- -- -- 
4.3.8.1.2 Station naintenance/payload placement 

Figure 4.3.8.1-8 shows representative locations and sizes of the 

external ly mounted pqyloads and servicing structures. The Space Station 

structure is removed from this figure for clarity. In Figure 4.3.8.1-9, 

the Space Station structure has been inserted. Also shown in this figure 

is the reach envelope of the MRMS as defined by the location of the wrist 

pitch joint. The envelopes are shown for various locations along the 

truss structure where MRMS assistance in servicing the station and 

externally mounted payloads could be required. With the URMS able to 

travel along both the front and back sides of both the upper and lower 

keels and the tracsverse boom, it is apparent that a1 1 areas of concern 

can be reached with an MRMS hrving the same dimensions as the current 

RMS . 

4.3.8.2 Orbital Flight Operat'ons 

The Space Stz'ion System consists of a manned orbiting Space Station, 

space platforms, free-flying sate1 1 i tes, orbital transfer vehicles, and 

orbital maneuvering vehicles that interact with the Space Station in 

orbit. In order to utilize all of these elements efficiently, many 

detached operations will int2ract directly with the station. Missio~ls 

involving two or more spacecraft will frequently require rendezvous 

and/or Proximity Operations (PR3X 3PS). Rendezvous is the act of 

bringing two vehicles, a target dnd a chaser, together. The amount of 

time requirea to rendezvous can range frcm ~C?IJ~S to days. This time may 

influence the launch/departure time. In general, tlie rendezvous 

maneuvers will leave the chaser vehicle at a specified offset point ahead 

of the target if the ch2ser is mawed or behind the target if the chaser 

is unmanned. At thls point, the -?ndezvous is completed. The chaser qy 

then stationkeep prior to init at:or! of PROX OPS involved with transition 

to the Farget and berthing. PROX OPS is concerned with that ?ortion of a 

mission when two vehicles are within approximctely 1000 feet of each 

other and one is performing maneuvers relative to the other. These 

maneuvers may :r may not be pe; formed ma:.iral?y by the active vehicle 

control 1 er; however, the results of these maneuvers will be mni tored 

closely through direct visual or electronic means. PROX OPS begins at 

that final stzte resulting from the rendezvous activities (normally on 

the target's velocity vector). PROX OPS maneuvers may include 

transitions to various offset points, statirikeeping, approachs, 

flyarouads, and the separation to a 'stand-c. , " position. Other 

activities such as departures, monitoring and servicing co-orhi ti ng 

satellites {satellites traveling in the same orbit as the Space Station) 

are a1 so inc!uded. 

Flrgura 4.3.8.1-9 - KRM8 eCIACH CAPABILITY FOR SERVICING EXTERNALLY MOUNTED 
4.3.5.2.1 Operational Control Zones for the Space Station 

The concept of Operation,~l Control Zones has been developed in an effort 

to provide an operational foundation for the Space Station System. 

Any concept for multiple detached pera at ions must address the following 

operational requirements. First, the concept must a1 low for standardized 

flight planning and operations. This means that the routine need for 

long lead-time preparation prior to execution of multiple detached 

operations must be avoided. Second, the concept must permit standardized 

crew planning and operations. 7!1i s will simp1 ify crew training and 

day-to-day crew activity planning. Third, the concept should allow early 

definition of requireinents. These include Comnunication and Telemetry 

requirements, as we1 1 as Comnand/Control /Tracki ng requirements. Knowing 

the requirements early will help to minimize over-design. Finally, any 

plan for conducting multiple detached operations should provide for 

collision zvoidance between spacecraft and hold Space Station disturbance 

and contami2ation from thr-uster firings to a reascnable level. 

In order to satisfy these operationdl requirements, tke current concept 

earmarks nine specific operational regions or zones in the orbital area 

of the Space Stdtion. Each of these zones is deeicated to a specific 

type( s ! of detachgd operation( s ). 

The fol oding cperat ional considerations and guide1 ines were taken intc 

account in defining the location and size of the Operational Control 

Zones : 

a. The dinoi.int of work on the Space Station will be he1C at a 

minimum by having the ground responsible for all flight planning, 

tracking, and ~o~trolperations where a high level of crew involvement 

would otherwise be required for final rendezvous and PROX OPS phases. 

b. Fw all flight operations in which a vehicle is approaching, 

stationkeeping, performing flyarounds, ana/or berthing with the Space 

Station, a crew member will be required to actively monitor and if 

necessary control the flight operations. Thus, if the active vehicle is 

unmanned, the capabi 1 i ty to actively take over Comnand/Control must 

exist. For most spacefl ight operations performed to date, this contro! 

us~ally takes over at a separation distance of approximately 37 km (20 

nmi ) between the two vehicles. 

c. Co-orbiting Satellites will be maintained in a desired orbit 

position relative tc the Space Skation and will be actively controlled 

from their groti,id control centers. The definition cf this relative 

position will be a :unction of the pay11 3d's reauirflments on the Space 

Station. If it requires no Space Station support other than routine 

servicing, its relative control position can De allcwed t~ be very large, 

possibly 
hundreds of kilometers along the Space Station's flight path. However, 

a1 1 co-orbi ti ng sate1 1 i tes wi 11 be maintained approximate1 y cop1 anar wi th 

the Space Station. 

d. Non-co-orbiting satellites will be positicned to avoid any 

possible recontact with the Space Station and its co-orbi ting sate1 1 i tes. 

The Command/Control of the non-co-orbi ting sate1 1 i tes wi 11 be maintained 

by their ground control centers. No attempt will be made to control the 

relative position of these vehicles, cvcept in instances of coll ision 

avoidance. However, some control may be exercised upGn the satellite's 

relative altitude in order to control the magnStude of the relative plane 

at the time of satellite servicing, dspecially if the satellite is to be 

serviced frdm the Space Station. 

e. Payloads returning from geosynchronous or other high energy 

orbits, and requiring Space Station support, will be initially targeted 

to a parkir~g orbit above the Space Stati~n. This will ~llow for relative 

phasing and planar adjustments between the returning vehicle's orbit and 

that of the Space Station. This parking orbit will also act as a 

hand-over interface for fl ight planning activities, allowing the descent 

from the high energy orbit to be decoupled from those activities 

occurring in low earth orbit such as final rendezvous with or servicing 

by the Space Station. 

A detailed discussion of each of these operational zones is out1 ined 

below. Refer to Figures 4.3.8.2.1-1 through -8. 

Zone 1: Prox~mity operations zone 

Zone 1 consists of the region of space enclosed by a 1-km (approximately 

3280 ft. ) sphere centersd on the Space Station, Within this zone, all 

nominal prorirni ty aperations including stationkeepi ng, flyarounds, final 

approaches snd berthing wi 11 take place. In addition, fiominal MMU/EVA 

activity will be performed within this zone. 

Zone 2: Control zone 

This concept assumes that tire Space Station will have the capability to 

acquire active Commnd/Control /Track of any unmanned vehicle that enters 

this zone. 

This zofie contains a "slab-li ke" vnlume of space that is curved and 

centored about th~ Space Station orbit track. (Note that this shape is 

comnon to all zones cenicr~d around the Station's orbit.) In the 

horiror~tal direction (curvilinear "x" 1, Z~ne 2 begins 37 km (20 nmi ) 

behind the Space Station and enas 37 km ahead of it. In the vertical 

directis; (curvilinear "z"), the zone starts 37 km below the Space 

Station's orbit and evds 37 km above it. Finally, Zone 2 extends + 9 kn 

(5 ri!~ii ) out of tne Space Station'z orbital plane (cwvi 1 ivear "y" )T 
In order to support its Comand/Control responsibility- the $pace Station 

is assumed to have several additional capabilities which will be used 

only when an unmanned vehicle enters Zone 2. 'he functions are: 

a. Mcnitoring: The Space Station will be capab'e of actively 

monitoring the "system health" portion of any unmanned vehicle's 

telemetry. This will support any safety question; such as system 

ma1 fu~ction? that must be considered while maneuvering the active vehicle 

within near proximity of the <pace Station. 

b. Tracking: The Space Station will be capable of 

trackiny/comnanding/cont~01l ing a1 1 unmanned vehicles within the Control 

Zone. Lfrniting Space Station tr$ffic cantrol authority to Zone 2 

provides two obvious advantages. The first is that it reduces and helps 

to define antenna requirements. The se~ond advantage is the freeing of 

valuable crew time that is better srent on things other than routine 

mnitoring and flight planning activities. 

Zone 3: Departure zone 

Most nominal departures from the Space Station will take place within 

this zone, after initial deployment and separation maneuvers are 

performed in Zone 1. Zone 3 is centered upbn the Space Station's flight 

path and begins at the Space Station, extending forward to approximately 

185 km (100 rlrni 1. In the verticdl dimensien it begins approximately 37 

km (20 nmil below the Stltisn'z srbi .iid ei,ds appr.axirnai;ely 51 Km above 

the Space Station's orbit track. TI , zone extends approximately - + 9 km 

(5 nni ) out of the Space Station's 01 lit plane. 

Zone 4: Rendezvous zone - 

All nominal rendezvous with the Sp- 2 Station will be targeted to arrive 

within this zone. Zone 4 is centered upon tke Space Station's orbit 

track, and extends rearward from the Space Station to approximately 185 

km (100 nmf 1. It is approximately + 37 km (20 nmi ) in the vertical 

dimension, and approximately + 9 km75 nmi ) in the out-of-plane 

dimension. This zone's location and size have been designed to be 

consistertt with the standard Stable Orbit Rendezvous Technique In 

Stable Crbit Rendezvous, the chaser arrives at an offset point some 

distance behind the target and performs its closing maneuvers from this 

point. 

Zone 5 & 6: Co-orbiting zorei 

These two zones are dedicated to co-orbiting satellite oqerations. Zone 

5 is the leading co-orbiting satellite zone. It begins approximately 185 

km (100 nmi) ahead of the Space Station. It is centered a;-ound the Space 

Station's flight path, and ektends forkard to the opposite side uf the 

orbit; i.e., 180" ahay fror the Space Station (apFroximately 
21,609 km or 11,668 nmi for an assumed Spdce Station orbit of 500 km 

(270 n. rn!.). Zone 6 is the Trailing Coorbiting Sateliite Zone. It 

begins approximately 185 km (100 n. mi.) behind the Space Station and 

follows the Space Station orbit track until it contacts Zone 5. Hence, 

the co-orbiting satellite zoaes are continuous arou~d the Earth. 

Zone 7 & 8: Non-co-orbitins zones 

These zoncs contain the con-co-orbiting satellftes. Both Zones 7 and 8 

are concentric spherical shells centered about the Earth. Zone 7 

begins approximately 37 km (20 nmi ) below the Space Statim. It 

extends downward to approximately 185 km (100 nmi ) altitude above the 

Earth. Zone 8 begins approximately 37 kin (20 nmi) above the Space 

Station. It extends to an altitude of approximately 352 kin (190 nmi 1 

abovfi the Spac? Station's orbit. 

Zone 9: Parking - orbit zone 

This zone surrounds the parking orbit used by orbital transfer vehicles 

(OTV's) returning from gersynchrorsous or other high energy orbits. It 

may also be used by splcecraft returning from lunar or planetary 

missions. Zone 9 is a spheri -a1 shell, centered upon the Earth, that 

encloses the Space Station and Zones 1-8. This shell begins 

approximately 352 km (190 nmr) above the Spxt Station orbit track, and 

extends upward to an a1 ti tude approximately 389 km (210 nmi above the 

Space Station. 

4.3.8.2.2 Rendezvous o~erations 

Future orbital missions involving two or wre spacecraft will 

frequently require rendezvous. A s!mplistic definition of rendezvous 

is the act of bringing two vehicles, a target and a chaser, together. 

The target vehicle is assumed to be in a known, stable orbit. The 

chaser performs a series of predetermined tnaneuvcrs designed to ensure 

that the rendezvous with the target will occur at a specified psition 

in the target's orbit and within a prescribed amount of time. The 

magnitude of these maneuvers is dependent upon many factors, but it may 

be significantly Influenced by the time of launch (~r departure, in the 

case of a Space Station based "launch" ). The target position at 

rendezvous completioa de~e~ds mainly on 1 ighting requirements, while 

the amount of time that is prescribed for the rendezvgu?. can range from 

hours to days. The prescribed rendezvous time may influence the 

launch/departure time, which in turn influences the lfghting in the 

fintl phases of the mission. In general, the rendezvous maneuvers will 

leave the chaser vehicle at a specified offset point (typicnl'y 1000 

ft ) ahead of the target if the chas2r is manned or behind + target 

if the chaser is unmanned. At this point, the rendezvous 

completed. A stable orbit rendezlous is assumed which takes the chaser 

to a close offset range (between 5 nmi and 29 mi 1. Manned vehicles 

bil continue t3 approximateiy 100 feet in a 
+ 

ZONE 33 ZONE 43 

ZONE 5' 

-8 

100 PIMI 

(1 85 km; 

-- 

L 

ZONE 72 1 

ZONE ASSIGNMENTS 

1 PROXIMITY OPERATIONS ZONE 

2 CCNTROL ZONE 

3 DEPARTURE ACTlVlTY ZONE 

4 RENDEZVOUS ACTIVITY ZONE 

5 LEADING COORBlTlNG SATELLITE ZONE 

6 TRAlLlNG COORBITING SATELL;TE ZONE 

7 LOWER NON-COORBITING SATELLITE ZONE 

e UPPER NON-COOABITING SATELLITE ZONE 

9 PARKING ORBIT ZONE 

NOTES. 

1 THlS DRAWING IS CURVlLLlNEAR AND 

DRAWN TO SCALE 

/ 2 THlS ZONE IS CONTINUOUS ABOUT T

3 THlS ZONE BEGINS AT THE SPACE ST

i 4 THE MINIMUM OPERATING ALTITUDE 

ZONE IS - 100 NMt (-185 kmj 

5 ZONES 2,3,4,5, 8 6 EXTEND -- f9 krn(f 

OUT-OF -PLANE 

6 ZONES 7.8 AND 9 %RE SPHEdICAL SH

ABOUT THE EPaTH 

1 

Figure 4.3.8.2.1-1 - OPERATIONAL CONTROL ZONES. 
y/ Lspace Station flight path 5 

Figure 4.3.8.2.1-2 - SCALE DRAWING OF OPERATIONAL iONTROL ZONES IN 3 DIME

1 ;) I I I ., Zone 3 

Zone 

Zone 

Figure 4.3.8.2.1-4 - OPERATIONAL CONTROL ZONES: 

NEARFIELD VIEW, SLIGHTLY OUT-OF-PLANE 
Ui{!G-i%AL PAGE' iS 

OF POOR QUALI'W 

Figure 4.312.1-5 - OPERATIONAL CONTROL ZONES: NEARFIELD VIEW, 

OCiT-OF-~LAP : i3RAWING TO SCALE) 
Flggure 4.3.8.2.1-6 - OPERATIONAL CONTROL ZONES: NEARFIELO VIEW; 

EYEPOINT SLIGHTLY OUT-OF-PLANE (VIEW A) 

(DRAWING 13 SCALE) 
ONG?.iAL P>.G.I .-i 

OF POOR Q3Ab11"f 

Figure 4.3.8.2.1-7 - OPERATIONAL CONTROL ZONES: 

VIEW OF ZONE 1 FROM WITHIN ZONE 2. 

man-in-the-1 oop mds while unmanned free-flyers wi 11 continue with an 

automated modz. A11 planning, targeting and navigation for the longer 

ranges of the rendezvous will be externat to the Space Station and 

therefore will impose no requirement on the Space Station. Once a 

vehicle is in the final rendezvous phase the Space Station will monitor 

a1 1 vehicles. Monitoring includes co~munication, tracking, and back-up 

targeting for manned vehicles and conmunication, tracking, targeting, and 

back-up take over of unmanned vehicles. These activities imply 

requirements on the Space Station for various hardware and software 

systems. 

4.3.8.2.3 Proximity operations 

In order to develop an efficient set of PROX OPS procedures, numerous 

operational considerations must be addressed regarding the final position 

achieved by the active vehicle and the PROX OPS techniques used to 

manebver this vehicle to the desired position. A major concern is the 

effect sf the RLS plume on the target. This plume impingement creates 

problems of contamination arid disturbance (over-pressure) experienced by 

the target. In addition, the active vehicle's RCS propellant usage is 

critical. Propel lant constimption influe~~ces the choice of braking 

maneuvers and stationkeeping techniques. Another concern for Orbiter 

operations is whether it is more efficient to compute required maneuvers 

using onboard targeting software or to have the crew execute the 

maneuvers manually using out-the-window data such as Crew Optical 

Alignmnt Sight (COAS) or Rendezvous Radar (RR) data. In any case, crew 

visibil ity and procedural simplicity have considerable impact on the 

selected maneuvers. A final major consideration is that of lighting 

during PROX OPS procedures. 

4.3.8.2.3.1 Lighting impact on proximity operations 

The PROX OPS procedures discussed here are extremely dependent on the 

1 ighting available for the crewman to view the activities taking place. 

The crewman who is monitoring activiti~s or actively involved in 

performing the PROX OPS maneuvers requires that lighting (real or 

artificial) must be of the proper intensity as well as being projected in 

the corr~ct direction. The primary source of natural 1 ight is the sun, 

and its position reiative to a vehicle is a function of time. Hence, 

operations which impose a constraint on lighting direction also impose a 

restriction on the place in orbit when the activity can be performed or 

the attitude of the viewing vehicle. With this consideration, the 

direction of lighting must be in favor of the viewer having the final 

authority in performing the activities. Therefor?, 1 ighting direction is 

a function of the active vehicle being manned or unmanned. 

If the active vehicle is manneg, the 1 ighting should be in a direction 

favorable to the inanned cctiue vehicle. For the situation of natural 
lighting, the 1 ighting direction can be controlled by scheduling the 

activities at the appropriate time of day. It is for this reason that 

mnned active approaches are initiated on the positive velocity vector (+ 

V-bar) of the target and designed to take place as soon after orbital 

sunrise as possible and be completed before orbital sunset. Conversely, 

utmanned approaches to a manned Space Station will be initiated from 

behind the Space Station; i.e., on its negative velocity vector 

(-V-bar). This provides the manned Space Station with the proper 

lighting needed to lonitor the vehicle's approach. 

4.3.8.2.3.2 Approach and berthing to the Space Station 

A typical returt, mission for an active vehicle which has already executed 

a successful return rendezvous profile placing it at an 8 mi offset 

point behind the Station begins with an auto-return maneuver sequence 

that brings the vehicle to a 1000-foot offset point. This point will be 

behind the Station for unanned vehicles and ahead of the Station for 

manned vehicles, consistent with favorable lighting. The final PROX OPS 

returns are considered to begin at this point. 

The final PROX OPS begin with a period of stationkeeping at 1000 feet. 

For returns to any port, a maneuver is perforred to establish a closing 

rate until a position TBD feet frm the Space Station is reached. At 

this range stationkeeping may be perfomed, if required. For V-bar 

ports, a final closing rate is established along the V-bar until the 

berthing range is reached. Returns along other axes involve a similar 

approach, if required, then util ize a constant range flyaround to a: ign 

the approaching vehicle on the appropriate axis for final closure to the 

berthi~g port. 

4.3.8.2.3.3 Separation from the Space Station 

A typical separation scenario for a vehicle docked to the Space Station 

begins with small separation rate (approximately 0.2 fps using either a 

llechanical system or a vehicle maneuver) to begin the vehicle movement 

away frol the port. The direction of the separation will be directly 

away fra the berthlng port to maximize an opening rate. A small 

Intermediate mneuver may be required to ensure a favorable geometry 

between the separating vehicle and the station for the large separation 

burn to follow. After a coast of approximately 15 minutes the separation 

vehicle will perform the large separation maneuver of approximately 3 

fps. This will ensure a separation range of approximately 10 mi at 

orbital transfer mneuvev ignition. This range is needed so that the 

station is outside the assumed explosion range of the separatf ng 

vehicle. It is at this point that the separaticn iias been ccupleted. 
4.3.8.2.3.4 Plume impingement/contamination 

Keeping plume impingement on all orbiting elements at the lowest possible 

levels is a very important aspect of PROX OPS procedures both in Shuttle 

and Space Station operations. Currently, the Orbiter uses its RCS jets 

(primary and vernier) to control d ts attitude and execute various 

maneuvers. Subsequently, vehicles in the near vicinity of the Orbiter 

will experience both disturbance and contamination on their surfaces. 

Disturbance is in the form of external torques and forces caused by 

pressure differences on vehicle surfaces while contamination results 

depend on the toxic qua1 ities of the impinging plume. 

The plume flowfield contours of the Orbiter in the Low Z and Normal Z 

modes and for the Orbital Maneuvering System (OMS) are shown in Figures 

4.3.8.2.3-1 and -2. These plume flowfield contours illustrate the 

expected amount of dynamic pressure to be experienced by a vehicle in the 

near vicinity of the Orbiter. Figures 4.3.8.2.3-3 and -4 illustrate the 

Orbiter plume flux contours or the expected amount of plume particles 

that will be incident on a vehicle's su' faces when it is near the 

Orbiter. 

In order to demonstrate the effects of plume disturbances on the Space 

Station, a series of Orbiter 1-secmd jet. firings have been made using 

the power tower configuration shown in Figure 4.3.8.2.3-5. Two PRCS jet 

firings modes were selected for this demonstration. The first mode was a 

Normal Z mode and the second was a Low Z jet mode where the + X-jets are 

fired to provide a +Z Orbiter body translation. These 1-sec5nd jet 

firings were made starting at berthing and at 50-foot intervals out to 

200 feet (Figure 4.3.8.2.3-6) along the Space Station plus x-axis which 

woul d represent a final berthing approach direction or initial separation 

direction. The result of these 1-second jet firings can be seen in table 

4.3.8.2.3-1. The resulting di sturbances indicate tbat the Orbiter 

thruster configuration set up in the Digital Auto Pilot has a very 

pronounced effect cn the resulting impingement forces and torques. More 

detailed analysis of the effects of the various vehicla;' plume 

impingement on the Space Station during PROX OPS must be considered. 

These 1-second jet firings give an initial example of some of the 

expected disturbance forces from these operations. However, some PROX 

OPS activities may result in many jet firings lasting longer than one 

second. The duration and relative range of these firings will be the 

result of future PROX OPS activity requirements. 

Figures 4.3.8.2.3-3 through -9 il lustrate a typical Orbi ter primary 

reaction control system (PRCS) thruster time history for a normal Z V-bar 

approach to the Power Tower Space Station configuration. The figures are 

plotted for the +X, +Y, and tZ thrusters respectively, with each set's 

total ontime beiiig pTotted aS a function of the Orbiter's range from the 

station. The approach begins at approximately 850 feet with a 3 minute 

period of stationkeeping. The approach is initiated vja two initial 
bums at about 810 feet; a 1.0 fps -Z and a .17 fps + X burn. Thf s 

coabinatfon places the Orbfter on a closing trajectory toward the 

Station. Uhen the range equals approximately 575 feet, the pilot 

initiates a series of braking maneuvers to control the approach velocity 

until all rates arc nulled at approximately 50 feet. For this case, the 

plots clearly illustrate the increased firings necessary fn the near 

proximity of t.lw station i n order to null the Orbi ter's closing rate. It 

should be understood that these results simulate a typical Orbiter V-bar 

approach and by no means rept-esent the total actual Space Statfon PRO% 

OPS envi ronmen t . 
Flgure 4.3.8.2.3-1 - ORBITER LOW Z AND NORMAL Z DYNAMIC 

PRESSURE PLUME CONTOURS. 

Figure 4.3.8.2.3-3 - ORBITER LQW Z PLUME FLUX CONTOURS. 



"' Z 1 SEC FIRING 

NOR Z (1 SEC FIRING) 

RANGE TOTAL TRANL, (JAL 

PLUME AV 

FORCE EI* (EPS) 

TABLE 4.3.8.2.3-1 ORBITE W5PACE STATION 

ESTIMATED PLUME IMPINGEMENT DISTURBANCES 

PLUME TORQUE 

IMPLUSE 

(ft-l bf-sec) 

X Y z 

14.20 -.391 E + 3 -101. 

10.40 .66E + 4 -38.5 

9.09 .74E + 4 -32.0 

8.16 .55E + 4 -18.5 

5.52 .36E + 4 -8.18 

I 

C (DOCKED) 

50 

100 

150 

I ** 

RESULTI

X 

1.90E-5 

1 .ll E-5 

9.6GE-6 

7.83E-6 

5.01 E-6 

5.60E-4 

2.31 E-4 

1.83E-4 

1.30E-4 

8.30E-5 

0 (GQCKED) 

50 

100 

1 50 

ZOO 

154.72 

63.83 

50.5 

35.40 

22.80 

741 .O1 

230.22 

175.59 

167.32 

156.14 

2.68E-03 

8.34E-04 

6.36E-04 

6.06E-041 

5.65E-04 

0.1 3E-1 -36300 -0.45E-2 

0.43E-3 -9000 0.97E-4 

0.1 5E-4 -4700 O.26E-3 

-0.1 0E-3 -2950 0.25E-3 

-0.14E-3 -551 0.2453 

1.00E-8 

0.0 

0.0 

0.0 

0.0 


TYPICAL ORBITER THRUSTER TIME HIST

VBAR APPROACH 

12 

11 

10 

9 

8 

7 

6 

5 

4 

3 

2 

1 

0 

1 

Initial closing 

2 velocity (1 fps) 

3 

810 806 805 802 688 600 455 277 200 132 91 

RELATIVE RANGE (feet) T/I ORB +Z THRUSTERS ORB -Z THRUSTERS 

Flgure 4.3.8.2.3-9 - Typlcsl Orbltert Z PRCS thruster time hlstory for 

a V-bar approach. 
4.3.8.2.4 Station reference design suitability for PROX OPS 

fl ight operations 

The Station, which maintain; an LVLH stabil ized attitude profile, has one 

docking port located on its positive V-bar, with an a1 ternate port 

located on the + R-bar. Two large booms project from the central truss 

structure in the out-of-plane direction and each contains a set of four 

solar arrays. The area of these arrays is approximately 2400 square feet 

each. Figure 4.3.8.2.4-1 shows the Station frum four different 

orfentations. Clockwise from the upper left, the Station is shown 

looking along its positive angular nmentum vector; alang the positive 

V-bar; up the positive R-bar; and finally in an arbitrary orientation 

essentially viewed from below (+ R-bar). The solar array pal.-Is, 

radiators, and various modules (habitation, laboratory, etc.) are clearly 

visible in the figure. 

This specific configuration has been reviewed and evaluated in terms of 

the following evaluation criteria: 

a. Procedur~ design/implementation 

b. Propellant usage 

c. Mission timeline impact 

d. Lighting 

e. Cl earances 

f. Plume impingement 

The results of this general review are discussed in the following 

sections: 

4.3.8.2.4.1 Procedure designlimplementation ease 

With the accessibility of a port on the positive V-bar, the Space Station 

configuration provides for procedural design of the simplest of approach 

scenarios. This is ideal from the standpoint of a manned vehicle 

appr~ach to the Station. The alternate port, located on the positive 

rddius vector, offers a reasonable approach for an unmanned vehicle. 

After an initial approach to the V-bar, the unmanned vehicle can initiate 

a flyaround (T :ded by the effects of orbital mechanics) allowing it to 

reach + R-bar. This flyaround and tk final approach to berthing (along 

the + R-bar) may all be executed with :he aid cf ideal 1 ighting 

conditions. If the flyaround is initiated near orbital sunrise the Space 

Station control lers should have "over-the-shoul der" 1 ighting for the 

entire scenario. This is not the case for a manned approach along the 

R-bar. Manned approaches are tdrgeted to the + V-bar. To flyaround to 

the + R-bar would be more expensive in terms of propellart cost (flying 

against effects of orb I tal mechanics). Second, the R-bar approach would 

have to be timed to occur near orbital sunrisa or orbital sunset in order 

to insure that the Sun would not be in the crews' eyes during the 

approach (not WI thin - + 20' of tbe LCS) . 
(a) Out-of-plane view. 

(c) View from below. Cd) View from below at6d out-of-plane. 

--. 

Fwn 4.3.8.2.4-1 - VARIQUS VIE #S OF SPACE STATION CONFIGURATION 

394 

(b) In-plane view. 

, 
An alternative, if the + V-bar port should not be available, would be to 

target the rendezvous directly for the + R-bar. This would require 

timing such that the approach to the port could be made with proper 

lighting (depending on whether the vehicle is manned or unmanned). If 

stationkeeping were required in order to achieve this 1 ighting, 

propellant usage would increase. 

In order to facilitate ideat operations it i - recomnended that the plus 

R-bar port be relocated or an additional port be added to the minus V-bar 

axis. This would be especially beneficial in the case of unmanned 

vehicles as it would aid In the reduction of propellant consumption by 

eliminating the need for a flyaround. Lighting would be essentially the 

same as those operations occurring near orbital su:,set. Also, better 

vehicle berthing clearances would be expected using ports on the + and - 

V-bar rather than the + R-bar. Adequate clearance could be a problem 

with the current conf i g~,*ation. 

4.3.8.2.4.2 Propellant usage 

Propellant usage for approaches to a V-bar port is less than for the 

other locations. This is expected since a11 approaches are along the 

V-bar up to about 200 feet. At this point, a manned + V-bar approach to 

the port would simply continue along the V-bar to berthing. However, a 

flyaround would be necessary to reach the + R-oar port unless the 

rendezvous were targeted to the + R-bar, as mentioned above. Also, an 

approach along the V-bar can be stopped at anytime and a minimum 

propellant stationkeeping position could be established. Establishing 

stationkeeping durin9 a flyaround (especially on the + R-bar) can be 

expensive in terms of propellant usage. The most practical method would 

be to stationkeep on the V-bar. 

4.3.8.2.4.3 Mission time impact 

Generally, approaches to the V-bar require the least amount of time. 

Approaches to R-bar ports require alignment with the port prior to 

closure. The main tradeoff between a flyaround from the V-bar to the 

R-bar and direct rendezvous to the R-bar would be the resulting 

propellant isage. However, the V-bar port locations allow for more 

mission flexibility since they nominally a1 low for a stationkeeping phase 

prior to starting the approach at a small cost, both in terms of 

complexity and propel l an4 usage. 

4.3.8.2.4.4 Lighting 

Activities during PROX OPS are very dependent upon 1 ighting conditions 

since it is mandatory that the crew observe all activities by visual 

means using the eye (preferably), television, or other acceptable 

electronic techniques. Therefore, the 1 ocation of the sun will impact 

the activity timeline for PROX OPS. The basic ground rule for 
day1 ight approaches is that the angle between the 1 ine-of-sight to the 

sun and the line-of-si#t to the whicle be greater than 20 degrees. 

Proximity operations, especially final approach and berthing, are also 

preferred to be done in daylight. 

The reference configuratfons's plus V-bar p~rt provides the best port 

location for manned vehicle approaches such as the Orbiter berthing with 

the Space Station. The! + R-bar port is a much more difficult port for a 

manned vehicle to accomplish final approach and berthing because the Sun 

my be near the line-of-sight to the Space Station during these final 

activities. 

The + V-bar port is undesirable for approaches of unmanned vehicles to 

the Space Station since the crew on the Space Station would also have a 

possible Sun lighting problem during dqylight operations. However, the + 

R-bar port provides an ideal port location for sun1 ight operations with 

only one basic problem: a flyaround would nominally Be required to reach 

this port. If the stationkeeping position prior to the flyaround was on 

the - V-bar, the iapact would be minimized. Therefore, from a lighting 

standpoint the refererxe configuration is excel 1 ent for manned approaches 

to the + V-bar port and undesirable for manned approaches to the R-bar 

port. For unmanned approaches, the + R-bar port provides a very good 

situation for final approaches and berthing whereas the + V-bar port is 

an undesirable location. Adding a - V-bar port would make this an 

excellent configuration for both manned and unmanned active vehicle 

activities. 

Lighting considerations for separations from the Space Station are rm:ch 

the sane as those discussed for approaches. Manned separations are 

preferred along the + V-bar and should be initiated early in the orbital 

dqy for liQlting. Similarly, unmanned vehicle operations are preferred 

from the V-bar early in the orbital day. The station has a + V-bar port; 

hence it is very satisfactory from the standpoint of manned vehicle 

separations. 

The Station also has a + R-bar port which allows both manned and unmanned 

separations from the preferred direction for each type vehicle. 

Therefore, it will experience minimum impact with regards to 1 ighting 

during separations. 

Approaches and separations during darkness are also possibilities. They 

would require, however, sufficient artificial lighting such as running 

lights on both vehicles or flood lights. The assessment of the 

configuration regarding this requirement cannot be made at this point. 
4.3.8.2.4.5 Transfer vehicle visibility 

Vehicles in the vicinity of the Space Station will require monitoring by 

the Space Station traffic controllers at all times. This could be 

accomplished by strategically locating cameras around the Space Station 

structure. The number and location of cameras on the station is TBD. 

4.3.8.2.4.6 Clearances 

The reference configuration appears to present no major clearance 

problems between the statiori elements and PROX OPS vehicles as large as 

the Orbiter. However, i f an out-of -pl ane f lyaround were ever required, 

the lower booms emanating from the keel extension and the radiator panels 

might present some clearance problems. Alsc, active vehicle-to-vehicleclearances may be required when trying to dock two vehicles at once to 

the Space St~tion, especially Go Orbiters. Adding a minus V-bar port or 

moving the R-bar port to the minus V-bar post tion would basically solve 

thi s problem. 

4.3.8.2.4.7 Plume impingement 

Vehicles located in the near vicinity of another vehicle actively using 

thrusters to perform maneuvers (translation, rotation, or 

stationkeekping) are subject to impingement of thruster exhaust. 

Impingement on vehicle surfaces can result in two undesirable 

consequences--di sturbance and contamination. Each of these is a function 

of the location of the impinged vehicle relative to the thruster 

center1 ine and the relative range. Contamination is a1 so dependent on 

the type of exhaust products emitted; for instance, combustion reactions 

produce a greater volume of contaminants than a compressed gas system. 

It should be realized that plume impingement cannot be totally eliminated 

during the PROX OPS period, but its effects can be reduced through 

vehicle and procedure design considerations. During separation, jets 

must be fired to move the vehicle directly away from the Space Station; 

this results in the jet plume being directed towards the Space Station. 

For approaches, the relative motion must be nulled at some time prior to 

berthing which again results in jets being fired toward the Space 

Station. Design consi'?rations could place Space Station elements in a 

position or attitudr -.elative to approac!; lanes to reduce impingement. 

The reference conf i g~r 3ti~ has several areas quite susceptf ble to plume 

impingement for both sepdrations and returns. For PROX OPS activities to 

and from the + V-bar pol-t, the radiator panels are essentially normal to 

direction of thrusting, although offset. Also, one cf the modules is 

directly in line with the thrust direction. 
The rotation feature of the solar arrays may a1 low for a reduction in 

plume impingement. Through a possible combination of timing the V-bar 

approach or separation to occur slightly before orbital noon and 

orienting the Orbiter in the most advantageous attitude, the arrays my 

have rotated to a position essentially "edge-on" with respect to the 

plum. Problems coul d arise, however, duri ng approach/separation 

operations if orientation is such that the plume flow is essentially 

normal to the arrays. 

4.3.8.3 Construction of Large Structures 

The Space Station will serve as a base for construction of large 

struct~ res. The Station was assessed relative to its capabil ity to 

support construction of a 1 arge suhi 11 iireter astronomical observatory 

depicted in Figure 4.3.8.3-1. This structure is essentially a 

100-by-100-foot cylinder, with a 30-foot extension at the Saoe, weighing 

20,000 lbs. to 60,000 lbs. The following areas relating to the 

construction of this observatory were considered as part of this 

preliminary evaluation: offloading and storage of component parts, work 

stations, construction methods, and impacts on the Station. 

The high structure to Station mass ratio from 1:10 to 1:5 must be 

considered with respect to attitude maintenance during offloading and 

subsequent related operatiow. The MRMS will be positioned on one of the 

keel extensions when accessing the Orbiter cdrgo bqy. Preliminary 

calculations indicate that an offloaded mass of approximately 50,000 13s. 

wwld shift the Station center of gravity three to six feet cut of the 

ot-bital plane, which should be considered when sizing the Station 

attitude control system. The observatory packaging, mass, and number of 

packages are unknown; however, a gross estimate of the volume of storage 

required is 20,000 ft3 or more. The storage area choice must consider 

the location of the construction area, MRMS translation and reach 

capabil ity, and the OMV/payloads servicing bays, a1 1 of which occupy or 

require space along the truss. Another important consideration is CG/CP 

offset, n,ecessi tating that the storage/construction area be within the 

orbftal plane, and located on the "back" of the lower keel as indicated 

by TDM2060 in the station configuration reference drawing. 

Several mthods of construction were examined before being discarded as 

impractical or unfeasible. The method appearing to have more advantages 

utilizes the support structure of the instrument unit itself during 

construction, and requires a turntable with the added ability to ti1 t 

towards the lower keel (see the station configuration reference 

drawing). The pivotal point must be approximately the radius of the 

lnstru~lent unit away from the lower keel. The bottom of the instrulnent 

unit would be mounted on the turntable and tilted so that the top of the 

unit is as close as possible to the MRMS for an adequate reach envelope. 

At- this point the MRMS can begin construction and the turntable can 

present the correct arc and tilt away from the lower keel as the 

cylindrical base grows. 

The scenslrio for the construction of the sunshade can take two forms. 

Construct two or three panels of the shade, mounting each as f t f s 

constructed; construct the secondary reflector and mount it; then 

construct and mount the remainder of the sunshade panels. The 

alternative method is to first construct the secondary reflector, wnt 

it, then one by one construct and mount each panel of the sunshade. In 

either case a construction area would be required for assedling the 

sunshade panel s. 

Construction itself of the panels could be a problem. They have not been 

designed and little is known other than their dilllensions of approximately 

40 by 100 ft. The Station affords no construction platform this size, 

but it cannot be stated categorically that only this size is required. 

Following are the main areas requiring further study in order to 

accommodate construction of this large observatory at the Space Station. 

The effect of long and short duration exposure of the covered facets to 

the Sun is unknown. The effect on the Station of a delay in deploymnt 

of the observatory should be assessed. Attitude control consunsables and 

CMG sizing mqy be affected if the observatory remains attached for a 

prolonged period of time. Plume impingement modeling should be done. 

Station RCS thrusters are in proximity to the construction site and the 

Orbiter will probably return at least twice during construction. 

Finally, pqyload view factors during the period of observatory 

construction may be adversely affected and wfll need a detailed 

assessment. 

4.3.8.4 Logistics/Rewpply Operations 

Logi stics for the srbital operation of the Space Station system will 

consist of the orderly planning and execution of the resupply of 

propel lants and consumables, del f very of sparelrepaf r parts, 

deliveryheturn of payloads, and any new or damaged ORU, return of waste 

and pl*ocessed material, and delivery of consumbles to support crew 

rotation. Logistics support i ncludes the packaging, hand1 i ng, storage, 

and transportation of Station consumbles and replacemnt equipment. 

Long-term activity plannf ng will provide the i ntegratfon of requfre#nts 

and schedules for the various logistics tasks. The STS will provide the 

means for del ivery to the Statf on or return to the ground on a 90-day 

resupply schedul e. 

The Space Station bas unique features relative to logistics and 

resupply. The STS has a short turnaround (1 week) when mated to the 

Station. The Logistics Module offers 1 fmi ted stowage capabi 1 i ty for 

equipment and gases/fluids on ar\y single flight, and lnqy have to be 

supplemented by a pallet for additional tankage and supplies. Yet the 

manned mission is very extenrive. Crew turnover wfll prinarily be on the 

same 90-day schedule as resupply, requiring vol ume/wei ght 
a1 locations to support unique crew requirements that might otherwise be 

used for Statlon or pqyloads, and, due to the small her of 

cmmembers, there will be no dedicated sn-board maintenance/logistics 

crder to perfom maintenance on an on-call basis. This mans that 

scheduled maintenance and logistics functions must be minimized. In 

order to minfmize the logistics tasks, consideration should be given to 

the level of system redundancy of Station elements, quantities of 

consumbles on-board, and mintalnabil I ty requirelents to reduce the 

frequency of required resupply or repair missions. 

Logistics/resupply operations at the Station consists, in a broad 

context, priaari ly of changing mdules, transferring crew, and properly 

mloying the crew maintenance capabi 1 i ties between fl ights. The 

scenario for changing nodules consists of reroving a fresh Logistics 

Module (LU) fron the Orbiter, berthing with the Station, connecting 

fluids and power/data, disconnecting the used U, and berthing the used 

LM in the Orbdter cargo bay for return to earth. Any other ORU's not 

part of the iJ4 or payloads would generally follow the same operational 

scenario. Following are general requirements derived from the above 

scenario: 

a. Rigid Orbiter attachment to the Station 

b. Adequate RMS reach and adequate staging area for all items either 

for removal from Orbiter or for replacement in the Orbiter 

c. Good visibilily for the R)IIS/#RMS c;?erator 

Station operation= during crew transfer must consider the fact that 

normal Station operations will be interwpted and there may be a maximum 

of twslve to sixteen people on the Station far short periods. 

Consulwbles such as food, clothing, spares, payloads, and pqyload 

materials nust be unloaded and taken to use areas. Trash and waste must 

be processed and stowed, and dl rty clothing, processed material , experiment samples, and ORU's wst be stwed. These considerations lead 

to reqL jremnts for good nodule traffic patterns and both internal and 

external storage ;pace. Section 4.3.8.1 assessed the MRMS capabil ity to 

reach payloads externally mounted, and the reach envelopes of the )IR)tS 

denonstrate adequate capabil ity to off-load the Orbiter and provide for 

external storage in the vicinity of the module or on the lower or upper 

keel . 
4.3.9 Safety 

4.3.9.1 General 

The reference c~nf iguration as presently conceived is intended to meet 

the safety design requirements defined in the proposal (RFP). These 

requirements are to be applicable to all Space Station systems, 

subsysten~s, and operations. These requirements apply under worst case 

natural and induced environments. The configuration desf gn is to tither 

el iminate by design or control a1 1 hazards to the maximum extent and in 

the most cost-effective manner possible. 

4.3.9.2 habitable Module Egress Capability 

The Habitation Modules and Lab Modules assembled in the manner described 

ill the reference "dnfiguration (i .e., the "race track") allows for two 

egress paths from any one of Chese modules. The Logistics Mi9du1e is 

generally considered to be exempt from the dcal egress path capabil ity 

imp1 ied by the safe haven requirements stated in the RFP. A1 though the 

Logistics Module has only one egress path, the risk to a crewmember 

occupying this volume during the occurrellce of an accident forcing 

evacuation of the volume can probably be reduced to an acceptable level 

by the proper location of equipment, adequate materials control, 

elimination of potential ignition sources, and maintenance of adequate 

traverse clearance. 

The previously mentioned capabilities are described as a part of the IOC 

phase of the reference configuration. Habitation and Lab Modules added 

to form additional "race tracks" must meet the egress requirements. 

Additional Lab (for international iind commercial users) and Logistics 

Modules, not included in the "race track," will have to be assessed on a 

case-by-case basis according to their hazard potential. 

4.3.9.3 Protection/Control of High-Pressure or Hazardous Fl uic Tanks 

It is required thdt potentially explosive containers be located outside 

of habitable areas, isolated and protected such that the failure of one 

will not propagate to others, and designed to "leak-before-rupture." The 

referewe conf iguratf on a1 1 ows for potential ly explosive containers 

(e.g., high-pressure tanks, tanks containing hypergol ics, tanks 

containing flammable and/or toxic flu'ds) to be located outside of and 

away from habitable areas. The reference configuration provides ample 

areas on the keel of the Station structure for high-pressure and 

explosive containers to be located away from the inhdbited regions of the 

Space Station. In addition, potentially explosive containers located in 

a group or cluster of similar containers are required to be protected 

such that the failure of one does not propagate to the others. 
4.3.9.4 Isolation of Modules After an Accident 

Individual modules within the module assembly of the reference 

configuration (i.e., the "race trackY) are capable of being isolated froa 

other modules in the case of an accident or potentially hazardaus event 

such as loss of pressure or fire. In addition, tilose modules which 

contain confined hazardous or toxic materials anG which are not a part of 

the "race trdck" will be capable of being isolated from the mi* of 

the Space Station habitable areas during emergencies and/or accidents 

(e.g., loss of preasure, fire, release of toxic materials and/or fluids, 

materials of fgassing) . 4.3.9.5 Reaction Time After the Occurrence of a Leak 

The hatches of each individual module will have the capabil iG of being 

closed rapfdly in the case of a module leak. In addition, the 

Env-iromental Control and Life Support System (ECLSS) is required to have 

the capabil ity to acccrrrnodate atmos~here leakage up to 2.3 kg/day (5 

Ib/dzty) for the total Space Station pressurized volume. 

4.3.9.6 Safe Haven Capabilities 

A1 though not required, the module arrangement ( i .e., the "race trach") of 

the reference configuration allows the Space Station to tolerate aw 

single credible failure, including the complete functional loss of a 

module. The "race track" modular arrangement allows any crerJAleRlber Cr,: 

access to habitable conditions within the Space Station without EVA. The 

reference configuration's safe haven concept provides for habitable 

condltions for the crew in the remaining modules including amsphere, 

food, water, waste management, health maintenance, personal hygiene, 

sleeping provisions, communications, conmand/control , and fire 

suppression. The reference conf igu.-at ion provides for redbndant cmand 

and control by allocating locations where a portable control "consoleu 

can be Installed as needed. All critical systems/subsystens necessary 

for implementing the safe naven concept (e.g., ECLSS) will provide 

eircrget,:y capabilities for a period of up to 22 days. 

4.3.9.7 EVA Operations 

The reference confiobration provides capabil i ties to facil i tate EVA 

operations. A singular concerr~ for EVA operations with regard to the 

reference conf iguriltion is the 1 arge distances frond the Station ai?l ock 

to potential EVA work sites. Due to the large dimensions of the Space 

Station struclure, a means of transporting and supporting the EVA crew 

fran the airlock to a potential work site may have to be developed. This 

my be accmpl ished by the projected travel ir!g RMS or extensive use of 

the Manned Maneuvering Uni t (FMU) . The antenna 1 ocations are apparently 

'such that radiation hazards to EVA cremembers are minimal. At present, 

the current locations of the Reaction Control Syste~l (RCS) packages do 

not seem to constitute a hazard to EVA crewmembers. 
4.3.9.3 Repair and Reactivation of nodules after an Accident 

A major issue to be resolved for the reference configuratton is the 

capabf 1 i ty to repair and reactivate a lodule or modules after an 

accident. Detailed conceptual mans of implementing this goal has not 

been Qveloped at this time. As a general design requireaent, all 

critical systems/subsysteas are to be fail-operational/fail-safe/ 

restorable. In addition, there is a design goal fcr the in;ier surfaces 

of the wdule walls to be accesslb!e for inspection and repair by the 

shirt-sleeved crew. 

Z.3.9.9 Orbiter Docking Ports 

The abiliQ to docks with and access the Orbiter fn tile various volumes 

of the Space Statfon is accoaunodated by the rvCerence configuration. 

4.3.9.10 Other Safety Considerations 

Several areas of safety concern with regard to the reference 

configuration will require additional study. These areas include but are 

not lim4ted to: (1) servicing of satellites and/or free-flyers either on 

or in proximity of the Space Station, (2) servicing of sol id and/or 

liquid upper stages (PAM'S, MV's, OTV's, etc.) either on or in proximi$y 

of the Space Station, (3) fl ight operations of upper stages in transit to 

and from the vicinity of the Space Station and, (4) assembly aperations 

associated with the change-out of a module either for return to Earth for 

repair or refurbishment or for relocatior: in the growth phase of the 

reference conf igurati on. 
4.4 SUBSYSTEM DESCRIPTIOU 

4.4.1 Electrical power 

4.4.1.1 Introduction 

The Space Station is a much larger orbiting vehicie thzn NASA has ever 

flown in the past. Design criteria different from those of previous 

programs must be met. These considerations are reflected in the power 

system design. in several wqys. The power system must support Station 

activities which are largely undefined and which will change as 

activities on the Station evolve. The power system must be flexible 

enough to meet changing demands and growth over the anticipated 30-year 

station life. Initial power demands of 75 kW are expected to grow to 300 

kW by the year 2000. The power system must be modular wi th we1 1 defined 

f nterfaces to accomRlodate repa: r- 5;'-ropl acement and growth without 

disrgption. It should also be able to accept new technology as it 

becomes available or as it is needed. 

The power system for the Space Station csnsists of three elements: energy 

conversion subsystem (ECS) , energy storap subsystem (ESS) , and power 

management and distribution (PMAD) . Ui thin each of these elements are 

several candidate technol ogies for consideration as options. Possible 

technol ogies are 1 i sted in Table 4.4.1.1-1. Both deployable and 

erectable structures are considered. 

A primary goal of the power system design is the development of a 

utility-type system as nearly independent as possible of outside support 

from the crew or another subsystem. 

For this study effort, two reference power system configurat' IO~S were 

considered for the initial Space Station. One was a silicon photovoltaic 

power system with regenerative fuel cell energy storage and hi gh-vol tage 

ac distribution. A second used a solar dynamic power system with thermal 

storage and high-voltage ac distribution. Each of these power system 

configurations has merit for the initial Station. Only the initial 

Station photovol taic power system was subject to analysis and integration 

with other subsystems. A sketch of the initial photovol taic power system 

is shown in Figure 4.4.1.1-1. A sketch of the initial solar dynamic 

power system is shown in Figure 4.4.1.1-2. These selections were 

representative of many applicable power system options. Each component 

of the power system confi gurati ons wi 11 be described in the paragraphs 

be1 ow. 
TABLE 4.4.1.1-1 - ELECTRICAL POWER SYSTEM OPTIONS 

EKERGY CONVERSION 

DYNAMIC 

Braytor~ 

Rank i ne 

Stir1 ing 

PHOTOVOLTAIC 

Silicon planar 

GaAs concentrator 

I 

POWER MhAGEMENT 

ENERGY STORAGE AND DISTRIBUTION 

Thermal 

Regenerative Fuel Cell s 

NiCd Batteries 

Flywheels 

NiH2 Batteries 

High-vol tage ac 

High-vol taw dc 

High-Vol tags Hybrid 

SIDE VIEW FRONT VIEM 

Flgure 4.4.1.1-2 - IOC REFERENCE CONFIOURATION (SOLAR DYNAMIC OPTION) 
4.4.1 .P Photovol taic System Generation and Storage 

Assumptions used in this study arc out1 ined in Table 4.4.1.2-1. 

4.4.1,2.1 Photovoltaic generation 

The photovoltaic system chosen for this study is the flexible, planar 

array utilizing the large area (5.9 cm x 5.9 cm) silicon cells. This 

reduces the total number of cells and the manufactu~ing and handling 

steps required in array fabrication. The cells would be attached to a 

flexible kapton (or simf lar material ) substrate and welded to the 

attached circuitry in order tc! provide better rss!stance to the them1 

cycling that will be experienced in LEO. Sinca the blanket will have 

little or no structural stiffness, a means of structural support must be 

provided. For this study, a nit weight for the solar array blanket and 

support structure of 2.5 kg/m I was assumed. Whether thi s 1 ightwei ght 

structure is adequate is subject to further study. 

A design 1 ife goal of 10 years wsc, chosen for the solar array.' Since the 

Station will last longer than the array 1 i fe goal, provision must be made 

to change out solar arrdy components. Roughly a 10 percent degradation 

in cell performance i s expected to occur over the 10-year 1 i fe of the 

array. The rate of degradation is also subject to further study. 

The solar array wing used in this study would allow both deployment and 

retraction so that the entire wing could be replaced, Other concepts 

might allow replacement of smaller units. Concepts requiring EVA to 

erect structure and panels were also considered. As the array wings get 

1 arger, erectable concepts become more attractive. Growth capabil i ty is 

a function of the specific array design. In this study, growth would 

require adding 2 or 4 wings. Several packaging concepts to transport the 

array to orbit were a1 so addressed in the study. 

Two-axis gimbals with rotary power transfer devices have been provided to 

orient the array normal to the sun. Because of the geometry of the 

reference Space Station configuration, some of the station (structure or 

experiments) will shadow portions of the solar array during part of the 

orbit. This must be recognized as part of the design of the arrqy. 

The solar array si re is a function of the power level, array orientation, 

storage sys'em efficiency, distribution system efficiency, expected 

degradation, and the orbital parameters. The initial arrqy area is 

approximately 19,2i~0 ft2. 

4.4.1.2.2 Energy storage subsystem (ESS) 

The ESS defined in this study consisted of four regenerative fuel cell 

modules with the required supporting hardware. The storage units will be 

mounted on the array support structure outboard of the alpha gimbal . Each 20 kW module would weigh about 1202 1 bs. and occupy about 75 ft3. 
4.4.1.3 Dynamic Sy5tem Generation and Storage 

The djnamic system chosen for this study is a generic representation of 

the three most applicable technologies: Brayton, Rankine, or Stir1 ing, 

each bring mlten salt thermal storage. Basic operation of the solar 

4ynamic system involves the use of a mirror to collect and cancentrate 

the Sun's energy on a heat receiver. A heat engfne is then used to 

extract energy from the receiver and convert it to mechanical energ to 

drive an electrical alternator. Excess heat is collected and stored in 

the thermal storage material within the receiver and is used to operate 

the engine during eclipse periods. Waste heat from the engine is 

rejected to space through a dedicated radiator. Because the storage 

element is between the Sun and the engine, the engine operates 

continuously around the orbit. The major assumptions used to size the 

4ynamic system for this study are shown in Table 4.4.1.3-1. 

Specific choice of heat engine cycle will depend on the result of trades 

concerning receiver temperature, performance, and development risk. The 

nuhers presented here are indicative of a Rankine cycle using Toluene as 

a working fluid or a low temperature Brayton cycle. Design results are 

given in Table 4.4.1.3-2 for both the initial and growth stations. 

With proper design, the same rotating machinery can be used on the 

initial and growth station. The radiator required as part of this power 

system fits behind the primary mirror. This radiator is an integral part 

of the power system and is not part of the station thermal system. 

The collector can be a single parabolic mirror or a two-element 

Cassegrainian. Either may be deployable or erectable a1 though erectable 

devices seem to package in smaller volumes. The mirror has high 

ref1 ecti vi ty and high geometric accuracy to provide the concentration 

that permits a small receiver aperture to be used, resulting in low 

reradiation 1 osses. 

The heat receiver absorbs, stores, and transfers the solar eirewgy t-n tho 

engine working fluid. In most concepts, the storage is in the kat of 

fusion of a salt with a fixed melting point. Therefore the engine pe.3k 

operating temperature will be chosen in conjunction with the choice of 

salt and containment material. Mirror accuracy will also be a factor. in 

the choice of operating temperature since high temperatures place more 

stringent requirements on mirror fabrication. 

The unit size was arbitrarily selected and should not be considered 

optimm. However, the qse of four units allows a relatively simple 

initial configuration which is amenable to growth. 

Two axis gimbals, identical to those used in the solar array, have been 

.provided with rotary power transfer devices to maintain alignnrent of each 

wing to the Sun. In addition, a small amount of fine pointing capability 

has been given to each unit to compensate for station pointing and gimbal 

errors. 
TABLE 4.4.1.3-1 - Dynamic Power System Design Assumptions 

Collector efficiency ................... 90% 

Receiver efficiency ................... 90% 

Engine efficiency ................... 20 . 30% 

Sunperiod ........................ 59min. 

Ecl i pse period ........,............. 35 mi n . 

TABLE 4.4.1.3-2 - Dynamic Power System Design Results 

Initial Growth 

Station power demand ....... 75 kW 300 kW 

Unit engine size ......... 20 kW 37.5 kW 

Number of installe3 units .... 4 8 

Mirror/radiator diameter ..... 13 m 18 m 

K i r r o r type ........... simple parabol ic 

Unit weight ........... 2100 kg 3410 kg 
4.4.1.4 Power Management and Distribution (PMAD) 

The PMAD subsystem transfers electrical power from the source to the user 

interface. It conditions and control s the power and provides individual 

protection for each user interface, power source, and PMAD hardware and 

wiring. The subsyste~n manipulates sources, loads, and buses as required 

for most effective utilization of power and provides the whanization 

for recharging the energy storage devices. If the PHII.9 system is 

properly designed, it can function with several di ffeyent power sources. 

DC power (or ac from a solar dynamic power system) from the generation 

and storage device; is converted to high frequency three-phase power for 

transmission via mu1 ti ple redundant buses throughout the station. Each 

bus consists of a siTbstation with overrides for selected switches, an 

automated power manaqemnt system (APMS) interface, a primary ac bus, 

secondary ac and dc ouses, and user utility outlets. Disconnects and 

crossties controlled in the substation are in the main bus structure to 

provide the flexibil ity to util ize to the maxim that portion of the 

priwary circuitry remainjag intact after a fault or during maintenance. 

Finally, the interface between the PMAD and the user is the utility power 

controller (IJPC), which is controlled by the AFMS via a link to the data 

bus. Although this design utilizes high-frequency ac distribution, the 

choice between ac or dc for primary power transmission will be determined 

by we4 ght, volume, cost, and growth potential considerations. The 

interface with the user will remain sine wave ac. The distribution of 

PMAD equipment around the station is shown in Table 4.4.1.4-1. 

TABLE 4.4.1.4-1 - DISTRIBUTION OF PHAD EQUIPMENT 

Central structure ................. 1983 1 bs 

Logistics module ................. 2121bs 

Habitat module .................. 471 lbs 

Laboratory module ................. 514 Ibs 

Control module .................. 652 lbs 
4.4.2 Guidance, Navigation and Control -- 

This system has the responsibility for managing the sensing and 

atqui si tion of information, computation, and actuation required to 

provide position and att.itude control for the Space Station and to point 

its ~olar arrays, radiators, and pay1 oad mounting surfaces. The GNLC 

system will interface with the Information and Data Management System 

(IDM), Cm~unication and Tracking System (C2T) and Propulsion System 

(RCS) to perform these functions. The GN&C system will also manage the 

traffic control function and proximity operations. GN&C support will be 

provided to the payloads attached to the Station and to the Station 

traffic. Satellites and vehicles berthed to the Station will have their 

control systems inhibited, and the Station wi 11 be responsible f~r 

attitude control . The GN&C system I in conjunction with the propul sion 

system) wi 11 have responsi bi 1 i ty for orbit a1 ti tude mainterlance and 

adjustment, collision avoidance and deboost in the event that disposal of 

the Zpace Station is required. The GN&C system will use the mass 

properties, propellants, an8 structural bending data provided by IDM to 

provide attitude control and orbit maintenance of the Space Station. The 

basel ine system architecture is shown in figure 4.4.2-1. 

4.4.2.1 Functional Description 

The GN&C is a major Space Station system which is critical for 

dtti tude control, orbital maintenance, solar pointing, traffic contro: , proximity operations, and experiment stabil i ty. These functions are 

vital to the success of the mission, efficiency of operation, and safety 

of the Station dnd crew. For this reason, the higt-,?st level of 

reliability must be considered from the outset, using error detection 

techniques, redundancy, standby equipment, and backup or a1 ternate means 

for achieving the most critical functicns. In contrast to rcl iability, 

the perf~rmance requirements such as sensor acccuracy, memory capacity , computation speed afid effector characteristics are not as critical to 

this progrdm, and cer tdi nly not beyond existing technology. Therefore, 

new devel opment shoul d be consi derea primarily to update the specified 

equipment in terms of efficiency, re1 iabil i ty and 1 ifetime. The 

equipment and software interfaces must permit the substitution or 

addition of elements for this purpose throughout the lifetime of the 

Station. 

4.4.2.1.1 udsel i ne GN&C approach 

The base1 ine control system (Figure 4.4 2 i j is 2 cent. ized systev of 

colocated attitude sensors and C;{G control effectors which provide 

inherent stirbil i ty of +,kt. flexible body. The ndvigation system util izes 

the Inertial S~llsor Assemblies (ISA) composed of rate gyros and 

acceleron~eters to provide attitude, attitude rate, position, and 

velocity. Star trackers will be used to u~date attitudes, and GPS 

satell itss will be used to update position and velocity. The navigation 

processing will be 3erformed in the Navigation and Traffic Control 

Computer. 
HEXAD 2 ($1 

STRAPDOWN I 

STAR (TRIAD) 

TRACKER 

I TORQUERS MAG H MAG TORC 

ELECTROF 

NUMBER OF UNIYS 

' PAYLOAD I & , , I POINTING 

FUNCTIONING * SYSTEMS 

UNITS L-,,,,,,,,~ 

2 (0) 2 (1) w 

* - 3 (0) 8 2 (1) I ? 

0 I U 

RADIATOR * POSITIONING RCS * S 

ACTUATORS 

ELECTRONICS ELECTRONICS 

I 

(COLD SFARES) 

10 (0) 6 (3) 

F,cKlN, F a c COMMUNICATION , ID c &TRACKING - 

LASER 

-0-4 

DOCKING - 

SYSTEM 

- rACTUATORS 

1 DATA t-4 INFORUATION I BASE MANAGEMENT 

" 

SOLAR ARRAY 

ELE:TSONICS 

Figure 4.4.2-1 - GN&C SYSTEM ARCKITECTURE 
The primary attitude control effectors will be CMG's and may be 

suppl emented by magnetic torquers to perform continuous mmntum 

G~anagement. The RCS will be used to perform all translatfon maneuvers of 

the Station. RCS will be used to backup the CK's when the momentum or 

torque exceeds the capabil i ty of the CMG' s. A1 so, the RCS wi 11 be used 

for primary attitude control with reduced pointing in the event th~t the 

CMG's become inoperative, and the RC3 will be used to perform momentum 

management in the event that the magnetic torquers fail. All of the 

Station's control systems (CMG, RCS, magnetic torquer, etc. ) wi 11 have 

the capabi 1 i ty to operate simultaneously or 1 ndivi dual ly . An attitude 

tranirer system will be used to determine the relative alignment between 

the navigation base and the mounting surface of the pointing instruments. 

4.4.2.11.2 GN&C equipment location - 

The GN&@ equipment locations are as follows: 

a. The CMG's and ISA's will be i~talled in the Attitude Control 

Assembly (ACA) on the ground prior to Space Station launch. This ACA 

will, in turn, be installed at the intersection of the transverse bcom 

and the main keel to optimize vehicle mass distribution for control 

dynamics. The accessibil ity of this equipment for maintenance will have 

to be addressed in more detail. However, it should be noted that a1 1 

CMG's are centrally located on one reference base and that the gyros 

(ISA's) are colocated with the CMG's. 

b. The star trackers have tentatively been located on +he ACA. 

An alternate location would be on the upper boom. A study will have to 

be made to deternine the optimum location. 

c. The magnetic torquer locations are TBD at this tire, but will 

1 ikely be on the main keel and/or one of the booms. 

d. The solar array and thermal radiator drive electronic units 

will most likely be placed at the locations of the drives themselves. If 

this proves impractical, they will be placed i n the ACA. 

e. The G&C processors and the interface devices (ID'S) ass~iated 

with flight corrtrol and its sensors and actuators will be located in the 

ACA. These processors and ID'S are required to provide attitude control 

during the unmanned buildup phase of IW.. 

f. The navigation/traffic control processors and the other 

miscellaneous GN&C electronics will be located in Habitat Module 2, The 

location of each of the ID'S that support the remote sensors will be at 

the sensor i tsel f. 

g. The philosophy on spares is to have flight critical spares 

installed in place. It may a1 so be desirable to have the spares for a1 1 

the remote GN&C sensors installed and in place to alleviate complex 

instal lation/al ignment tasks. 
4.4.2.1.3 A1 ternate control system approach 

An innovative approach is the pointing and control system [PAC) which 

would colocate the control sensors and effectors with the pointing 

payloads and dynamically couple this system to the remaining Space 

Station. This would pernit precisely oriented pointing mounts that would 

be loosely connected to the Station, thereby i solatif,, some of the 

Station's motions from the pointing instruments. The S5ation's control 

torques would be effected through a spring-coupled system between the 

Station and the PAC. 

The Station cauld typically have mu1 tipic PAC elements, each of uhdch Is 

internal ly redundant and completely indepe-dent of the other in terns of 

sensors, conlputers, and effectors. Each element would be capable of 

operating in a slave mode, in which the effectors are camanded by the 

master element to support the momentuln management task of the entire 

Station. The elements would be connected by an optical link or star 

trackers that would establish the relative a1 igmnt of one to the 

other. The master element would be the one which has the best viewing 

direction to the stars, unless there is catastrophic fall ure of one 

element, in which case the other assumes command. The master element 

always establishes the reference coordinate system. Because of the rigid 

structure and the close proximity of sensors dnd actuators, the most 

stable part of the Station would also be the best pointing part of +.he 

Station. Therefore. the experiment pointing mounts would be attached 

directly to PAC elements. 

A single PAC elemetlt would be configured to fit easily wi thiil the Shuttle 

payload bay for replacc~nt and return of intact elements in case of 

major ma1 functions. For this reason, the attachment of the elements to 

the Station would be sim~l ified. During operation there would be an 

option for rigid connection to the Station structure or sprl ng-coup1 ed 

connection. One advantage of the PAC is that transient disturbances on 

the Station such JS man motion and rotating machinery are filtered and 

attenuated before reaching the stable element. Another advantage Is that 

the bandwidth of the controller can be increased substantially without 

int~racting with the Station structure. This makes the PAC element a 

very stable structure, capable of holding the Station in an orientation 

to minimize momentum a~c~~cnulation and simul tanewsly providing a htgh 

degree of isolation, st ,, ility ad pointing knowledge to the 

experimenters. This trans1 ates into much sihipl er and 1 ower-cost mounts 

for pointing the v'wious instruments. The mounts can be built with low 

technology gimbals, since their main function will be instrument 

orientation rather than a high level of stahil ity. The other two major 

problems of conventior~al mounts, namely, the constraint on experiment 

center of mass relative to the gimbal point and the transfer of 

electrical and fluid lines across the gimbals, are greatly reduced. 

4.4.2.5.4 Traffic control and proximity operations 

The GM&C system wi 11 have mandgement responsi bi 1 i ty for traffic control 

and proximity operations of station traffic which is operating within the 
Station's area of influence, in order to assure thht the operational 

orocedures meet safety and mission constraints. The area of fnfluorlce 1s 

defined to be a 300 cone about the plus and minus velo~i ty vectors of 

the Station out to 2000 km and 130 percent of a sphere with a radius of 8 

kin around the Station. The GN&C system will also have the capabil lty to 

monitor and control vehicles which lie outside this area, but whose 

trajectories wi 11 eventual ly cause them to enter this area of influence. 

The responsibil i ties for traffic co.i!:rol and proximity operations will be 

shared among .he GNLC, IDM, CaT, and ground zystems. Assurance of the 

safety of the Station will be the responsibility of the GN&C. The ground 

systems have responsibil l ty for monitoring space systems while they fire 

beyond the area of influence of the Statio~t; however, the GNK system 

assumes responsibil i ty for the final approach and berthing of Statfon 

traffic. 

Traffic control incl ec'ss tracking. monitoring, a114 control 1 ing of the 

space systems which are performing operations withic the Station's area 

of influence. Proximity operations inclsde stationkeeping (formation 

flying), berthing, and departure froa the station. Using CAT arid IDM 

system, the GN&C svstem will determine and maintain the :lative state 

vectors I posi tion and velc;: ty) of stctian traffic. Th~s information 

wi 11 be used to establish the velocity maneuvers required to maintain 

safe positions for stationkeepinj and for berthing. Tracking of statlon 

traffic a::; detectable debris will be yovided, and trajectory 

extrapol ations will be performed to predict tile poss~ bil i ty of 

collision. The GN&C system will haw command, coiltrol , monitoring and 

abort capabilitjes to achieve safe berthing of active vehicles with the 

Statioa. An override capahil icy to wave off any rendezvolls which is 

deemed to present a safety hazard to the Station will be provided. 

Remote pilcting of unmanned tra~sportaticn systems which perform 

proximity operations with other systems in the area of influence will be 

suppr~rted. This wi 11 requi re capabi 1 i ty for cor, ,inuous cmar.d, control , monitoring, comnunication, and tracking of the transportation system and 

monitoring and tracking of the target vehicle. 

The rriative state vectors of tethered sate1 li tes will be prov'ded and 

csntrol of the tethel system (e.9. reel control) to meet both safetj and 

operational requirements will olsc be provided. The capability to 

retrieve, stow or jettison the tethered satellite wf 11 also be provided 

if tile safety of the Statior is in jeopardy. 

4.4.2.2 Equipment 

4.4.2.2.1 Inertial sensor - assembly ( ISA) 

4.4.2.2.1.1 Traditional approach. 

The ISA will consist of gyros and accelerometers ~t:.-~ited in a redundant 

strapdown configuration, The ISA will measure attitude, rate. and 

accclerations. The ISA outputs will include atti ti~dtt, rate, velocity, 

and acceleration iq v3r1,i.i~ reference systems that are TBD at this time. 
The ISA will also provide subsystem status, i.e., malfunction, 

configuration (mode), cormandlcontro? , at~d data routing. The ISA's will 

be located witn the CMG's in order to avoid structural instdbil ities. 

Accessibil f ?y, modularity, and raintainabil ity considerations wi 11 have 

to be recognized as significant design drivers. The hexad (6 skewed 

sensors) configuration was selected because it a1 lows complete fault 

isolation and recovery for the first two failures and partial for a third 

far 1 ure. A1 so the hexad is tantamount to having spares in place so that 

failed components can be replaced later when Orbiter brings up 

replacements. 

Laser and fiber optic gyros will be evaluated as csndidate components. 

Laser gyros which employ dither will be evaluated within NASA to insure 

that the dither vibraticns and/or its related noise do not adversely 

effect star trackers, fl ight control or low acceleration experiments. 

Particular attention will be given to fine pointing, low acceleration and 

high resolutionllow noise rate data. Accelerometers must be able to 

measure reboost accelerations and mni tor 1 ow 1 eve1 accelerations in 

support of experimentation. 

4.4.2.2.1.2 Optical approach 

An optical ISA approach is the same as the traditional ISA approach 

except that the rate gyros are eliminated. Currently under development 

by JSC/ASa is state-of-the-art star tracker technology for the direct and 

continuous determi nation of inerti a1 attitude by continuous tracking of 

two or more stars. Spacecraft located in orbit can use star trackers to 

measure attitude and angular rctes directly. Large angle excursions 

require that standby gyro assemblies be available to provide attitude 

infornation during these events. Each tracker assembly will have three 

tracker heads with fields of view arranged in a triangle having 10 to 30 

degrees between the lines of sight of the star trackers. Any 90 star 

trackers are sufficient to determine attitude dnd attitude rate. If two 

star trackers become inoperative, stars in only one tracker will define 

the attitude of two body axes accurately but the third body axis will be 

very coarssly defined. Built in microprocessors will manage the 

acquisition, identification, tracking and data conversion functions of 

the trackers. The design goal is 5 arcsec (1 sigma) per ax's. Lab tests 

show 10 arcsec accuracy is easily achievable. The design goal is to 

support angular rate prxision of .001 deglsec at 10 HI update rate. 

4.4.2.2.2 Star tracker 

Star trackers will be used to update the GNBC jttitude reference for the 

traditional ISA system or for direct attitude reference and rate for the 

optical ISA system. The FOV is bounded at the upper end by view 

blockage, attainable accuracy, imaging 1 ens collecting aperture and 

scattered 1 ight shielding and at the lower end by the number of stars 

with known coordinates, angular rates and practical upper size of image 

lens. A 2-degree FOV is recommended for the star track r as the best 

canprtnllise, since it yields maximum flexibility for location relati ve to 

the stsuc ture, compact 1 i ght shades, a reasonably small imaging aperture, 
sensitivity sufficient to guarantee at least one star per FOV using 

existing star ca'alog, and capability to measure angular rates of up to 

0.1 deg!sec. 

The largest currently available precision star catalog contains 

approximately 256,000 stars. About 100,000 of the cataloged stars are 

redundant or not usable, because they are too faint or too close to other 

brighter staras. The distribution of stars in the catalog is non-uniform 

and can vary over a 10:l range, thus requiring that the star catalog for 

the star trackers contain at least 156,000 stars in order to guarantee at 

least one star per FOV. A catalog of this size and precision requires 

about 10 megabytes of ROM storage which is achievable with bubble memory 

chips or a miniature laser disk drive in the near future. Only one 

catalog storage device and one backup device would be required for a11 of 

the star trackers, since each would bse only a small subset of the 

catalog at any one time. 

Light shading is primarily driven by FOV and imaging optics aperture 

size. If a relatively modest FOV and lens size is selected, a shade 

length of one foot can provide 10 degrees bright source shading for 

Earth, Moon, and spacxraft and 20 degrees for the Sun. The sensor head 

that is proposed is lmune to damage by any sources sxcept the Sun. A 

combination of blocking filters and automatic gain control could be used 

for protection against the sunlight. 

A single triad assembly of three FOV would provide fail safe capability. 

Addfng a s~3nd triad assembly provides fail-op, fail-op, fail-op, 

fail-safe capability provided that the two triads are mechanically tied 

together. Some addi ti onal redundancy capabi 1 i ty i s 1 ost with tracker 

head failures due to reducing tlie number of fields available which 

increases the probability cf occasional blockage by the Moon, Sun or 

co-nrbi tin3 objects. 

Installation options are driv2n primarily by maintenance and 

accessibility. The only sensor-related driver is that the light shade be 

in the vacuum enwironmert ahead of the tracker's inlasing lens or window. 

Mounting configurations can range from the entire assembly being in a 

vacuum to the assembly being inside the modules, viewing through a port. 

The first option requires EVA for maintenance but offers the maximum 

flexibility for sensor pointing. The last option requires no EVA but has 

minimum pointing capability. Also, the sefisor head can be separated from 

the processing electronics with head located outside and the electronics 

inside or the lens and light shade can be located outside with the head 

and electronics inside with a viewing port. 

------- Computers and data interfaces 

The overai' design will bz compatible with the Informatior and Data 

Manageme; (IDM) system. The IDM design philosophy is distributed 

processi:~~. GY&C will utilize the IDM data network to communicate both 

intra- a? well as intersystem data (see Figure 4.4.2-1 1. The GN&C system 

will utSlize IDM standard/common components as dtctated by the results of 
trade studies to be conducted in support of the IOU architecture and 

topology. The corputational requireaents will be partitioned into Go 

processors: (1) navigation and traffic, and (2) guidance and control. 

The sfdance and control and the navigation and traffic control areas can 

be considered as individual subsystems within the overall GW area of 

responsi bi 1 i ty. Each of these subsystems requires its own standard/cmn 

processors and interface devices (ID). The associated redundancy for 

each of these devices will be determined as a result of the technology 

level and design capabil ity chosen as a result of trade studies. The 

system will be designed to meet a mininun of fail operationallfail 

safe/res+,orable. The system will be designed to accept technology 

developsrent so that improvements can be made to the GN&C system wfth 

little or n? impact to the remaining system. The requirement of three 

computers for fault isolation will set the mininuln. Because fast 

reaction is not required on the Station, consideration will be given to 

operating only two computers normally in order to reduce the power 

requirement. In the ewilt of a corputer failure, two computers can 

recognize that a failure has occurred. lhe control system can hold last 

output until the third computer is brought on line and the failed 

computer is identified. The two banks of computers (NAV 4 traffic 

computers and guidance & control computers) will be capable of taking 

over the othor's functions in order to fulfill the redundancy 

requirements. 

4.4.2.2.4 Control moment gyros (0%) 

The CMG's are devices that store momentum by changing spin vector 

orientation of a wheel which is spinning at a constant rate. The CM6's 

have a double gimbal which permits momentum storage about any axis. The 

double gimbal approach provides for redundancy and flexibility of 

operation. A combination of double gimbal CMG's and magnetic 

desaturation devices is recommended for the GN&C baseline system. The 

NASA TM-82390, Steering Law for Parallel Mounted Double Ginbaled W'S - 

Revision A, by Kennel (Reference 34 j, is recotmended as a guide for the 

implementation of CMG steering 1 aws . 

Momentum storage requirements are defined here as the peak cyc;ic 

momentum plus secular accumulation over cne orbital period. This is 

equivalent to taking the maximum PSS value of the momentum components 

about each axis. The full range of the momentum storage device is 

considered in the sizing, assuming a prior knowledge of the mottenturn 

profile or the ability to initid1,te tc a state which takes maximum 

advantage of momentum capacity. Double gimbal CMG's will be oversized by 

50 percent to account for uncertainty of the analysis. 

Momentum management is necessary to preven? saturation of the CMG's. A 

. storage device will eventually become saturated by secular (bias) torques 

which act in one direction. Momentum desaturation can be achieved by 

applying torques that reset the storage devices to their f nitial state. 

These toques can be generated in various wqys such as realignmnt of - - 
station principal axes in the gravity Field (gravrty gradient 

desaturation), utilization of aerodynamic moments, generation of magnetic 

fields which react against earth's field, or thrusters which produce 

reaction control torques. Ori tnti ng the station in the gravity gradient 

field is the preferred momentum management technique; however, magnetic 

torquers may be required in addition to supplement gravity gradient 

momentum management. Large area magnetic coil s wi 11 a1 so be studied. 

4.4.2.2.5 Magnetic torquers 

If a device is required to desaturate the CMG's, a magnetic torquing 

device is a low power, re1 iable, and nonconsumable means of providing 

torques or momentum into the station. Magnetic torquing coils have flown 

on numerous spacecraft over the years with excellent results. These 

systems, along with systems currently under design for use in the late 

1980ts, proviae torque atitputs as large as .088 ft-lbs. Depending on 

design, it is reasonable that magnetic torquing requirements may be 

greater and several bars may be required. Previous coil applications 

have been di rected heavi 1 y toward 1 i near control ; however, increased 

magnetic dipole requirements may make non-1 i near magnetic torquing s 

viable option. As a result of the large magnetic dipole being created, 

significant magnetic fie1 d disturbances will be produced in the vicinity 

of the magnetic torquers; however, the field decreases kith the square of 

the distance, and the earth field predominates at moderate distances from 

the bars. 

An a1 ternati ye technique for producing control torques using magnetic 

torquers is the large area coil. Current flowing in a large loop with 

multiple turns creates an magnetic field which reacts with the Earth's 

field to produce desired torques. The coil loop may be attached to the 

perimeter of a large structure or deployed using booms. Although these 

produce less concentrated magnetic fields than torquer bars, 

consideration must a1 so be given to not locating magnetically sensitive 

devices near the large coil loop; however, moved six feet from the coils, 

the Earth's magnetic field predominates. 

While magnetic torquing is considered primarily for desaturation of 

CMG' s, other appl ications are avail able. Magnetic torquers can be used 

to provide a supplement to the primary control in the event CMG control 

authority is reduced or unavailable. Magnetic torquers provide a simple, 

low power, nonconsumabfe torquing a1 ternative which will increase overall 

space station redundancy and re1 iabil i ty. 

Reaction Control Sys tem (RCS) 

Reaction control will be achieved by means of appropriately placed 

thrusters which are capable of producing torques about three orthogonal 

body axes and translational foiaces. The thrusters will maintain attitude 

and translational control by time modulated commands as generated by the 

control algorithm. P:acement of the thrusters should be such that 

impingement and contamination effects will be minimized. In order to 
keep the control laws simple and to avoid transporting pro,pellent across 

rotating joints, the thrusters should be placed on a sectiot~ of the 

structure which allows 3-axis control in translation and roti~tion. 

The greatest reboost efficiency from the fuel consumption point of view 

can be achieved by firing only those thrusters which accelerate the 

Station along the velocity vector. The required thruster geometry would 

be for the center of mass to be inside the thrusttr envelope in or* ta 

produce the required toques. Attitude and translation cmtrOV (ln 

decoupled by thruster mdulatfon. The thruster envelope includes the 

effects of unequal thrust and lever ams, thrust misaligrrrent and body 

flexing. 

A force vr-.ctor arrangement will be used to demonstrate a possible 

propulsion system geometry. Control redundancy has not been considered; 

however, the final design should include contingencies which inswe that 

control will be maintained in the event of thruster failure. Thruster 

redundancy will be addressed in Section 4.4.5 of this rcport. The 

thruster arrangement for the Station is shown in Figure 4.4.2-2. The 

thrusters are represented by heavy arrows which indicate the force 

direction. Time modulation of those thrusters shorn in the side view 

will provide for translation. Toque abwt all body axes is provfdec? 

while the translational forces are parallel to the x-axis only. A total 

of 12 thrusters will be needed to satisfy the torque and force 

requirements. Since the flight path is always along the x body axis, no 

reorientation wi 11 be required for reboost. 

The RCS will be used to transfer the Station to a circular orbit 20 mi 

higher qr lower than its present orbit in addition to a 90-dqy reboost as 

required due to atmospheric drag. The Station has a ~~~inimm operating 

a1 ti tude of 250 mi and maximum operating a1 ti tude of 300 mi with 270 

mi as nmical. The reboost and orbit altitude transfers will be 

performed imneaiately after Orbiter departs in order to maintain .00001 

g's or less until the Orbiter returns. Because the 20 mi transfer will 

not be utilized after every Orbiter visit, this propellent will be used 

as reserve pr~pel 1 ent for other contingencies that wi 11 be discussed 

later. Reboost propellent will be based on an altitude transfer at 

Orbiter departure which will be sufficient for 90 days. The Orbiter is 

expected to rendezvous with the Station at the same altitude. The 

propelient is sized on the greater reboost at 250 mi with nominal 

atmospheres or 270 mi with two s?gma atmospheres. Analysis has shown 

that 270 mi with two sigma atmospheres is greater. 

The RCS will be used to avoid collision with space debris, space station 

traffic, or natural space objects. Coll ision avoidance will not size the 

system; however, a propel'lent budget sufficient to impart a 5 ftls 

velocity change -in the station has been provided. If additional 

propellent is required. it may be obtained from the 20 nmi orbit transfer 

propellent, provided it has not been used. No propellent will be 

provided for station disposal (deboost), because it is assumed that the 

enti re propel l ent 1 oad can be made avai 1 abl e. 
9 CUBE (TYP) 

'NOTE SOLAR DYNAMICS SYSTEM A PRIME OPTION (NOT SHOWN) 

Figure 4.4.2-2 - RCS THRUSTER LOCATION 
The RCS will be the backup system for attitude control in the event that 

the CMG systera becomes inoperative and will provide momentum management 

in the event that the magnetic torquers becore inoperative. Failure of 

both the Cffi systm and the magnetic torquer system such that both 

systems are in the aalfunctioned status at the same time is considered to 

be too remote to size propellent for both. Consequently the propellent 

reserve will be bas?d on whichever requires the most propellent. Failure 

of the entire CMG system is considered an emergency, and the Orbiter i s 

expected to bring up the necessary parts/personael for repairs in 22 

dqys. Consequently, the propellent will be sized for reduced pointing 

for 20 dqys awaiting Orbiter arrival, 2 days of noninal pointing for 

Orbiter berthing and 2 days after Orbiter arrival (with reduced pointing) 

for repairs. Sufficient propel lent is provided for CMG spin up. 

The magnetic torquer failures will not be considered an emergency 

situation, and the RCS will be sized to perfonn momentum management for 

90 days realizing that the Orbiter may return early since the .00001-g 

environaent will be violated with RCS in use. The RCS will also be used 

to perfonn attitude control augnentation in the event that the torque or 

Aloraentua exceeds the capability of the CMG-magnetic torquer systems. No 

propellent has been provided for excesses except in the cases listed 

below; however, it is assumed that this propellent would cane from the 20 

mi transfer budget providing that it has not already been used. All 

available control effectors (magnetic torquers 2nd function1 ng CMGs) will 

be used to conserve propel 1 ent . Some of the evpnts that are expected to e~ceea the torquer/nunnentum 

capability of the CMG-magnetic torquer control systems are the Orbiter 

capture ~nd subseqkent terthing/docking and module berthfng and/or change 

out; consequently, propellent has been budgeted for these events. The 

velocity difference between the Orbiter and the Station was assumed to be 

.1 ft/s in three orthogonal directions and .2 deg/s about three 

orthogonal axes. Two of the velocities were assumed to act on a 75-foot 

moment am, and the third was assumed to be zero. The manipulator am on 

the Station was assumed to require 5 accelerations and 5 decelerations 

along and about two axes for berthing the orbiter and each module with 

each acting on a 75-foot moment am. 

The required thruster force is sized by the maximum disturbance torque. 

The manipulator am has been assumed to provide the highest torque. 

Since the Station arm has not been defined yet, it was assumed that the 

am would be capable of a 25-lb force which would be 100 feet from the 

Station's center of grzvity. Using two thrusters, each 50 feet from the 

center of gravity, a 25-1 b thrust would be required. 

4.4.2.2.7 Solar array actuator interfaces 

There are two types of solar array actuators. The Alpha (OC) actuator is 

required to maintain the sol ar panel 1 ongi tudi nal axi s perpendicul ar to 

the sun line. Motion in tn~s axis is due to the orbit of the Space 

Station about the earth. The oC actuator turns at one revolution per 

orbit. The beta (PI actuator maintains the solar panel transverse axls 
perpendicular to the Sun lint. Motion in this axis is due to both the 

Earth's rotation about the Sun and the nodal regression of the orbit. 

This actuator cycles approximately eight times per year. 

The apparent load on these actuators is rssumed to be largely an inertia 

load. In addition to inertia, these actuators must overcome stiction and 

rot 1 ing friction. Torques due to aerodynamic loads, plum impi ngemnt, 

dynamic motions of the Space Station, and the solar panels themselves are 

assumed to be negl i gibl e. 

The 1~: actuators will also incorporate four slip rings or roller rings to 

pass the electrical power generated by four solar panels, three to 

provide for data links to thep actuator drive electronics and three to 

provide for conditioned power to the &3 actuator drive electronics. 

The / actuators, in addition to their inertia loads, must overcome 

torques due to power and coolant lines. 

The prime mover for both actuators :?ill be brushless dc motors. The 

motor stator will be divided into three segments. Each segment will 

contai rr a coinplete three-phase winding to provide electromagnetic torque 

summing, The motors will be sized such that any two segments can drive 

the load. This technique provides a match between the three channel 

input from the GN&C system and the nonredundant output of the actuator. 

A1 ternate approaches include veioci ty sumnation of three independent 

motors using differential gears or other unique motor designs not 

presented herein. 

4.4.2.2.8 Core radiator actuator interface 

The alpha oC actuator(s) for the core radiators will be approximately one 

foot in diameter. These actuators will maintain the radiator transverse 

axis normal to the Sun line. The radiators are unwound on the dark side 

of the orbit to eliminate the need for continuous rotation of a fluid 

joint. 

The design of this actuator will be similar to the actuators described in 

paragraph 4.4.2.2.7. 

4.4.2.2.9 Payload ~ointina systems 

Table 4.4.2-1 1 i sts selected payloads which have pointing requirements 

and shows the suggested location for the payloads. Two payloads (Pinhole 

Occulter, SAA0009, and Star1 ab, SAA006) have been recommended for the 

28.50 platform rather than the Space Station because of their severe 

pointing requirements. Also, the Pinhole Occulter Facility (50 meters in 

length ) would prefer to be 1 ocated on an inertial ly-oriented platform, 

If the Station were pointed to within 1 degree, only the Earth 

Observation Instrument Techno1 ogy Pay1 oad ( TDM2260) would require a 

pointing system to meet the pointing requirement of 0.1 degrees. An 

inexpensive pointing system could be used to meet this requirement. It 

is also in the realm of possibility that the Station may be able to point 

that accurately, in which case, no pointing system would be required. 
Two solar experiments (Space Plasm Payload, SAA0207, and Material 

Performance, ~010) mounted on the transverse bow require a single 

axis gimbal for angle correction. With the 4 angle gimbal, the Station 

pointing accuracy is sufficient to meet their pointing require~lents of 1 

and 2 degrees respectively. 

With the presctlt payload list (Table 4.4.2-1 1, only Inexpensive pointing 

systems are required if two payloads (SAA0006, SAA0009) are relocated to 

the 28.50 platform. It is possible that future pgyloads which can be 

acca~podated by the Station will require a paylcmd pointing system in 

sJhich case the Dornier IPS will be a good candidate. It appears that a 

new development for a pointing system for the station is unwarranted at 

this tlme. 

The payload pointing mount will be interfaced to the GNLC system using a 

payload pointing mount interface electronics unit. The function of this 

unit will be to detect faults in the incoming corrnands from the GMC 

system, is01 ate the fault, and reconf igure such that the payload pointing 

mount receives a valid coarse pointing cmnd. 

4.4.2.2.10 System sizing 

The sensor and processors have been chosen such that at least fail 

operational /fai 1 safe/restorabl e redundancy exists for a1 1 components. 

Cold components will be used where practical to reduce the power 

requi reinents. Control effectors have been sized 50 percent greater than 

the number of units estimated to be necessary while making certain that 

sufficient redundancy exits. 

a. IOC station. Table 4.4.2-2 contains the quantity, 

size and weight of ae components selected for the IOC Station. Table 

4.4.2-3 contains the electrical power requirements. 

b. Growth station. The growth Station has not been 

sufficiently designed and analyzed' to determine the system size of the 

control effectors. Depending on the design of the growth Station 

relative to the IOC Station, the number of control effectors could either 

be reduced or increased. It is anticipated that the growth Station will 

require about the same or a slightly greater number of control 

effectors. The sensors and processors are anticipated to be the same as 

the IOC Station. 
TABLE 4.4 -2-1 - PAYLOAD ACCMODATIONS LOCATION 

orientation ID Code Pal oad 

Station Zenith 

- FOV 

Pointing 

Accuracy 

Heau i red 

Anti -Earth SAA00005 

Station Nadir 

Transition Radiation & 

Ion Calorimeter (TRIC) 

Earth SAA0201 

Earth SAA0207 

Earth TOM2260 

Earth TDM2510 

Transfer Boom - 1 Solar Array ) 

Sol ar SAA0207 

Solar TDM2010 

Inertial TDM2410 

Inertial TDM2420 

- 28.5O Platform 

Solar SAA0009 

Inertial SAA0006 

LIDAR 

Space Plasma Payload 

(Tether) 

Earth Obs. Instr. Tech. 

Environmental Effects 

Space Plasma Pzy 1 oad 

Material Performance 

Attitude Control Tech. 

Figure Control Tech. 

Pinhole Occulter Fac. 

Star1 ab 

10 arcsec 

2 arcsec 
TABLE 4.4.2-2 - SPACE STATION GNLC EQUIPMENT LIST 

Dimensions Vol . Vol . Wei ght Wei ght 

Per Unf t No. Unit Total Per Unit Total 

Unf t Description (In.) Units (Cu.Ft.1 (Cu.Ft.) (Lbs.) (Lbs. ) 

Actuators 

CMG Assembly 42 Sphere 6 22.45 134.70 420.00 2520.00 

Magnetic Bar 2.25 Dia.x98 6 0.23 1.35 109.00 654.00 

Total Actuators 

Sensors 

Star Tracker Triad 13x13~12 2 1.17 2.35 20.00 40.00 

Hexad Strapdown 12x12~12 1 1 .OO 1 .OO 50.00 50.00 

Spares for above 12x12~12 1 1 .OO 1 .OO 50.00 50.00 

Total Sensors 4.35 140.00 

Electronic Support 

Magnetic Torquers 9x9~9 

RCS Cont System 12x 18x8 

MC Processors 7.5xlb.5~15 

NAV/Traf f i c Proc 7.5~10.5~15 

Interface Devices 7.5~10.5~15 

Solar Array Elec. 8x9~6 

Spares for above 8x9~6 

Radiator Elec. 8x9~6 

Spares for above 8x9~6 

Pqyload Elec. 2x 5x6 

Total Electronics 

Grand Totals 
TABLE 4.4-2-3 - SPACE STATION GN6C POWER REQUIREKNTS 

Avg Pwr Avg Pwr Peak Pwr Peak Pwr 

No. Per Unit Total Per Unit To ta 1 

Unit Description Units (Watts) Matts) (Watts) (Watts) 

Actuators 

CMG Assdlj 6 100.00 600.00 125.00 750.00 

Magnetic Bar 6 3.50 21 .OO 7.00 42.00 

-- 

Total Actuators 

Sensors 

Star Tracker Triad 2 10.00 20. OQ 12.50 25.00 

Hexad Strapdwn 1 100.00 100 .OO 125.00 125 -00 

Spares for above 1 0.00 0.00 0.90 0.00 

Total Sensors 120.00 150.00 

Electronic Support 

Magnetic Torquer 

RCS Cont System 

G&C Processors 

NAV/Traff ic Proc 

Interf'ace Devices 

Solar Array Elec. 

Spares for> above 

Radiator Elec. 

Spares for above 

Payload Elec. 

Total Electronics 

Grand Totals 2101 .OO 2707 .OO 
4.4.3 Cmunications and Tracking System 

4.4.3.1 Comnunications and Tracking System Gefini tion 

The Communications and Tracking System will be designed to provide 

communications and tracking services between the Space Station and 

ground. This is done via a relay satellite system, various space 

vehicles interoperating with the Space Station, as well as internal 

communications services. Interoperating vehicles for the initial phase 

will include the Space Shuttle Orbiter, EVA crewmembers, co-orbi ting 

free-flyers, orbital maneuvering vehicles (OMV's), and a space platform. 

For the growth phase an orbital transfer vehicle (OTV) will be included. 

In translating the general comnunications and tracking system 

requirements into a Space Station System Concept, it has been necessary 

to formulate some basic assumptions. These are summariz~4 be1 ow: 

a. Free-flyers will rendezvous and dock with the Space Station 

under control of an OMV. It is assumed thcit the Space Station will 

communicate with free-flyers out to 1080 nmi. 

b. The co-orbiting space platform will be treated as a 

free-f lyer for communications to/from the Space Station. 

c. Space Station/OMV comnunications and tracking are required 

only out to 20 nmi. 

d. Co-orbi tal vehicles locatcd beyond 20 nmi. of the Space 

Station along the velocity vect~r will provide position data to the Space 

Station. 

e. Tracking data will be updated on a near continuous basis. 

f. Low data rate users will be required to provide at least 10 

db of EIRP at 1080 nmi . 

g. High data rate users will be required to provide at least 23 

db of ElRP at 1080 nmi. 

h. In addition to some duplicated hardware, redundancy 

requirements wi 11 be met by either electronic design, functional 

considerations and/or operational considerations at some reduced 

performance capabil 'ties. 

i . Selectively locating comnlrnicatioos and tracking hardware 

throughod- the Space Station will satisfy the requirement of safe haven. 

Each link will provide a variety of services and will entail unique 

operational requirements. Multiple simultaneous 1 inks will also be 
required. Figwrc. 4.4.3-2 shows an overall Space Station comnunications 

and tracking system block aiagram. 

The Space Station/ground uplink and downlink channels wfll operate 

through a relqy satellite at S-band and Ku-band. The com~unication 1 inks 

between tb:: Spare Station and Opbitw will operate at S-band 

frequencies. The links betwen the Space Station and space platforms, 

free-flyers, EVA, OMV, and/or mnned/umnned OTV's will be at I(-band. 

The comnuni ca t ion system wi 11 be capable of transmission, reception, arid 

processing of voice, telemtry , conmands, wideband data, television (TV), 

and text and graphics. The system will include the capability for 

private comnunications including Canmunications Security (COMSEC ) 

requirements. Relay of interoperatf ng vehicle data to/from the ground 

will be provided through the Space Station via a synchronous satellite, 

The internal C&T system will provide for Space Station intermodule and 3.-A-..-- 

,IIL, ~~~~dule CET. Services provided by he internal CbT system include 

video, audio, commands, telemetry, data text jlrrd graphics, and CbT 

rrl~nagement/c:ontroI/distribution. The C&T System harrlware will have 

embedded processors that are software controlled for system 

confi guration. 

The C&T system will be developed and implemented incrementally to support 

assembly, initial and growth phases. The phases will incorporate a 

logicdl progression in C&T system capabi 1 ity ~natching ttie growth in Space 

Stati on CLT requirements. 

A more detailed description of the comnurlic~tions and racking functional 

areas (RF comnunicaticns, audio system, videa system, tracking systems, 

signal processing, data acquisi tion/telemetry/comnandr , systems 

mnitoring and control, and antenna systems) is contained in the 

following sections. 

Comnunication 1 ink characteristics are show in Table 4.4,3-1 and 

tracking 1 ink characteristics are shown in Table 4.4.3-2. Performance 

requfrements both external and internal are shown in Tables 4.4.3-3 and 

4.4.3-4. 

4.4.3.2 RF Comnunications 

The Space Station Program RF system corrsists of dedicated RF links to 

TDRS, Orbiter , MFWS, and RF 1 inks with free-flyers, co-orbi t lng platform, 

EVA'S and OMV. 

An ini tlal S-' low data rate TDRS, GPS, and SSIOrbiter S-band 

capability wf, re provided durfng the assembly phase when the Space 

.Station is unmanned. The S-band RF systen~s hardware may require 

relocation if after ass-ly there is significant antenna blockage or it 

is necessary to optimize performance by minimizing cahle runs. 
TO DMS 

CENTRAL r------ CREW INTERFACES 

INTERMODULE OISTHIBUTION Chs INTkRNAL 

INTERFACES 

--------------__ 

PROCESSOR 

CdT SPACE 

STATION 

---- ---- 

EXTERNAL DETACHED 

INTERFACES PAYLOADS 

Figure 4.4.3.-1 - SPACE STATION COMMUNICATIONS AND TRACKING SYSTEM 
TABLE 4.4.3-1 - COMMUNICATION LINK CHAMCTERISTICS 

VEHICLE 

LIMKS 

SS-TDRS 

- 

SS-ORBITER 

P 

w 

VI 

SS-EVAjMMU 

SS-FREE 

FLYER 

PLATFORM 

XMIP 

KuBMD 

S-BAND 

S-BAND 

S-BAND 

K-BAND 

K-BAND 

- 

K-BAND 

- 

FREQUENCY 

RVC* 

KuBAND 

S-BAND 

S-BAND 

S-BAND 

K-BAND 

K-BAND 

K-BAND 

K-BAND 

K-BAND 

NO. OF 

VEHICLES 

(SIMULTANEOUS) 

ONE 

ONE 

OHE 

ONE 

ONE 

TWO 

TWO 

ONE 

(EIGHTI** 

ONE 

(EIGHT)** 

CAPABILITIES RANGE 

(MAX 

(NU1 ) 

- 

23,000 

23,000 

23,000 

20 

20 

0.54 

0.54 

1080 

1080 

DATA RATE (MAX) 

XM I T* 

VOICE, TLM 

TV ,DATA 

XHIT* 

300-MBPS 

3 MBPS 

96 KBPS 

16 KBPS 

16 KBPS 

48 KBPS 

400 KBPS 

- 

48 KBPS 

- 

RC V* 

VOICE, CMD 

TV, DATA 

TEXT 6 

GRAPICS 

RC Vw " 

25 MBPS 

300 KBPS 

32 KBPS 

2 KBPS 

16 KBPS 

64 KBPS 

25 MBPS 

64 KBPS 

5 MBPS 

(25 MBPS)** 

I 

VOICE ,TLM 

TLM 

TLM 

VOICE 

VOICE,CMD 

TV 

- 

CMD 

- 

VOICE ,CMD 

CUD 

CMD 

VOICE 

VOICE ,TLM 

TV 

TLM 

TV OR 

DATA 
lABLE 4.4.3-1 - COM14UNICATION LINK CHARACTERISTICS (CONTINUED) 

* REFERENCED TO SPACE STATION. 

** A TOTAL OF EIGHT VEHICLES WILL BE 

SERVED BY THIS LINK AT HIGH OR LOW 

DATA RATES OR A COMBINATION OF THt 

TWO RATES IN THE GROWTH CONFIGURATION 

RAN

(MAX

(NM

NO. OF 

VEHICLES 

(SIMULTANEOUS) 

-. - 

VEHICLE 

LINKS 

- 

ONE 

ONE 

bNE 

ONE 

DATA RATE (MAX' 

20 

0.54 

SS-OMV 

SS (MRMS) - 

ORB I TER 

P SS-MIWS W 

m 

.em.-. 

FREQUENCY T-~TCIPILIT ~ES 

XMIT* 

- 

0.05 

0.05 

XMIT* RCV* 

48 KBPS 

- 

16 KBPS 

(4.5 MHz) .. 

2 KBPS 

(4.5 MHz) 

K-BAND 

- 

S-"4NL 

S-BAND 

64 KBPS 

25 MBPS 

2 KBPS 

- - 

16 KBPS 

- 

RCV* 

K-BAND 

K-BAND 

S-BAND 

S-BAND 

XMIT* - RCV* 

CMD 

- 

TLi4 

TV 

CMD 

TLM 

TV 

CMD 

TLM 

TV 
TABLE 4.4.3-2 - TRACKING LINK CHARACTERISTICS / 

CWME NTS 

GPS RCVR 

REQ'D ON 

EACH VEH, 

GPS RCVR 

ON EACH 

mu 

RANGE 

20 NMI. 

1080 NMI. 

3280 FT 

N/A 

SS-DOCKING 

L 

VEHICLES 

# OF VEH. 

ONE 

ElGHT 

TWO 

FOUR 

NAV-STAR 

DATA RATE 

1/SEC 

DURING 

TRACK 

1/10 SEC 

1/SEC 

1/10 SEC 

VEHICL E 

LINKS 

SS-ORB 

SS-OMV 

. 

SS-PLATFORM 

SSFREE-FLYER 

SS-EVA 

SS-GPS 

RANGE 

RANGE-RATE 

AZIMUTH 

ELEVATION 

ATTITUUE 

OPTICAL 1000 FT WE FOR 

EACH 

DOCK1 NG 

PORT 

CAPABILITIES 

RANGE 

RAN GE-RATE 

AZIMUTH 

ELEVATION 

ANGLE RATE 

GPS NAY 

POSITION DATA 

ON COMM LINK 

GPS POSITION 

ON COMM LINK 

NAY POSITION 

DATA 

FREO. 

Ku BAND 

K BAND 

K BAND 

L BAND 
TABLE 4.4.3-3 - EXTERNAL CMUNICATIONS AND TRACK ING 

PERFORMANCE REQU IREMNTS 

FUNCTION PERFORMANCE REQUIREMENT REMARKS 

TELEMETRY WIDEBAND BER - c 10-5 

AND COFMAND DATA 

BER MEASURED AT EARLIEST 

BIT SYNCHRONIZER OUTPUT 

COMMAND AUTHENTICATION 

REQUIRED 

VOICE DIGITAL BER - < 10-5 BER MEASURED AT EARLIEST 

VOICE RECOGNITION BIT SYNCHRONIZER OUTPUT 

REQUIRED, VOICE VOICE INTELLIGIBILITY 

INTELLIGIBILITY MEASURED USING HARVARD 

9i) PERCENT PHONETICALLY BALANCED 

UORD LIST 

TEXT AND GRAPHICS BER < 10-5 - 

TELFVTSION BER - < 10-5, NTSC 

DIGITAL STANDARD DIGITAL TO 

ANALOG PROCESSOR, Spp/ 

Nrms 35 Db RESOLUTION = 

250 LINES 

BER MEASURED AT EARLIEST 

BIT SYNCHRONIZER OUTPUT 
TABLE 4.4.3-3 - EXTERNAL CWUNICATIONS AND TRACKING 

PERFORMANCE REQU IREMNTS ( CONTINUED ) 

FUNCTION PERFOrZMANCE REQUIREMENT REMRKS 

TELEVISION Spp/Nms - > 35 Db, NTSC MEASURED AT THE N 

ANALOG STANDARD TERMINAL 

RESOLUTION = 250 LINES 

COMPUTER DATA BER < 10-9 - 

LONG RANGE MAX RANGE - 1080 nmi 

TRACKING COVERAGE - LIMITED TO 

COMM DATA LINK COVERAGE 

ACCURACY - (GPS POSITION) 

+/- 15M (49.2 ft) 

SHORT RANGE MAX RANGE - 20 NMI 

TRACKING COVERAGE - 4 PI STERADIANS 

ACCURACIES: 

ANGLE - +/- 10 MRAD (0.57 

DEG) 

RANGE - +/- 100 M (328 FT) 

OR 1% 

VELOCITY - a3 M/SEC (1 FPS) 

OR 1% 

BER MEASURED AT INPUT TO 

CO)((PilTER TERMINAL 
TABLE 4.4.3-3 - EXTERNAL COWUNICATIONS AND TRACKING 

PERFORMANCE REQUIREMENTS (CONTINUED ) 

FUNCTION 

-- PERFORMANCE REQUIREMENT REMARKS 

PROXIMITY MAX RANGE - 3280 Ft 

OPERATIONS COVERAGE - LIMITED TO COW DATA 

TRACKING COVERAGE 

ACCURACY: GPS POSITION- +/-1M 

(3.3 FT) 

DOCKING SENSORS MAX RANGE - 1000 Ft 

COVERAGE - 20 DEG CONE 

ACCURACIES: 

RANGE - +/- 0.5CM (-02 FT) 

ANGLE - +/- 2 MRM (0.1 DEG) 

VELOCITY - 1.0 CM/SEC (0.03 FPS) 

ATTITUDE - +/- 10 MRAD (0.57 DEG) 
TABLE 4.4.3-4 - INTRA-SS CWUNICATIONS 

PERFORMANCE REQUIREMENTS 

FUNCTION PERFORMANCE REQUIREMENT REMARKS 

TELEMETRY COMM~WD BER - < 10-9 BER MEASURED END-TO END 

AhD W IDEBAND DATA OVER THE INTENDED 

CMUN ICAT ION PATH 

VIDEO HIGH- BER = 10-7 

RESOLUTION AT OUTPUT OF TV DIGITAL 

DIGITAL TO ANALOG PROCESSOR, 

Spp/Nms = 48 dB 

RESOLUTION = 700 LINES 

NTSC DIGITAL BER < 10-7 

AT ODPUT OF TV DIGITAL 

TO ANALOG PROCESSOR 

SPP/Nrms > 40 Db 

RESOLUTI(m = 350 LINES 

HIGH RESOLUTION Spp/Nrms > 48 dB 

ANALOG RESOLUTION = 700 LINES 

NTSC ANALOG Spp/Nrms >40 dB 

RESOLUTION = 350 LINES 

VOICE DIGITAL BER < 10-5 WORD INTELLIGIBILITY 

VOICE INTELLIGIBILITY MEASURED L'SING HARVARD 

> 90 PERCENT. VOICE PHONETICALLY BALANCED 

KECOGN I T I ON REQU I RED YORD LISTS 

99% 

CONTROL/ BER < 10-9 

MONITORING/ - 

STATUS DIGITAL 

DATA 

BER MEASURED END-TO-END 

OVER THE INTENDED 

COMMUNICATIONS PATH 
The Space Station RF system will provide spherical comn~inications 

coverage out to 20 nmi. for near range operations using low gain 

hemisphere antennas. Figure 4.4.3.2 describes this multiaccess RF 

comnunication concept. For proximity operations out to 3280 feet a low 

data rate (LDR) mu1 tiaccess system will transmit simul taneously to two 

EVA'S, commands, voice and freeze frame TV and receive volce and 

telemetry from both. To receive TV from an EVA a four channel high data 

rate (HDR) receiver is provided. 

The system a1 so provides LDR comnunications to/from an OMV transmitting 

commands and receiving telemetry out to 20 nmi. In addition, the OMV can 

transmit TV to the Space Station over the HDR channel within 3280 feet 

during docking maneuvers. 

The Space L .tion will have two-way comnunications out to 20 nmi. with 

the orbiter. The system will be compatible with the present Orbiter 

S-band payload interrogator system providing telemetry and comnand 

capability. A voice capability will be provided utilizing the Space 

Station EVA 1 ink in conjunction with the upgraded Orbiter EVA system. 

An RF system on the Space Statiotl that provides commands to and 

telemetrylTV from a mobile remote manipulator system (MRMS) is shown in 

Figure 4.4.3-3. The system design is compatible with the Orbiter for 

operations during assembly and a1 so will provi de control capability when 

the Space Station is manned. Return TV to the Orbiter will use an FM 

link which will be compatible with the Orbiter's EVA TV S-band FM system. 
11 I I 

1 1 Ill I 

1 

ZONE^ ZONE2 1 ZONE3(4) I ZONE

PROX OPS LOW GAIN HEMI /G = 13d~ 

LOW DATA RATE ElRP = 

MULTl ACCESS EIRP = 5dB 13dB I 

DIPLEXER 

G = 5dB ElRP = I 

I b a I 

i RTN 1 I PROXOPS 

I 

I I 

HIGH DATA RATE I I I 

RETURN LlNK 

+ I - ME^ GAIN CONICAL I 

j (MULTIPLE BEAM) I L~R 2 

1 

FAR RANGE I 

I LOW DATA RATE 

MULTl ACCESS I 

P 

I P -' I I 

W I 

I 

I I 

I I FAR RANGE 4 

G=SWB HDR I 

HIGH DATA RATE RTN I 

RETURN LINK 1 

HlGH GAlN 

I 

I 

((NARROW BEAM) 1 

I 

-SPACE STATION- 

,' ""100 NMI--------------r( I 

Flgun 4.4.3.-2 - MULtl ACCESS RF COMMUNICATIONS FOR GO-ORBITAL ZONE COVERAQL. 
-------- 1 

I , S-BAND 

I 

I CMD'S-* 

' ' - I 

I 

I TLM - XCVR RCVR -TV I 

I i I 

I 

I 

1 

I > ) FM !* 

I XMTR S-BAND 

I TV I 

I I 

I 

I 

I 

I 

I 

I CMD'S S-BAND 

I XPNDR - TLM 1 (ORBITER) I I I 

I I 

I TRANSVERSE BOOM I L ,,,--,,--,-,,--,-----

A 

'ORBITER COMPATIBLE LINKS 

Flgure 4.4.3-3 - MRMS RF SYSTEM CONCEPT 
For the initial Space Station configuration a far range, single access, 

system will be used for coveraqe out to 1080 nmi. This system will 

transmit comnands to co-orbiting free-flyers and platforms and receive 

telementry and data or slow scan TV from these co-orbiting free-flyers 

and platforms. Since this system utilizes a single beam steerable 

parabolic dish antenna, approximately 2.5 feet in diameter, only one user 

at a time is serviced. 

This single access system will require the HDR users to provide an EIRP 

of + 23 db for coverage at 1080 nmi . for data rates of 5 mbps. The LDR 

users will be required to provide a minimum of +lo db for coverage at 

1080 nmi. for a data rate of 64K bps. 

The growth RF systems far range configuration as shown in Figure 4.4.3-2, 

will utilize a medium gain, multibeam, offs. : reflector antenna which 

will provide four beams out to 1080 nmi. This will allow f@r the 

simultaneous conmunications with eight free-flyers (four farward/four 

aft), one of which may be a co-orbiting platform. This multia~cess 

system will have a low data rate capabil fty for transmitting comnds and 

receiving telemetry and a high data rate capabil: ty to receive wideband 

scientific data or digital TV. HDR users will be required to provide a 

minimum of + 30 db of EIRP for data rates of 25 mbps at 1080 nmi. 

This will be a four channel system. The LDR portion of the multi access 

system will require the users to provide + 10 db of EIRP at 1080 nmi. for 

a 64K bps data rate capability similiar to that of the single access 

system in the initial configuration. 

'The HDR system will be used in combination with the medium gain multibeam 

antenna to provide up to eight high data rate channels simultaneously 

( four-Fwd/four-Aft ). 

The respective EIRP's and receive antenna gains for the LDR multiaccess 

system were sized for a tlme division modulation scheme which requires a 

large received signal power. However, this design will support other 

modulation schemes such as frequency hopping, direct sequence, or 

frequency division. This could be more desirable If some other 

performance characteristics were considered primary. The HDR 1 ink, 

because of the wide bandwidths required, may be 1 imfted to a frequency 

di vi si on modul ati on technique. 

The Space Station TDRS 1 ink will include both S-band and Ku-band single 

access links. The S-band and Ku-band links will have the capability to 

use maximum TDRS data rate capacity. A low data rate capabi 1 ity at 

S-Band is provided as a backup. 

.Power amp1 ifiers, low noise amp1 ifiers, and up/down converters will be 

located at the antennas to obtain optimum performance. All electronics 
4 

\ ., mounted on or near the antennas will be redundant because of the 

difficulty of repairing and/or replacing components. IF signals from a 

module will be routed from their respective modulation sources to the 

antenna electronics for up conversion to the specified RF frequency. The 

dowr converted signals will be routed to their respective demodulators 

I ~:ated in a module. Df stributing the C&T hardware in such a manner will 

permit the system to be improved, reconfigured or expanded without 

modifying the entire RF subsystem. 

4.4.3.3 Audio System 

The audio system processes and distributes various sources of crew voice 

and other audio signals throughout the Space Station modules, berthing 

ports and airlocks. Intercom and paging channels are provided along with 

duplex voice channels to the ground, and to EVA/WU's and Shuttle 

Orbiter (s ). Speech recognition (voice control 1 ed commands and speech 

synthesis is provlded along with the capability for re:ording, for CCTV 

usage, for recreational purposes and private 1 ines. Distribution of 

caution and warning tones is also furnished. 

Each Space Station module will contain a largely self-contained all 

digital system for audio generation, processing, switching, and 

distribution. Duplex, mu1 tiaccess TDM channels are used to accommodate a 

crew size of slx for the initial phase an3 16 crewmembers in the growth 

configuration. 

Bandwidth compression techriiques will be used to limit external 

transmissions to 16 Kbps per voice channel (Orbiter channel to remain at 

32 Kbps) while still retaining good "toll quality" voice fidelity. 

Capabil ity for private voice transmission (local ized encryptior: will be 

provided for medical, family, or proprietary conversations (compatib'e 

decryption devices must be provided by each listener on the ground). 

A crew-mounted wireless comnunicator will be used to transmit crew vofce 

and to receive onboard and ground audio signals. Val 1 -mounted 

speaker/mic units will a1 so be suppl ied for hands-free operation. 

An emergency comnunications ;cni t provides for intermodul e voice 

communications (short-range RF system) during an audio circuit loss via 

the primary system. 

Local audio terminals in each module contain vofce recognition units 

which can decipher voice commands and di ctri bute encoded digital words to 

the display and control (D&C) system. A recognition accuracy of at least 

99 percent will be met. Voice synthesis uni ts will also be provided such 

that digital comnand words from the D&C system can initiate a good 

qua1 ity synthesized voice output to the crew, 

Since the audio system wi'll be frequently reconfigured by the crew 

(channel selection, transmi t/recei ve modes, volume levels, privacy 

control, etc.) various switching functions will be localized at the c- 
particular local audio terminals in the innu iate vicinity of the working 

crewlaenrber . Due to on-orbit pernranency of the Space Station, tkt, audio systeln is 

designed to maximize flexibi i ity in hardware and system capabilities for 

distributed control and processing. Real ti= adaptaticn and on1 ine 

reprogrami ng capabil i ties will be provided. 

Interfaces with the DLC system for caut!on and warning tones and voice 

c~nd/synthesi s signals, with the onboard recording system, wi th the 

CCTV system, aoc! principally with the CbT network signal processors and 

associated RF transnission/reception systems are handled throur'- 

interface hardware external to the audio system. 

4.4.3.4 Viri.:o System 

Video service will be provided in the appropriate Space Station elements 

for such operational tasks as docking, trackfng, construction, and other 

monitoring tasks; to provide recreational/entertainmnt/lei sure, to 

record and store video, and for training. 

Tlre configuration, safety, and functional requirements of the Station 

call for c~ntrol stations in eacj Fabitable module so that different 

crewmelnbers can perform their required tasks with minimal or no 

interruption to or from others. The system will be >inrple to operate 

since there wi! i ;he a large number of users specialized in many different 

fields, and specla1 training for Space Station equipment is kept to a 

mininum. 

These requirements drive the design of the television system to a 

d,i stri buted control system where camera rontrol s, video swirching, and 

other system functions will be control led frtm any workstation or 

monitoring location. This distributed control station concept bill allw 

continuous operation even if parcs of the Station become uninhabitable. 

These workstations wi 11 incorporate user-friendly input devices such as 

touch screen sensors, joysticks, and voice control inputs used in 

conjunction with color graphics generated menus, and displays. The 

capabilitj to move TV mnitors from one lccation to another will be 

1 ncorporated. 

Video signals, both black and white arid NTSC color id! ll be d: stributed 

aboard the Station and for transmission to/from the yound. The fowt 

for video signals wil l te both digital and analog. Large video bandwidth 

capability will be available onboard for high definition television. 

Solid state video cameras will be located in each module; some with 

.pan/tilt capability. External cmeras will be located in appropriate 

locations for viewing and visual tracking uf external activities. Two 

cameras will be provided at each docking port to aid the lafer docking 

system. 
4.4.3.5 Tracking artd BerthinglDocki ng System 

The tracking and berthing/docking system hardware will primarily be based 

on function by opera'ional zones in accordance with the table Lelow: 

Far Rarlge 0-1080 nmi. 

Near Ranye 0-20 nmi . Proximity Range 0-3280 ft. 

Docki ng/Berthing Range 0-1000 ft. 

The far range ~osition information will be obtained by providing a GPS 

receiver/processor on each vehicle to be monitored, and transmitting the 

navigational data to the Space St3tion utilizing the comnunications 

return 1 ink. The C&T processor wiil combir!e this ddta with the onboard 

GPS information to provide relative position with the Space Station. 

The near range zone tracking will be accomplished with a radar consisting 

of a transmitter/receivep RF assembly, a signal processor, and a 3-'foot 

parabolic dish antenna. Two complete rystems will be required to provide 

adequate coverage. The primary purpose of this radar is to track 

approaching vehicles as they enter the trear range zone, through the 

proximity operations zone, to a position within the acquisition region of 

the docking system. It will also monitor departing vehicles in a similar 

manner. 

The proximity operations wi! 1 involve simultaneous tracki ~g of two 

EVA/EMU/MMU1s. A CiPS receiver/processor on the MMU will supply position 

information to the Space Station via the conmunications data link. A 

small ranging radar will also be a part of the MMC to give the astronaut 

range and range-rate information to any object in front of him. 

Each docking/berthing port on the Space Station will include a laser 

docking system that will provide range, range-rate. angular position, and 

attitude of the docking vehicle. Small pdssive retro-reflectors will be 

r jquired on each vehicle in a standardized pattern to facilitate the 

attitude d3ta. 

A mu1 tichannel GPS receiver/processor w;' 1 be included on the Space 

Station to provide positional data to the Spac? Sticion navigation 

system. 

A Ku-band radar transponder will be included to provide augmented 

tracking for the Orbiter Rendezvous Radar. 

Structural stabil i ty monitor sensors, using 1 aser ranging techniques, 

will be placed at appropriate locations on the Space Station structure 

with corresponding retro-reflector: . This will provide flexure and 

bending data for structural integrity monitoring, re1 ative structure 

alignment wit" the Nav-base, and for fine-tuning of antenna pointing, ?S 

required.. 
4.4.3.6 Signal Processing 

Signal processors to interface with TV, text and graphics, digitized 

audib, command, telemetry, and data acquisition systems for the 

processing and protocols necessary for optimum operation of these system 

will be provided. These processors will proviG2 the equipment necessary 

for synchronization, encoding and decoding, error detection and 

correction, routing and distribution of signals throughout the Space 

Station and Stat: 9n elements. 

Coimnand authentication will be processed for cmnd channels and 

encryption/decryption and key distribution for secure 1 inks will be 

provi ded. 

The capabil ity to receive and transmit text and graphic hardcopy will be 

available. 

Dedicated system processors for the various elements (free-flyers, OMV, 

TDRSS, etc.) will be provided for efficient handling of these links. The 

processor hardware will be distributed throughout the comnunication and 

tracking system to optimize overall performance. 

4.4.3.7 Data Acquisiticn, Telemetry, and Comnand 

The command/data acquisition system processes and val idat.es up! i nk 

corrmands, processes commands to detached vehicles, and processes Space 

Station, detached vehicle, and experiment data for onboard distribution, 

down-1 ink telemetry, or onboard recording. These functicns will be 

accomplished by the use of local data bus networks to interconnect the 

elemeqts and transport information to various users. 

Uplink operational cmmnands authentication processing and formatting for 

use in autonomous controi and management when the station is unmanned and 

f7r use by the facility management system when manned, is Yerformed. 

Capability to encode and process onboard originat-ed comnands and wide 

band data for transmission to detached vehicles, including the Shuttle 

orbiter shall be provided. 

Uplink science comnands will be pcocessed for transmission to onboard and 

detached vehicle experiment packages. This information may be 

time-shared with other up1 i nk data if advantageous. 

Operational , engineering, and science data from various onboard and 

detached vehicle sources will be acquired, processed, formatted, and 

interleaved as necessary for distribution onboard and down-link 

transmission via TDRSS. 

Provisions for packetizatio of data sources (either at the source or at 

the point of mu1 tiplexi ng) saall be provided to optimize end-to-end data 

hand1 ing and distribution to experiment dat? users at widely dispersed 

locations. 

449 
To minimize hardware and software impacts fra requimnts changes and 

equipment upgrading distributed processing and control shall be 

eqhastzed. kiuw flexibil ity in hardare and system capabil ity should 

be a major goal in the design of the vstem. Realtim adaptation range 

chanss capabil i ties, reprograming of data fomats including deletions 

and addition of data sources will be required for the system. The 

ability to azcmdate di fferent redundancy requirents for numrous 

systems by the si.ple addition or deletion of hardwarelsoftuare should be 

itxluded in the design. 

4.4.3.8 Systers Monitoring and Control 

The basic monitoring and control concept is for hardware (internal and 

external 1 to be control led and monitored by a dedicated C&T processor1 

control ler. The CAT processor will interface with the Space Station 

regional data bus. The CLT processor wild perform system level tascs via 

a local system data bus dedicated to the C&T qystxm. The local data bus 

will be routed thrwg3wt the Space Station because of the wide 

distrfbution of C&T hardware. The processor tasks include LRU 

configuration control and verification, LW status, fault monitoring and 

distributing control information to the au&o and video distribution 

control centers. The CiiT processor will forrat information to be 

distributed to the data management system, such as LRU operational 

status, system configuration, etc. 

Local processing within each LRU will be rquired to decode and pmess 

crnds fn the CiiT processor. Status/fault detection will be 

perfor.ed at the LRU level. 

4.4.3.9 Antenna Systems (see Sections 4.3.6.2 and 4.3.6.3) 
4.4.4 Information and Dhta Management 

4.4.4.1 Introduction 

4.4.4.1.1 Overall concept 

4.4.4.1.1.1 Distributed processing approach 

The Space Station Information and Datz Management Subsystem (IDMS) will employ 

a distributed architecture that allocates processing services resident in each 

subsystem. The IDMS will consist of two elements: a core (housekeeping) 

services element providing data distributim, display and control , time and 

frequency reference, and short-term core data storage services to subsystem and 

payload users in the space and ground environment; a user services element to 

receive and transmit user data, without preprocessing of this data, in the 

space and ground env i ronment. 

The IDMS concept will center around an Optical Data Distributi~n Network 

(ODDtiet) through which each subsystem or payload transfers infomation to/from 

its owr! effectors and sensors and to other subsystems (e.g. Communications and 

Tracking) ( see Fig. 4.4.4.1-1). Each subsystem will use standard Subsystem 

Data Processors (SDP's) runniog applications programs developed by each 

subsystem. The SDP's will interface to the LDDNet through Interface Devices 

(ID'S) which will provide all inter-and intradata transfers between and within 

the modules. The Connection Device (CD) ai lows the ODDNet to be distribbted 

between modules. Subsystem data may be input into the IDMS through the ID if a 

SDP is not required. 

4.4.4.1.1.2 Growth 

The distributed processing approach nl lows for the evolutionary growth of the 

IDMS to support the incremental expansion of the Space Station. The addition 

of new modules should easily be accommodated by adding more SOP'S and ID's to 

the ODDNet. As new technology increases the capability of the IDMS. elements 

may be replaced with higher technology elements which will offer greater 

capability with decreased size, weight, and power requirements. 

4.4.4.1.1.3 Commonality 

The IDMS concept will employ comnon elements in the form of ID's, SDP's, and 

executive software overhead distributed through all subsystems and modules. 

This level of commonality extends to both the co-orbiting and polar platforms, 

as well as the Spa-e Station and the ground. 

4.4.4.1.2 Environmenta; considerations 

The IDMS distributed concept is based on the assumption that cosmic rays and 

charged particles will affect the performance and longevity of the various IDMS 

re1 ated lata hand1 ing devices. This, together with the Earth-based, and Space 

S tatio~ operations re1 ated, induced EM1 environment drives the redundancy, 

maintainability, and processing concepts. Specifically, error protection, 

measures to protect against error and fault propagation, and diversity of 
location will be fundamental parts of the distributed IDS concept to protect 

against the anticipated environmental effects. 

4.4.4.1.3 Data protection 

The IDMS concept will be to protect data by partitioning subsystem processing 

to individual, subsystem resident devices. The ODDNet will prevent 

unintentional radiation of Space Station data and precl ude interference by 

intentional and unintentional sources. Relying then, on encryption/decryption 

techniques for protection of the RF link input/wtyut interface with the IDUS, 

the IDnS will be assured of operating in a hostile enviroment without upset. 

Those users requiring protection of proprietary or classified Qta/functions 

will be required to provide the essential barriers that preclude unauthorized 

data access. 

4.4.4.2 IDMS Architecture 

4.4.4.2.1 Conwan element 

4.4.4.2.1.1 Optical data distribution network (ODDNet) 

411 information and data transfer between and within Space Station modules will 

be niade via the ODDNet, a network of optical fiber cables convieniently 

accessable everywhere in the Space Station. The architecture of the ODDNet 

will be designed to i ncl ude sufficient redundancy, or fault tolerant qua1 i ties, 

to be consistent with Space Station requirements. This high capccity data 

transfer medium will handle all the information needs of the Space Station 

Initial and growth configurations. 

ODDNet cmpatibility and connection will be accomplished by ID's and SOP'S. 

4.4.4.2.1.1.1 Interface device (ID) 

ID's will be used to make all connections to the ODDNet providing all interand intrasubsystem/payload data transfer between Space Station modules and 

within nodules where short cable runs are obviated. IDS may also be used where 

the ODDNet is available, to make subsystem connections to remotely located 

sensors and effectors. Connection to the ID will be by electrical connector 

from the SDP's. Users with special high speed data requirements may be 

provided with optical ccnnections to the network. Such users would then be 

provided with an ID to be located within their equipment enclosure. 

Each ID will contain an optical transceiver and the necessary processing 

electronics to provide coup1 ing to the ODDNet consistent with its fault 

tolerant design. IDS are commonality items, each designed for a single 

subsystem/payload user. Internal fault to1 erame within the ID will facilitate 

redundant system connections. 

4.4.4.2.1.1.2 Subsystem data processor (SDP) 

Each subsystem/payload user will be responsible for its own computational , logic and storage needs using a SDP composed of standardized modules or slices 
\ fl 

comnonali ty items). A standard sl ice (to be a part of each subsystem element 

with direct connecti~n to the ODDNet) will format, code and/or sequence data 

for transfer compatibility with the network. Additional comnon element slices 

will be dzvoted to subsystem peculiar processing, computation, storage etc. as 

necessary. Specialized processing outside SDP capabil i ties will be the 

responsibil ty of the user. 

4.4.4.2.1.2 Multi-purpose appl ications consoles (MPAC1s) 

The Multi-Purpose Applications Console (MPAC) will provide the madmachine 

interface onboard the Space Station. These MPAC's shall contain 

multifunctional display screens and controls to assist the crew in their duties 

and to alert the crew of any catastrophic failures. 

There will be two types of MPAC's, fixed, and portable (see Fig. 4.4.4.2-1). 

The difference between these Go types is that fixed MPACs will be used for 

routine operations whi 1 e the portable MPAC 's wi 11 be designed to handle 

operations located away from fixed units, such as maintenance or operations 

where direct outside viewing through a window is desired. Both MPAC types will 

have mu1 ti functional display screens and programmable controls. There should 

be no dedicated displays on the Space Station MPAC. 

Resident in the fixed MPAC kill be the capabilty to print data and graphics 

from the display screen. The crew will have the capability to plot timed 

events data which will be selected from the MPAC. Tne operator will also be 

able to choose between raw and processed data. In addition, a method for 

recording video images will be provided. 

The design of the Space Station MPAC must ta~e into account the zero-g 

environmental effects and astronaut positions. Granted that a local \.-?rtical 

is desired, a one-g rigidity in the design may not be desired. For example, 

the d-isplay screen may be positioned to any astronaut orientation. 

4.4.4.2.1.3 Software 

4.4.4.2.1.3.1 Network operating system (NOS) 

The IDMS NOS will consist of a set of common software provided to all 

subsystems to perform the distributed protocol, formatting, data handling, 

control, and display functions common to a'll subsystems as determined by the 

IDHS architecture. The NOS basis will depend on a common High Order Language 

(HOL) and common processor (slices) to be successful. The approach will be to 

base the NOS on a highly structured concept that partitions f~nctions based on 

comnuni~ation between functions at a high level. 

4.4.4.2.1.3.2 Operational 

Software for operations, conceptually, will consist of a subset of the NOS 

where routine data acquisition, formatting, health and status reporting, 

control I/O, and maintenance functions occur as common elements of all 

subsystem software. 

453 
4.4.4.2.1.3.3 Applications 

Application software commonality will exist through the IDMS concept of a HOL 

interpreter and progr,m writer. Each subsystem will be able to write 

appl ication software in the standardized HOL. The interpreter will convert the 

application software to the structured HOL to be caapatible with the NOS. 

4.4.4.2.1.4 Standards 

The use of standards will be consistent with the established cwnonality 

requirements. Standards will minimize the types of spares used on Space 

Station as well as the types of interfaces. This will allow the use of 

standard test equipment as well as a "remove and replace" philosophy. At all 

possible times, standards wi 11 avoid mu1 tiple, functionally equivalent types of 

hardware. 

For example, one area of standardization will be MPAC's. The same hardware 

will be used on the fixed MPAC's for displqys and controls, even thcugh the 

functions of the displays and controls will vary. In addition, the syllbology 

of the displays should be standardized, such as a flashing red number would 

indicate a sensor out of limits. Another area to benefit from standards will 

be software. Reducing the number of languages will allow consol idation of 

training, compiler development, and compiler maintenance. 

4.4.4.2.1.5 Government-furnished equipment (GFE) 

Subsystem and Payload users will be provided a portable MPAC with the necessary 

ID's, SDP's and a GFE test set. The purpose of this GFE will be to allow 

system level access and control to the Subsystem/Pqyload at the Vendor's plant 

and to assure ODONet and IDMS interface compatibility. The test set will 

provide only those services necessary to functionally simulate the IDMS. These 

simulations will include the ODDNet, Time and rrequency functions, Data Storage 

and the NOS protocol. 

4.4.4.2.2 Core subsystems (housekeeping) 

4.4.4.2.2.1 Subsystem partitioning 

The objective of subsystem partitioning is to mi~imize the nunber of interfaces 

by grauping (partitioning) like, interrelated functions into subsystegs. IDMS 

partitioning will be consistent with the distributed, stand alone processing 

approach described In section 4.4.4.1.1.1. Subsystem users will interface wfth 

the ODONet by way of ID's and SDP's (sections 4.4.4.2.1.1.1 and 

4.4.4.2.1.1.2). Access and principal operation will be accomplished using 

MPAC ' s . 

4.4.4.2.2.2 Core services 

4.4.4.2.2.2.1 Data storage 

Adequate mass storage capabil ity will be provided by the IDMS for the 

collection of Station 1 eve1 status concrrning subsystems, inventory, and 

proximity activities. General Space Station capacities, margins and lir~iits for f 
the purpose of records, trend analyses and distibution will be stored as 

needed. Periodically, stored data will be down1 inked to the ground for 

archival purposes. 

4.4.4.2.2.2.2 Time and frequency reference 

The IDMS will provide time and frequency references, distributed to all 

subsystems simultaneously. The NOS time tags all subsystem 1/0 transactions 

using the d; stributed time and frequency reference to assure time correlation 

of all data. The distributed frequency reference provides a stable source for 

a1 1 subsystems an4 pay1 oads. 

4.4.4.2.2.2.3 Facilities manaaement 

Facil ities Management is an IDMS service that will be available to all 

subsystems and payloads. This service will be responsible for the manaagement 

and sequencing of IDMS data storage and access and Station level data analysis 

and ODDNet management. In additicn, this service will be responsible for 

Station level configuration management incl uding logistics, resource 

a1 1 ocation, and schedul ing data for the crew's use. Tutorials for refreshing 

the crew on procedures will also be provided from t5is service. 

4.4.4.2.2.2.4 Data Base Mdnagement 

A Data Base Management System (DBMS) will be provided to support the convenient 

and effective storage, exchange, manipulation, and retreival of data by all 

appropriate Station subsystems and users. This DBMS will operate within the 

IDMS NOS architecture so as to provide multiple, concurrent access to onboard 

data with bandwidth and response time characteristics sufficient to support 

Station requirements. The system will require minimal manual administrative 

support, thereby providing such automatic features as recovery from machine jnd 

user errors; up-load and down-load of data with appropriate ground facil i ties; 

elimination of redundant and insignificant data; and data partitioning and 

integrity protections. Additional features to be provided by the OBMS include 

mu1 tip1 e views of data; temporary workspaces for data manipul dtion; a natural 

language mechanism for on1 ine query from the MPAC's; and a HOL interface for 

operations appl ications usage. The system wi 11 incorporate advanced data 

base managemer,t techniques, such as the use of the relationai data model, which 

support data element types necessary to scientific, engf new ing, process 

control, Station housekeeping, and other tasks which must be supported by the 

IDMS. In particular, the system must support a full set of geometric entities 

for use in the MPAC display of three-dimensional objects, animation, CAD, and 

two-dimensional graphical datd represen tdtions . 

4.4.4.2.2.3 ODDNet management 

A distributed processing concept will be used for the management of the 

ODDNet. Protocols and priorities may be preprogrammed either in the ID or a 

common module of the SDP. Override of selected functions may be cmnanded by 

the Facil ity Management unit or manl~ally via a MPAC. 
4.4.4.2.2.4 User interfaces 

As in subsystem interfaces, user systems will interface with the ODONet by way 

of ID's and SDP's. Access a;td principal operation will be accmpl ished using 

MPAC ' s . 

4.4.4.2.3 User accommodations 

The IDMS will provide a set of standard terminals and displays (MPAC'S) and 

access devices (ID'S and SDP's) which will satisfy the requirements of mst 

users. However, some users will have special requirements beyond that provided 

in the standard system, e.g., unique, dedicated processors and sensors, 

encrypt ion devices, research instrumentation and preprocessors, 

payloads/experiments, special-purpose controls, etc. In cases such as these, 

the user is expected to qrovide for his specialized needs, and he is encouraged 

to use the comnon module and ID's wherever possible in developing his system. 
THERM SDP D 

A 

-II 

SDP c(t, A 

ID 

D ' I 

GN&C SDP S 

T 

- R . . MECHA I 

NlSMS U 

TIME AND 

SDP N FREQUENCY E . T 

1 

SDP 

TRACK 

DATA 

ORBITAL 

AND 

Figure 4.4.4.1-1 - INFORMATION AND MANAGEMENT SUBSYSTEM 

SPACE STATION 

457 

+-

WRKST'N 

. FACIL.ITIES 

MGMT AND . CONTROL 

I 

W 

SDP 0 

R 

, 

PORTABLE 

WRKST'NS 

COM & CNTRL 

WRKST'N 

K 

SDP SDP 

SDP 

- 

LAB 

1 

WRKST'NS 

LAB 

2 

WRKST'NS 
YULTO PURPOSE r\PPCICATIONS CONSOLE WORK STATION 

BASIC TeRMlNAL 

24 FLAl PANEL 

SCREEN 

MULTl WINDOW 

TOUCH TONE PAD 

EIGHT BALL 

MULTl PURPOSE APPLICATIONS TERMINAL 

(PORTABLE TERMINAL) 

[lo LBS.] 

(20 WATTS] 

Figure 4.4.4.2-1.- WORK STATIONS 

9 MULTI WINDOWS 

REPLICATION FOR EXPANSI

@ VIDEO/DlQlTAUTEXT/QRAPH

FLAT PANEL SCREEN 

LOW POWER 

HIGH RESOLUTION 

9 MAIN CREW INTERFACE 

9 BUlLDlNQ BLOCK 

CAN BE USED BY CORE OR 

VOICE I/O 

DETACHABLE KEYBOARD 

PORTABLE TERMINAL 

MlJLTl WINDOWS 

SAFE HAVEN 

MAINTENANCE 

CREW TRAlNlNQ 

ENTERTAINMENT 

FLLXIBILTY IN LOCATINO T

SUPPORT 0PERATIO)URISE

O OFE TO SUBSYStLY CONTR

IN DDTIE, SOFTWARE DEVE

USERSIPAYLOAD OLVCLOPE
(-. . 4.4.5 Propulsion 

-~ \ fl 

4.4.5.1 Introduction 

The propulsion subsystem concept selected for Space Station conf!yuration 

evaluation was monopropel 1 ant hydrazine, which represents state-of-art 

design, low initial cost, and modest performance. This concept is 

described herein for the reference configuration IOC Space Station. 

The impact on several Space Station configuration a1 ternatives for thi s 

base1 ine propulsion subsystem was examined. The propulsion subsystem was 

developed only to the extent necessary to estimate physical 

characteristics (weights, volumes, etc. ) and to identify significant 

design and opsrational features that would bear on the relative merit of 

Space Station configurations. Configuration impact comparisons focused 

on Space Station buil dup and operational scenarios, propel lant resupply 

strategy, hardware commonal i ty and maintainability . 

Propulsion subsystem impulse requirements for the Space Station , 

configuration were provided by the GN&C group. For these comparative 

studies it was imp1 ici tly sought to minimize the use of propulsion for 

control. Propulsion was assumed required only for altitude 

mai ntenancelreboost, re1 ief of torques that 1 ie outside the capabi 1 i ty of 

the momentum management system, coll ision avoidance, and backup to loss 

of prirnary attitude control (CMG's). Primary attitude control was 

assumed to be maintained with control moment gyros and magnetic torque 

devices. No detailed evaluation was perfcrmed at this time on propulsion 

evolvability for growth station, propulsio~ strategies other than those 

provided by the GN&C group, or propulsion s~bsysi~~~s other than bl owdown 

hydrazine thrusters (75 to 25 1 bf 1. 

Other propulsion concepts may be more benefic .a1 to the overai 1 Space 

Station program by offering reduced resupply costs, integration with 

other subsystems, the abil i ty to efficiently dispose of unneeded 

consumabl es, the enabling of more efficient strategies for forn,?tion 

flying with platforms, reduced contamination, ana reduced overal i 

1 ife-cycl e costs. Representative of such subsystems are hydrogen-oxygen, 

warm hydrogen, and mill ti fuel , 1 ow-thrust resi stojets. Additional 

comments are provided in paragraph 4.4.5.5. 

4.4.5.2 Subsystem Requirements 

The propulsion subsystem speci fic requirements are intimately re1 ated to 

numerous other Space Station subsystems and functions. Typical of these 

requirements are those arising fraom the GN&C, operations, resupply, 

commonality, re1 iabil i ty, maintainability, pay1 oad accommodations, 

contamination, and safety. 'ihese requirements appear throughout the RFP 

and are col 1 ec ted here for c~nveni ence. 
Propulsion reauirements -- 

The propulsion s~ubsystem Is required to perform the a1 titude maintenance 

and attitude control maneuvers. However, for the reference 

configuration, the CMG's and magnetic torquers are assunred to be the 

primary control effectors for performing the attitude control function. 

Therefore, the propulsion requirements were assumed t o f a l l into tuo 

major categories: those that can be done only by the pcopulsion 

subsystem, and tttose in which the propulsion subsystem operates in the 

event that another subsystem fails and a backup capability is reqirired. 

The primary requi: rements are 

( 1 ) reb~os t and orbi t adjustment 

( 2) col 1 i si on avoidance 

(3 ) accomnodation of di sturbances resulting from docking, 

berthing, and movement of objects by the ~~nipulator 

(4) maintenance of an adequate propellant reserve margin 

The backup requirements are 

(5) three-axis stabilization in the event of failure of 

momentum exchange devices (in this event, control is 

provided at a reduced level ) 

(6) desaturation of the mntentum exchange devices 

Guidance, Navigation and Control requirements 

The GNK requirements and strategy dictate the propulsion subsystem 

design and operation. The input for the GNM: comes from the operational 

requirements dnd the payload accomnoda tion needs .. Requirements 

applicable to propulsion are 

(7) a1 1 credible failures of critical systems shall have 

fail operational/fail safe/restorable levels of 

redundancy except for pressure vessels 

(8) selective or total inhibits shall be available for EVA 

operations 

(9) propellant usage data base shall be provided 

The operational requirements pertain t3 restrictions and needs based on 

the overall Space Station. Those requirements pertinent to propulsion 

are 

(10) the Space Station shall not perform ary maneuvers &ring 

docki ng/berthing operations 

(11) maintenance operatiuns including resupply shall consider 

EVA a 1 imi ted resource 

f 

460 
(12) maintenance operations shall minimize the need for 

specfal skills such as soldering and welding or time 

con:dmi ng procedures 

(13) removal, repair, or replacement of equip-lent shall be 

dt the ORU level without the need for special fixtures 

( 14) subsystems shall be as functional ly Independent as 

possible to facilitate maintenance 

(15) subsystems shall be automated to the fullest extent 

possible 

(16) the program shall support the rapid assimilation of new 

techno1 ogy without requiring major redesign or 

reval idation 

(17 1 the station will be resupplied by the Orbiter on a 90- 

day basi s 

4.4.5.2.4 Customer accommodation reql~i rements 

One critical customer accommodation requirement that impacts the 

propulsion subsystem i s 

(18) the steady state gravity level of the station shall be 

designed for .00001 g's or less 

4.4.5.2.5 Svste~ reauirements 

Requirenents ?laced on all subsystems are 

(19) SSPC's shall have t!-e ability to remain operational 

indefinitely through periodic maintenance 

(20) onboard spares shall be provided 

(21) maintenance of ORU's shall tlot introduce hazardous 

condi ti ons 

(22) contami ndnts from ext~rnal sources shall be 1 imi ted 

Safety requirements 

Safety requirements that have a particular impact on the propulsion 

subsystem are 

(23) materials in the habitable volumes shall not outgas 

toxic constituents 

(24) toxic fluid storage or lines shall +e external to the 

pressurized volumes 
;.4.5.3 Propu ;ion Subsystem Physical and Functional Description 

The propulsion subsystem consi sts of four thruster clusters, multiple 

prcpsllant storage tanks and a propellant supply system. A schelaatic of 

the propulsion subsystem is shown in Figure 4.4.5-1. 

The propellant tanks util ize a 3 to 1 blowdwn pressurization scheme. 

The nitrogen pressurant gas is separated fmn the hydrasine by a ysitdve 

expulsion diaphracp. The four thruster clusters are located in a pattern 

that accmmdates the varying locations of the Space Station center of 

gravity. Thruster cl us'ler 1 ocat;nr,s are shown Cn Figur~ 4.4.5-1. 

Tnruster firing directions, shown in Figure 4.4.2-2, provide toques 

about the three axes of the Space Station and translational capability 

along tte x-axis. 

The fa1 I -operational/fail -safe red~ndancy requirement dictates three 

tnrusters in each firing direction on each of the four thruster cluster 

modules. This results in a total of 36 thrusters. The 1 ine rooting and 

isolation capability fn. these multiply redundant thrusters and a method 

of stacking three identical cluster modules for triple redundancy are 

shown in Fiyre 4.4.5-2. 

In addition to installation c;f multiple thruster. in Eacn c:' .-a for 

redundancy, considerat~on must be given to thruster life and .ssion 

for spares. The x-axis thrusters used for peridic reboost will likely 

reach eild of life conditions first. One wqy to accommodate this 

circurnsta?cz would be to provide in-place spares, that is, addi tion~l 

thrur? rs :n tile x-direction at each 1 ocation. 

iquise requirements and propel lant quantities are shown in Table 

4.4.5-1. The impulse requireme;~ts for reboost and for Orbiter docking 

diAurbances must be met at 90-day intervals; thus propellants are 

resuppl led bt each Orbiter visit. This ql !ntit; i: approximately 2,300 

pounds. The contingency impulse requirements for a1 ti tude transfer, 

-01 lision avotdance and backup attitude control must be mt only if the 

sl tuation a* iscs. Thus propellants must be maintained onboard for that 

purpase. This quantity is approximately 4,700 pounds. The 90-dqy 

resupply propel 1 -st quantity sized tka resupply tank volume; the 

contingcr~cj prop,llant quantity sited the onboard tan^ volume, as shown 

in Table 4.4.5-2. 

Pro~a1s:on subsystem performance is a1 so shown i~ Table 4.4.5-2 a1 ong 

with other signifi -ant. system features and characteri stlcs. A single 

thrust level of 25 1 bf minimu~n was selected based on control. This 

provides siln~l ici ty of the control syscem design, mjnimum hardware 

{nventory and sPirropr I ate performance for tile varied requd rents. The 3 

to 1 propellant pressurization ;,lowdown ratio resAt?tc in a maximum thrust 

of 75 lbf per tkru~te~. 
LO'rNER LEFT 

CLlJSTER 

UPPER 

LEFT 

UPPER 

RIGHT 

CLUSTER 

-- 

TOP VIEW OF UPPER 

RIGHT CLUSTER 

SHOWING LINE RO'JTING 

(TYPICAL 4 PLACES) 

KEEL MOUNTED 

HYDRAZINE SUPPLY 

INTERFACE 

LOWER RIGHT 

CLUSTER 

Figure 4.4.5-1 - REFERENCE CONFIGURATION PROPULSION SYSTEM WITH 

APPROXIMATE EQUIPMENT LOCATIONS 
FEATURES SHOWN: 

TRIPLE REDUNDANT INDEPENDENTLY r.EPUCABLE MODULES 

PLUMBlh!G ARRANGEMENTS 

DISTRIBUTION SYSTEM DISCONNECTS 

DISTRIBUTION SYSTEM ISOLATION VALVES 

THRUST 

CHAMBER (nP)<P r D~STR~BUTION SYSTEM Is01 

DISTRIBUTION 

ATION VAL 

X SYSTEM DISCONNECT 7 

THRbSTERSI 

TOP VlEW 

SINGLE CLUSTER MODULE I 

)WING INTERN4L PLUMBING 

SIDE VlEW 

CLUSTER )$, STACK OF THREE CLUSTER 

E It1 

MODULES MODUL, .. ., 

MODULE #2 

CLUSTER 

MODULE #3 

DISTRIBUTION SYSTEM #1 

DISTRIBUTION SYSTEM #2 

+% DISTRIBUTION SYSTEM #3 

STACd OF THREE CLUSTER MODULES 

Figure 4.4.5-2 - TYPIC ' THRUSTER CLl ER 

464 
A detailed weight sunmary is shown in Table 4.4.5-3. These data are 

presented separate,, to show the propulsion subsystem component weights 

which are located on the station lower keel and those located in the 

Logistics Module (LM). Also on this table, total dry system weights and 

wet weights are shown for both the station and the LM. 

A summary of volume and average power usage for the propulsion subsystem 

is shown in Table 4.4.5-4 for those system components which occupy 

signi ficant volume and/or demand quanti fi?bl e arnounts of average power. 

A shrouded cluster of six propellant tanks, located cn the lower keel, 

was zstimated to require an average power of 35 watts for thermal 

stability. The same value was used as an estimate of the thermal 

requirement for the cluster of three LM tanks. An average power of 50 

watts was estimated as required for the catalyst beds on ~ach of the four 

thruster clusters. Distribution line hedting requirements were 

estimated, based on Orbiter experience, to be 640 datts average power for 

the 1,600 foot Station run, and 20 watts for the 50 foot run on the LM. 

4.4.5.4 Propel 1 ant Storage, Di stribction, Management and Resupply 

4.4.5.4.1 Prooellant storase 

The total quantity of propellant required for the Space Station 

propulsion subsystem reference configuration is 7,050 1b of hydrazine 

(NzHq). The propellant requirement is split between contingency 

propellant, 4,730 1 b, and 90-day resupply propellant, 2320 1 b. The 

pr,pulsion subsystem will also contain a total of 57 cubic feet (390 It) 

of gaseous ni trogen (GN2) pressurznt. 

The tanks selected for the reference configuration are spherical 

elastomeric diaphragm tanks 4G inches in diameter. The selected tanL 

she1 i material is 6A-4V titanium. Six of these tanks contain the 

contingency propel 1 ant and three addi ticnal tanks contain the 90-day 

resupply propel 1 ant . 

The contingency tankage (six tanks) wlll remain with the Space Stdtion. 

These tanks will be located near the station center of gravity envelope, 

at the interface between the lower keel and the keel extension. The 

contingency tankage will be packaged within the envelope of one of the 

9-foot cubes of the lower keel truss structure, near the X-axis 

centerlir~e. The tankage is packaged such that the diagonal brace struts 

of the Space Statfon truss structure will not preclude removal and 

replacement of individual tanks or tankage modules for maintenance or 

contingency ;.'rposes. Tanks are also arranged as a pair of three-tank 

modules to facilitate maintenance, as required. These modules should 

incorpot-ate Space Station Mobile Manipulator/Orbi ter RMS grapple fi ttings 

to facilitate handling and assembly. Individual tanks within a 

three-tank module incorporate propel 1 ant/pressurant quick-di sconnec t 

fittings. Structural /mechanical connections shcul d also minimize 

connectionldi sconnection time. As an a1 ternate, propel 1 ant tanks could 

be located closer to thrusters to facil i tate Space Station 

launch/assembly . 465 
Table 4.4.5-1 REFERENCE CONFIGURATION IMPULSE REQUIREMENT

CONTING

IMPUL

(LB -SE

61,50

831,00

- 

1 47,00

- 

- 

- 

- 

- 

1,039,50

TOTAL AVAILABLE 

90-DAY 

RE-SUPPLY 

IMPULSE 

(LB-SEC) 

- 

- 

483,000 

- 

26,000 

1 

509,000 

- 

ANGULAR 

MOMENTUM 

(FT -LB -SEC) 

- 

- 

- 

10,000,000 

150,000 

1,500,000 

600,000 

-- 

- 

1,548,500 LB-SEC - 7,050 LBS-N2H4 

TRANSLATION 

REQUIREMENTS 

* COLLISION AVOIDANCE 

AV = 5 FT/SEC 

ALTITUOE TRANSFER 

20 N. MILES 

RE BOOST 

90 DAYS AT 270 NM (LAIN) 

ROTATIONAL 

REQUIREMENTS 

ATTITUDE C0NTRC.L ' 24 DAYS - Z AXIS ONLY 

* ORBITER CAPTURE 

ORBITER BERTHING 

- 

MODULE BERTHING 

N2H4 RE SUPPLIED EACH 

90 DAYS (LOCATED IN 

LOGISTICS MODULE) 

N2H4 CONTINGENCY 

(LOCATED ON STATION) 

90-DAY 

RE-SUPPLY 

N2H4 

(LBs 

- 

- 

2,200 

- 

120 

I 

2,320 

- 

MOMENT 

ARM 

(FEET) 

- 

- 

- 

68 

88 

88 

88 

- 

- 
Table 4.4.5-2 - REFERENCE CONFIGURATIQN PHOPULSION SYSTEM FEATURES 

PROPELLANT 

'SPEFIC IMFULSE (LB -SECILB ) 

TDTAL IMPULSE (LP -SEC) 

>234 (LBs 

N2Y4 VOLUME (FT 3, 

PRESSURANT VOL (FT 3jZ~~ 

'TOTAL TANK VOLUME (FT 3, 

NUMBER OF 40" DIA TANKS REQ'D. 

'VOLUME CF ONE 40" DI.4 -ANK (FT 3, 

NUMBER OF THRUSTER CLUSTERS 

NUfL?BEF( OF ; HRUSTERS PER CLCJSTER 

TC!TAL NUMBER OF THRUSTERS 

THAUST LEVEL (MIN-MAX) (LBS) 

TANK/SUPPLY PRESSURE (MAX-MIN) (PSI) 

PRESSURiZATlON SCHEME 

THRUST MOMENT ARM-Z AXIS (FT ) 

' THRUST MOMENT ARM-X AXIS (FT ) AVG. 

THRUST MOMENT ARM-Y AXIS (FT ) AVG. 

STAT I0 N 

'"254 

220 

1,039,500 

4,736 

7 6 

38 

114 

6 

19 

4 

- 9 

36 

25-75 

300-1 00 

BLOW DOWN 

68 

162/2 

162/2 

LOCATION 

LOG MODUL

N2E4 

220 

509,009 

2,320 

38 

19 

57 

3 

19 

0 

0 

o 

- 

- 

- 

- 

- 

- 

k 

SSO' 

- 

3A09V 

31~I91193N 

31fl191103N 

33s. 

- 

.OZO' 

-0- 

.S60' 

3inao~ 31~~1007 

SL8' 

- 

3AOBV 

319191193N 

318191193N 

33s- 

- 

*Otr9' 

(~4) U3MOd 

S'Z 

Z'P8 

- 

2' 

1 

1 

1 

(€I=/) 3VVfllOA 

€6 1 

€ 

1 

09 

1 

6 

\ 

. 

.OOZ' 

SEO' 

~33~ u3~o-1 

S

- 

Sl

3tlnl3nHlS

S3A

NOll

lOHl

~3~

911 

91 

733~ -- H~MO~ 

L 

-0- 

S 

ilnaovu 311str301 

t331S

3E)V)1NV
The 90-dqy resupply tankage (three tanks) will be located on the LM, 

which is relatively close to the contingency tankage. The LM will remain 

connected to the Station between Orbiter resupply misslons. The nitrogen 

purge gas, if required, will be located within the LM. 

The 90-day resupply tankage will be packaged within the end section (aft 

skirt) ot the LM, which is the end opposite fm t;,e docking adapter. 

Ideally, this resupply three-tank module will maximize the nutnber of 

components which are common with the contingency tawcage thr.ee-tank 

mdules. The resupply tankage module a1 so incorporates 

propel 1 ant/press~lrant resupply quick-di sconnec t interfaces. These 

interfaces a1 1 ow connection to the propulsion subsystem propel 1 ant 

distribution system. 

Propel lant Storage Interfaces: The propel lant storage 8 d i s t r i b u t i o n 

systems will interface with the following SS subsystems: 

Structural idechanical ; Propel 1 antlpressurant Mani fol ds; Power; Guidance, 

Navigation and Control (GN8C) ; Data Management System (DMS) ; and the 

Thermal Control Subsystem. Thi s reference co~fi guration (Figure 4.4.3-1) 

provides the option for interconnection with other compatible hydrazine 

propel 1 ant tankage, such ds the OMV or sate1 1 i te reservicing propel 1 ant 

tankage. 

4.4.5.4.2 Propellant dl stribution 

The propellant distribution system of the Space Station reference 

confi gtiration propulsion subsystem is designed to provide and control the 

flow of propellant from the prope; 1 ant storage tanks to the thrusters. 

It is also required to provide the fluid system for resupply of the 

propulsion subsystem storage tanks. The propel 1 ant di stribution system 

will also provide propelllnt interconnect capability with other hydrazine 

storage systems onboard the Space Station. Propellant feed 1 ines are 

routed from the centralized propellant storage tanks to the four thruster 

clusters at the end of each cluster boom. A resupply propellant fill 

1 ine system connects the Space Station central tankage to the LM 

propellant interface via the central Space Statign propellant manifold. 

Triple redundant lines (fail operational/fail safe) have been selected 

for this reference configuration. The three line runs will be separated 

by an optimum distance to preclude sim~ltaneous damage to lines. The 

reference configuration has been base1 ined to utilize hard1 i nes for 

propellant distribution. An a1 ternati ve would be to use convoluted 

stainless steel flex-hoses. However, it is felt that the addirional 

pressure drop and additional weight would be unacceptable, considering 

the relatively 1 ong 1 ine runs. Tef :on-1 i ned (smoth-bore) fl ex-hoses 

could be utilized, but long term propellant permeation could be a 

problem. 

The selected line is 3/4 inch O.D. stainless steel tubing. The total 

propel 1 ant dl stri bution system I !ne length is apprax4mately 1600 feet, 

including redundancy. The longest thruster feed 1 ine run is 

approximately 150 feet. The selected system pressure range is 300-100 
psia for the three-to-one bl owdown systea~. Propel 1 ant distribution 

system pressures wi 11 be slightly higher to account for pressure drops. 

A1 so, re1 oadi ng of the di aphraqm propel 1 ant tanks wi 11 require pressures 

in t,he 6CO-700 psia range. The line safety factor is in excess of 4. 

The propulsion subsystem reference configuration propel 1 ant di str ibution 

system has been base1in.d to have heaters azd insulation wrap applied to 

the propellant lines to preclude hydrazine freezing. An alternative 

would be to utilize waste heat from other subsysten~s as a heat source. A 

trade study is required to determine the optimum method. 

The reference c o n f i g u r a t i on has a total of 36 propel 1 ant di sconnec ts. 

The LM has 6 and the Space Station has 30. The disconnects are of the 

quick-action type where dppl icable. They are designed to incorporate 

features to preclude mi sconnection. The reference configuration 

incorporates a total of 33 propellant i solatlon valves. The LM has six 

dnd the Station has 27. The valves are of the clectromechanical solenoid 

tv~e. 

4.4.5.4.3 Pro~ellant utilization and manaaement 

The propel 1 ant util ization and management system is des~qned to gage 

propel 1 ant quantities, control intertank propel 1 ant usage, and directly 

or indirectly control and monitor propel 1 ant resupply operztions. The 

operational strategy for the propul sion subsystem reference configuration 

involves feeding propellant to the thcusters from only one prapellant 

tank at a time, with the rest of the propellant tank isolation valves 

closed. When the first tank is depleted, its isolation valve will be 

closed and the yecond propel lant tank's isolation val ve will be erlened. 

In this manner, some cf the disadv,ntagos of the blowdown system kill be 

minimized. The propellant utilization and management system has 

temperature and pressur- transducers located at key points within the 

propulsion subsystem. Since propellant q~anti ty gaging is accomplished 

by thz pressure/volurne/temperature technique, only one tank at a time 

will suffer the inherent propellant gaging inaccuracies. The other tanks 

will have accurately known propellant quantities either empty or full. 

This in turn wi I 1 facilitate resupply operations. Components of the 

Space Station propulsion subsystem will interface directly with the IDMS. 

4.4.5.4.4 Resupply 

The 90-day resupp;y three-tank mcdule will be instilled on the LM aft 

skirt, and the propellant will be loaded at ground facilities. Since the 

90-day resupply tankage is based on the LM, on-orbi t propel lant servicing 

operations are not required for this tankage. Three-tank modules for 

reloading of Station-based contingency tankage wi l! be instal led, on an 

as-required basis, in the LY aft skirt on the ground and loaded 

consistent with the amount of contingency propellant used by thc Station 

between Orbiter 90-day resupply missions. Thus, a LM could contain as 

nany as line tanks f~r a WO~S~ cose re~uppiy mission. Six oi these tanks 

are for Statiun-based cootingency zankage reloading, and three tanks 

fullfill the 90-day resupply tankage requirement. The purge gas, iC 

required, will be installed in the LM aft skirt and loaded at ground 

facilities. 
4.4.5.5 Concerns and Issues 

Space Station: If only +he reference Space Station hydrazine subsystem f s 

considered, the components that mgy be exactly repl i cated include 

tankage, thrusters, valves, and instrumentation. It is more 1 ikely that 

most of these components would be part of a module-1 ike subassembly that 

could be replicated except For mounting. The reference propulsion 

subsystem shown earl ier includes repl ication of tanks and thruster 

modules, with sow exceptions for mounting left and right modules. 

Platforms: Component comnoriality between the Space Station and Platform 

propulsion subsystems is desirable from an initial and operational cost 

perspective. In the level of description available on the two platforms, 

no unique propul sion subsystem requirements were identi f ied that viol ate 

the conmonal i ty possibi 1 i ties discussed. However, requirements are not 

sufficiently defined to assess thoroughly the potential impacts on 

propulsion subsysten~ requi rements. 

OMV: The Orbital Maneuvering Vehicle (OMV) is anticipated to use 

hydrazine and/or monomethyl hydrazinelni trogen-tetroxide propulsion 

subsystems. The OMV may provide an alternative option for Space Station 

reboost and other translational functions. The propel 1 ant storage 

facilities provided at the Station might also serve iis a commn depot for 

the station propulsion subsystem needs. And component cormonal ity night 

be available between the OMV and the station. 

OTV: The Orbital Transfer Vehicle (OTV) and its propellant storage 

facilities may also provide a component and propel 1 ant storage 

comnonal i ty, depending on the propulsion subsystems selected. 

Pqyload: Platform and satellite consumables include hydrarine, liquid 

hydbogen and liqdid he1 ium. The quantities of these fluids, which are 

anticipated to be stored on the Station for resupply of payloads, 

significantly exceed the Station propellant requi renents. Use of these 

fluf ds fcr Station propulsion might enhance resupply logistics. 

4.4.5.5.2 Hardware 1 i fe/replacement - 

Very long life monopropellant hydrazine thrusters must be designed to 

meet the requirements for maintainability, limited crew attention, and 

overall Spdce Station 1 ife. The reboost requirement wiil resalt in a 

very large throughput of propellant for the catalyst bed of those 

thrusters when compared to current fl ight propulsion subsystems. 

Several alternatives are available to meet the life requiremnts. These 

include the installation of in place spares that cwld be functionally 

switched into the system, or the development of improved designs of 

catalyst beds through early t-echn~! nnxy progrmz. 
4.4.5.5.3 Station buildu~ 

The rdass properties (magnitude and distribution) of the Space Station 

will undergo significant change throughout the 1 ife of the stsation, first 

during the assembly of the IOC station and then with the addition of 

attached payloads and servicing facil i ties. Propulsion subsystem 

accomnodation of these changes is an issue. Component selection and 

packaging, station scarring, and ;?stallation and resupply technfques 

will be affected. 

4.4.5.5.4 Growth station 

The growth version of the Space Station will be significantly larger, 

notably as a result of increased power capacity and the expansion of 

servicing and other facilities. There will be an increased demand for 

accommodation of payload and platform servicing, of OTV and OMV 

activities, and of onboard operations. These may lead to control, 

propulsion and propellant management subsystem requirements for greater 

capability and flexibil i ty than tliose anticipated for the initial , 

station. 

4.4.5.5.5 Advanced and a; ternate ~ro~ul si on o~ti ons 

Other propulsion subsystem options must be c0nsiaer.d for both the 

initi?; and growth station to obtain the most effective balance of 

capacil ity and cost. Consideration must be gSven to: 

(a) ability to respond to variations and uncertainty of 

required impulse bi th station growth and expanded 

mi ssion defSni tions 

(b) sirnul ttlnelus a1 ti tude maintenance and CMG 

desatura *.Ion 

!c ) reduced resupply requi rements acd mani festing 

variations 

(d) increased operations flexibility, particularly in 

coordinating act'ivi ties with platforms 

(el reduced operational complsxi ty 

(f) increased component life and maintainability 

(g) integrated fluid systems management 
(h) disposal of unwanted Station fluids 

(i) initial and life cycle costs. 

Options, in addition to the referent? blowdown hydrazine subsysteln, 

include Earth-storable bipropell ant thrusters, hydrogen-oxygen thrusters, 

raul ti fuel resistajets, and thermal ly-augmented subsystems. The 

bipropel lant subsystem offers the opportuni Q to make use of OMV and 

Orbiter propellants, and possibly of cmmon components. The H.0 

subsystem offers clean exhaust and high specific i~ulse, and could use 

cryogen H-0 storage boil off, external tank scavenged-flui ds, or 

electrolysis of water with potential interface with the ECLS or other 

subsystems. The resistojet with a thrust level of less than 1 lbf could 

acconwdate a1 ti tude maintenance and CMC desaturation for the Station and 

platforms rhile maintaining acceleration continuously below .00001. g. 

Platfonn ephemeris could be maintained in spite of the highly variable 

atmospheric perturbations. The resistojet would operate at ?ow 

temperature to achieve long life and could operate on a broad range of 

fluid, inc I uding hydrazine, water, hydrogen, carbon diaxi de. 
4.4.6 Environmenta'l Control and Life Suppart System 

4.4.6.1 Introduction 

The ECLSS described herein provides for partial sy;tc,:t closure for the 

initial Space Station. Partial closure of the ECLSS is based on 

regenerative processes for clobing both the metabol ic oxygen and water 

cycles. The only resupply materials are nitrogen for module and airlock 

leakage makeup and extravehicular activity (EVA) equipment servicin;, and 

food for crew consumption. A completely closed ECLSS is not viable at 

this time because of the high cost and risk of developing food 

regeneration techpol og2. Nitrogen generat ion via hydrazine di ssociat4dn 

could be considered for the Growth Station. The partially closed ECLSS 

is selected because it offers tlie optimum balance between development 

costs, techno1 ogy risk and resupply penal ties for Space Station 

appl ication. 

The partially closed cycle ECLSS shall contain suffic.;er+ mc>dulari,ty an3 

flexibility in design to accomrnodate on-orbit maintainability, repair, 

and evolutionary growth capability. These characteristics are important 

in that the initial Space Station FCLSS must allow for future 

'ncorporation of addi tional capabil i ty ( increast 'n crew s;ze and habi tat 

area) without incurring additional devel oprnent costs. The ECLSS design 

must also be able to adaot to advances in state-of-the-art corlcepts that 

offer iricreased efficiency, process yield, reliability and/or 

tnaintainabil i ty without requiring major a1 terations to e;;i sting ECLSS 

flight hardii:-e. 

4 / .6.2 System Functional Requ i rz-.>nts 

The functions of the partiall;/ closed ECLSS include atmosphere presstire 

and ,omposi tion control , module cercperat.ure and hirmidi ty conti-01 , atmspher~c revi tal ization, water managemeEt, metabol ic waste m~nagenient, 

and EVA equipment servicing. Detailed expansion of th fse requiremeO,ts is 

as follows: 

a. Atmosphere Pressure and Composition Control 

(1) Control and monitor module tatal pressure within 

speci fied 

1 imi ts 

(2) Control and monitor oxygen (02) and ri trogen (Nz) partial pressures 

(3) Provide fire detection and suppression 

b. Module Temper ature and Humi di ty Control 

(3) Coii trol ambient temperature 

( 5 Control humidity 

(6) Provide air ventilation 
f 5) ad mmitor true cortrriauts Md 

r vfpc 

t' (231 1 ec ,, pretmat and process uastz water 

- liask uater (dish uasher .nd clothes washer) 

(2) Post*mat recovered water 

(3) n#litnr wter quality 

(4) Store and deliver recovered wter to water use 

intltfxes 

(1 Col lect urine for water recovery 

(2) Collect and store fecal utter 

f. EVA Equipent Servicing 

(11 Provide 02, water (He), and heat sink regeneration servics to Extrawehicular Mobility Unlt (EMI) 

(2) Provide Hp recharge capabtl ity for Manned Naneuverirg Unit (MI 

(3) Provide 1 ife support services to airlodt/hyperbaric 

frci1iQ 

In fulfilling these requirements for the initial Station, the ECLSS will 

use a remati ve C02 mval pracess for collection of retabol ic 

COp and subsequent delivery to the C02 reduction process. The CO2 

reduction process will reduce C02 to water and carbonaceous products. 

Overboard venting of these carbonacws products will not be permitted. 

After appropriate processing, the C02 reduction effluent water and 
condensate frm *jre Station teqerrture and hurietiw contml process will 

be used as :duble wter for food preparation and dritking purposes. 

Additional water for reprocessing uill be obtained fm hygiene 

(hbmbdash, shmer, ~rine/flush) and housekeeping wter ( di sh/clothes 

washing) sou-xes. A phase change process and pre- and posttrea-nt 

processes will be used to recover usable wter fnr these hygiene water 

wnces. Effluent waste water frar dish and clothes washing dl 1 be 

reclaid by a filtration process. Gxy- ui?l be provided by 

electrolysis of the ;recessed hygiene uater. Yitrogen for the initial 

Space Station will be supplied fnr either cr)o~ic or high pressure gas 

supage. A zontaairiant control process will be incluQd in the EUSS to 

mve airborne trace contaminants.. In addition, a solid waste 

ranagent process will be included for hand1 ing and trea-nt of human 

fecal matter. 

4.4.6.3 Reference Configuration 

4.4.6.3.i System sizina criteria 

The fo:lwifig criteria *re used in sizing and develo~ing the reference 

ECLSS for Space Station: 

a. Ctesrr Size 

I~itial: 6 persons 

Growth: 18 persons 

b. Space Station leakage: 5 lb/day, maximum 

c. Resupply Interval: 90 days 

d. Safe Haven Emergency Capability 

Duration: 22 dqys 

Location: 2 modules, ainimm 

e. EVA equipment servicing 

EVA duration: 16 person-hours/EVA; 5 EVA' s/week 

WU Rechavge: 115 lb N2/week 

f. No dumping of gases (goal 1 

g. Fail Operational /Fail Safe (Restorable) 

Basic ECLSS definition 

A conceptual design of the partially closed ECLSS satisfying the 

requirements identified in Section 4.4.6.2 for the initial and growth 

versions of the Space Station is depicted in Figure 4.4.6.1. The mass 

balance for speci fic regenerative process techno1 ogies is depicted in 

Figure 4.4.6.2 The mass balance numbers are for a 6-persor! crew and all 

flow rates are in lblday. It should be emphasized that specific ECLSS 

REsuPmr 

WATER 

1"'"'"' UNPREPARED FOOD rooo (Bone DRY) 

0.0 

REIUI'PLY 

0.0 7.3 -- -------------- --------------- ---- ------------------.-----

wA"n'" 1 I I 

METABOLIC LATENT 20.0 FECES IOLIDS 1.8 

I CLCEI WATER 1.3 

Figure 4.4.6-2 Reference Partially Closed ECLSS Mass Balanc
* 2 - Three Person Units ** 1 - Three Person Snit (Module) 

1 - Three Person Unit (EVA) 

o Dump and Relief Valves 

TENPERATURE/EIUKIDITY CONTROL 

EARW VIEYIK~ o Heat Exchangers 

o Fans 

Pire Detection/Suppression o Noise Mufflers 

Tap. /Humidity Control ATMOSPHERIC REVITALIUTIO?J 

o Cop Concentration* 

o C02 Reduction* 

o 02 Generation** 

o Catalytic OxidieerlFilters 

o Monitor 

WATER WGEPIENT* 

o Hygiene Water Recovery 

o Potable Water Recovery 

o Storage and Dispensing 

o Monitoring and Collection 

WASTE MANAGEKEN'S' 

o 1 Cde 

o 2 Urinals 

EVA SUPPORT 

o EMU 02. H20, Beat Sink Recwery 

o FQ4U - N2 Supply 

o Suit - Servicing, Cleaning 

o ~irlock/~yperbaric Facility 

Support 

SAFE HAVEN 

o 6 Portable Oxygen Supply (POS) 

o 02 and Hz0 Recovery 

o Ventilation 

o Temperature/Eumidity Control 

o C02 and Contamination Control 

o waste Collection and Storage + 

I ,

I 

! 

! 

I 

l 

I 

j 

1 

1 

I 

I 

I 

1 I

! i

i. i 

Figure 4.4.6-3 PARTIALLY CLOSED ECLSS EQUIPMENT/SPACE STATION LAYOUT 

t 

I 
rocesses were chosen only for the purpose of illustrating ECLSS mass 

k~he baseline processes for U)Z removal, U), reduction, & generation and waste water reclamation have not been selected. 

The Space Station habi tab1 e module configuration used in this reference 

ECLSS is depicted in Figure 4.4.6-3, and the recommended centralized/ 

distributed functional a1 location of the ECLSS is given in Table 

4.4.6-1. A1 though a totally distributed ECLSS (i .e., independent 

subsystems in each Space Station module) initially appears lore 

supportive of the "common module" philosophy, replication of a1 1 ECLSS 

functions in each module involves excessive hardware. Furthemre, such 

extensive rep1 ication of hardware does not use the available module 

volumes effectively. As an example, it is impractical to locate waste 

management and waste water generation and recovery equipment in each 

module. It is also prudent to co-locate the waste water generation and . recovery equipment to minimize interface plumbing and to enhance 

reliabil i ty and maintenance. Therefore, the centralized/distributed 

approach was selected for the reference ECLSS configuration. This 

approach also supports the "conmon module" by providing the following 

basic ECLSS equipment in each module: ventilation fans and ducting, 

temperature and humidity control, module atmosphere dump and pressure 

re1 ief valves, fluid transport plumbing, and fire detection/ 

suppression. Table 4.4.6-2 contains a similar functional allocation for 

the Growth Station. 

Distribution of the ECLSS equipment for each module is defined in the 

Crew Accommodations section, Figures 4.3.2.1-1 through 4.3.2.2-6. Two 

3-person capacity ECLSS units that provide the functions of CO2 

removal , C02 reduction, trace contaminant control and waste water 

reclamation will be centrally located in each of the two Habitation 

Modules. These Go modules, which interface with the EVA airlock/ 

hyperbaric facil i ties, were selected for the centralized location of the 

ECLSS processing equipment to consol idate equipment and minimize 

replicdtion and interfaces. One of the ECLSS units in each of these 

modules will be in an "unpowered" standby mode and the other unit will be 

"operational . " Each 3-person capacity ECLSS unit will be capable of 

operating at a 3-person capacity in normal mode and at a 6-person 

capacity off-nominal mode. The 02 generation function in each of these 

Go modules will consist of a 3-person capacity (low pressure) unit for 

module and metabolic usage and a 3-person capacity (high pressure) unit 

for EMU recharging. Each of these electrolysis units will be capable of 

generating 02 for 6 crewmen (at a higher current densi ty) for 

off -nominal mode operation. 

Air revitalization in the remaining modules will be accompl ished with air 

distribution ducts. Humi di ty and temperature control wi 11 be local ired 

in each habitable module. The ducting and plumbing in each module will 

a1 so include provi sf ons for bypassing each module and be capable of 

isolation from each module. 

The ECLSS weight, volume, and power requirements for the initial Space 

Station are presented in Table 4.4.6-3. The power levels specified fn 
TABLE 4.4.6-1 INITIAL SPACE STATIOll ECLSS FUlCTIOSlAL ALLOCATIW 

Function 

Fire. kt. 6 Supp. 

Cabin Thermal 

Control 

C02 Reduction 

Atmspheric Control 

Potable Yater 

Recl amati on 

Potable Water 

Treatment/Storage 

Waste Managemnt 

Hygj enefiash Water 

Reclamation 4 Storage 

Safe Haven Provi sf ons 

EVA Equipment Servicing 

HM2 

- LAB 1 - LAB2 

- 

X X X 

X 
TABLE 4.4.6-2 

GROWTH ( 18 PERSON 1 SPACE STAT ION ECLSS FUNCTIONAL ALLOC

Function HMl HM2 LAB1 LAB2 LM1 HM3 HM LAB3 LAB4 

- - --- -- 

Fire. Det. & Supp. X Y X X X X X X X 

C02 Removal 

Cabin Thermal 

Control X X X X X X X X X 

C02 R&duc ti on X X X X 

'42 Supply X 

Atmospherf c Control X X X X X X X X t # 

Potable Water 

Reclamation X X 

Potabl e Water 

Treatmen t/S torage X X 

Waste Management X X 

Hygi enefiash Water 

Reclamtion & Storage X X 

Safe Haven Provisio,?~ X X 

EVA Equipment Servicing X X 
TABLE 4.4.6-3 INITIAL SPACE STATION ENVIRO))ENT# CONTROL AND LIFE SUPPORT SYST

Subsystem Func t I on Yelght (LBSI 

Atnospheric Pressure/ 685 

Compost tlon Control 

AGospherfc Revital I- 2506 

zatlon, Ventilation 

Hum1 dl Syflemperature 500 

Control Ventilation 

Water Management 2526 

Waste Management 500 

EVA Servlclng 84 

Safe Haven 420 

Tankage (Dry) 20% 

Flulds 6000 

Volume (FT~) - Power (Wat

Tok? Dry 9351 

Wet 15,351 

~rlnals/cs (intermittent) If a11 used at once. 

-wjj& - ' 

. -. 

z* ' .* I.' t 
Table 4.4.6-3 are for continuous ECLSS operation throughout the orbital 

period. However, cyclic light sideldark side operation of the EUSS .qy 

be a more effective approach in power snapent. Cyclic operation could 

take advantage of the availability of the inexpensiven direct solar 

power during the sun1 it (light side) phase of the orbital period. 

4.4.6.3.3 EVA equiplent servicing 

EMU and W servicing will be an integral function of the basic Space 

Station ECLSS. Replenishrent of EMU and MJ expendables and the recovery 

of waste products will be provided. The ECLSS will supply oxygen and 

water for the EMU'S and reprocess the collected waste water and CO2. 

The EUSS also will provide nitrogen for W propellant system 

recharging, and generate make up gases for air1 ocklhyperbaric facility 

leakage and pressurization as well as atmospheric revitalization and 

cooling for pre-EVA activities. A detailed description of the EVA 

systems is contained in Section 4.4.9. 

4.4.6.4 Conclusions 

Partial closure of the ECLSS using regenerative processes offers the -st 

cost effecti w approach and enhances evolutionary growth flexibil i ty for 

Space Station. Additional developrent costs should be offset by the 

substantial reduction in resupply requirements. Furthemre, locating 

the ECLSS equipment in the central izedldi stributed manner described above 

minimizes the awnt of hardware while accmdating the "c-n mdule" 

philosophy, and fulfills the required redundancy and "safe haven" 

requirenents without addittonal backup equipmnt. For example, the two 

3-n C02 removal subsystems in the two separate mdules el iminate the 

need for a backup 1 i thium hydroxide CO;! mval system. Similarly, the 

redundant electrolysis units eliminate the need for backup emergency 02 

tanks. Only additional -rgency N2 for mdule ledrage lake u: and 

portable 02 supplies and masks for use in mdule evacuation are 

required in achieving a "safe haven". Centralizing the ECLSS also 

localizes noisy process equipment and minimizes waste fluid 1 ine 

1 engths. The dual central ired configuration a1 so enhances system 

maintenance operations and readily accommodates temporary crew overload 

as during crew change out. 
4.4.7 Them1 Control 

4.4.7.1 Introduction 

The fol 1 owing requirements were used for thi s evaluation: 

a. Base station electrical power character1 stics (IOC and 

growth 1. 

(1) Useful electrical power at bus: 75 WE (IOC). 

300 WE (growth). 

(2) Orbital average waste heat load: 95.8 ~WT (IOC). 

393.9 ~WT (growth). 

(3) Sunside waste heat load: 

(a) Customer: 50 ~WT (IOC) and 250 klJ~ (growth). 

(b) Housekeeping: 25 ~WT (IOC) and 50 ~WT 

(growth). 

(c) Power subsystem: 13 ~WT ( IOC) and 53 ~WT 

(growth). 

(4) Dark;ide waste heat load: 

(a) Customer: 50 ~WT (IOC) and 250 ~WT (growth). 

(b) Housekeeping: 25 ~WT (IOC) and 50 ~WT 

(growth 1. 

(c) Power subsystem: 46 ~WT (IOC) and 182 kW7 

(growth). 

b. Manned core waste heat distribution and allowable heat 

source interface temperatures for IOC (Tab1 e 4.4.7-1 ) . c. Manned core waste heat distribution and allowable heat 

source interface temperatures for growth (Table 4.4.7-2). 

The major design assumptions used for this sizing study are as follows: 

a. Heat acquisition and transport by a pumped Go-phase 

"thermal-bus" approach with b ~ter as the transport fluid in manned 

modules and ammonia outside. 

b. Heat rejection by high capacity heat pipe radiators. 
c. No heat load sharing bebeen Station modules (1.e.. a module 

with excess heat rejectlon capablllty fn Its body-mounted radiators 

dries not reject heat from other modules). 

d. IOC housekeeping pomr dfstrlbutfon by Statfon dule. 

(1) HMl - 274. 

(2) H)12 - 27%. 

3 LAB1 - 29%. 

(4) LAB2 - 12%. 

(5) LH - 5%. 
TABLE 4.4.7-1 - MANNED CORE WASTE HEAT DISTRIBUTION AND ALLOWABLE 

SOURCE INTERFACE TEMPERATURES FOR 10C 

*ORBITAL AVERAGE LOADS BASED ON AN ELECTRICAL BUS LOAD OF 75 WE 

**VALUE IN [ 1 IS PASSIVELY hADIATED (CO).MUNICATIONS) AND IS NOT PART OF ACTIVE WASTE 

PO

S

24

24

LAB 

2 

8.5 

0.61 

2.0 

- - 

11.11 

H AE 

1 

2.24 

3.45 

5.69 

LOG 

0.56 

0.5 

1.06 

CUSTOMER 

HOUSEKEEPING 

POWER SYSTEM 

h AB 

2 

2.24 

3.45 

[3.61** 

+[3.6]** 5.69 

SPACE STATION TOTALS* 

10.5 kW @ 35OF 

16.5 kW 8 68 - 77°F 

23.0 kW @ 68 - 122OF 

- 

9.6 kW @ 3S°F 

11.8 kW @ 68 - 77°F 

+[3.6]** 

24.4 kW @ 115°F +S°F 

- 

i 

LAB 

1 -. 

10.0 

13.5 

3.96 

2.4 

29.86 

I/si.BLI-- +[3.6]** -- 

4.4.7.2 Heat Acquisition and Transport 

4.4.7.2.1 Sizing approach 

The isothermal requireaents for custaner thermal control and the 

capability to adapt to highly variable thermal loads have led to more 

advanced them1 management system concepts than have previously been 

utilized on spacecraft. These concepts take advartage of the properties 

of --phase fluids which include an enhanced thermal capacity (latent 

heat of vaporization), improved heat transfer coefff cf ents, decreased 

pumping power requirements, and reduced fluid in\ entory. The reference 

configuration Active Tneraal Control Subsystem ( ATCS) is composed of (a) 

two-phase amnia external thermal transport locrps, or 'them1 buses', 

which transport the heat load from the station modules, power subsystem, 

and attached payloads to the central radiators for heat rejection and 

(b) two-phase water loops internal to the manned modules which transport 

heat from the equipmnt and experiment coldplates to the module/external 

them1 bus interface heat exchangers. 

The external and internal thermal transport loops for the manned modules 

are segmented into three separate temperature levels (3S°F, 70°F, .and 

90°F; to .educe required central radiator area and to enhance the 

isothermal characteristics of the thermal nranagelnent system. The 35OF 

loop services ECLSS and some customer heat loads; the 706F loop services 

equipment for "housekeeping" and customer heat loads; and the 90°F loop 

services customer heat loads. Thermal control of power subsystex 

equipmnt (regecerative fuel cells and power conditioning hardware) is 

accomplished by an autonomous 115OF loop. 

The analysis sizing the thermal acquisition/transport system for the 

reference configuration was accomplished utilizing a computer code for a 

parallel flow thermal bus system. The program was configured to size 

single-phase as well as two-phase transport systems and to incorporate a 

variety of fluids with temperature dependent properties. 

The program optimizes the sum of the fluid weight, line weight, pump 

package weight, pump power penalty weight, and valve weight for a 

parallel flow system. Line wall thickness is calculated from stress and 

mi crometeori od considerations and i s then compared to standard pipe 

sizes. Assumptions used in the analysis are as follows: 

Pump efficiency : 25 percent 

Pumping power penalty: 250 1 bs/kW 

Col dplate weight: 20 1 bs/kW 

Heat exchanger weight: 

3.2 lbs/kW for two-phase to single-phase 

6.4 lbs/kW for two-phase to Go-phase 

V3lve weight: 6 lbs each 

FIinimum pipe ID: 0.188 in. 

Module inter~~al bus plumbing: stainless steel 

External bus plumbing: aluminum 

Heat loads and loop temperature per Table 4.4.7-3 
TABLE 4.4.7-3 

IOC SPACE STATION 

THEM LOAD DISTRIBUTION 

35F(2C) 7(K(21C) 90F132C) ll!iF(lbC) TOTAL 

- ELEIIEMT =mo - 

HAB 1 2.24 3.45 0.00 0 -00 5 -69 

HA02 2.24 3.15 0.00 0.00 5.69 

LA01 13.96 2.40 13.50 0.00 29.86 

LA02 0.61 10.50 0.00 0.00 11.11 

PAYLOAD 0.50 8.00 9.50 0.00 18.00 

LOGISTICS 0.56 0.50 0.00 0.00 1.06 

PMR m)OULE 0.00 0.00 0.00 24.40 24.40 

- - 

IOC TOTK 20.11 28.30 23.00 24.40 95.81 
Further evaluation of the thermal acqui si tion/heat exchanger components 

(i .e., col dplates and heat exchangers) was not undertaken because of the 

greater level of detail required and the time constraints present. 

4.4.7.2.2 External central ized thermal bus sizing 

Fluid transport lines for the centralized system are external to the 

modules, requiring the tube wall thicknesses to be sized to reduce the 

probabil ity of micrometeroid penetration. The external them1 loops 

will interface with the modules across a heat exchanger. However, the 

heat exchanger weights are not included in the central bus but are 

instead included in the module heat transport weights. There is assumed 

to be a distance of 30 feet between the Station modules which are 

connected in parallel, and 50 feet between the first Station module and 

the central radiator condenser. A schematic of the external thermal bus 

(excluding the power subsystem) is given in Figure 4.4.7-1. Althougb the 

logistics module i s shown to require some thermal rejection capabil ity, 

it is assumed to be self-contained and independent of the central 

transport system. 

Results for a two-phase amnonia external bus capable of transporting 

70.35 kJ from the modules and attached payloads (excluding the power 

subsystem) to the central radiator are shown in Table 4.4.7-4. The vapor 

lines are sized to minimize pressure drop in order to maintain a +5"F 

temperature band for isothermality across the heat loads. It shoad be 

noted that the central transport lines will ultimately be sized by the 

expected growth potential of the Space Station. This is especially true 

of those lines which cross the radiator gimbal mechanism. These lines 

will either be significantly oversized for IOC or the plumbing and fluid 

joints will be changed out at the time the thermal load surpasses the 

system capability. The results presented for the IOC external bus 

represent minimum capabil i ties. 

4.4.3.2.3 Module/element internal thermal transport sizing 

The individual modul e/el ement thermal 1 oads, as given previously in Tab1 e 

4.4.7-3, were further subdivided into several thermal quantities to 

represent coi dpl ate 1 ocations within the modules. Each module heat load 

at a particular temperature level was divided into three to six 

individual loads, with the larger module loads being divided into the 

iarger number of individual loads, all of which are connected in 

parallel. As is appareqt, this is a rough representation of what the 

actual heat load distribution would be, but is considered sufffcient for 

determining approximations of system line sizes and weight. A schematic 

of the internal thermal transport system for each module/element is shown 

in Figures 4.4.7-2(a) and 4.4.7-2(b). 

Results for each of the station module elements, using two-phase water 

with stainless steel plumbing f~r thermal transport,are given for IOC in 

Tables 4.4.7-5 through 4.4.7-7. The att.ached payload and power subsystem 

transport sizing results are summarized by Table 4.4.7-8. They util ize 

the same fluid as the central thermal bus, i.e., two-phase amnia, not 

water . 
RADIATORS 

@ FLOW MODULATION VALVES 

Flgun 4.4.7-1 - EXTERNAL THERMAL BUS SCHEMATIC 
TABLE 4.4.7-4 

TWO PHASE AMMONIA EXTERNAL BUS 

THERMAL TRANSPORT CHARACTER1 STICS 

35F(2C) 70F(21C) 90F(32C) 115F(46C) 

EXTERNAL BUS- IOC 

TOTAL Q 19.55 27.80 23.00 0.00 

MASS FLOW (LB/HR) 123.46 186 79 160.96 0.00 

LIQ LINE ID (IN.) 0.19 0.22 0.23 0.00 

VAP LINE ID (IN.) 0.63 0.54 0.42 0.00 

PUMP POUER (WATTS) 25.00 35.30 23.50 0.00 

PMP PWR PENALTY ! LBS) 8.80 12.40 8.20 0.00 

PUMP PACK WT (LBS) 19.40 23.30 21 .OO 0.00 

LIQ LINE WT (LBS) 28.10 32 .GO 10.90 0 .OO 

VAP LINE WT (LBS) 86.20 72.90 19.60 0.00 

FLUID WT fLBS) 1.40 1.80 0.90 0.00 

VALVE WT (LBS) 30.00 30.00 12 .OO 0.00 

COLDPLATE WT ( LBS) 0 .OO 0 .OO 0 .OO 0 .OO 

HEAT EX WT (LBS) 0.00 0.00 0.00 0.00 

- TOTAL 

- 

TOT TRANS WT (LBS) 173.90 172.40 72.60 0.00 418.90 

WET WEIGHT (LBS) 165.10 160.00 64.40 0.00 389.50 

TOTAL DRY WT (LBS) 163.70 158.20 63.50 0.00 385.40 
HAB 1 MODULE 

TABLE 4.4.7-5 

TWO PHASE UATER 

THERMAL TRANSPORT CHARACTER1 STICS 

TOTAL Q 

MASS FLOW (LBIHR) 

LIQ LINE ID (IN.) 

VAP LINE ID (IN.) 

PUMP POWER (WATTS) 

PMP PWR PENALTY (LBS) 

PUMP PACK WT (LBS) 

LIQ LINE WT (LBS) 

VAP LINE UT (LBS) 

FLUID WT (LBS) 

VALVE UT (LBS) 

COLOPLATE WT (LBS) 

HEAT EX UT (LBS) - 

TOT TRANS WT (LBS) 101.10 137.30 0.00 0.00 238.40 

WET WEIGHT ( LBS) 101.10 137.20 0.00 0.00 238.30 

TOTAL DRY WT (LBS) 100.70 136.80 0.00 0.00 237.50 

TWO PHASE UATER 

THERMAL TRANSPORT CHARACTERISTICS 

35F(2C) 70F( 21C) 90F(32C) 115F(46C) TOTAL 

HA0 2 MODULE - 

TOTAL Q 

MASS FLOW (LB/HR) 

LIQ LINE ID ( IN. ) 

VAP LINE ID (IN.) 

PUMP POWER (WATTS) 

PMP PWR PENALTY (LBS) 

PUMP PACK UT (LBS) 

LIQ LINE WT (LBS) 

VAP LINE UT (LBS) 

FLUID UT (LBS) 

VALVE WT (LBS) 

COLDPLATE UT (LBS) 

HEAT EX UT (LBS) 

TOT TRANS WT (LBS) 

MET WEIGHT (LBS) 

TOTAL DRY UT (LBS) 
BODY-MOUNTED RADIATOR 

HABITAT MODULE 

BODY-MOUNTED RADIATOR 

LAB 1 MODULE 

BODY-MOUNTED RADIATOR 

1 I I 

LAB 2 MODULE 

Figure 4.4.7-2 (a) - INTERNAL THERMAL TRANSPORT SCHEMATICS 

496 
-- -- 

LOGlSTICS NODULE 

FUEL CELL U 

u 

ATTACHED PAYLOAD 

Figun 4.4.7.4 (b] - INTERNAL THERMAL TRANSPORT SCHEMATICS 

497 
TABLE 4.4.7-6 

TWO PHASE WATER 

THERMAL TRANSPORT CHARACTERISTICS 

35F(2C) 70F(21C) 90F(32C) 115F(46C) TOTAL 

LAB 1 MODULE - 

TOTAL Q 

MASS FLOW 

LIQ LINE 

VAP LINE 

I ( LB/HR) 

ID (IN.) 

ID (IN.) 

PUMP POWER (WATTS) 

PMP PWR PENALTY (LBS) 

PUMP PACK WT (LBS) 

LIQ LINE WT (LBS) 

YAP LINE WT (LBS) 

FLUID WT (LBS) 

VALVE WT (LBS) 

COLDPLATE WT (LBS) 

HEAT EX WT (LBS) 

TOT TRANS WT (LBS) 456.50 101.70 422.70 0.00 980.90 

WET WEIGHT (LBS) 456.20 101.70 422.40 0.00 980.30 

TOTAL DRY WT (LBS) 455.80 101.30 422.00 0.00 979.10 

LAB 2 MODULE 

TWO PHASE WATER 

THERMAL TRANSPORT CHARACTERISTICS 

- TOTAL 

TOTAL Q 

MASS FLOW (LB/HR) 

LIQ LINE ID (IN.) 

VAP LINE ID (IN.) 

PUMP POWER (WATTS) 

PMP PWR PENALTY (LBS) 

PUMP PACK WT (LBS) 

LIQ LINE WT (LBS) 

VAP LINE WT (LBS) 

FLUID WT (LBS) 

VALVE WT (LBS) 

COLDPLATE WT (LBS) 

HEAT EX WT (LBS) - - - -- -- 

TOT TRANS WT (LBS) 47.60 345 -60 0 .OO 0.00 393.20 

WET WEIGHT (LBS) 47.6; 345.40 0.00 0.00 393.00 

TOTAL DRY WT (LBS) 47.20 345.00 0 .OO 0.00 392.20 
The internal thermal transport system were sized for an IOC capability. 

However, scnne excess capability will ultimately be designed into the 

systems. This growth potential will be a function of the growth 

projections for the Space Station, 1 .e., expansion of existing mdule 

capabi 1 i ties or addition of supplemental modules. The results presented 

for the IOC thermal system represent minimum capabilities. 

A sumuary of significant results for the reference IOC configuration are 

given in Table 4.4.7-8(a). As is evident, pumping power requimnts an 

quite low and the fluid inventory relatively small. It should be noted, 

however, that fluid accumulator quantities are not included in the 

analysi s. 

4.4.7.2.4 Design sensitivities 

In order to evaluate the sensitivity of the thermal acquisition/transport 

systew weight and pumping power to di fferent design paraeeters, alternate 

worki ng f 1 uids were incorporated into the analysi s . Sing1 e versus 

two-phase transport systems were a1 so evaluated. Subsequently, 

integrated transport systems were composed of the various fluids and 

fluid phases. 

4.4.7.2.4.1 External central ized them1 bus a1 ternatives 

Two-phase amnia was chosen as the working fluid for the centralized 

thermal bus because of its attractive thennal characteristics such as its 

large heat of vaporization, and low freezing temperature. Ammonia 

operiting strictly as a liquid would also be a possible consideration for 

theml transport because of its high specific heat, so this was 

evaluated as well. Additionally, Freon-114 was assessed in a one- and 

two-phase operation for the external bus. Figures 4.4.7-3(a) and 

4.4.7-3(b) il lustrate the comparabil ity between the fluids and their 

working phases for pump power, system weight, and maximula 1 ine size at 

IOC. In all cases, an isothermality of +5'F was maintained within the 

transport loop. As is shown, the two-phzse ammonia requires 

significantly less pumping power, lower total weight, and smaller line 

sizes. 

4.4.7.2.4.2 Internal theml transport a1 ternatives 

Two-phase water was selected as the working fluid for the nodule internal 

thermal transport systems because of its nontoxicity, high specific heat, 

and extremely large heat of vaporization. However, the dl sadvantaps of 

water are its high freezing point and low vapor pressure. Consequently, 

if toxicity requirements were to be relaxed or redefined, there is a 

passibility that another fluid such as one in the Freon class could be 

considered as an alternative. To illustrate how this would affect module 

thermal transport system pump power, system weight, and 1 ine sf zes, an 

analysis was accomplished for a single and two-phase Freon system as well 

as for a single phase water system. 
CENTRAL BUS 

HA8 1 

HAB 2 

LAB 1 

LAB 2 

PAYLOAD 

LOGISTICS 

POWER MODULES 

TABLE 4.4.7-8(A) 

SUMMARY OF IOC SPACE STATION 

THERMAL TRANSPORT CHARACTER I STICS 

PUMP PWR 

(WATTSI 

FLUID WT 

-EST 

DRY WT 

TEI 

TOT TRANS KT 

11851 

IOC TOTAL 116.50 10.00 3613.70 3623.70 
EXTERPJAL THERMAL BU5 PUMP POW/ER 

e 

EX.TERN.AL THERMAL F3 iJ 5 VdEIGHTs 

TRAIJSPORT FUJlCI 

r?) UPT %T FLUID YqT 

FIGURE 4.4.7-3(a) 

503 
EXTEFiP4AL THERMAL J MAS 

TRANS PI3 RT FLU 113 

FIGURE 4.4.7-3(b) 

504 
Fim 4.4.7-4Ia) and 4.4.?-4lb) present a caplrison betueen the one- 

.J -st ma- and tk ocre- and --phase Freon-114 systems. As 

-, tm-phse Fr~orc-114 thermal transport lines mild be considerably 

wile Olrrr the tro-pbse uakr lines. Also, syster weights would be 

wt li- vie Freon due to alwinu rather than stainless steel 

beiag riltd in ZLt p'laiq system. Pup pare+ uauld be higher with the 

Fraar sysDa, Lawrrr. Thc single phase water and Freon them1 

sjsts mirr .uch hiwr pup parer with higher system wf ght 

of tht inr~eised flujd inventory and flaw rates. 

44.7.. Integrated them1 +zansport altenati ves 

Vaiow aWainrtirms of We extemal and intwnal them1 ti ansport 

vcre caqored to assess their viability for optimizing the 

-1 systee for +A reference configuratioc. Hybrid systems 

Jidc wre cmted for analysis are as follows: 

Case Mo. FJtemal Thermal &rr/Interml Therral Transport 

Single-gkse Frei~b-114/Si~gle-phirce Freon-114 

Single-pMse ~ia/Siqle-phase water 

Tuo-phzse ;nia/Single-phase water 

Tw+bse aumniaflw-phase water 

Tuo-phase amonia,flw-phase water (70°F 6 90°F loads cadined) 

Tro-pirase Freon-114/Tuo-phase water 

Tw-pbse wnia/Two-phase Freon-114 

Two-phase Freon-114/Tvo-phase Freon-114 

fi- 4.4.7-5(a) and 4.4.7-5(b) compare these thermal transport system 

in ttra of total system ueight, maxim 1 ine diameter, and wing 

parer. All of the two-phase systems are corparable in tiem of weight, rltha~gh the total Freon-114 system is the !ightiest because interface 

heat exchangers betueen the central ioop and module loop would not be 

mi*. The all single-phase systeas woulc' be ccrlsiderably heavier 

mth h~gkr wing mer requireaents. 

4.1.7.2.5 External thermal bus/mdule them1 transport syster interconnects 

W:e heat loads will be collected by the central thew1 bus to be 

rejected by the central radiators. Conceptually . the station nodules 

rharld be connected in paralle: to the central bus in order to ainiafze 

pressure drop, to uinirize two-phase flow in a single conduit, and to 

minioize system coaplexi ty for accmdating load sharing. There are, 

mmr, a variety of concepts for connecting these mdirles to facilitate 

system growth and to miniuize single-point failures. A ninimfzation of 

fluid disconnects and reduction of system complexity is also required. 

Additionally. EVA activity for connection of the mdule/pqyload them1 

loaQ into the central systea should be minimal. 

In consonance rith the previous requirements, a nuraber of interconnection 

corcepts have been proposed. Functional diagrams of these, with positive 
2P F114 1P H2O 1P r!l4 

TRANS PORT FLU I0 

D TRAPIS PO FUJ ID GRI. MT. GIJ FLUID WT 

FIGURE 4.4.7-4(a) 
FIGURE 4.4.7-4(b) 

INTERNAL THERMAL BUS MAX LINE SIZE 

2F F1lJ IF t120 1P F114 

TRANSPORT FW ID 

FIGURE 4.4.7-4:b) 

507 
1 2 3 4 5 B 7 

CASE CJUMEER 

CUE NUMBER 

URl' 'f'fr. 1'7UlO vr7. 

FIGURE 4.4.7-5(a) 

508 
FIGURE 4.4.7-5(b) 

INTEGRATED THERMAL BUS MAX UNE SIZE 

4 

1 2 3 4 5 6 7 E 

CSE NUMBER 

FIGURE 4.4.7-5(b) 

, , 

j 

: 5 i :- I:. 

0 
ORIGIMAL FA(lE -16- 

OF POOR QUALIW 

and negative attributes denoted, are shwn in Figure 4.4.7-6(a) and (b). 

As shown, a flow control assembly (FCA) consisting of a pump, 

accumulator, and regenerative heat exchanger is included in each 

concept. This can either take the form of an autonomous, modular package 

or as module-integrated hardware. Module interconnects to the centrsl 

bus in each concept can be accomplished with fluid coup1 fngs external to 

the habitable volume, with contact heat exchangers (also external to th 

habi table volume), or with an intermediate nontoxic loq wrtll) #bdd 

couplings internal to the habitable volume. The lightest mh wid 

utilize external fluid couplings and the heaviest would use contact kt 

exchangers. For this study, a fluid coupling approach was ass& for 

sizing purposes. However, a1 though weight is an important crlkrlon, 

factors such as EVA support and re1 iahility prevent fluid couplings from 

being the most desirable approach. 

Therefore, an indepth review of the proposed connection schemes Is 

required with a trades analysis being accampl ished to establish UK 

optimum candidate. This was not accomplished during this brief study, 

though the criticality and complexity of this task was identified. 

4.4.7.3 Heat Rejection 

4.4.7.3.1 Sizing approach 

The reference configuration heat rejectlon system consists of high 

capacity, monogroove, heat pipe radiator panels, as shown !n Figure 

4.4.7-7. In this approach, the radiators wi! 1 be erected in space with 

an MRMS as fllustrated by Ffgure 4.4.7-8. An individual radlator element 

can be removed and replaced if damaged. The heat pipe radiator elements 

are "plugged in" to contact heat exchangers as shown by Figure 4.4.7-9 

whi ch i 1 lustrates one concept under development. These heat exchmg~rs 

provide a loose fit for the heat pipes when they are initially plugged 

in. A clamping action is then provided by the contact heat exchanger, 

thus giving the contact force needed for good heat transfer contact 

conductance. 

The heat rejection system illustrated by Figure 4.3.5-1 was divided into 

four functional segments: 

a. A central fzed system which rejects the majority of the 

station heat load except for power system, logistics module, and 

passively rejected heat 1 oads 

b. A power system radiator which rejects the fuel cell and 

power conditioning heat loads with a llS°F heat load interface 

temperature 

c. A body-mounted logistics module rddiator system which 

rejects the entire logistics module heat load 

d. A body-mounted radiator on each of the remaining modules 

which interfaces with the 35°F central transport loop to provide 

autonomous safe haven thermal control for each module 
CONCEPT $2 n 

CENT. 

Lpi$jHq 

PROS 

TOTAL CENT. RADIATOR 

AVAILABLE IF STATION MODULE 

FAlLS 

a TOTAL CENT. RADIATOR 

AVAILABLE IF STATION MODULE 

FAILS 

VULNER

FAILURE 

a REQUIR

(i.e., FCA

DURING 

9 MAY P,E

a VULNER

FAILURE 

REQUIR

MODIJLE 

VCILNER

FAILURE 

PLUMBI

a REQUIR

COMPLI

PLUMBI

a TOTAL CENT. RADIATOR 

AVAILABLE IF STATION MODULE 

FAILS 

NOTC: FCA - "FLOW CONTROL ASSEMBLY" CONSISTS OF PUMP/ACCUM. AND REG

Figure 4.4.7-6 (8) - CENTRAL BUS PLUMBINQ CONCECrS 
CONCEPT #4 

CENT. 

R AD. 

PROS CO

PROVIDES DECENTRALIZED LARGE NO. OF P

CAPABILITY WITHOUT BODY- THRU GIMBAL 

MT. RADIATORS TOTAL CENT. R

ACCOMMODATES GROWTH AVAILABLE (WIT

EASILY ADDIT!ONAL MA

VALVES) IF STA

FAILS 

REQUIRES ADD

TOTAL CENT. RADIATOR VULNERABLE T

AVAILABLE IF STATION MODULE FAILURE (i.e.. FC

FAILS MAY REQUIRE A

HEQUIRES ADD

CONCEPT #5 MODULE BYPASS LINES I \ 

PLUMBING & CO

HOOKUP 

CONCEPT t6 MODULE BYPASS LINE 

1 

PROS CO

TOTAL CENT. RADlAf OR COMPLICATED 

AVAILABLE IF STATION MODULE HOOKUP 

FAILS 

cn 

c-. I 

1 I*; 

CENT. 

RAD 

VULNERABLE TO 

FAILURE (i.e., FC

MAY REQUIRE A

PROPER FLOW D

MAY BE PROBLE

EITHER DIRECT

CENT. 

RAD 

b , 

I 

- - 

, FC A H AB H AB 

Figure 4.4.7-6 (b) - CENTRAL BUS PLUMBING CONCEPTS 

-* 

- 

I 

t I: - 

LAB LAB 

FC A HAB H AB LAB LAB 

P 
VERY HIGH CAPACITY - DUE TO SEPARATE, LOW 

PRESSURE DROP LIQUID FLOW CHANNEL 

VERY LIGHTWEIGHT - ESSENTIALLY AN AL TUBE 

VERY SIMPLE CONCEPT - NO WICKS, NO VARIABLE 

CONDUCTANCE CONTROLS, NO ION PUMPS 

RELATIVELY EASY TO MANUFACTURE - PRIMARILY AN 

EXTRUSION HP 

CAPABLE OF BEING TESTED AT NEAR FULL CAPAClTY 

IN ONE-G 

RELATIVELY EASY TO ANALYZE - DUE TO SIMPLICITY 

OF DESIGN 

Figure 4.4.7-7 - MONOGROOVE HEAT PIPE RADIATOR 
FIGURE 4.4.7-8 - SPACE CONSTRUCTABLE RADIATOR 
EVAP SECTION 

(HEAT PlPE) 7 

i - 

MECHANICAL PRESSURE CUMPINO CONCEPT 

THERMAL AeTUATlON 

EXPANSION BOLT 

MULTI-LEG 

EVAPORATOR 

FIGURE 4.4.7-9 

MECHANICAL CONTACT HEAT EXCHANGER 

515 
Central radiator system - sizl np 

A central radiator system using a radiator panel fin thickness of .020 

inches for the articulated radiators was chosen giving a weight per unit 

pro2ected area of 1.4 lb/ft*. The panels were assumed to be 50 feet 

long and 1 foot wide. Other panel lengths and widths could be used to 

meet f aunch packaging and assembly or deployment requirements without 

si gni ficantly degrading thermal performances. 

Each panel has approximately 2 feet of additional length for an 

evaporatcr section. However, this section's weight was lumped with the 

contact heat exchanger. A detailed sizi~g analysis of the contact heat 

exchanger was not performed. Based on prototype designs, the heat 

exchanger was assumed to weigh 20 pound per panel and have an overall 

conductance of 461 BTU/HR-OF-PANEL. 

To provide maximum hardware comnonality and to simplify the analysis, the 

same radiator panel and contact heat exchanger design was used for all 

operating temperdtures. Future studies should include optimizatiofis for 

each temperature to allow an assessment of the reduced thermal 

performance associated with comnon hardware. 

The central radiators were sized with orbi t-averaged incident heat fluxes 

and view factors to space at fl =OO. The incident heat fluxes and view 

factors were obtained from a TRASYS computer program analysis which 

included the orbital varying radiator/station geometr3 and multiple 

diffuse reflections. Station surfaces were assumed to be at the radiator 

temperature. The radiator view factor tu space was found to vary from 

0.9 to 0.98 twice an orbi t. Additional analysis must be performed to 

determine if a thermal capacitor or larger radiator is required. 

The solar absorptivity was assumed to have an end of life value of 0.2. 

While the UV degradation rate is expected to be small for the central 

radiators, the effects of atomic owgen and contamination have not b?en 

quantified and require further analysis. 

4.4.7.3.3 Power system radiator sizing 

The power system radiator and heat exchanger construction was assumed 

identical to the central system. These radiators are located on the 

transverse boom and were found to have a view factor orbital variation of 

0.75 to 0.91. The radiators were sized with an orbit average view factor 

of 0.80 

Due to the different efficiencies between fuel cell operation and 

electrolysis, the IOC power system waste heat load varies between 13 kW 

on the Sun side to 46 kW on the dark side. The 10C power system radiator 

was sized for the orbit average heat load of 24.4 kW. An 11.9 kWhr 

thermal capacitor was used to average the varying heat load, ignoring any 

inherent thermal capacitance in the system. A detailed sizing analysis 

of the thermal capacitor was not performed. The capacitor was assumed to 

weigh 50 pound/kWhr. Of thi s 50 pound/kWhr, approximately 34 pound/kWhr 
Is requl red by the phase change wax material. Since the equl va? ent 

radiator weight s about 68 pound/kWhr, and the equivalent radiating area 

is about 76 feet 1 /kWhr, the cost of developing such a capacitor rqy not 

be worth the 214-pound system weight reduction. 

4.4.7.3.4 Logistics module radiator sizing 

To elllrinate a thermal interface between the station and the logistics 

module, the logistics module radiators were size3 to continuoi~ly reject 

the entire module heat load. 

The logistics mdule body-munted radiators were assumed to be 

constructed with an ,010-inch sluminurn fin. These radiators are not 

space constructable and were assumed to use a brazed heat exchanger 

weighing 3 pound/panel. Since these radiators also serve as a meteoroid 

bumper shield, a weight credit for the meteoroid protection was taken in 

the radiator weight. 

The radiators were located on the windward side of the logistics module 

and assumed to be 11 feet long. The required width was determined based 

on orbit-average 4 =5Z0 incident heat fluxes. A 180" segnent of the 

nodule circumference was used to reject the nodule heat load. Gue to the 

incident heat flux orbital variation, a thermal capacitor ~ay berequired 

in this system, but has not been included in this study. 

4.4.7.3.5 Safe-haven body radiators 

Since safe-haven heat rejection requirements have not been defined, the 

radiatorr were assumed to cover a 180' segment of the nodule I t 

circumference. The heat rejection capability of the radiators was then I 

determined with orbit-average incident heat fluxes. 

4.4.7.3.6 Heat rejection system results 

The radiator area and radiator and heat exchanger weights are sunarized 

in Table 4.4.7-9. The power system capacitor weight of 594 lb is not 

included in the weight shown in Table 4.4.7-9. These results are 

presented for the IOC configuratfon and represent minimum capabilities; 

however some exce i capability should ultimately be designed into the 

systen. 

4.4.7.3.7 Design sensitivities 

a. Heat pipe radiator fin thickness Previous heat pipe 

radiator fin thjckness optinizations minimized only the radiator an el 

weight. The heat rejection system weight also indudes the heat ' 

exchanger weight, which is dependent on the fin thickness through the 

radiating area requi red. In addition, 1 aunch packaging, on-orbi t 

operatfons and reboost propellant requirements favor reduced radiator 

area at the expense of weight. To assess the effects of the fin 

thickness on the heat rrjettion weight and area, a system was sized for a 

range of thtcknesses from .010 inch to .030 inch for two-sided and 

single-sided radiators as sumrized by Figure 4.4.7-10(a) and 

4.4.7-lO(b). 
TABLE 4.4.7-9 - HEAT REJECTION SYSTEM RESULTS FOR IOC SPACE SlATION 

WSTW SUMP IRP lWlA EPS F TO FIN EFF 01M-E6 F.EFF SIM T RAD TMP HUT LMD MDIRTIWG PA!

OE6 F DIFT2-HA BIFT2-HR SPACE -PAUL DE6 F DE6 F KY MU, Fl2 REB'D 

Ulf 35.00 99.00 21.00 0.20 0.80 O.P5 0.79 1000.00 -24.84 34.73 0.56 79.03 i.0

70.40 49-00 21.00 0.20 0.80 0.95 0.75 1000.00 -19.15 6.57 0.0 43.21 5.9

TOTAL M3Y 1 121.3 11.

SINGLE-SIDEG 

RADIATOR 

E.XTRUSlON WT = .49 LB/FT 

PANEL WIDTH = 1 FT 

FIN THICKNESS, INCHES 

Flgure 4.4.7.-10 (b) - HEAT PIPE RADIATQR WEIGHT 
The ~~inilrmm heat rejxtion system weight of the combined central and 

power systems was ,ound to occur at a fin thickness of aC-^" .015 inch 

as illustrated by Figure 4.4,7-11(a). Since the sensit ry of the 

1 aunch packagf ng, on-orbi t opcrhtions and reboost costs ane radiator 

area are unknown, the referewe configuration fin thickness was 

arbitrarily taken to .020 inch rather than the minima. This thIcker fin 

results in a 3 percent system weight increase while providing a 6 percent 

area reduction. Figure 4.4.?-.ll(b) illustrates similar information for 

the single sided body-mounted radiators. Future studies should attempt to 

quantify the station cost ssnsi ti vi ties to radiator area. 

b. Fixed vs articulated radiators Since there is sow risk 

and develcprnent cost associated with the rotating joint required by the 

articulated radiator concept, a fixed central radiator system was sized 

for the reference configuration. The fixed central radiator was assured 

to be edge to the earth and perpendicular to the orbit plane. Both 

system had the same heat load distribution, fin effectiveness, and 

radiator/heat exchanger wei@t. Since the fixed radiator will be 

affected to a greater degree by solar absorptance ( ac] degradatior, 

rate, ac va:ues of 0.2 and 9.3 were used. 

The results, normalized by the articulated system, are shown below. 

- Total - Central Radiator System Weight Factors 

. - d = 

c. Battery vs. fuel cell 3f the 24.4 kW orbit average fuel 

cell heat rejection, about 16.3 kW are required by the fuel cell itself. 

One alternative to the fuel cell is a rechargeable battery. During power 

generation on the dark side, tke battery was assumed to produce 13.6 kY 

of waste heat. During recharge, only a very small waste heat is 

produced, resulting in an 5.2 kU orbit average heat rejection 

requirement. In addition, the battery requires a .3S°F cool lng interface 

tengerature, while the fuel cell system #as assumed to use llS°F. The 

results, excluding the heat rejection requi remnts for the power 

conditioning equipment comn to the two systems are: 

FUEL CELL BATTERY 

Radiating Area, ~t2 819 

Radiator plus H/X Weight, L9 737 

Capaci tor Weight, LB 594 

Total Heat itejection Weight, LB 1331 

d. Solar absorpti VI ty sensitivity Since some uncertainty 

exists in the solar absorptivity degramrdte, the heat loads and 

environmnts shown in Table 4.4.7-9 were applied with a range of 

absorptivities from the reference configuration value of 0.2 to the fully 

degraded value of 0.8. The required radiator area relative to the 
DOUBLE-SIDED RADIATORS 

20 LEVPANEL HEAT EXCHANGER WEIGHT 

FIN THICKNESS, INCHES 

Flgure 4.4.7-11 (a) - HEAT REJECTION SYSTEM SENSITIVITY TO FIN THICKNESS 
SINGLE-SIDED RADIATORS 

20 LB/PANEL HEAT EXCHANGER WEIGHT 

t 

REF 

CONFIG 

FIN THICKNESS, INCHES 

Figure 4.4.7-11 (b) - HEAT REJECTION SYSTEM SENSITIVITY TO FIN THICKNE
reference configuration is shown in Figure 4.4.7-12 for both the 

logistics module radiators and the combined central and power system 

radiators. Also shown is the total body-mounted safe-haven radiator heat 

rejection capacity for a constant surface area. 

Due to the dii*ect solar flux, the body-mounted radiators cannot tolerate 

an absorptivity degradatlon much greater than 0.2. 

e. Elimination of 90°F thermal loop Since the heat loads 

rejected by the VO'rloop1 imlts of 68OF (see Table 

4.4.7-I), they could satisfnctorily be rejected by the 70°F loop. 

aejectlng these heat loads at a lower temperature wouid impose a penalty 

In net increased central radiating areas and weights, but would eliminate 

extra transport hardware and reduce Interfaces for another loop. For the 

reference conffguratlon, the central radiating area and weight 

requirements increased by 5 percent when the 90°F loop was el?minated. 

These small rejection system Increases should be traded off against the 

decreases i n transport system complexity and wel@ts. 

4.4.7.4 Passive thermal control/protection 

The basic approach to the passive thermal design of the Space station is 

to desensitize, as far as practical, the station elements and subsystems 

to the external environment and surface thermal optical property 

degradatlon whi le minimizing makeup heat requi rements and thermal load 

excursions on the Active Thermal Control System. 

The thermal surface treatments and insul atton performance are an integral 

part of the overall active and passive design and must be selected to 

obtain an acceptable heat balance that will not result in local 

condensation on internal surfaces of pressurized volumes, excess1 ve heat 

1 eaks, or viol ate structural and subsystem temperature 1 imi ts. 

All pressurized modules will be provided wf th a high performance 

mu1 tilayer insulation (MLI) with an effect1 ve emi ttance of approximately 

.01 and external surface (with the exception of body mounted radiators) 

solar absorptance, a~ , to emittance, € , ratio approachdng 1.0. The 

MLI will consist of approximately 20 layers of organically coated 

aluminized fil m with dacron mesh separaters weighing approximately 0.25 

pounds per square foot. The outer package layers are porolated to allow 

for adequate viscous and molecular venting. The selection of an oC /E 

of 1.0 with individual magni tudes of approximately 0.8 should negate 

surface refurbi shment requirements. A simil ar approach appears 

appropriate for the power system and external airlocks. 

The following sumnarize~ the major station elements insulation weights: 

a. Habi tation, Laboratory, Logistics modules - 400 pounds 

each , 

b. OTV hangar - 2340 pounds 

c. Power system - 130 pounds 
RELATIVE HEAT REJECTION CAPACI

OR 

RELATIVE RADIATING AREA REQUIR

-L 

Z o in 6 
d. Prapulsion (stat Ion only) 

- statfon contingency tanks - 120 pounds 

- propellant 1 ines - 15 pounds 

- engine modules - 15 pounds 

e. External airjocks - 50 pounds each, 

f. Satellite thermal protection shrouds - 795 pounds each. 

4.4.7.4.1 Pressurized modules 

Typical module heat leaks are presented for Habitation modules 1 and 2 in 

Table 4.4.7-10 for a range of surface thermal optical properties and 

insulation performance characteristics. The maximum heat 1 eak out (1191 

BTU/HR or 345 watts) of the modules cccurs for an insulat~an effective 

emittance of 0.02 and an W/ < ratio of 0.3/0.8 while the maximtrm heat 

gain (-363 BTU/HR or 106 watts) occurs for an insulation effective 

mi ttance of 0.02 and an a / t5 ratio of 0.8/0.8. These represent a 

small portion of the total internal heat load and do not effect the 

performance requirements of the Active Thermal Control Subsystem, 

Therefore, an external module surface treatment such as anodizing can be 

selected it1 l ieu of a painted surface thereby eliminating the need for 

surface finish refurbishment. The module heat leaks were determined from 

a simple cylindrical math model with four external circumferential nodes 

(micrometeoriod shield) separated from the module wall by a gap and 

insulation on the module wall. The internal air node was held at a 

constant 70°F. 

4.4.7.4.2 OMJ servicing and storage 

Since no requirement is identified to provide an active or passive 

thermal interface for the QMV, the power requirenents for thermal control 

of the OMV are dependent on the final thermal design of the OMV. 

Therefore, the OMV becomes a thermal integration consideration for 

definition of the OMY thermal environment and resulting OMV thermal 

system power usage. 

4.4.7.4.3 Satellite thermal considerations 

For satellites not requiring a thermal protection shroud, a cylindrical 

model divided into four circumferential nodes and top and bottom nodes 

(ends), assumed adiabatic on the backside, was used to define maximum and 

minimum sink temperatures that would be experienced by a satellite for 

various thermal optical properties. In addition, maximum and minimum 

transient temperatures for a 0.05-i nch alumi nun wall were a1 so 

determined. These are presented in Table 4.4.7-11. Figure 4.4.7-13(a) 

and (b J show the Orbital variation of sink tempeature for a satellite 

located in the -y servfce bay with an oC/ € of .8/.8. 
TABLE 1-4-7-10 

HABITABLE rDDULE PASSIVE HEAT BMANCE 

(ORBITAL AVEME ) 

MD0lJl.E INWATIQN HEAT GAIN/LOSS 

r EFF BTU/HR 

NOTE: HUT GAIN DENOTED BY NEGATIVE VALUE. 
TABLE 4 -4.7-11 - UNPROTECTED SATELLITE TEMPERATURES 

TEMPERATURES ON SATELLITE SURFACES - +Y SATELLITE BAY, BETA = 52' 

- - a/€ =. &F = .3/. ock = . . d-'&&T~* SINK TRk* 

TEMPS .05" WALL TEb!PS .0511 WALL TEMPS .05" WALL TEMPS .05" WALL 

200 162 lbs 125 60 - - - 

TOP -460 -118 -46Q -125 -460 -140 -460 - 190 

145 102 125 85 7 0 45 30 20 

BOTTOM 5 25 5 2 5 5 15 5 10 

270 25s 230 223 120 9 5 0 - 

CYLINDER -140 -75 -140 -80 -140 -100 -140 -120 

TEMPERATURES ON SATELLITE SURFACES - -Y SATELLITE BAY, BETA = 52" 

o( 16 = .8/.8 -= .69/86 d/G = .3/.8 oC/C = .1/.8 - 

SINK TR~NT SINK TRA~EEFK~-TT!K TRAK'IBT 

TMPS .05" WALL TEMPS .05" WALL TEMPS .05" WALL TEMPS .05" YALL 

160 125 130 85 25 - - - 

TOP -460 -120 -460 -130 -460 -1 50 -460 -200 

165 110 140 90 80 --T!rBOTTOM 5 3 5 5 30 5 20 

285 2 70 245 - 5 12 

22 5 10 - 

CYLINDER -145 -85 -145 -95 -145 -110 -145 -125 

TEMPERATURES ON SATELLITE SURFACES OF REFUELING BAY - 8ETA = -52' - 

d/E = .8/.8 ad/€ = .69/86 d/€ = .3/.8 d/~ = .1/.8 

SINK TRANSI~NT SINK TRA~I~NT SINK TRA~ SINK TRA~ 

H 

TEMPS .05" WALL TEMPS .05" WALL TEMPS .05" WALL TEMPS .05" WALL 

230 180 115 40 - - 

TOP -250 -50 -250 -70 -250 - 90 -250 -145 

h 50 6 0 3 5 15 0 - - - BOTTOM -40 - 10 - 40 - 10 -40 -25 -40 -35 

245 215 210 110 I00 - -ZO 

CYLINDER -145 -95 -145 -100 -145 -1 10 -145 -120 

1 

TEMPERATURES ON SATELLITE SURFACES OF REFUFLING BAY - BETA = +5Z0 

- ocg.: . fj/g . d/C = .69/86 d/€ = .3/.8 = .1/.8 

SINK TR~ SINK TRANSIERT SINK ~TNX Tm 

TEMPS .05" WALL TEMPS .05" WALL TEMPS .05" WALL TEMPS .05" WALL 

235 7 5 200 40 90 

TOP -220 -60 -220 -75 -220 -115 -220 -165 

-75 135 50 - 6 5 0 0 - 

BOTTOM -45 0 -45 -10 -45 -25 -45 -35 

280 270 245 2 30 130 100 10 ~. - 

CYLINDER -135 -100 -135 -100 -135 -1i5 - 135 -120 


For satellites rqulring a them1 protection shroud, a 15 feet dlamettr 

by 60-feet long deployable nultilqyer insulation (eEFF. = .O1) shroud 

was assumed wi th thermal closeouts at each end. The average Orbital 

internal temperatures for various external surface W/ € ratios were 

determined as a function of internal heat generation to represent payload 

and service operations. The +Y storage bay and -Y service bay 

temperatures are shown in Figure 4.4.7-14. Note that a small variation 

in temperature exists between the two locations. For an =/& of 

0.8/0.8 and no internal heat generation, an acceptable temperature range 

between 50 and 64OF can be maintained. However, internal heat generation 

would be limited to approxinately 475 watts in order not to exceed 

10Q°F. On the other hand, an W/G of 0.1/0.8 would result in an 

approximate temperature of -83OF with no internal heat generation and 

would a1 low approximately 1.25 kY of internal heat generation before 

exceeding 10O0F. These values are presented to point out the sensitivity 

of the internal shroud temperatures to OC/~ and internal power 

generation and the necessity to undertake a detailed review of satellite 

requirerents and servicing operations in order to define the them1 

design envelope for the thermal protection shrouds. 

Similar data for the refueling bay is presented in Figure 4.4.7-15 for 

its extreme environments of +5Z0 beta anqle and -52' beta angle. For 

hfghW/€ ratios, thfs location runs 10 to 20°F wamr than the storage 

and service bays. These results indicate that low oC/€ them1 coatings 

will have to be selected which might requi re periodic refurbishment. 

4.4.7.4.4 OTV Facility 

The CTV facility, TIM 2570, was assuaed to have multllqyer insulation on 

the inside surface of the hangar wal is with an effective mittance of 

0.1. Figure 4.4.7-16 presents the Orb!tal average internal hangar 

temperatures for various OC/ € ratios as a function of internal power 

generation. Because of the 1 arge surface areas of the hangar, an =/€ 

of 1.0 can be applied and still maintain an acceptable temperature range, 

25.F to 100°F, for 0 to approximately 2 kW internal heat generation. 

Hard anodized aluminum (W = .69, E = .86) can maintain temperatures 

above Q°F and could be a1 lowed to degrade from contamination sources 

thereby deleting arW refurbishment requirements. Active them1 

interface requl re~ents have not been f denti fled. However, support 

equlpnent pose the potential for active thermal control. 'These 

requirements wst be identified during the definition phase. 

Propulsion System Thermal Control. 

The OW, OTV, satell i te refuel ing and station contingency propel lants are 

located and thermally protected as described in Sectioo 4.3.5.3. Since 

the exact location of each type of propellant has not been identified, 

the hottest and coldest locations for a beta angle of 52" were considered 

for this study. These locations are the +Y and -Y corners of the upper 

keel extension. The analysi s is based on a fully 
NOTE: 1 - (+) Y (HOT ENV.) 

2 - (-) Y (COLD ENV.) 

I 1 I 1 I 1 

0 .5 1 .O 1.5 2.0 2.5 3.0 

INTERNAL POWER, kW 

Flgure 4.4.7-14 - PAYLOAD SERVICE BAY SHROUD INTERNAL TEYmRA
MOTE: 1 - 8 = +52" 

2-a=-52" 

Flgure 4.4.7-15 - REFUELING BAY INTERNAL SHROUD TEMPERATLJRE 

insulated 9-foot cube assumed to be adiabatic at the wall adjacent to the 

next 9-foot cube. Average Orbital tank temperatures and makeup heat 

required for various W/€ ratios and an insulation effective 

emittance of 0.01 are as fo:lows: 

-- Qc/& +Y LOCATION 

TEMP WE UP TEMP MAKEUP 

(OF) ( MATTS) (OF) (MATTS) 

The above analyri s appl ies on1 y to the bi propel 1 ants and ncnopropell ants 

with an assumed operation temperature range of 40 to 100°F. As can be 

seen, this range can be acc~lrnodated with minimum makeup heat 

requirements by applying an L '/€ of .8/.8 external surface treatment. 

The makeup heat is supplied by electric heaters control led by a software 

themstat logic thereby avoiding the use of fai lure-prone thermostats. 

The station contingency mnopropell ants are located within one truss 

cube. Therefore, the maximum makeup heat is 54 watts. Other 

bipropell ant and monopropel 1 ants will require approximately the same! 

amount of makeup heat per truss cube. However, cubes located in the 

center portion of the keel extension should require less heat since two 

sides of these cubes are adjacent to other propellant tanks. Design of 

these areas and cryogenic propellants are left open for the definjtion 

phase. 

Approximately 1600 feet of propellant transfer and distribution lines are 

required in both sides of the lower keel, keel extension, and lower 

boom. The lower keel and both sides of the keel extension were analyzed 

for a beta angle of -52" to determine acceptable thermal optical 

properties and makeup heat to maintain the propellant between 40°F and 

100°F (Table 4.4.?-12). Since the lines are small, 0.75-inch diameter, a 

nu1 ti layer insulation effective erni ttance of 0.02 was assumed to account 

for performance degradation which results from wrapping. The two sides 

of the keel extension represent .the extremes of environment encountered. 

An K/c of 0.8/0.8 results in m4nimum makeup heat requirelpents, 0.1 

watts per foot of line for the cold condition, but slightly violates the 

upper liait of 100°F. Assuming that propulsion 1 imit can be increased, a 

propellant line makeup heat of 160 watts is required. The maximum makeup 

heat of 0.27 watts per foot or a total of 432 watts occurs for an =/€ 

of 0.110.8 which is representative of aluminized Teflon film. The film 

could be all w. I to degrade over a period of ti= (represented by 

the %I€ = .3/.8 in Table 4.4.7-12). However, the allowable degradation 

might be limited if the makeup heat required for the initial 4./€ values 

result in over heating as a result of local degradation. The insulation 
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and external surface material mist be flexible to acr:onmodate packaging 

an4 deployment of the flexible lines. This will be the prime driver in 

selecting a material that best meets the overall desfgn requirements. 

This is an open issue for the definition phase. In addition, every 

effort should be made to co-locate the central thermal bus transfer liner 

with the propell~nt lines to facilitate use of a common insulation wrap 

thereby reducing thc makeup heat required f rm electrical heaters. 

4.4.7.4.6 Cmunication system 

Antenna thermal design is completely passive. Each electronics box will 

be of minimal thickness with the electronics conductively coupled to a 

radiator plate on the surface of the box for heat dissipation during 

standby and y-rational modes. Electrical heaters, control 1 ed by 

software 1 r,gic, can be appl ied for "off" or contingency conditions. 

Operation of thtse heaters is minimal and does not add to the power 

requireiiiints of 1 he operational and standby modes. The electronics box 

i s insulated and ,> radi ator-type coating, rqui ri ng periodic 

refurbishment, i s ,.?pl ied to the exterior surfaces. Standby heaters and 

appropriate thermal sb,-f :s, treatments are appl ied to the antenna gyros 

and gimbals. 

Since the actual thermal requirements of each antenna are unique to the 

particular design, only an estimate of electronics package radiator areas 

and "off" or contingency makeup heat requirements was determined. 

Radiator areas were determined by ratioing the Shuttle Orbiter Ku-band 

antenna power and radiator ar,d. Contingency makeup or standby power was 

estimated by assuming that the Shuttle Orbiter Ku-band electronics box 

was maintained at 100°F at the radiator in a cold condition when the box 

was on. The standby power required to maintain a minimum temperature of 

O°F was estimated assuming that the ratio of the standby power to the 

operational power is equivalent to the ratio of the fourth power of the 

two temperature levels. Radiator area and standby or makeup power 

estimates are presented in Table 4.4.7-13. 

4.4.7.4.7 Solar array temperatures 

Figure 4.4.7-17 shows the variation in solar array temperatures for a 

beta angle of 0" and various back face thermal optical properties. The 

data indicates that the temperature (138OF to 166OF) does not vary 

,,preciably for the common range of materials (crossed hatched area) 

generally appl ied to sol ar arrays. 

Since the solar cell efficiency is inversely proportional to the cell 

temperature, it is desirable to apply a back face material or surface 

treatment with the lowest K/E ratio that is practical, in order to 

minimize the array temperature. The sensitivity of the solar cell 

efficiency to temperature is Czpicted in Figure 4.4.7-18, For the comnon 

back face materials, the efficiency can range for 0.112 to 0.121. 

Obviously, degradation of the thermal optical properties can result in 

large impacts to the array power producing capabil i ty. Therefore, 

materials or surface treatments must be selected to minimize 
TABLE 4.4.7-13 - COmWICATIOtlS AND TRACKING 

TOTAL 

PUR. WATTS 

TOTAL RADIATOR 

AREA, FT2 

i 

TOTAL MAKEUP 

EQU I PME NT QUANTITY HEAT, WATTS 

ANTENNA: 

m/Ku-band Ant. 

Co-Orbi t Med. Gain Ant. 

Co-orbit High Gain Ant. 

Low Gain Multi-beam 

Prox. Ops Ant. 

Near Hemi sph. cov. 

Prox. Ops Ant. 

Orbiter Low Gain (10db) 

S-band Ant. 

Orbiter Low Gain (0-3db) 

S-band Ant. 

GPS Low Gain Ant. 

Med. Gain Rendezvous 

Radar Ant. 

MRMS Low Gain S-band Ant. 

GPS low gaSn 

S-BAND low gain 

Radar Xpndr low gain 

MRMS S-band low gain 

S-band low gain MRMS-TV 

R.F. SYSTEMS: (1) 

-it/Rec. 4 

RF Assy fot Co-orbit 

Med. Gain Ant. 

RF Assy for Co-orbi t 

High Gain Ant. 

RF Assy for Low Gain 

Prox. Ops. Ant. 

RF Assy for Near Hemi. 

Cov. Prox. Ops. Ant. 

RF Assy for Orbiter 

S-band Antennas 

S-BAND xpndr (TDRS) 

S-BAND xpndr (Orbiter ) 

S-BAND pa/preamp (T/O) 

XCVR (MRMS) 

RCVR (MRMS) 

XMTR (ORBITER) 

TRACKING: 

RF Assy for Rendezvous 

Radar Antenna 2 4 50 7 2 04 

Laser Docking System 1 20 -** 9 

Structural Stab. Moni t. 4 400 6 182 

GPS rec. /proc. 2 180 2.7 82 

Radar XPNDR 2 20 -** 9 

* MAKEUP HEAT DEPENDS ON MATERIAL LIMITS, ANTENNA CONFIGURATION AND 

LOCAT ION AND LOCAL STRUCTURE. ** POTENTIAL FOR DESIGN WITHOUT DISCREET R.4DIATOR AREA. 
degradation from ultraviolet radlatlon and atomic oxygen and thcy wst be 

protected from contamlnatlon sources to ~inl~lze or negate the sizeable 

task of refurbishment. 
Flgure 4.4.7-17 - SOLAR ARRAY TEMPERATURES 
-10 0' I I I 1 I J 

70 )O 90 100 110 120 1m 140 160 1w 1'10 180 190 

OPERATINO TEMPERATURE, OF 

Fm 4.4.7-18 - SOUR ARRAY EFFICIENCY 
4.4.8 Structures and Mechani sms 

4.4.8.1 Introduction 

This section of the report deals with various aspects of the main 

structural truss, the pressurized modules, and the mechanisms of the 

operational IOC Space Station. It is not meant to be a mathematical 

paper on structural design or analysis, or a presentation of novel 

structures or mechanical devices. It is, however, meant only to transmit 

concepts, thoughts, and ideas as well as preliminary analysis. Each 

subsection will be a condensation of the more extensive White Papers 

(pub1 ished separately). The topics to be covered are as follows: 

a. Space Station truss structures and construction 

considerations 

b. Preliminary structural design and analysis of a Shuttle 

launched Space Station manned habitable module 

c. Module and payload truss attachments 

d. Module - module and berthing interface 

e. Sol ar panel depl oyment mechani sms 

f. Conceptual rotary joint design for Space Station appl ication 

g. Structural deployment verification 

h. Conceptual design of a Mobile Remote Manipulator System 

4.4.8.2 Space Station Truss Structures and Construction Considerations 

4.4.8.2.1 Introduction 

Over the past several years, a considerable effort has been expended on 

structural design for large space systems. Most of the effort has been 

generic in nature and directed towards large reflector appl ications. The 

concept or the Space Station represents a new set of structural 

requirements that were not considered for the previous appl i cation. 

However, the experience gained on these past structural studies provides 

a wealth of knowledge and insight that can guide the selection process 

toward an optimum or near optimum Space Station structural foundation. 

Prior studies for the construction of large space systems indicate that 

there are three major approaches for packaging, transporting, and 

construction of those structures in orbi t. These approaches involve 

structural members that are classified as deployabl e-single fold, 

erectable, and depl oyable-double fold. An overview of these approaches 

and the identified structural requirements for the Space Station will be 
presented herein. A mre detailed stu* is presented in the White Paper 

entitled "Space Station Truss Structures and Construction Consideration." 

For a comparison study of the three approaches, overall dimensions of a 

reference Space Station \?ere used. These dimensions are shown in Figure 

4.4.8.2-1 which depicts the deployable-single Pol d structural truss 

arrangement. The overall size is representative of a gravf ty gradient 

stabilized station with a solar array that will Be1 iver 75 kW continuous 

power. 

4.4.8.2.2 Truss reauirements 

The following list contains some of the requirements that have been 

identified for the Space Station truss system: 

a. Provide a structural foundation for construction of the 

Space Station 

b. Provide a surface area for the attachmnt of payloads and 

utility 1 ines 

c. Provide a structural stiffness that will minimize Space 

Station control probl ems 

d. Provide a road bed for a track system that wil I allow the 

use of a Mobile Remote Manipulator System (MRMS) to aid in Space Station 

construction and transportation of payloads 

e. Provide a redundant structure that will offer alternate 

load paths if a member of the truss is damaged so that strength integrity 

of the Space Station is not impaired 

f. Provide structural repair capabil ity without the loss of 

structural integrity 

4.4.8.2.3 Deployable-single fold beams 

Due to the need for redundancy in the structural arrangemnt, a box beam 

with four 1 ongerons and diagonals between the 1 ongerons was considered. 

In determining the size of the beam, it was considered desirable to make 

the bean cross-section as large as possible to provide bending stiffness, 

but confined to the dinensions of the Shuttle cargo bay. In addition, a 

large cross-section will provide a wide track for the MRMS and the 

payloads it must carry, and also provide a large open area inside the 

packaged bean for incorporation of uti 1 i ty 1 ines. 

The Shuttle cargo bay can accept a 14-foot dlameter payload to transport 

to orbit. This diameter establishes a maximum upper limit on a square 

cross-section of 10 feet. For the current study, a maximum cross-section 

size of 9 feet has been established which will allow additional space in 

the cargo bay for external beam fitting attachments for installing solar 

arrays, radiators, and etc. The beam chosen for this study is an 
orthogonal tetrahedral design, having longerons that fold inward and 

diagonals that telescnpe to effect a compact packaging configuration. In 

addition, a provision was made to incorporabe quick-attach joints at the 

sf de of each bay to pe-ni t add1 tion of additional beam structures. This 

configuration is shown in Figure 4.4.8.2-2. 

The packaging characteristics of this beam are such that each bay 

compacts to a dimension equal to 2-longeron diameters. For this 

particular example, a 216-foot long beam can be deployed from an 8-foot 

long package using 2-inch diameter longerons. 

Deployment of the beam is a sequential and control led operation in which 

one bay unfolds at a time so that instabilities normal to the deployment 

direction may be avoided. A deployment cani ster can accomplish this task 

but will have to be discarded after the deployment because it would 

interfere with the MRMS movement. Enerw schemes to provide deployment 

can include precompressed springs in the member joints of each bay, 

electric motors with lead screw or chain drive mechanisms, or use of the 

MRMS . 

Space Station assembly analysi s indicates that it will take a minimum of 

two Shuttle flights to provide the structural foundation required for the 

reference station configuration condi dered in this study. The basic 

limitations are the size of the beam required, the folding 

characteristics of the bean, and the size of the Shuttle cargo bay. 

4.4.8.2.4 Erectable beam 

The erectable beam concept has the advantage over the deployable beam 

just discussed that the maximum Sh~ttle cargo bay diameter constraint is 

eliminated. Erectable beam members do not rely on folding techniques and 

can be packaged as individual pieces which will all ow a greater packaging 

dens1 ty. The only limitations placed on this type of construction is the 

individual member length which should be compatible with the equipment 

storage and function of the Space Station. Tn this section, two 

different sizes of erectable beams were considered. The first has a 

9-foot cross-section identical to the deployable beam of the previous 

section. The second has a 15-foot cross-section which will not only 

increase the beam structural stiffness but will also be compatible with 

the attachmnts for a 14-foot maximum diameter Shuttle payload. In both 

cases, the bays should be square so that the MRMS can move in orthogonal 

directions as required. 

The principal difference between the erectable and deployable beam of the 

previous section is that the beam has to be constructed while in orbit. 

This task will require a large amunt of EVA for the :onstruction 

personnel. In addi tioil, the utility 1 ines that could be incorporated on 

the deployable beam will have to be added to the erectaole beam, However, 

the erectable beam has the advantage of requiring only one Shuttle flight 

to provide the struct~ral foundation for the reference Space Stati~n 

configuration. 
4.4.8.2.5 Deployable-double fold truss 

It Ss generally understood that a dwble fold structure is a more 

efficient mans of packaging truss structures for transport in the 

Shuttle cargo bav than the single fold concept. Because of the high 

packaging efficiency of the double fold concept, it would be cansidered 

feasible to deploy a larger than required truss in orbit to minimize 

subsequent add-on structure for future Space Station growth. 

The double fold truss considered for this stu@ was made of a tetrahedral 

configuration. iGgure 4.4.8.2-3 shows the reference Space Station using 

the tetrahedral truss as the foundation structure. Figure 4.4.8.2-4 

shows a single tetrahedral element of this truss configuration. The 

melabers of this element were considered to be 10-feet long for thd s 

study. It should be noted that ttre inherent characteristics of the 

tetrahedral eleraent requires that all ders have the same length 

thereby reducing nanufacturing costs and spares requi resents . In 

addition, this truss system is highly redundant which will allow 

alternate structural load paths and preserve strength integrity in case 

some are the members are damaged. 

The upper and lower surface members hinge at their rfdpoints for cmpact 

folding. As a result, the packapd truss becomes a function of the truss 

node fitting dimensions (or member diameter) and the length of the 

individual truss member. It was shown in this study that this packaging 

concept will allow the entire Space Station truss system to be installed 

in the Shuttle cargo bay leaving morn for addi tiondl packages such as 

solar arrays, radiators, etc. This implies that the complete Space 

Station structure can be delivered to orbit in one flight. 

Two methods of deploying the truss were considered. In both methods :he 

deployment energy was considered to be prestressed springs in the hinged 

joints of the top and bottom surface menbers. Ttd flrst method is a free 

deployment in space that required the package to be removed fra the 

cargo bay using the RMS. While supporting the package with the MS, 

aciivirte the control led depl oymnt lnechani sm that allows the truss to 

expand at a controlled rate. The second method utilizes a set of rails 

that can be attached to the Shuttle cargo bay longerons. The packaged 

truss is then inserted in the rail system by the RMS. The rail system 

can then serve as the control led deployment device; and also stabilize 

the package, freeing the &!S fnr other duties. 

4.4.8.2.6 Comparison study 

The three approaches studird are be1 ieved to generally represent the 

major techniques for constructing the Space Station foundation. 


, Folding longcron 

Telescoping dia

End view Side view 

Flgure 4.4.8.2-3 - SCHEMATIC OF DEPLOYABLE BEAM SHOWING ONE BAY B

AND DETAIL OF JOINT 
I 3 stowed bays 

I I 

Bay-alze payload element 

Packaged structure 

First bay erected for 

g deployment stablllty 

ua 

Flrst bay deployed 

fully deployed structure 

I 

Flguro 4.4.8.2-4 - TRUSS DEPLOYMENT SEQUENCE AND RAIL DETAIL 
Although the three different constructions ai-e not exactly comparable, 

they are similar enough that general observations can be made. The 

parameters selected for characterizing each const.ruction approach are 

part count, structural weight, and structural stiffness because of their 

relationship to the structural design, fabrication costs, and performance 

of the Space Station. In an attempt to achieve commonality between 

constructions, one- and three-bay wide versions of the 9-foot deployable 

beam, a 15-foot erect-able beam, and a six-bay wide deploydble tetrahedral 

truss were examined. 

The results showed that the 15-foot bay erectable beam resulted in a 

structure which has about half as many parts and weighs half as much as 

the +foot deployable beam, yet possesses 3 times the stiffness. The 

deployable tetrahedral truss is seen to have 50 percent or more parts 

than the 9-foot deployable beam but has only a slightly higher weight. 

In addition, the tetrahedral truss is over twice as stiff as the 9-foot 

deployable beam but only $1 ightly stiffer than the three-bay version. 

Some appropriate papers can be found in Reference 9 through 33. 

4.4.8.3 Prel iminary Structural Design and Analysis of a Shuttle-Launched 

Space Station Manned Habi tab1 e Modul e 

Previous studies have shown that generally the structural sizing of a 

Shuttle-launched S~ace Station module is governed by the loads 

experienced for the Shuttle flight environments rather than the activity 

that the mudtile is to perform while attached to the Space Station. As a 

result, it could be considered feasible to design a single structural 

module shell for Shuttle flight environments that would serve a 

general -purpose function for a1 1 manned and unmanned activities planned 

for the Space Station. This concept would permit an assembly line 

production of module shells requiring only one set of tool ing. Interior 

equipment installation at the end of the assembly line is all that would 

be required to transform the structural shell into a special identity as 

a habi tatation, service, or 1 aboratory mdul e. 

At the time this Space Statiirn module study was initiated, the preliininary definition of the module was out1 ined as a cylindrical pressure 

vessel 34 feet long and 15 feet in diameter. Tne general structural 

configuration chosen for thi s study intended to optimize the structural 

zomponents so that they could perform a multifunction purpose. 

The following sections summarize the structural design requirements that 

were defined at the time the strldy began ard the preliminary design work 

that was done to establish the module integrity to meet those requirements. Detailed information can be found in the White Paper of the same 

tit?c as this section. 
4.4.8.3.2 Module design requirements 

The fol lowing list specifies general requirelnents that were identified 

for thi s study. 

a. The module should be lade of materials that will provide a 

service 1 ife of 10 years or mre wi thwt intermediate refurbishment 

b. Module gross weight and overall dimensions will not violate 

shuttle and pqyl oad bay constraints 

c. Provide strength and life integrity to sustain a manned shirt 

sl eeve-envi ronntent of 14.7 psi a 

d. Provide adequate internal attachment stv~cture for module 

function configuration 

e. Provide neteoroid/debris protection at a 95 percent 

probability of not havins a penetration for 10 years 

f. Provide docki ng/berthing capabil i ty to other mdul es an,d to 

the Space Shuttle 

g. Provide dual ingress and egress capability 

h. Provide windows for observation 

i. Impose structural ultimate factors sf safety for structural 

design and analysis 

I. Factor of safety of 2.0 Tor pressure loading 

2. Factor of safety 1.5 for mechanical acd therurl 

1 oading, 

4.4.8.3.3 Proposed module construction 

Ilplied in the requirement for a module service life of 10 years or 

better is the selectioq 04 a structural material that will not suffer 

erosion from the space environllent as we1 1 as mintain sufficient 

strength to endure lasltiple repetitions of pressure and them1 cycling. 

Also iaplied in the requirent to maintain an internal pressure of 14.7 

psia is an acceptable leak rate in the pressure vessel. It is expected 

that it would be difficult to maintain an acceptable leak rate for a 

module made of convantiona? riveted and mechanically fastened 

skin-stringer construction for 10-years tenure; therefore, to satisfy 

both the life and material requirenents, it is proposed that the module 

be constructed of all welded integrally machined skin-stringer panels 

using 2219-1851 aluminum plate. This material has good strength, a high 

fracture toughness, a good resistance to stress corrosion, mod 

weldability, and good machineability. It would be expected that 
structural certification could be minimized by devising a simple proof 

pressure test on each module to verify the 10-year strength and life 

integrity requirement. 

Shuttle cargo dimensions dictate a maximum 15-foot diameter module. 

However, there are restrictions to this dimension due to deflections of 

the payload in the cargo bay. These restrictions are defined in the 

Space Shuttle System Pay1 oad Accommodations Speci fication and usual ly 

restrict the maximum payload diameter to a 3-inch clearance between the 

pqyload and the Shuttle structure. The maximum diameter chosen for the 

module design was 14.5 feet. Also, maximum gross weight for the Shuttle 

pay1 oad is 65,000 pounds for ascent but restricted to 32,000 pou;lds for 

landing. However, the 32,000-pound landing weight may be improved by 

applying suitable restrictions to reentry maneuvers and 1 andi ng sink 

speed. A representative 43,000-pound gross weight payload was chosen for 

this design study. Figure 4.4.8.3-1 shows the overall module geometry 

used ia this study. 

Internal attachment structure for equipment and experiments will be 

provided by the ring frames inside the module. With the exception of the 

berthing segment, the module cylinder will have 5 equal bays of ,55.6 

inches that contain heavy ring frames for the attachment of equipment, 

experiments, floors, and other hardware required for the specific nodule 

configuration. The ring frames should be designed for in-plane loading 

only since out-of-plane loads will cause twisting of the frames and 

produce a heavy weight design. Any out of plane loads should be sheared 

to the skin-stringer panels of the nodule. 

Meteoroid/debris protection of the module must be designed to provided a 

95 percent probability of not having a pressure skin penetrated in 10 

years. The concept chosen for this design was the double wall bumper 

which has a .a45 inch thick aluminum bumper located 2.0 inches away from 

the module pressure skin. The 2-inch gap is filled with multilayer 

insulation to provide thermal control and may a1 sq improve the 

meteoroidldebri s protection slightly. Since there is a requirement to 

prevent crushing of the multilayer insulation, there will be nonthermal 

conducting stand-offs between the two aluminum sheets that will keep the 

bumper from deflecting into the insglation. 

Docki ng/berthi ng capabi 1 i ty wi 11 be desi gned into the module conical end 

domes as well as a cylindrical segment in the side of the module. These 

various berthing positions will allow greater flexibility in generating 

several Space Station configurations from a common module design. It is 

expected that the module will be securely anchored to the Space Station 

by means other than the docking/berthing ports. Therefore, these ports 

should be designed for li~ht berthing loads only and not be required to 

support the Space Station i n t e g r i t v . Thc current port design includes a 

berthing mechanism support structn~rz thdt has a 70-i nch clear diameter. 

It has been estimated that the actual hatch diameter is 50 inches with 

the remainder of the area being dedicated to disconnects for plumbing, 
electrical, air conditioning, efx. The cylinarical segnent in the side 

of the module contaf ns four berthing ports located 90' apart for the 

reference Space Station configuration. Thi s would allow the nodule to be 

rigidly attached to the truss structure by the same trunnions used to 

support the module in the Shuttle cargo bqy and pernit a var-iety of 

choices for berthing wfth other modules or the Shuttle. 

The cylindrical portion of the module has been designed wfth four 16-inch . 

diameter windows that can be used for observation. A double pane concept 

allows a redundant pressure pane to save mdule pressure integrity f f the 

primary pressure pane should break. 

4.4.8.3.4 Module-to-Shuttle attach points 

Design loads for the module are strongly dependent on the module center 

of gravity location within the Shuttle cargo bqy. For this study, the 

mdule will use the Shuttle attach di ttings for deployable pqyloads and 

an active keel fitting. Available fitting locations for the mdule !as a 

pqyload) are given in the Space Shuttle Sy?~tem Pqyload Accw~lodatfons 

Interface Control document. A fi w-point attachnent system (four Shuttle 

longeron attachraents and one-keel attachment) wi 11 be statically 

indeterminate. However, it i s possible to make certain simp1 i fyi ng 

assunptions that will render the problem statical ly determinate. 

It will be advantageous to locate the fomard and aft longeron attach 

points as far apart as possible to reduce the module reaction loads 

entering the Shuttle longeron. For this study, the trunnions will be 

located at the forward and aft ring frames of the cylindrical module 

sectior; approximately 560 inches apart. The location of the mdule 

within the cargo bay was determined by the load capability of the Orbiter 

longerons. The large 1 iftoff and landing loads that were detemined in 

the next section will require that splitter beans be used at both the 

fomard and .1,Zt trunnion locations to divide the mdule reactions between 

adjacent Shuttle frames. Figure 4.4.8.3-2 shows the mdule positiorr in 

the cargo bay that maximizes the Orbiter load carrying capability. A 

forward splitter beam is needed to distribute the high z-direction loads 

while the aft splitter beans is needed for x-direction reactions. 

Module design loads -

Shuttle ascent and descent load factors were obtained for the module 

which were based on prior Shuttle payload experience. It was noted that 

the angular accelerations appeared to be rather high particularly for the 

landIng condition. However, these accelerations could not be confirred 

until a math model of the mdule could be constructed and processed by 

the corlputer program. This work is still in progress. Using the load 

factors as gi ven, ultimate module-to-Shuttle attach reaction loads were 

pnerated assitming the mdule to be a simple beam and making certain 

assu~tions to render the probl stcitlcal?y determinant. Th.2 result4 ng 

redction lcrads were used to size the atsch ring frames at stations Xo 

862.5., Xo 1055.5, and Xo 1222.5. Cirimate shear and bend',ng mmnt 
diagrams about the module centerline were also generated for the critical 

load conditions and used to size the integrally-stiffened module skin. 

An ultimate factor of safety of 1.4 was used for the module strength 

integrity afid a factor of 2.0 was used for the mdule pressure vessel. 

For manned habitation, the module was considered to contain a "shirt 

sleeve" environment of 14.7 psia. Pressure conditioning for astronauts 

will be reserved for airlock modules and are not a part of this study. 

The maximum thermal environment for the module was considered to occur 

while the module is in orbit and was determined to varying between t15U°F 

and -20°F; however, thi s structural loading was considered negligible due 

to the anticipation of applying thermal insulation and coatings over the 

module surface. 

On-Orbi t module 1 oads considered combined inrernal pressure and berthing 

loads. Berthing loads considered a worse case condition which was based 

on the assumption that the Shuttle would be berthed to the Space Station 

and cantilevered from one module. Using a Shuttle (plu pqyload) mdss of 

269,000 pounds and a berthing acceleration of 3.8 X 10-a g's at the 

Shuttle and a 40-foot moment arm, the bending and axial l~ad on the 

module berthing mechanism is 49,100 inch-pounds limit. The axial load is 

lC2 pounds 1 imi t. 

4.4.4.3.6 Meteoroidldebris protection 

The mete~roidldebri s protection criterion for the manned habitable module 

has been established as 95 percent probability of not having a 

pene~ration of the module pressure skin for a 10-year 1 ife in the space 

environment. Using this criterion and the meteoroid/debri s flux, the 

estimated bumper thickness should be .045 inches for a 2-inch spacing 

between the bumper and the module skin. There is some question that the 

.070-inch thick module skin will be thic~ en~ugh to resist the impact 

loading caused by the limited 2-inch spacing between skins. However, 

this criterion will be used until the flux can be better defined. 

Due to the large unsupported arzas of the bumper, nonconductive standoffs must be used between the bumger and the module skin to maintain the 

2-inch spaci~g and not crush the multilayer insulation. Crushing the 

mu1 ti 1 ayer insulation will cause degrading of its thermal characteri stics. 

4.4.8.3.7 Primdrv rina frame desian 

The module reaction loads are transmitted to the Orbiter structure by the 

trunnion and kecl fittings which are attached to the primary ring frames 

of the module. Three primary rings are proposed for the module: the 

forward ring at station Xo 862.5, the midmodule keel ring at Xo 1055.5, 

and the aft ring at station Xo 1222.5. Since the trunnims are offset 

from the ring section neutrdl axis, the reaction loads will result in a 

combination of thrust and shear loads as well as bending moments that 

m.~st be used to size the r'ngs. 
A 1 iyhtweight ring design would certainly adhere to a varying load 

distribution around the periphery of the ring and dictate thin webs and 

beam flanges in areas where the loads are low. However, for this 

exercise it was assumed that the ring would maintain a constant crosssection that would a1 low addition,,: stiffness for attachment of equipnsnt 

inside the module. These ring sections will have to be checked as soon 

as the mass cf the installed equipnent becomes known. Preliminary s?zirig 

of the ring frames assumed that they would appear as a simple I-beam 

section. 

4.4.8.3.8 Intermediate ring frame design 

The intermediate ring frames within the module will be requited to have a 

multipurpose function of providing a sol id support for the module skin 

panels and providing a solid support structure for the equipment mounted 

inside the module. Experience has shown that the ring frame stiffness 

required to insure skin panel stability is much less than the stiffness 

requirer to support the in-plane forces generated by the mass items 

mounted the frame. Therefore, the sizing of the intermediate ring 

frames should concentrate on the forces generated by the qufpmnt mass 

i terns. However, since these forces have not yet been defined, the 

prel imiaary sizing will focus on the forces applied by the window and 

berthing ports and the results assessed when the equipment loads and 

their iocations become available. 

Preliminary sizing of the ring frames will consider that all intermediate 

rings afe identical and that the sectional properties will remain 

constant around the circumference of the ring. This will add material 

where it is not needed for the present design load conditions, but this 

additional material may be needed for the equipment loads to be 

determined 1 ater. 

Maximum intermediate ring loads were determined from the combined 

internal pressure and berthing loads that would occur at one of the 

c y l i n d r i c a l segment berthing ports. These loads are transmitted to the 

ring frame by means of the berthing support structure which spans the 

primary forw~rd ring and the first intermediate ring at statio~~ Xo 

944.5. 

Cylinder design 

In this study, the ,witnary load carrying structure is considered to be an 

integraliy stiffened skin with ring frames. The skin will resist the 

pressure loads for habitability and also be stiffened by strinqe-s 

equally space around the inner circumference to resist bodv bending and 

axial loads from the Shuttle flight environment. 

Placement of the stringers on the inner surface of the cyl inder skin wil l 

provide a clear area between the meteroid shield and the skin for the 

insulation installation. Ring frames designed in the preceedfng sections 

will be used internally to stiffen the thin-walled shell and also provide 

-n -- a;
material for the attschment of equipment and bulkheads for the module 

function. Additional structure will be required at the in~:ersectlon of 

the module and aft trunnions to distribute the Nx shear loads and prevent 

large torsional moments from being induced in the aft primary ring frame. 

4.4.8.3.10 Conical end dome design 

A truncated conical she1 1 has been proposed as the end closure 

configuration for the pressurized module. The primary load carrying 

structure for the domes is considered to be an integrally stiffened skin 

with ring frames. The skin will be required to resist the internal 

pressure 1 oads of the module and the integrally machined stringers will 

be required to resist the bending and axial loads from the Shuttle flight 

environment and Space Station berthing environment. 

Loads analysis indicated that the critical design condition for the cone 

occurs for combined interndl pressure and berthing. Structural 

arrangement for the truncated cone resulted in placement of 20 integrally 

machined stiffeners equally spaced around the inner surface of the cone 

skin to react the axi a1 and bendi ng 1 oads. Skin thickness to resist 

internal pressure loads was variable depending on the location of welds 

due to the reduced allowables for an as-welded ~naterial. However, a 

minimum thickness of .07 inches was retained in the piire membrane areas 

due to the meteoroi dldebri s requi rernent. 

4.4.8.3.11 Window segment design -

The aft cylindrical segment of the module will contain four 16-inch 

diameter double pane windows located 90" degrees apart around the 

perimeter of the cylinder. The structural requirements for these windows 

are to contain the modul e pressure and provide meteoroi d/debri s 

protectio~ for the crew and equipment. It is also desirable to be able 

to replace these windows while in orbit if they should become damaged. 

The outer window pane will be required to act as the redundant pane and 

also the meteoroid/debris shield. This pane will be considered to be 

unloaded as 1 ong as the inner pane is in place. However, it will be 

required to carry all the pressure load if the inner pane becomes 

broken. The air gap between the two panes should be vented to the space 

environment in order to keep moisture from forming on the glass 

surfaces. The venting should be control led such that it will produce 

only a slow leak if the inner pressvre pane should break and can be 

sealed off completely while the broken pane is being replaced. 

The design of the pressure pane should contain a preloading mechanism 

that will keep the glass in a compression state of stress through its 

thickness. This compressed state of stress will keep any flaws in the 

glass frorn growinn when pressure loads are applied. Feasi bil ity of such 

a preloading mechanism has been studied and approved Ly structural 

engineers at JSC. It is also recormended that both panes of window be 

made identical to reduce manufacturing costs and replacement complexity. 
The structural arrangement of the module cylinder segment containing the 

four windows will be identical to the other stiffened cyl inder sewcts 

presented in Section 4.4.8.3.9 with the exception of the additional 

stiffening required for the window cutouts. Module bending and axial 

loads will not be reacted by the windows. 

4.4.8.3.12 Berthing segment design 

Four berthing ports are located in the cylindrical body of the mdule to 

provide flexibil i ty in Space Station construction and configuration. It 

was fwnd that due to the large cut-outs required to fit the berthing 

mechanisms, the berthing seguent should be located at the forward end of 

the mdul e to preclude interference with the mdule-to-Shuttl e attach 

structure. In addition, the module design loads are much lower in the 

forward end for the flight environment. Therefore, the critical design 

loads for the berthing segment emerged as the on-orbit conditions of 

combined internal pressure and berthing 1 oads . 

4.4.8.3.13 Module weight summary - 

Primary structure weight for the module containing the four berthing 

ports is summarized below. It should be noted that the primary load 

carrying structure weight is approximately 25 percent of the total 

assumed gross module weight which is in keeping with earlier studies 

performed on modules of different configurations. It should also be 

noted that the weights calculdted for the meteoroid/debri s bumper system 

and the thermal protection system are not charged to the primary 

structure system. 

Cy 1 indrical Sidewall Assembly: -. 

(1) Skin/Stringers ........... 2561 1 bs 

(2) Primary Ring Frames ........ 1508 1 bs 

(3) Intermediate Ring Frames ...... 957 lbs 

(4) Berthing Segment .......... 3489 Ibs 

(5) Window Segnent. ........... 1101 lbs 

(6) Trunnion Longerons ......... 667 lbs 

TOTAL 
Primarv Structure 

Cylindrical Sidewall Assy ........ 10283 1 bs 

Conical EndDomes ............ 630 lbs 

Total Primary Structure ......... l(lm37 bs 

- Subsystems 

Meteoroid/Debri s Shield ........ 946 I bs 

Thermal Insul~tion .......... 180 lbs 

Remaining Subsystems ......... 28044 1 bs 

Totdl Subsystems .......... .24171Jlbs 

Secondary Structure - 

10% of Subsystems ........... 2917 lbs 

Total Secondary Structure ....... xms 

Total Launch Gross Weight .... 43000 Ibs 

a;

-. --- * 


4.4.8.4 Module and Pqyload Att;~h.~lent 

4.4.8.4.1 - Introduction 

Successful construction and operation of the manned orbit'rng Space 

Station is dependent on the mthods thtit are used for the attachmnt of 

mdules, pqyloads, and other equipllent YO the truss 5trul;ture. In this 

section, the issues associated with attac-hinq these cqonents to the 

truss structure are discussed. 

4.4.8.4.2 Basic concepts 

Two scenarix for attaching modales and payloads to the truss have been 

studied. Both methods use the Orbiter cargo bay tmnnion hardware which 

will allow this hardware to serve a multipurpose function. The purpose 

of these scenarios was to investisate methods for attactnent that can be 

easily m~ipulated by personnel in the space envirom2nt and also provide 

adequate structural load paths. 

Figure 1.4.8.r-1 shows the first scenario in which a uadule (or pqyload) 

has heen located in its app-oximate position on a p'ianar truss by the 

Orbiter RMS. Next, structural tripods and bipods &re instal led on the 

trunni~ns and attached to appropriate triss nc:&5. Finally, the trtpod 

and bipod strut assemblies are adjust?a by a 7m;k/unlock telescoping 

feature designed into the struts. After st final adjusttwnts, the 

struts are locked in place and the RMS is rt;lk~,ad. 

Figure 4.4.8.4-2 shows the second scenar.io which uses foul tripod strut 

assemblies which are mounted to appropriate truss node points to fom a 

pseudo-Orti ter cargo bqy. The module (or pqload) is then removed from 

the Orbiter cargo bay by the RMS ctnd placed into the trunnion fittings of 

the tripod fixtures. Final adjustment is made for any aisalignnent, the 

trunnions are locked in place, and the RMS is removed. 

Both these scenarios indicate that attachmnt struts makirlg up the bipod 

and tr~pod assemblies should be designed for adjustment in length to 

all ow for unplanned distortions due to manufacturing to1 erances and/or 

thermal expar.si~n and contraction. Both also advwate using the sane 

trunnion attachnrentr yovided for plpyl oad storage in the Orbiter cargo 

bay. Finally, both require attachment to truss mdes rather than 

attachment to a truss mbcr. The following sections will elabarate on 

these strut/attachraent requirements. 

4.4.8.4.3 -- Strut a4justment -

Studies have shown that there will be unplanned misal igmnts between the 

pqyload .*nd the truss node gennetry which wi 11 require adjustlmt by 

Spce Statuon personnel prior to locking the pqyload to the Space Station 

scruc core. 6zsical ly , small mi sal i gnnrents will be caused by 

wnubccluring i': srances in the hardware and :hem1 distortions due to 
thermal gradients and di ssimi?ar. materials. However, camnonal ity of 

hardware will require that the attaching struts have large lengtn 

adjustments to accomnodate various ~ay load sizes and shapes. 

Figure 4.4.3.4-3 shows an adjustable strut configuration developed for 

this study. The telescoping mechanism in the middle of the strut, a1 lows 

for large edjustments in length. The turnbuckle at the top of the strut, 

allows for fine adjustments. Details of the teiescoping mechanism are 

shown in Figures 4.4.3.4-4 and 4.4.3.4-5. The screw-jack turnbuckle 

mechanism is shown in Figure 4.4.8.4-6. Also shown in this figure is a 

proposed trunnion fitting tnat will cradle and lock the pqyload trunnion 

pin to the strut mounting structure and a lanyard attachment that will 

a1 low quick release at the trunnion fi ttinglstrut interface. 

4.4.8.4.4 Trunnion attachments 

Payloads trarsported in the Orbiter cargo bay contain attachment points 

that have been designed for the Orbiter flight 2nvironment. Previous 

studies have shown that the flight environments are more critical for the 

design of these fittings than the Space Station environmect. Therefore, 

these fittings are more than adequate for Space Station attachments. 

4.4.8.4.5 Truss node attachments -

Due to the inherent characteristics of truss desiqn, all loading should 

be introduced at the truss nodes where there is a collectio~ of 

structural members to share the load. In addition, ccmfiguration of 

trusses studied in this report dictate that all node joints are 

sssential ly pin-ended which will react concentrated loading but have no 

resistance to bending moment,. For this reason, all truss loading should 

be introduced at the node points in such a manner that there &ill be no 

induced bending. 

Figure 4.4.8.4-1 :,bows a design concept fcr a tetratruss node fitting. 

The payload attzchrllent strut is pi;~ned to a swivel in the center of the 

fitting that elimirlates bending in $40 planes. In addition, the fitting 

was designed so that the strut load path would pass through the 

intersection o' the load patiis of the truss members. This arrangement 

will eliminate induced bending in the remaining plane. 

4.4.8.4.6 Attachmen: to a squarc -- truss 

Proposed concepts forb mount4nq ,mdul,!s to 9-:00t and 15-foot square 

trusses are shown in Figure, 4.4.8.4-8 through 4.4.8.4-11. The strut 

arrangements shown ill these figures are somewhat different from the 

arrangements depicted in the scenarios due to the variation in truss 

configurdtizns. Hobever, these attachments still form a statically 

stable system. 
4.4.8.4.7 Concludina remarks 

The tnrss structure being proposed for the Space Station presents a 

unique design feature of presenting a "peg board" to which various 

pqyloads an+ modules can be attached. This section presented preliminary 

design concepts that take advantage of this inherent truss capabili Q. 

Hawever, these concepts do not represent a closed set of attachent 

designs but rather present a starting point for future designs. 

Versatility, rigidity, stability, structural load paths, corronalitj of 

hardware, and ease of crew instal lation are of prime iaprtance !n desigr 

of this system. 

4.4.8.5 Module to Mule Interface, Docking/Pc thing 

In the interest of collllronali ty of hardware and . ;ximum operational 

flexibi 1 i ty, the module-to-mdul e interfaces ana docki ngberthing 

interfaces are geometrically identical. That is, any interface will mate 

and latch to arW other interface, including completion of a pressurized 

transfer interface. Tk per-forpance requirements for the 

module-to-mdule interface are based on (1) its use as an assembly 

interface in an essentially static a1 ignnent, acljustmnt, interconnect 

operation and (2) its use as a passive port for docking or berthing sf 

the Orbiter, which will carry the active &cking/berthi~o mchani sa. 

Alternately, the station interfaces identified as berthing ports could 

also include active docking/berthing mechanisms but that is not the case 

for the reference configuration. The Orbiter must be given the 

capabi 1 ity for total active control of the rendezvous, approach and 

dnrkina/berthing task routinely because only the Orbiter has significant 

translation capability and for contingency cases because the Station 

systems may not be woric:ng or control of the Station systems mqy be 

inaccessible to the crew. 

The module-to-module interfaces will be identical throu@out the Station 

in the reference configuration. The interfaces feature a 6-inch 

extension capabil i ty to faci 1 i tate the mdule assembly and raodule change 

out processes. In the IOC mdule pattern, side-wwnted interfaces are 

mated orlly to end-mounted interfaces, so a clear o~tion exists to place 

all the extension capability in the end ports. This would significantly 

increase the internal useable volullw! in the vicinity of the side ports. 

4.4.8.5.1 Module-to-module interface - 

The pressurized modules will be joined Lo each other using either the 

Orbiter RMS or the Station Manipulator in a highly controlled, law 

velocity berthing process. There should be no need tor active shock 

attenuation within the interface mechanisms. Currently developed 

assembly scenarios involve hunting of the mdules-to-truss structure 

prior to conr:rticn of the pressurized interfaces, so the latter 

operation will be basically static. Mounting of the mdules-to-truss 

structure will be a berthing process quite similar to berthing a payload 

in the Orbiter pqyload bay. The design drivers in terms of loads 
Cn 

01 

1. POSlTlON MODULE IN PLACE 

f. 

2. INSTALL TRIPO

IN PLACE 

3. ADJUST POSI TION, LOCK STRUT

REMOVE RMS 

Figure 4.4.8.4 -1 - SEQUENCE OF INSTALLING MODllLE ON TRUSS-M
1. INSTALL TRIPODS IN PLACE 2. POSITION MODULE 

3 F 1 NAL I NSTALLAT I ON 

Ftgu?~ 4.4.8.4 -2 - SEQUENCE OF INSTALLIN0 MODULE IN PLACE-METH


/-- 

ECCENTRIC CAM LEVER 

n (UNLOCKED POSITION) 

\ f f ECCENTRIC 

(LOCKED P

-- CI AT CD

L~~~l -ROTATION PIN 

LONQITUTIONAL SLOT IN INNER TUBE 

SECTION 8-8 

Fiauta 4.43 4 ,5 - CROSS-SECTION THROUGH TELESCOPING 
T CAM LOCK LEVER 

r MODULE TRUNN 1 O




v'fiir;c?.,qL P$.t:;. ,.F 

''F POOR QL'ALITY 
and mechanical flexibility for both the trussinounting and pressurized 

inter-faces will be those associated with alignrent and adjus-nt both 

for assembly per at ions and for control of the load paths for the 

assenbled configuratf on. 

The general design requirements for the aodule-to-adule interface are 

. Physical interface capable of mating with any other 

adule-to-mdule interface or to any docking/berthing port 

b. Alignment and adjustment capabf li ty consistent with 

develop~rent of the module pattern specf fied 

c. Capabil ity for retraction of interface to provide 

*- zarance for dule reroval 

d. Provision of pressurized clear passage uqy consistent with 

a 50-inch diamter hatch opening 

e Provi sion of space for util i ties interconnects 

f. Provision of a latch system capable of accmmdating 

pressure loads and dynamic joads in the mated configuration. 

One design concept which meets these general requiremnts is sharn in 

Figure 4.4.8.5--1. The physical interface incorporates the ring-finger 

guide concept as used f n the Apollo-Soyuz Tebi Project system. The 

*ides are only 4 inc;;os high, consf stent with the misal iglllents 

associated with berthing and assdly operations. Euen in the relrti vcly 

static assellbly process, the guides will assist in achieving proper 

alignmnt of the modules prior to latching of the interfaces. Outcr 

diamte:. of the interfacing structural ring is 80 inches. 

With a system of four pairs of electromechanical actuators, the 

structural interface ring is provided with an axial stroke of 6 inches. 

The primary purpose of this stroke capabf lily is to facilitate the 

assedly and disassembly process. The mdules lrqy be transported with 

interfaces retracted, which benefits the pqyload bv installation 

process. 

After latching of the structural interfaces, the actuators will be fully 

extended. There is considerable room for designer's choice in the 

latching concept. The reference concept includes the structural ring 

mounted latch shown in Figure 4.4.8.5-2. Mated loads are ~arried through 

the 1 ocked actuators. 

Once the structural interfaces are in the extended and latched position, 

separate small actuators extend telescoping pressure tunnel el emts from 

either side of the interface. Tunnel elements are shown in the retracted 

and extended positions in Figure 4.4.8.5-3. Redundant pressure seals are 

incorporated throughout the pressure tunnel system, as shown in Figure 

4.4.8.5-4. Extended lffe positive sealing is the driving requlrewnt, so 
a totally passive seal is preferred. For leak detection, the ability to 

aver-pressurize the cavi ty between redundant seals is highly desirable. 

The t~nnel system will carry only radi a1 pressure loads. Axi a1 pressure 

loads and dynamic loads will be carried through the structural latch 

system. These loads are defined in Paragraph 4.4.8.5.3.3 of this 

section. Note that the telescoping pressure tunnel offers li~sited 

lateral or angular a1 ignment capability. If increased alignment 

capability should prove to be a requirement, an extendible flexible 

pressure tunnel may be a more desirable solution. 

Utilities interconnects will form a major part of the interface tci-een 

modules and between individual modules and the truss structure. Major 

service util i ties, (electrical power, thermal control fluid and vapor, 

and data systems) will be installed on the truss structure in bus 

fashion. The thermal control fluid and vapor (amnia) system will be 

routed to 1 iquid-to-liquid heat exchangers mounted outside the mdule 

pressuriz~ volum. Power and data busses will be routed from the trvss 

to umbilical plates on the sides of individual modules. Within the 

module string, fluid, gas, and air circulation interconnects must be 

included across the module-to-module interface. In addition, the 

electrical power and data busses will also be carried thrwgh the 

mcdule-to-module interface. A detailed listing of the utilities supplied 

to each module is shown in Table 4.4.8.5-1. 

The preferred method for completion of tlti li ties interconnects across the 

modul e-to-m~dule interface is manual connection with1 .I the pressurized 

tunnel. Thi s reduces the rleed for sophi sticated auto ma ti^ t~mbi lical 

mechanisms and renders these interconnects avail able for IVA i~;spection 

and servicing. As an a1 ternati ve, the electrical power interconnects 

be*aeen modules could be mounted on the berthing interfaces rings and 

thereby automatically connected as the structural interface is completed. 

4.4.8.5.2 Androgyny, indexing 

The optimum geometry for the dtilities interconnects depends heavily on 

the requirements for androgyny and indexing of the interfaces, hatch size 

and shape, and the inner diameter of the pressure tunnel. 

With the system of four alignment guides, the modules may be connected in 

any of four orientations, 90" apart. The location and design of internal 

elements such as hatches and util i ties interconnects will influence the 

,umber of possib:e orientations for which sl ignment of all elements can 

be achieved. For example, if a conventional D-shaped hatch is used, the 

flat part of the riatches will be aligned for only one modulelmdule 

~rientation. Thp sole criterion for alignment of internal elements is 

that the elements be symnetrica! about an axis that passes through the 

edges of the alignment guides. See Figure 4.4.8.5-5. By arrangement of 

the utilities interconnects such that they are symnetrically located 

about one, two, or four axes, an equal number of 
module/mdule orientations will give total a1 ignlsent of i nbrnal 

elenents. Note that the hatch also must be designed with the same degree 

of syr~etry for perfect alignment. Two orientations require a hatch wf th 

two flats, or rectangular or elliptical shape. Fwr orientations require 

that the hatch be a square or any other poly#n divisible by 4. 

As aentioned, the favored concept for util ities interconnects f s manual 

coapl etion of the connections wi thi n the pressurized envi romnt. For 

thls concept, the intradule uti lf ties will be terminated on the 

interface base ring (to which the hatch is mounted). The area around the 

hatch will be used for manual installation of jullper connections between 

interface base rings, which in the mated position are about 30 inches 

apart. With flexible jumper assemblies, it is not mandatory that all 

util i ties interfaces be a1 igned for a1 1 module/module orientations. (Nor 

has it been established that perfect aligrment of hatches is mandatory.) 

The use of inductive couplings (cone pairs) for electrical power 

interconnects offers the option of automatic connection as the structural 

interface is completed. For the option, the cone pairs are mounted on 

the berthing interface ring outside the pressure tunnel and are joined as 

these interfaces are brought together. This creates a special case for 

interconnect geometry in that the cone pairs comprise male/female sets. 

Since a polarized connector cannot be located on an axis of symnetty, 

the rninilnurn number of connector positions is twice the desired nuer of 

wdulej~lodule orientations. That is, eight connectors (four male, four 

female) are required to accolnaodate all module orientations provided by 

the four guide berthing system. 

4.4 -8.5.3 Docking/berthing interface 

The Shuttle Orbiter will join the Space Station at regular intervals for 

resupply, exchange of crew, and return of manufactured items and waste 

products. The Orbiter will also be used ds the base for assaly of 

early Station elements and must join with those elenents alreedy in orbi t 

as it delivers a new element. The concept of loosely coupling the two 

craft through a flexible pressure tunnel and accomplishing transfer of 

materials and elements by RMS to manipulat~r hand-off (with EVA assist as 

required) has been considered. The advantage of this concept is that i t 

eliminates the need for a rigid structural coupling and contact forces 

fm berthfng or docking that might be an order of magnitude higher than 

any other operational forces applied to the station. Stay ti- of the 

Orbiter in close proximity to the station will range froc, hours to dqys. 

Atti tude control of both the Station and the Orbiter must be maintained 

during close proximity operations. In addition, the Orbiter must 

translate as required to maintain the station keeping pos! t!nn uithout 

risk of collision with the Station. Orbiter attitude control and 

trans1 ation maneuvering requires the use of RCS thrusters. Periodic 

firing of RCS thrusters in the vicinity of the Station cannot be 
permitted because of contamination and plume impingement forces. 

Therefore, joining the Orbiter to the Station by a structural interface 

through which the Station can provide attitude control of the Orbiter is 

considered mandatory. 

Concepts for bringing the Orbiter stqctural interface into contact kith 

the mating station interface must be evaluated against two primary 

constraints: (1) impingement of Orbiter RCS plumes on station elements 

must be minimized and (2) interface contact forces must be minimized. 

Recent studies have typically adopted the berthing concept, wherein the 

Orbiter RMS gri,pples the Station and maneuvers the interfaces into 

contact at extremely low velocities (0.1 Fps or less). However, the 

Station mass will far exceed the design capability of the RMS, which is 

32,000 pounds. Considerably higher masses can be handled by 1 imi ting the 

rate of movement. Further gains are possible with minor to moderate 

changes in RMS hardware and control software. More extensive system 

modifications will be required to gain full capability for RMS controlled 

berthing with the Station especially with respect to initial capture of 

the Orbiter. Th2 RMS contrdctor is currently examining the effects on 

dynamic performance of doubling the joint brake torque and equalizing the 

torque capabili ty of the shoulder and wrist j3ints. 

A practical a1 ternati ve to RMS control 1 ed berthing i s docking, whereby 

the Orbiter translates directly into interface contact using the RCS 

thrusters. Recen i f 1 i ght experience and simul ations have shown that the 

current Orbiter systems can place the interfaces into close proximity at 

rates of 0.1 fps or less, provided that the pilot has adequate visual or 

sensor coverage from which to determine control input requirements. 

Recent developments in laser sensor technology can be applied to provide 

precise final approach control information. 

The selected approach for the Station reference configuration is to 

provide an interface mechanism that is fully compatible with either RMS 

control led berthing or docking operations. Detailed evaluation of 

developing RMS capabil i ty, man-i n-the-1 oop simulation GI the Orbiter 

final approach to RMS grapple or docking, and detailed study of interface 

mechani sm desi gn and performance will be required to establ ish the 

technique that offers the best capabil i ty for achieving 1 ow-force mating 

with minirum RCS plume impingement for both normal operation and for 

design failure modes. 

4.4.8.5.3.1 Design concept 

The docking/berthi ng mechani sm wi 1 i be mounted to a telescoping module 

mounted in the front of the Orbiter payload bay as shown in Figure 

4.4.8.5-6. In the extended position the interface ring is 27 inches 

above the Orbiter outer mol dl ine. With actuators fully retracted, the 

interface remains 15 inches outside the mold line. The alignment guides 

and the interface ring are identical to the module-to-module interface. 

This is so that the Orbiter can be berthed to any module port, though the 

end ports are preferred because of the gredter clearance provided. 
The basic concept for design of a1 1 elements of the docking/berthing 

mechanism is to achieve positive mating of interfaces with mininum 

forces. The heart of the system is a laser sensor which will provide 

precise information on the relative positions and rates of the interfaces 

as they are brought into contact. A system of fwr pairs of 

el ectronechanical attenuator/actuators control the position of the 

docking interface relative to its base ring. The actuator system 

provides an overall stroke ca, ability of 12 inches. 

The structural latches are identical to those shown in Figure 4.4.8.5-2. 

An a1 ternate capture latch concept with extended redch capability is 

being evaluated. The extended reach latch (see Fig. 4.4.8.5-7) can be 

used to achieve low force control of the approaching interface while the 

interfaces are still several inches apart. This concept can expand the 

capture envelope by reducing positioning accuracy requirements, thereby 

making possible either reduced contact wloci ty or reduced RCS th~ ster 

firings, or both. 

The docking/berthing mechanism will incorporate an extendible pressure 

tunnel which mates with the module-to-module tunnel interface. Manual 

s ompletion of required utilities interconnects will be accomplished as 

for the module-to-modul e assembly process. 

4.4.8.5.3.2 Operation 

The docking/berthing interface will be brought into close proximity to 

the statioi~ berthing port, either by using a modified Orbiter RMS within 

the limit of its capabilities or by directly approaching with the Orbiter 

under RCS control. The laser range/rate/attitude sensor system will be 

active during the rendezvous and approach phase, providing the data 

required to minimize approach velocity as the interfaces are brought into 

proxi mi ty . 

After latching of the interfaces and strokicg of the actuator system to 

eliminate re1 ati ve velocities and angular rates, the actuators are slowly 

driven to the retracted position. The pressure tunnel is then extended 

from either the Orbiter or station side of the interface. Required 

util ities interconnects are manually completed within the pressurized 

interface. 

4.4.8.5.3.3 Performance 

The aligmnt guide height of 4 inches can accommodate lateral 

misalignments somewhat greater then + 3 inches in combination with pitch 

and yaw misalignments of about + 5 d'yrees. Roll misalignment of + 5 

degrees can a1 so be tolerated. The I di~--~J i;?i .ral i qament capabi! i e? 

approaches + 4 . . .-k ~..~.~rd. inches as the pitch and yaw m:,,i . . 

Thi s capdbiTi ty is consi stent with RHS design bertriiiry per-i urmance and 

with projected Orbiter docking capabil i ty if adequate information on the 

relative position of interfaces is available to the pilot. An increase 

in guide length to gain greater misalignment tolerance would impact the 
useable internal volume of the modules unless folding guides were 

provided. An a1 ternati ve i s to provide 1 onger guides on a 1 imi ted number 

of identified berthing ports and of course, on the Orbiter 

docking/berthing mechanism. 

Interface contact velocity for RMS berthing operations can be 1 imi ted to 

0.1 fps. The major problem with berthing of very large masses is that 

tip velocity rmst be severely restricted because of arm flexibility/ 

control interactions and joint dri ve toque and braking capabil i ty . Current dynamic studies indicate that it may be necessary to 1 imit the 

velocity for large masses to 0.05 fps or less. This is not a problem for 

the berthing interface mechanism. The real difficulty lies in the RMS 

track and capture maneuver. The Orbiter must position itself under RCS 

control within the working track, capture, and deceleration envelope of 

the RMS, at a very low velocity relative to the Station. For the current 

RMS, this working envelope and associated relative velocity may shrink to 

zero as the Station mass increases. As previously mentioned, significant 

improvement in RMS capability to handle large masses is possible. 

However, during evaluation of potential RMS capability it became clear 

that the requirements for 2ositioning the Orbiter for RMS grapple might 

be more severe that those being used for design of the berthing ' 

interface, so the obvious question is, "khy not simply dock?" 

Orbiter flight experience and recent flight simulations show that the 

Orbiter can fly to a re1 ati vely precise position with a residual velocity 

of about 0.03 - 0.04 fps using the normal Z RCS mode and with 3 velocity 

less than 0.1 fps using the low Z RCS mode. Flying in to dock versus 

flying in to grapple with the RMS will place the Orbiter 2d-30 feet 

closer to the station while the RCS system is still activated and a 

detected RCS failure could lead to an abort with extensive RCS thnisting 

away from the station (plumes toward the station ). The choice of soft 

docking versus RMS controlled berthing must be based on detailed study 

and simulation of 

a. Improved RMS capabil i ty to track, capture, and maieuver 

large masses 

b. Orbiter approach to RMS grapple and to closure of docking 

interfaces, speci fical ly with regard to resulting impingement of RCS 

plumes on the Station 

c. Failure modes and effects 

d. Operating capabil i ty of docki ng/berthing mechanisms 

More detailed discussion of Orbiter proxi mi ty operations and the nature 

and magnitude of the RCS plume impingement problem may be found in 

Paragraph 4.3.8.2.3.4 of the Operations Section of this document. 

Loads applied to the Station during the docking/berthing process must be 

defined through detailed simulation of the process, including simulation 

of the mechanism attenuation charac teri sti cs. The 1 arge offset between 
the Orbiter center of gravity and the docking/berthing axis renders crude 

simulations subject to significant errors. Estimates predict that the 

Orbiter berthing or docking maneuvers can be completed with an applied 

load to the station of less than 500 pounds. 

The loads across the mated module-to-module and docking/berthing 

interfaces are dominated by the pressure load, which for the reference 

tunnel diameter is 48,800 pounds. The maximum predicted dynamic 

environment, at the berthing port with Orbiter mated is an axial force of 

102 pounds and a moment o 4,089 foot-pounds. This is based on an 

acceleration of 3.8 f x 10- g applied to the Station, an Orbiter weight 

of 269,000 pounds with the Orbiter c.g. displaced 40 feet from the 

berthing axis. The reference desi~n includes eight latches which gives a 

design latch load of 6,100 pounds to accomnodate pressures and about 

6,800 pounds if the worst case dynamic load is applied. 

9" RETRACTED 

1- HEIGHT 

\- ALIGNMENT GUIDE 

4" G

.- < .'? 'f =,7;. 

,/., $-I 15'' E

+qun 4.4.8.5-1 - MODULE TO MODULE INTERFACE 
HOUSING 

Figure 4.4.8.5-2 - DOCKING/BERTHING CAPTUREISTRUCTURAL LATC
T-- TUNNEL CROSSOVER R

DOCKINWBERTHING RING 

m 

% 

50" HATCH OPENING 

ACTUATOR ATTACH POINTS 

Flgura 4.4.8.5-3 - TELESCOPINQ PRESSURE TUNNEL 

A. ONE INDEXING POSITION 8. TWO INDEXING PO

Flguro 4.4.8,s-5 - INTERFACE INTERNAL SYMMETRY FOR INDEXINO 

.. - - 
kt, 
I 

LATCY 

Flgun 4.4.8.5-7 - EmENDED REACH CAPTURE LATCH CONCEPT 

589 
4.4.8.6 Solar Panel Depioyment Mechanism 

The packaging and deployment concepts for the Space Station solar array 

el enlents are based on mi nimurn extension of the techno1 ogy developed over 

the last fw years to support various spacecraft and studies of advanced 

platforms and the Space Station. 

The heart of the solar panel deployment system is an automatically 

deployable boom structure which forms a central mast that supports array 

panels on either side. Several deployable boom concepts khich meet the 

general packaging and stiffness requirements have been devel oped. The 

reference Space Station system is not based on a specific lesign. That 

choice is left to the designer based on a more detailed study. 

The automatically deployable boms suitable for this application are not 

redundant, ei ther structuritl ly or i the depl oyment mechani srn and 

ppqcess. Further, the boom structut-2s have 1 i ttle damage to1 erance. 

Damage to a singlc element c?n cause buck1 ' _ of the total boom 

assenkly. In the reference concept, damage tolerance and J degree of 

redundant capability I s gained through overall design of the solar arrqy 

system wherein, for the IOC power 1 eve1 , eight independently deployed 

booms are included. This distribution of power systems will minimize the 

effect of boom deploy~rent failures or operational damage, and will 

simplify replacement of d boom a~ses~oly, should that be required. 

4.4.8.6.1 Packaging 

Individual 960" x 1@' solar drray panels are accor~ian folded into a 

186" x 12" x 9" rectangular Sox. Cne is mounted to either side ~f a 

cylindrical bcorn deployment canister (see Fig. 4.4.8.6-1). The boxes are 

rotated 90° about the boom aepl oyment axis such that the two boxes are 

adjacent and extend perpendicular to the boom deployment canister. In 

the reference configuration, the bases of bJo depl oyment canni sters are 

mounted to either end of the bets rotation mechanism, which is centrally 

attached to elements of the power boom truss. The deployment canisters 

are 30 inches in diameter. The total len9th of two deployment cdnisters, 

including the beta rotation interface, is 20J. inches. 

4.4.8.6.2 D~D: ovment 

In the favored deployment concept, the power boom is deployed with the 

array system in the packaged configuration just described, yiel ding the 

configuration shown in Figure 4.4.8.6-2. The indi vi duai array boxes are 

then rotated to the deploy position with an array blanket box on either 

side of the mast deployment canister as shown in Figure 4.4.8.6-3. 

Lastly, the transverse beam which forms the top of the array panel boxes 

is released and the mast extends, unfolding the array blankets (Fig. 

4.4.8.6-4). Note that the individual array blankets are attached only to 

the transverse beam at the deployed end, and to the array box on the 

canni ster e- J. The accordian-fol ded array blankets are flattened by a 

negator spring tension system fol 1 owing deployment of -;he mast. 
4.4.8.6.3 Deployable mast design 

As previously uectioned, several concepts for automatical ly deployable 

booms have been developed or advocated. These include cylindrical booms, 

which are fomd by one or more mtallic, cylindrical, thin, shells. 

Also included are coilable lattice booms, bear structures whose longeron 

and batten members are usually fiberglass rods and whose diagonals are 

typically steel cables. Another type is articulated lattice booas, whose 

longeron and batten leillbers are segrents of letall ic tubing that are not 

deformd when the boa is retracted. 

A detailed evaluatio~r of existing boom concepts, and others that might 

develop is aeeded to ev-luate deployaent reliability wi th the 

consideration of incorporating redundancy in the deployaent process. 
LL 

"\ 

%= 0 

.f! 2 u 

L XI- \ i)BdOWAL 3F POOR 
% 'J

? 

. . 
Figln 4.4.8.6-3 - PANEL STOWAGE BOX ROTATION 

Fro

4.4.8.7 Conceptual Rotary Joint Design for Space Station Appl ications 

The reference Space Station configuration has i! functional requirement 

Pcr three sets of rotation devices to providc pointing of the solar 

arrays (alpha and beta axes) and the thermal control system rsdiators 

(alpha axis only). Of the three drives, only the alpha-axis drive on the 

power boom is required to provide a CO~~~~UOUS rotation capability. The 

beta axis solar array drives (at the base of each of the array masts) 105 

degrees at a very low rate and the alpha-axis radiator drives (at the 

nadir end of the lower keel drive approximately 260" and then rewind to 

an initial starting position while the Station is on the dark portion of 

its orbit. 

All of the drives must conduct a significant portion of one of the major 

Station uti 1 i ties across the ratating interface. In addition the solar 

array alpha joints function as primary structural elements on the Power 

Boom. 

A 3-week study was undertaken at the Langley Research Center in support 

of the Space Station Project Office to develop a "strawman" design 

concept of a single continuously rotating solar array Alpha joint. The 

purpose was to point out major system issues in the design of such a 

joint and provide a guide to the areas requiring further study. 

At the time of the study inception, some configurations being considered 

required both the thermal and electrical utilities ta be carried across 

the rotating interface. T:;,.--fore, a "worst case" configuration which 

transfehSred electrical power and thermal fl uid/vapor across the joint was 

chosen for this effort. The structural issue was addressed by imposinq a 

design constraint requiring the sti ffness di stribution of the attdched 

truss structures be maintained across the joint. A more detailed 

presentation of the design requirements for the study is located in the 

specific White Paper deal ing with this effort. 

4.4.8.7.1 Joint description 

The rotary joint concept developed consi sted or three major parts: 

a. A fluid coupling 

b. A power coupling 

c. A mechanical drive coupl ing 

Figure 4.4.8.7-la and Figure 4.4.8.7-1b show the joint concept using a 

sl ip ring assembly or an inductive power -oupl ing. The fluid coupl ing 

and power coupling are located ,in tandem on the center1 ine of the jcint 

and are suspended from the mechanical drive housings. The mechanics' 

drive forms the link between sections of the power boom truss elemt,~ts. 

The fluid coupl ing consists of a ported rotary shaft inside a mating 

stationary housing. Eight fluid lines pass through the rotary shaft. 
carrying liquid and gaseous amnia for four separate circuits. A 4-inch 

diameter cylindrical condui t i s provided through the center of the shaft 

to allow fluid and electrical connections to the pwer coupling. The 

shaft extends through the housing and attaches to the power coup1 ing 

elenrents whicl- rotate with the shaft. 

The power coupling is sized to carry 100 kW of electrical power through 

either a slip ring assembly or an inductive coupling. A separate cool ing 

loop would have to be provided for the inductive coup1 ing. 

4.4.8.7.2 Fluid coup1 ing design overview -

The fluid coupl ing design concept is capable of handling the amnia flow 

required for 159 kW of cool ing, and of providing for continuous rotation 

of one side of the joint relative to the other. Figure 4.4.8.7.2-1 

provides d cross section view of the coupling. 

The coupl ing has an approximate overall length of 31 inches including the 

shaft extension to mate with the power coupling. The larger diameter is 

approximately 14.5 inches, and the small diameter about 9.7 inches. The 

overall weight is estirwted at 500 pounds based on the use of 316 

stainless steel as the primary structural material. It is possible that 

the weight could be reduced significantly, but this would probably be at 

the expense of addi tional manufacturing complexi ty. The toque required 

to drive the coupl ing (overcome seal sliding friction) is estimated at 

8655 i nc h-pounds. 

The coupling is basically a rotary shaft. Fluid will enter the coupling 

through the end of the rotating shaft, flow to annular passages on the 

shaft outside diameter, and then through tgbes in the outer housing. 

Rotary seals provide the isolation between flow paths at the 

shaftihousing interface. A radi ator/scavenger system is used to collect 

any gas or liquid which may leak past the seals. Fluid flow 

characteristics and the heat transfer bebeen circuits in the coupling 

are summarized in Tables 4.4.8.7.2-1 and 4.4.8.7.2-2 respectively. 

Materials proposed for the fabrication of the fluid coupl ing were 

selected based on the compatibility and known prior use in ammonia 

systems. Amnia is corrosive, in one form or another, to maw of the 

commonly used engineering materials. Ammonia manufacturers and suppliers 

were contacted for recomnendations and a 1 imi ted search was rdde of the 

1 i terature. The metals which have been most common::/ used with ammonia 

are carbon steel and 304 or 316 stainless steel. There are a nunber of 

nonmetal1 ic materials which are compatible with ammnia including TFE 

Teflon. The conceptual design utilized these materials since the time 

available did not permit a more detailed investigation in this area. 

The seal5 which were evaluated for this application are rotary Teflon 

seals which contain a support spring. The seals are loaded against the 

sealing surfaces by both the spring and the internal pressure. A number 
TABLE 4.4.8.7.2-1 - SUWRY OF FLOW CHARACTERISTlCS 

CIRCUIT TOT HEAT FLOWRATE PRESSURE DROP THROUGd JOINT 

NUMBER REJECTION 

k W 1 blhr psi 

Liquid Side Gas Side Total 

TABLE 4.4.8.7.2-2 - SUMMARY OF HEAT TRANSFER ANALYSIS 

CIRCUIT AVG JOINT FLUID FLUID COWD CONDUCTANCE CIRCUIT HT 

NUMBER TEMP TEMP PER LENGTH GAI N 

OF OF BT U/ BTUIIN. WATTS 

HR SQ IN.OF 


of seals of this type are available comercially. The following 

information ruppl ied by the Fluorocarbon Comparly i s typical. 

Material - Virgin TFE Teflon provides cmp=?ibility 

with MH3 and has good seal ing 

characteristics. 

Springs - CRES 316, CRES 304, or Haste11 oy springs 

Lifetime - At the rotation rate anticipated, on% 

revolution every 90 minutes, the 1 ifetime 

of the seals should exceed 10 years. 

Leakage - "Near zero" 

Surface Finish - The surfncc in contact with the seals 

shobl d be pol i shed tc a 6-8 mi croi nch 

finish for best results. 

The seal configuration which appears ~mst promising has a "C" shaped 

cross cection with a metal spring inside the section,. Figure 4.4.8.7.2-1 

shows the fluid cwpling with 18 seals in place. Each flow passage is 

isolated with a separate pair of seals. An additional seal is provided at 

each end of the coupling to reduce the possibility of arly ar,:mnia 

leakage. The rad:ator/scavenger system will condense aw leakage to the 

outboard seal cavities irr the scavenger reservoir. 

4.4.8.7.3 Scavenger system -

Aw dynamic system under pressure will leak some small amount past the 

seals. A s-ystem whfch would "pump" most of the leakage to a storage tank 

is relatively simple in concept. A small flow passage would be !ocated 

between the seals which define the annular flow passages. This aassage 

will be connected to a storage tank outside of the truss supporting 

structure. The storage tank would be: oassfvely mintairred at taperature 

of 0°F or lower by covering the external surfaces with a material such as 

silvered Teflon. Any ammonia leaking into the area between the seals 

would be "pumped" tg the t,rx and condensed on the walls. The talk couid 

be sized to accept all expected leakage for the lifetime of the station. 

Or it cculd be designed with appropriate valves and a bladder to reclaim 

the condensed 1 i qui d. 

4.4.8.7.4 - Power coup1 ing 

The power transfer techniques used in this design were derived from 

similar concepts developed by the Space Divi~ion of the General Electric 

Company under contract tu NASA Lewis Research Center. The concepts were 

pub1 i shea as NASA CR-165431 and CR-168021. Since i t was beyond the 

scope of this study to deterwine the ovcrdll configuration of the power 

di stri bution system, Go basic configurations of the power cwpl ing were 

used to establ f sh trial lqyouts for the rotary joint. 
4.4.8.7.4.1 Design overview 

Sther studies have established that power distribution throughout the 

Space Station can be most economical ly acLieved using (re1 aLi vely ) high 

volt,age AC at 5 to 30 kHz. Tra~sformers for this range of frequencies are 

1 ightweight in comparison t~ thoce used in the 60 Hz commercial 

di stributi ~n network, and cen be made quite efficient by use of modern 

maynetic materials technology. Sol ar arrays generate DC power; inverters 

and transformers are reqtlired for an AC system. Integration of rotary 

transformers into the power distribution syqtem may result in overall 

power system economies, even though it introduces considerable 

complication into the design of a rotary joint. Therefore one design 

examjned in thi 5 concept devei opinent incorporated a rotary transformer as 

a cylindrical assembly about 17.3 inches in diameter and about 9.2 inches 

between attachment flanges witti an estinated weight of 500 pounds. 

The a1 ternat i ve technique which was examined assumed conventional power 

trdnsmission !in,ofar as the rotary joint is concerned) via slip rings. 

Slip rings may transmit either AC or DC, and are not aFfected (within 

reasonable 1 imi ts) by the voltage 1 eve1 being carried or by ~21 tage 

variations. They may be high maintenance i tems ?el ative to 

transformers. Selection of material s and specific design parameters are 

critical for space appl ications. particl~larly with the very slow rotation 

rates antici~ated for tni s design. They are, however, very efficient, 

quite light, and have few packaging constraints. The unit examined in 

this study is adapted rrom a design referenced in the reports noled above 

as a cylinder about 5.7 inches in diameter and 11.0 inches between 

flanges with an estimat~d weight of 30 pounds. 

4.4.8.7.4.2 Configuration 

The transformer reported in the CR's noted above, wss designed to operate 

at pow,r leve?s similar to those required for this study, but at a higher 

voltage and at 20 kHz. However, other considerations may lead to a 

frequency of 10 kHz. Since the volume of nagnetic material required in a 

transformer is almost exactly in~trsely proportional to cperating 

frequency, ali other parameters being held constant, the size of the 

transformer was increased for conservatism from that of the noted study 

to accommodate the doub; ed magnetic volume. The traqsformer of the noted 

study had redundant windings. Since that wa; not a requirement or this 

design, no effort was made to exactly size the winding space. 

F~rthermore it was determined that the excess space could be used to 

util ize nnre efficipnt winding techniques, thereby reducing capper 

1 osses. 

The exact configdration of the transform~r is described in the noted 

reports and need not be repeated here except in genera! terms. The unit 

is of pancake cofistruction, consi sting cf fcur independent circular 

transformer systems, each with a different diameter. These are mounted 

concentrically on a common axis. The suppcrt structure is arranged so 

that the primary windings and part of the mayneiir: circuits rotate with 

one attachment flange. The remainder of tne magnetic circuits rotate 

with the other flange, at the opposite end of the unit major axis. The 
structure incorporates bearings to locate and mintain alignnent of the 

parts which have re1 ati ve moveaent. 

The slip ring assernbly used in the alternative &sign was taken directly 

fm the noted study regarding envelope, weight, and form. Cetails of the 

internal design were not presented in the original study. 

4.4.8.7.4.3 Electrical characteri stics 

The power transfer device for this stuQ was speciCied as being capable 

of passing 100 kU at a frequency of 20 kHz or less, at a voltage of 200 

volts or less. The power was to be in four separate circuits. Althou* 

not specified, it was assulned that the four circuits would divide the 

power equally, i .e., each circuit was designed for 25 kW. As stated 

above, the transformer was sized for the required power at 10 kHz. For 

constant power, frequency is the only variable which has an appreciable 

affect on the size and weight of the unit; the winding remains 

approximately the same size (volun~) regardless of the voltage, within 

the range of probable Space Station distribution voltages. No attempt was 

made to optimize the windings for a specific voltage or frequency in this 

design; the a1 located volume or the transformer of the nated report Mas 

used. 

The efficiency of the transfomr is expected to be greater than 98 

percent. This should be easily achieved, but still rill require that 2 kU 

be di csip?tzd in the transformer. Because of the small volume and 

radiative surface, supplemntal cool ing wst be provided. In this 

design, a portion of the cool ing capacity being passed through the fluid 

loop of the rotary joint is used to cool the transfomr, versus heatpipe 

cooling in the transfomr of the noted report. 

For the design of the alternative slip ring assembly, less than 100 watts 

will be dissipated in the assembly. Supplemental cooling will probably 

not be required. 

1.4.8.7.5 Mechanical drive coupling 

4.4.8.7.5.1 Structural design overview 

The structural design of the mechanical drive coupling is basically two 

concentric structural cylinders maintained in alignment and held together 

by two sets of preloaded angular contact bearings. Attached to the inner 

cylinder would be a gear driven by a WfJRrl pir,isn/ gear reducer/ motor 

drive assembly mounted on the outer cy? inder. The fluid coup1 fag and the 

power coupling are supported from tlte cylinders using truss elearents. The 

cylinders are, in turn, connected tc ihe Space Station Transverse Power 

Boon tmss using a transition structure which for this study will be 

assembled witn 'snap connectionm fittings similar to those used for 

erectable truss methers elsewhere on the Space Station structure. The 

siring of structural elements was driven by stiffness conriderations 

based on the requirement to maintain the stiffness distribution through 

the joint. 
Table 4.4.8.7.5-1 sumrizes the axial, bending, and torsional stiffnesses for trusses transitioning to several djarneters and for a 42-inch 

outside diameter bearing assembly. 

The transition truss concept proposed provides a structural tie between 

the basic 9-foot cubic truss structure and a bearing housicg having a 

nominal diameter of 42 inches. In order to maintain the required 

stiffness it is necessary to fabricate the 12 members of the truss of a 

graphi te-epoxy composite having an el astic modulus in tension of 

40,000,000 psi. The maximum cross-sectional area of each member must be 

12.75 square inches. Therefore a tube with a 6 inch C3 and a wall 

thickness of 0.79 inch wa; selected. This suggests that particular 

attention must be paid to the design of the joint and the local 

supporting structure at both ends of these tubes. The weight of a single 

tt-ansition truss assembly is estimated at 680 pounds. The bearing 

assembly stiffness is governed by bendina. The results are based on the 

use of a preload of 1021 pound-inches on 2 pairs of back to back, Kaydon 

type A, 42 in OD bearings spaced 32.4 inches apart. 

The housing is fabricated of graph i te-epoxy ccmposi te with its wrapping 

biased to yield a tensile modulus of 40,000,000 psi and a shear riiodu'~~ 

of 10,000,000 psi. The worm gear was mounted with its longitudinal axis 

perpendicular to the longitudinal axis of the bearing assembly. In doing 

so, the torsional stiffness of the bearing asseably through the drive 

mechanism is essentiallv a function of the EA of the worm gear. 

Particulw attention should be paid to the design and fabrication of the 

worm gear assembly to minimize backlash and the resulting ncnl inear 

behavior. Weight of the drive housing and bearing asseably is estimated 

to be 383 pounds. 

The internal cylitlder of the pouer couplings is driven relatiale to the 

outer cyl inder bsi ng a worm gear/reducer dri ve/motor assembly mounted on 

the outer cylinder (Fig.4.4.8.7.5-1). The worm gear wi 11 be mounted 

between bearing assemblies which are designed to react both tnrust and 

radial loads. The worm gear is driven by a stepper motor though a reducer 

drive. The total reduction through the system is approximately 2400: 1. 

During rotation, the motor will step about 200 increments per revolution 

with rotaticnal speed of 26.7 rpm. The torque requirement for the motor 

is calculated to be 1.757 foot-pounds. The motor sized for this design 

concept has a capability of 3.33 foot-pounds at the described rates. 

Motor drive and gear mechanism weight is estimated at 40 pounds. 
TABLE 4.4.8.7 -5-1 - S;;{UCTVAAL STIFFMESS WWRY 

TRUSS TRANSITION 

- 2' OF X .1 WALL CIEHBERS 

CASE 42' 01, UP OD 60' 00 REQUIRED 

AX I A1 

LBIIN . 9.102 EOS 9.882 E05 1.162 E06 1.627 136 

BENDING 

IN. -LB/RAD 1.088 E08 1.636 E08 3.364 €08 3.748 E09 

TORSIONAL 

IN .-LB/RAD 4.212 EOS 5.820 E08 r, noo ma - ,.-.A - -..w Lw ~.tl>~ EO7 

AXIAL 

CB/IN . BENDING 

IN .-LBIW 

TORSIONAL 

IN. -LB/RAD 

DRIVE HOUSING 

42 IN. OD BEARING 
4.4.8.7.6 Conclusions 

The rotary drive described in -~nis report meets all of the requirements 

that were defined at the inception of the study. The design of the 

drives, fluid feedthrough, electrical feedthrough, and structgre are all 

well within the current state of the art and should require only the 

normal desi gn/devel opment effort for a subsystem of thi s type. 

Several areas were identified that require a more "indepth" design and 

trade-off study than could be performed in the limited time available for 

this effort. The most obvious area for additional work is weight 

reduction. The primary weight offender is the structure which provides 

the transition from the "square" truss platform to the cylindrical drive 

housing. The root of the problem is the reduct'on in truss stiffness as 

the sectioc is r2duccd to the 48 inch drive diameter. Three possit:, 

areas ,for study seem evi:1eqte 

a. Increase the drive holrsing dii:~??,". 

b. Utilize a different type of aesi9r, for the transition 

structure. 

c. Refine the stiffness requirement used fo; the joint. 

The size of the drive in this design is esssnfially controlled by the 

bearing configuration. For this study, the 42-inch diameter bearing was 

chosen because of it: re! ative "off the shelf" availability and weight. 

The required stiffness and load carrying capacity are readily met with 

the size bearings currently in the design. Larger diameter thin section 

bearings are avail able but the s?ctions are significantly thicker with 

correspondl ngli hi aher weight, rotating inertia, and torque. 

The tabular data presented shows ik resulting effective loss in axial 

and bending stiffness in the transition si;':!rture as the diameter is 

reduced and truss member cross section is kept consi~nt. The final 

weight values in the design are a result of meeting the stiffness 

requirements by increasing the transition truss member diameter o?d wall 

thickness. 

With the truss tyix of transition d~ci gn there is a1 so a problem of local 

deflections at the attachment points between the drive cyl inder and the 

truss. This was recognized as an issue chring the study, but not 

addressed irl detail other than making an allowance for additional 

structure in the weight estimates. 

The requirement that the joint maintain the stirfness distribution of the 

trl~ss was chosen as arl initidl design limit. The rationale was sinply 

not to reduce the basic stiffness of the overall structure. A trade 

study examining the effects of such a local stiffness reduction on the 

affected Space Station systems seems warranted. 
Materials for the a~~onia transport system are an obvious concern. 

However, with reasonable care in selection and implementation it should 

not be a significant problem. Leakage to the outsfde can be controlled 

to acceptable levels with a scavenge system. Aside from 

materials-compatibility, seal life is predominantly a function of 

rotational speed and abrasion of the seal elements. Yi th the 1 or; 

rotational speed and the use of good finishes (6 to 8 micro-in. nns) on 

the seal running surface 1 i fe should be acceptable. 

The torque required to overcome system friction is substantial. However, 

the average power required i s small due to the 1 ow rotational speed. The 

friction torque level itself is not of great concern since it is an 

internal force control system. Of nore concern is the issue of 

rotational smothness or the tendency of the drive to exhibit a 

"stick-sl ip" friction characteristic. 

The choice of an electrical feedthrough approach is not driven by the 

joint design, but 1 ies elsewhere in other Space Station system 

requirements. Either approach, sl i prings or a rotary induction coup1 ing, 

can be accomnodated for power transfer. 

4.4.8.8 Depl oyment Yeri fication 

4.4.8.8.1 Introduction 

Space Station studies of a decade ago envisioned Space Station as various 

coinbinations of modules attached together in clusters. Current 

requirements for ststion design are far- more extensive and include 

satellite servicing. large space construction and payload attachments 

that can only be accmdated by providing large service areas on the 

station. This is best accomplished by lightweight truss structures. At 

the present time, virtually every Space Station concept uses large 

trusses as one of its major components. 

These truss structures can be erected on orbit by EVA cra operating in 

the paylo23 bay vicinity of the Crbiter. The erectable truss is 

assembled in a strut by strut mode, and is an EVA intensive operation. 

The truss structure can be a deployment design in lieu of the erectable 

type. The deployable truss is automatically deployed frm the payload bay, with minor EVA assi st from crew. 

In a deployable truss structure, such as envisioned for the Space 

Station, verification that the structure can be successfully put in place 

is of prime importance. In the past, deployable structures, such as were 

used for the deployment of sate1 lite solar cell s, thermal radiators, etc. 

were designed for high re1 iabil i ty . To obtain the high hardware 

re1 iabi 1 i ty , hi storical ly , it was necessary to overdesi gn the device and 

perform an extensive amount of testing and analysis. This led to long 

development times and high cost. In the construction of the Space 

Station the trusses are deployed from the payload bay of the Shuttle 

Orbiter with the asststance of the crew. The presence of the crew during 

the deployment phase makes this program unlike any previous program in 
th~t now we can obthin high Space Station re1 iability without the need 

for excessively high hardware rcliabil ity. In other words, the crew 

provides backup procedure if a deployment malfunction were to occur. 

4.4.8.8.2 Parameters 

4.4.8.8.2.1 Types of structure 

Current Space Station concepts include three basic types of deployable 

truss structure: 

a. Tetrahedral , planar truss 

b. Box beam 

c. Continuous coilabl e longeron mast (CCLM) 

The tetrahedral truss design includes hinged and fol ding members that 

enable compact packaging in t4e payload bay. The box beam design 

includes hinged, folding, and telescoping members for compact packaging. 

It may also include a deployment mechanism that is a permanent and 

load-bearing part of the beam after deployment. The CCLM design includes 

coiled longerons, hinged strcts, and cable diagonals. It also includes a 

deployment canister that is part of the load path after deployment. See 

Figure 4.4.8.8-1 for views of truss structures. 

The tetrahedral truss is a planar truss. It follows that the deployment 

process involves two dimensional change in size, in width and 

length. The box beam and CCLM deployment undergo only a length change 

only during deployment, while theit- cross-secticn remains constant. 

4.4.8.8.2.2 Deployment means 

The truss deployment process requires energy for mcving the members in 

position and to operate the mechanisms that lock che truss into the 

deployed state. The energy source may be from seored energy in springs 

integral to each join: of the deployable truss. An a1 tetanate energy 

source is an electric mctor driven mechanical e. The tetrahedral 

truss can be deployed ~nly bymeans of stored energy witrlin ti~r juiiits. 

The box beam deploymnt 'an be of the mechanical drive or the stored 

energy type aes i gn. However, the depl oyaent relirlbil i ty cf telescoping 

members with stored energy joints is not we1 1 establi shed. Thz CCLM can 

be deployed by means of a mechanical drive type energy source only. 

The ideal deployment process would provide sufficient energy to each 

joint to sffect depl~ynent and proper lockodt of the joint, with a 

minimun amount of residual energy left over after deployment. This can 

be more easily acccmplished with a stored energy type system where the 

required deployment energy is contained within each joint. This affords 

an even distritlution of depl~yment forces within the structure. The 

mechanical drive type energy source transfers energy to the deploying 
TETRAHEDRAL 

BOX BEAM 

CONTINUOUS COILABLE LON QROW HAST (CCLM) 

FIG. 4.4.8.8.1 VIEWS OF DEPLOYABLE TRUSSES 

609 
beam at a definite number of points. The dynamic load distr;bution 

within the deploying beam is always a function of the stiffress and mass 

distribution of the partially deployed portion of the struct~ire. The 

share of the deployment energy that each deploying joint receives is a 

function of a :oad path that behaves In a nonlinear mode and varies in 

stiffness as r function of tine. This inevitably results i:i uneven 

deployment. 

The deployment process for a deployable structure may be ccntrol led in 

order to slow deployment rate, and thereby minimizing dynamic effects of 

deployment. In a stored energy system, this cart be accomplished with 

simple tether-type mechanisms that are unreeled at a control led rate. In 

a mechanical drive system adequate control of the deployment process is a 

function of avail able information about the 1 oad di stribution within the 

deploying structure that may be used as feedback information to the 

control device. A closed-loop control system would be difficult to 

design due to the inherent limitations of being able to obtain feedback 

data. An open-loop type deployment control of a mechanical drive is 

inherently unreliable, because it can overload and fail truss joints 

without any indication (acquired data or visual observation) that it is 

doing so until the failure occurs. 

A unique problem in truss deployment occurs when testing is attempted 

under one-g loading on the members. The energy thiit is required to 

activate each member varies as a function of orientation of the member 

with respect to local vertical direction. Members that open "down" are 

assisted by grzvi ty . Members that open "up" are retarded by the weight 

of the members. For certain truss orientations a large portion of the 

truss is lifted, i.e., energy is needed above and bgond that of normal 

deployment. Thus, ground testing of these structuraes is a difficult 

task. 

4.4.5.8.2.3 Verification rationale 

The go:! sf tk6 verification activities is to assure with maximum 

reliability that the truss will successfully dep10.y on orbit. The 

process of deployment certi fication is based on a logical combination cf 

tests, analysis, and deployment backup procedures by the crew. 

The requirements for analysis and ground test activities are 

significantly impacted by the planned Ilse of on orbit EVA backup 

procedures by the crew. The availab ty of such backup procedures is a 

function of the design of a particular deployable truss. If the truss 

members and joints are designed so that crew EVA activity with the use of 

on-orbit equipment can assure full deployment, then the requirements for 

certi fication data from ground testing and analysi s are considerably 

reduced. 
4.4.8.8.3 Verf fication procedures 

4.4.8.8.3.1 Ceploymnt analysis 

The analysis which mist be performed to complement the deployment tests 

conoi sis of kinem~tic/dyr,amics, stress, and deploynent failure mode. 

Kinemtic/dynamic deploymnt analysis studies the re1 ati ve motion and 

internal loads of each component as the entire structure deploys. rhis 

analysi s nust consider the various node and joint tolerances ks well as 

the usual bending, torsional, and external stiffness. Inertia load and 

most importantly, the friction loads need to be included to insure that 

the structure have a positive deployment margin during all phases. 

Special emphasis should be placed on the final locking phases of 

deployment. The subsequent prel oadi ng of the joints (where required) 

should be thoroughly studied to guarantee that all joints are 

sufficiently preloaded so as to remain in compression (that is, clearances are all in one direction at a joint or node) under the highest 

expected operational loads. 

The stress analysis as usual is used to determine the adequacy of the 

components to withstand the maximum-induced 1 oads. 

The deployment failure analysis is a very detailed analysis aimed at 

determining all possible modes of fai 1 ure. These fai 1 ure analyses should 

include mechanism jaming due to higher than expected friction loads and 

the inadvertent presence of foreign objects. The analysis should include 

the loads /stress included in the members as a result of the mechanism 

jamming as well as any subsequent failures. A further study should 

include a loads and stress analysis of the application of contingency 

devices such as small jacks, strut replacement aids, etc. 

The analysis activity will vary in difficulty according to the type of 

energy source that is used for deployment. In the case of stored energy 

systems the analytical mdel will be more amenable to simplifying 

assumptions since the distribution of deployment energy within the model 

is known. On the other hand, for the mechanically powered system special 

consideration will be given to the nonlinear behavior of the deploying 

truss in the dynamic modeling. Correlation of analysis results with 

ground test data will be extensive in either case. 

4.4.8.8.3.2.1 Test. $19 in one-g environment 

Deployment viri fication tests for deployable trusses may be performed on 

Earth. However, the deployment tests performed In a one4 environment 

offer formidable challenges. The deployment mechdni sets for any structure 

(tetratruss, box team, etc. must provide deployment energies far in 

excess of those requised for zero4 deployment on orbit. It is further 

complicated by the fact, that the required deployment energy for each 

joint will vary as a function of lnember orientation with respect to local 

vertical. Ground tests will have to be designed such that gravity 
effects on deployment are minimized. Due to the sizes of the deployable 

trusses under consideration the test acti vi ty will require 1 drge 

f a c i l i t i e s . 

There are several options for ground testing of deployable trusses. One 

method involves the suspension of the truss from long cables and allowing 

deployment to take place in the horizonal plane. This method is 1 imi ted 

in size 1.y the height. of existing facilities. The length of the 

buspension cables should be long enough to minimize horizontal loads fram 

the suspension system. A second method involves the use of rollers at 

each node and deploying the truss 0'1 a large flat surface, such as a 

parking lot. The results of this test I:,ay not be viable if the casters 

offer excessive resistance to truss nodal joint translation. A third 

method involves the flotation of the truss in a sufficiently large body 

of water. This may be accomplished by varying the effective density of 

each truss member by fi1;ing it with foam plastic to the point where the 

truss would be neutrally buoyant. The truss then would be submerged to a 

sufficient depth and the truss deployed. This approach would require the 

fabrication of a test article that would not corrode or otherwise 

malfunction in a water environment. A fourth method of testing could be 

var'ous drop tests. However, this would most likely be limited to 

smaller test art'cles and may very well result in the destruction of the 

test article. 

4.4.8.8.3.2.2 On-orbit tests 

Forb an) type of deployable structure, consideration should be given to on 

orbit testing for verification of truss deployment. Testing of this type 

would offer the exact environment in which the Space Station will be 

deployed. The trusses are designed for deployment from the payload bay; 

thus, on-orbit test hardware and deployment dids would be identical to 

the actual Space Station hardware. A demonstrated successful depl oyment 

would provide complete certification of hardware deployment aids dnd crew 

procedures. In order to minimize costs, the exp~riment could be 

manifested on a space avail able basi s. 

4.4.8.8.3.2.3 Box beam deployment 

The reference Space Station concept has a 3 feet square deployable box 

beam for the keel structure of the station. Figure 4.4.8.8-2 shows the 

station concept and associated box beam structure. 

The box beam depl oyment schematic is shown in Figure 4.4.8.8-3. The beam 

is deployed d bay at a tiine ir, d sequential fashion. 

The box beam could b? tested in arbi I. for deployment re1 iabil i ty in the 

same fash-ion as the tetrahedral truss. Such test would be conducted 

under the actual operating environment and would provide an adequate 

verification of deployment. The box beam is currently envisioned without 

a retraction feature. For an on-orbit test, either a retraction means or 

some form of control led deorbit would be required. 
Ground testing of the box beam would have lilaitations quite similar to 

the tetrahedral truss. Gravity effects would increase the required 

Geploymnt energies for the strut joints. To minimize the gravity 

effects tests would have to be conducted where the beam i s suspended from 

cables or on rollers on a flat supface. Figure 4.4.8.8-2 shows a 

schematic of such test arrangewnts. Due to ,;he long length of the keel 

bean it would be difcicult to find a till; enough structure that could 

sene as the tower to suspend the bzam for the cable-supported deployment 

test. The test option with the rollers would be difficult to accomplish 

due to the amount of rolling resistance as a larger and larger portion of 

the beam was deployed. 

The box beam could well be tested underwater with neutral buoyancy 

flotation mec..ns. The design of the deployment motors and associated 

electrical devices for the test would have to accomnodate submersion in 

water . 

4.4.8.8.3.3 Deploymnt backup 

It is the incorporation of crew backup into the deployment re1 iabil ity 

which can render extensive cost savings. Crew hackup to deployment 

simply mans that if during the automatic deployrrlent of the Space Station 

structural element a malfunction were to occur, a crew member with small 

tools could fix the malfurictioning cornpollent and thus result in totally 

successful deployed structuibes. The cost savings mentioned earlier will 

be realized in a greatly decreased +testing and design program. In 

essence, this phi1 osophy can accepx lzss hardware reliability because of 

the additional crew supplied re1 tabil ity . 

Since the Space Shuttle Orbitcr has repeatedly proven itself as a 

"spacetruck", the aerospace iomnunity must begin t.o treat the pqyloads 

(espei; ially deployable st~~cture) as cargo trot as another space vehicle. 

The crew, who provides the additional backup in case of a malfuncti~n, 

must have the tools onboard to fix any deployment problem. As the most 

probable failure modes are identified and tools would then be designed so 

that thes; failures can be fixed. 

Figure 4.4.8.8-5 shows a worse case jammed tetratruss .;'oint due to a 

foreign matte obstruction. Also shown is a crew member performrng 

Man7ed Maneuvering Unit (WU) fly ir, .,un 4 inspection. Figure 4.4.8.8 -6 

reveals the details of the deploymitt jack components. The crew l~en ,r 

is shown (in Fig. 4.4.8.8-7! jacking dowr! the jamned strut after the tool 

has been attached tc the truss cell. In figure 4.4.8.8-8 the crew her 

is shcz; removing the obstruction while the jack tool retains the strut. 

in Figure 4.4.8.8-9 the crew amber is shown jacking the strut to its 

full deployed position. Tools of this kind with simple modification can 

be utilized to remove and replace a broken strut, hinge on even a node. 

Similar tools can be designed to accomnodate any fajlures of a deplayable 

truss as long as the deployable structure had been originally designed to 

be repaired, such as by providing for removable sthquts, removable joints, 

etc. 


Flgure 4.4.8.8-4 - BOX BEAM GROUND TESTS 

QUIC

QI 

G 

CLAMP & UNIVERSAL JOINT 

Flgure 4.4.8.8-6 - DEPLOYMENT JACK COMPONENTS 

CREW 

OBST 

q?J .t 

-,A'. I.' 

* -.- 
Ch;Gurhl PA^& *IS* 

OF POOR QUALITY 
4.4.8.9 Conceptual Design of a Mobile Remote Manipulator System 

Space Station studies have identified the need for a Mobile Remote 

Manipulator System (HRMS). Such a 'logistics or utility device is 

envisioned to be wtfi tted with a spacecrane capability (i .e., Shuttle 

RMS) and astronaut-foot restraint arms. The system is required during 

ici tial Station construction activities, to r~sition astronauts for EVA 

functions, to transport modules and/or paylrads from the Shuttle cargo 

bay and posi tiori them for attachment to the truss structure. Subsequent 

to the initial Space Station construction activities, a MfMS is 

considered necessary for rwintenance or repair activities and to provide 

a construction capability for future station growth or large spacecraft 

assembly and servicing. 

4.4.8.9.1 MRMS mobility requirenrents 

The squarc bay truss structure of the Space Station configurations shown 

in reference 1 suggests the need for an MRMS vhich cari move in two 

orthogonal directions. This capabil i ty pemi ts movenrent (1) a1 ong the 

Space Station keel structure between the modules and the solar array 

support structure, and (2) perpendicularly a1 ong the sol ar array support 

booms. An MRMS with only undirectional dili ty theoretically could be 

rail-mounted to accomplish this function, but would probably need to be 

detached and reattached to additional orthogonal rails to move in a 

perpendicular direction. This is an operational drawback which probably 

should be avoided. Mounting rails onto the Space Station truss structure 

introduces addi tional mass and significant design complexi ty which a1 so 

must be corsi dered. 

The lwer mass alternative of positioning rails on the MRMS (instead of 

the truss) which "ride" on the truss hard points is possible. However, 

endless tracks (chains or belts) which provide mobil i t y i n this case must 

completely span two truss bays to ensure stabiiity of the MRMS during 

motion. Such an arrangement avoids the increased mass and complexity of 

rails attached to the truss structure and provides for a "smooth', 

continuous unidirectional motion capability. However, mvement in a 

perpendicul ar direction i s not simply accornpl i shed and the undesirabl e 

feature of an MRMS which must be bqys in length is introduced. A 

2-bay-1 ong MRMS presents Shuttle packaging problems. It also and 

degrades mneuverabil i ty and, therefore, usefulness for maintenance and 

construction acti vi ties, parti cul arly in CI ose proximi ty to the modul ar 

habitats or surface-attached equipment. An MRMS for movement on a 

triangular gridwork truss is presented in the structures and mechanism 

White Paper entitled "Space Station Truss Structures and Construction 

Proceduresu. A robotic walker "spider" conceptually could serve as an 

HRMS and accompl ish the necessary functions, but would quire extensive 

development of a device which is not considered state-of-the-art. 
4.4.8.9.2 Reference MRHS design 

A conceptual design for a bidirectional HRHS which is only 1-bqyl square 

and avoids a truss-rounted rail system is illustrated in Figure 

4.4.8.9-1. This design, which is a modification of a 2-bqy long device, 

consists of three basic elements or Iqyers. The bottom or track layer 

consists of a square track arrangelrent which rides on structural guide 

pins attached to the truss nodes. The four tracks are arranged in a 

single plane and connected at the corners of *switchesu which can be 

aligned to permit lotion over the guide pins in either of Wo orthogonal 

directions (see Fig. 4.4.8.9-2). The track layer does not rotate 

relative to the truss structure. The four corner switches rotate 90°, 

but only when centered over the guide pins. 

The central elelsent consists of a pushlpull drive mechanism which also 

has 360" rotation capabi lily (see Fig. 4.4.8.9-3). This feature permits 

platform movement in four directions by either a push or pull operation 

and greatly enhances mneuverabil i ty wi thout requiring addi tional 

structure for trans1 ation. It a1 s2 pemi ts changing mve~lent direction 

without rotating the logistics platform or attached pqyloads. The 

push/pull motion is envisioned to be powered by an electric stepper motor 

through a rack and pinion drive. Mounted on the driubar ends are "drive' 

rods which are aligned with, and electrically inserted into, the nodal 

guide pins and locked. The switches are a:inned appropriately and the 

drawbar is actuated to push or pull the MRMS in the desired direction. 

The drawbar extends to span a conplete bay, such that four-point support 

of the MiWS is maintained at a1 1 times. Translation of the MMS is 

accoapl ished by operation of the pushlpull drive mechanism to mow the 

platform longitudinally in an "inch-worm" fashion. This sequence of 

events is illustrated in the upper half of Figure 4.4.8.9-4. Transerse 

translation involves use of the pivoting as well as pushlpull feature of 

the mechanism, and is illustrated in the lower half of Figure 4.4.8.9-4. 

Sketch (A) shows the MRMS pivoting 90" from the direction of travel. 

Sketch (B) shows a translation to construct an adjacent truss cell which 

is in the next row. Sketch (C) shows the MfW sliding onto the cell just 

constructed. Sketch (Dl shws a 90" rotation into a position parallel to 

the original, but on an adjacent row. Sketches (El and (F) show 

longitudinal motion and construction of added cells to complete a 

platform. The corner switch illustrated in Figure 4.4.8.9-5 corner 

switch il lustrated in Figure 4.4.8.9-5 shows tke open top mechanism 

feature which permits the drawbar to lock onto a guide pin which is also 

occupied by a track switch. 

The top element of the MRMS is the logistics platform which is envisioned 

to rotate with respect to either the track or drive elemnts (see Fig. 

4.4.8.9-5). This platform wmld serve to transport pqyloads and cargo 

over the Space Station surface. A central feature of this element would 

be the capability to operate a transposed Shuttle W.S, which is shown in 

Figure 4.4.8.9-6 mwnted on a moving carriage. Also sham in Figure 

4.4.8.9-6 are Mobile Foot Restraint (MFR) positioning arms. 

Pressure-suited astronauts attached to the MFR's are positioned within 

their work enuelope by the movable positioning arms. The MFR arms should 

not be 
considered to be miniature versions of the Shuttle RMS , wi ttl ; is 

precision control requirements. Rather, each MFR arm should be 

control lable by the astronaut who can adjust its position in a manner 

similar to a util i ty servicernan cperating a "cherry pickerM bucket. The 

degrees of freedom required by the MFR ams are determined by the extend 

to which EVA is utilized to perform various future Space Station 

functions. 

The MMS should have a self-contained, rechargeable power suppiy which 

does not require umbilicals or power rails. Control of all features of 

the MRMS should reside with the EVA astronaut(s) to avoid hard1 ine or RF 

control links to a central station. Transport cradles or similar devices 

must be provided to support payloads being moved abwt the Space Station 

surface by the MRMS. 

4.4.8.9.3 MWrtS plane change concept 

The proposed MMS can operate over both the "top" or "bottom" surfaces of 

the reference configuration, if required. It uses the rotary joicts and 

solar array boom rotation as a turntable to translate betaeen the two 

paral lel surfaces. For generality and versatility , however, it is 

desirable to operate the MRMS in a plane which is perpendicular to these 

paral 1 el surfaces. Two concepts for rotating the operational plane of 

the MRMS 90' from its original position are showfi in Figures 4.4.8.9-7 

and 4.4,8.9-8. The concept shown in Figure 4.4.8.9-7 uses a tilting 

frame approach to rdtate the MRS 90" and enable transl ation a~d 

operation onto a perpendicular plane. The tilting mechanism is 

envisioned to be self-contained and installed as a truss cell unit intc a 

beam or a1 ong a pl atfo~m edge. Thi s uni t shwl d be capable of rotating 

90" in both the "left" and "right" hand directions. 

A second approach is illustrated in Figure 4.4.8.9-8 which is 

operational ly more complex, but mechanical ly simpler and probably more 

compact than that of Figure 4.4.8.9-7. This concept consists of a planar 

guide pin frame which is attached to the orig~rlal structure at the center 

of an element which replaces a truss strut. A center sectior; of this 

element contains a "T" fitting which has two perpendicu'rar rotational 

degrees of freedom. Operational ly, the MRMS transl ates 1 ateral ly onto 

the attached guide pin frame. The MRHS and frame are the? rotated 180' 

around the frame centerline attachment as showr in Figure 4-t.8.9-8 into 

an "upside down position". The MRMS and frame are then rotated 90' 

around the strut element center1 ine to a position which permits the MRMS 

to translate onto the plme which is perpendicular to its original 

operational plane. 

Two devices such as those just discussed or ancther appropriate design, 

placed on opposite faces of the truss structure would permit rapid and 

convenient transl ation of the MRMS between to^" and "bottom" surfaces of 

the station without interrupting the rotatiorr ,f the solar wing for this 

purpose. 
4.4.8.9.4 Concluding relaarks 

A conceptual design for a Mobile Rmte Manipulator System has been 

presented. This concept does not require continuas rails for mobility 

(only guide pins at truss hardpoints) and is very conpact, being only 

1-bay square. The MIUS proposed is highly maneuverable being able to 

move in any direction along the orthogonal guide pin arrqy uooer complete 

control. The proposed concept would greatly enhance the safety arrd 

operational capabil i ties of astronauts perfomi ng EVA functions, such as 

structural assembly, pqyload transport and attachment, Space Station 

maintenance, repair or mdi fication, and future spacecraft const~~ction 

as servicing. 






1 

Figure 4.4.8.9-7 - MRMS OPERATIONAL PLANE CHANGE-CONCEPT I 
Figure 4.4.8.M - MRMS OPERATIONAL PLANE CHANGE-CONCEPT II 
Close inspection of the reference Space Station configuration indicated 

several areas requiring detailed technical supgort of thp structures and 

mechanics team. 

The basic structural foundation for the Space Station is the truss 

system. The truss configurations investigated for this station are 

clearly not a closed set but are only representative of trusses that can 

be successfully used. The best, truss for the prime structure will be 

determined after trade studies considering deployment reli abil i ty, EVA 

time, cost, structural redundancy after damage, and ease and desirabil i ty 

of integrated utilities. 

The heart of the solar panel deploynrent system is an autoiitical ly 

deployed mast structure. This mast is required not only to support the 

solar array but a1 so to extract the solar panels from their containers 

during deployment. The masts presently available for solar arrqy 

dcployrnent are not redundant, either structurally or in the deploywent 

process. Furthermore, darwge to a single elelsent, caused by debris 

inpact or manipulator contact, can cause buckling of the total mast array 

and possible additional damage to neighboring structure. In the 

reference concept, eight independently deployed masts provide sow degree 

of redundancy to the' power capabil i ty . Numrws module and payload attachment designs have been studied. The 

design emphasizes versatility, rigidity, aod ease of crew installation. 

Design concepts were discussed for attaching payloads and modules to the 

9-foot square truss beams, the 15-foot square truss beam and the planar 

tetratruss. It appears that the attachwnt of various iteins to the truss 

configurations presents little challenge. The most difficult was 

attaching the 14-foot modules to the 9-foot deployable keel. 

It was concluded that the module to module interface should conform to 

the foll owing generz; requi renents: 

a. Capability of mating with any other module interface or to 

any docking/berthing port. 

b. Capabil ity of aligrment adjusttent that is consistent with 

the devel opment of speci f ied mdul e patterns. 

c. Provision for retraction capabil ity of the interface to 

all ow clearance for module instal 1 ation and removal. 

d. Provision for a pressurized passage that is consistent with 

a !%-inch diameter hatch opening. 

e. Provi sion for util i ties interconnect (where necessary). 
f. Capability of an interface latch to acc-date pressure 

loads and dynamic loads in the mated configuration. 

After the second orbiter buildup flight, the RMS will not have sufficient 

strength to move the station to the orbiter. This seems to necessitate a 

docking system be attached to the Station during the Station buildup. 

Docking, rather than berthing, requires the design of a docking. system 

that induces higher loads into the Station, thereby requiring stronger 

and stiffer interfaces. 

The preliminary structural design and analysis study of a Shuttlelaunched common module has been completed. It is clear that the primary 

structural shell presents no new technology requirement either in the 

fabrication or in the analysis phases. The largest loads are experienced 

during launch and landing (assuming the mdule to be returnable). The 

weight of the primary structural shell was determined to be approximately 

ten thousand pounds. Areas requiring additional development are the 

effects of 1 ong-term meteroid/debri s protection and the precompressed 

window design. 

Since the reference configuration uses numerous rotating joints, a 

detailed study of these joints was performed. The results indicate that 

the rotating joints can be desjgned to have continuity of stiffnes; 

across the joint and also Qunction mechanically and electrically. The 

remaining concern of the rotating joint is its tolerance to meteoroid and 

space debris impact. A failure mode effect and replacement study for 

each part of the joint must be performed in the near future. Further 

joint sti ffness studies considering the wear of the bearings (thus 

relieving preload) must a1 so be made. 

In the past, deployable structures, such as those used for the deployment 

of sate1 lite solar cells, thermal radiators etc. were designed for high 

reliabil i ty. To obtain this re1 iabil ity, it was necessary to over-design 

the device and perform an extensive amount of testing and analysis. In 

the construction of the Space Station, the crew provides a backup 

procedure to be incorporated if a deployment malfunction was to occur. 

Overall Space Station construction re1 iabili ty can be just as high as in 

the past, but it comes from a combination of hardware reliability and 

crew backup procedure re1 iabil i ty. 

The MRMS design employs the previously developed Orbiter RMS mounted to a 

mobile base that moves from truss node to truss node by means of short 

tracks hand mounted to the base. It util izes rechargeable batteries $at 

can support full operations for approximately 6 hours. The crew ride on 

and control s the maniplrl ator during a1 1 operations. This coV,iept i s 

workable and requires only concept verification in the area of 

node-to-node traveling. It should greatly enhance the operational 

capabilities of the Station crew in performing such tasks as pqyload 

transport, Space Station repair/modification and structural assembly. 
Two concerns shculd be mentioned. The first is that the primary truss 

structure, the solar array booms, and the rotary joint strut structure 

have little damage tolerance to impact by such things as space debris, or 

the Station manipulator. The second concern is the durability of the 

rotary joints, as well as procedures in case of joint failure. 
EVA System Configuration - 

4.4.9.1 Introduction 

The Space Station provides an EVA capability for assembly, maintenance, 

servicing, and repair of the Space Station and other large space 

structures; servicingand repair of sate1 1 i tes; maintenance, servicing 

and repair of the 0)9V and OTY; and maintenance, servicing, and rcpair of 

payloads and experiments. EVA systems can be broken into ?auo areas: the 

GFE which includes the EMU'S, HU's and miscellaneous universal EVA 

equipmnt and the Space Statf on dedicated hardware which includes the EMU 

and WU automatic service stations, the airlock, and Space Station 

dedicated EVA support hardware. Thi s section presents a reference 

configuration for the Space Station dedicated hardware. It a1 so presents 

weight and volume estjmates for preliminary concepts of the GFE. 

RFP-9BC72-72-4-36P describes a 1-year study effort that will define EVA 

hardware requirements for the Space Station era. 

4.4.9.2 Space Station Dedicated LYA Systems COII~ iguration 

The Space Station provides the hardware necessary to support routine EVA 

aboard the Space Station. This hardware includes the EMU 

servicing/checkout subsystem and the MU servicing/checkout subsystem, 

the airlocks, various EVA support equipment outjide the Space Station, 

and the EVA communicati~ns and data management system. 

The Space Station provides for simul taneous EVA'S (8 hours per dily , 5 

days per week) of two crew mnbers during IOC and for four crew members 

during the growth phase. The following is a list of the EVA requirements 

which impact the Space Station systems configuration. These requirements 

are organized into the subsystem areas. 

4.4.9.2.1 EMU servSce/checkout subsystem 

The EMU reservice/checkout subsystem wi 11 be partially located in the 

Habitability or the Cormand Module and partially in the EVA airlock. 

Table 4.4.9-1 lists the functions of the EMU reservice/ check& 

subsystems. Figure 4.4.9-1 presents the reference EMU service sutsysten 

configuration. The power, mass, and volume penalties for the EMU 

service/checkout station are presented in Tab1 es 4.4.9-2a,c ,d,g. 

The above reference confi guration is prel imi nary. The EVA hardware 

requirements are a subject of an independent RFP effort. The results of 

this independent study wi 11 affect the EMU reserv~cz/checko~t subsystem 

configrlr'ation. The EVA study results will be provided to the Phase B 

Contractors both prior to and at IRR. 

The EMU service station will interface with the ECLSS. The EMU'S will 

contain regenerative, nonventing life support modules. Tne service 

station will regenerate tr~e EMU life support mdules. The EMU will 

norinally be reserviced as an assembly. The EMU modules which require 
more than one hour to be reserviced will be serviceiible at the modular 

level to accommodate contingency reservicing requirements by replacemerpt. 

The EMU's will support 8 hotirs of EVA/dcqy (cumulative time dt reduced 

pressure) at an average metabolic rate of 1000 BTU/Hr. The 8 hours EVA 

will generally be divided hy a break in the Space Station. EMU 

subsystems may be revitalized during this break The EMU's are 

regenerated in 12 hours without human intervention to provide next day 

EVA capabil ity. Rapid emergency EVA egress (within 10 minutes) is 

possible with minimal EMU functional checkout. The EHU service station 

provides for automatic servicing, checkout, and IV operations of the EMU 

while in the airlock. 

Automatic servicing and performance checkout of the EMU's includes 

expendables replenishment such as 82 and Hz0 resupply and the 

regeneration of time dependent processes such as C02 and Hz0 removal, 

heat rejection and power storage. The service station provides for the 

automatic checkout of the EMU components and system performance. The 

service data may be retained by thc, Space Station data system. The EMU 

software automatically upddtes to adjust for small EMU instrumentatio;l 

shifts to reduce the need for frequent instrumentation adjustments. 

Performance trend data shall be used to define the need for maintenance 

of the EMU. The service station will automatically dry the suit. EMU 

servicing capabilities are based on 10 reservices per we& for IOC and 

for 20 reservices per week for the growth station. 

4.4.9.2.2 MU reservi ce/checkout subsystem -

The MMU will be used for free-flying proximity operations around the 

Space Station. 

The MMU uses an expendable cold gas propellant which must be resupplied 

after each use. The MJU also requires a recharge of its batteries. The 

MU will be automatically checked out after each use. The reference MMU 

service station is presented in figure 4.4.9-2. Power, mass, and volume 

penal ties are presented in Tab1 es 4.4.9-2a,c,d,g. 

Automatic servicing and performance checkout of the MMU's includes 

expendables replenishment such as N2 resupply and the regeneration of 

time-dependent processes such as thermal and power storage. The MU 

service station provides for the automatic checkout of the MU components 

and system. The service data is retained by the data system. 

Performance trend data is used to define the need for maintenance of the 

WU. The entire normal servicing will be accomplished in 12 hours 

without human intervention. Propellant reservicing will be accanpl ished 

in 10 minutes. Servicing capabilities are based on 10 reservices per 

week for IOC and for 20 reservices per week for the growth station. 
4.4.9.2.3 EVA airlock 

The EVA airlock provides the means for transferring an EVA crewmember 

from the Space Station to space, the stowage for the EMU's inside the 

airlock and for the MU'S outside the airlock, and provides a hyperbaric 

chamber for treatment of rap id decmpression i 11 ness. 

The EVA adrlock provides a controlled rate df depressurization and 

p~essurization. The nominal rate does not exceed .B psi/sec. The 

maximum rates are not to exceed 1 psi/sec. Control of depressurization 

and pressurization is possible from inside the Space Station, inside the 

airlock and from outside the airlock by a single individual. An airlock 

gas recovery subsystem shall punq airlock gas back Snto the Space Station 

during the initial depressurization to conserve consumables. The airl ock 

design accommodates the transfer of a standard laboratory experiment rack 

or the return of an incapacitated EVA crewmember.. Life support umbilical 

connectors is available outside the airlock to allow umbilical EVA 

opera ti ons. 

Stowage of the EMU's inside the airlock versus in the Space Station is 

required to allow for automatic checkout of the EMU's during 

depressurization and for reconnection of life support for contingencies 

while at vacuum. Stowage of the MU'S outside the airlock is required to 

centralize the EVA servicing equipment and to local ire the EVA hardware. 

This localization also allows for easier relocation of the EVA equipment 

for flexibility for growth phases. The power, volume, and mass penalties 

for the airlock are shown in Tables 4.4.9-2b,d,f,g.. The airlock systm 

consists of the basic airlock structure, the airlock gas recovery 

compressor, the hyperbaric equipment and the lights. The basic airlock 

structure is 1 arger than the Shuttle airl ock to provide addi tional space 

for the servicing hardware and additional space for the crew to maneuver 

for rapid egress capability. The structure is heavier because of the 5 

atmosphere pressure requiremnts of the hyperbaric chamber and the 

weights of comnon docking adapters. 

The airl ock, the EMU, and the EMU service system have the capabil i ty of 

serving as a hyperbaric chamber for two crewmembers (one assisting). The 

airl ock pressure is raised to as high as 5.;) atmospheres above the 

aqbient cab1 n pressure (88 psi ) . The crewmembers breath a1 ternately 

between 02 off a mask and airlock gas with the 02 pressure maintained 

above 3.0 psi. The 02 concentration does not exceed 30 percent at any 

time. A small airlock between the EVA airlock and the Space Station is 

provided for passf ng medication, equipment, food, etc. The unmanned 

EMU's provide the life support fcr this contingency. Provisions will be 

provided for decontaminating the EMU after a chemical spill event. 

Verification of acceptable contamination levels shall be made prior to 

opening the airlock/(Space Station) hatch. 

4.4.9.2.4 Eaui~ment airl ock 

The equipment airlock serves as an equipment and tools transfer airlock. 

It is provided to minimize the consumables and time required to pass 
items between the Space Station arid the EVA em. The equipment airlack 

can be located at any convenient location on the Space Station. The 

dimensions, mass, and power penal ties are presented for the reference 

configuratior~ in Table 4.4.9-2b.d.f.g. These dimensions are arbitrary 

and are to be the subject of an independent EVA systa, requirements study 

to be performed to support phase 8. 

4.4.9.2.5 - Mi scell aneous EVA support equipment 

Miscel lanmus EVA support equipment is to be provided by tht. Space 

Station. Tools, tethers, foot restraints, hand holds, slide wires, 

equipment storage boxes, etc. are required for routine EVA. Weight and 

volume penalties for the reference configuration are presented in Table 

4.4.9-2. The final values will be dependent on the Space Station 

configuration and the nature and amoubt of EVA required. 

The EVA crewmember productivity is enhanced through the use of powered 

tools. Battery powered tools kill be utilized for activities at remote 

sites. Plug in power tools will be provided for use at locations where 

an extensive amount of EVA work i s expected to be performed ( i .e., at the 

end of a WS). Electrical crutlets and flood li yhts will be provided at 

these 1 ocati ons. 

Sufficient food and drink for 8 hours (38 oz. ti20 and 753 calories of 

food) will be available for use by an EVA crewmember while EVA. For 

EVA'S with a scheduled break, only H20 need be provided. EVA water end 

food containers will be cleaned and refill ed with gal 1 ey subsystms. 

The LCVG ( 1 iquid cool ing and ventil ati on garment) wi 11 be rwti nely 

washed or replaced. Provisions for EVA equipment and spares stowage are 

provided inside the Space Station and on the outside the EVA airlock. 

Maintenance of all EVA equipment is performed inside the Space Station. 

4.4.9.2.6 Comnunications, data, and TV 

The EVA crew will be provided with data by use of full page freeze frame 

N transmission. EVA procedures, sate1 1 i te servicing procedures, payload 

servicing procedures, Space Station assembly procedures, etc. wi 11 be 

provided in this manner. An audible instructions channei is provided to 

give audible instructions corresponding to a particular data page. The 

page displayed is controlled by voice comnand of the EVA crewmember 

eliminating the need for an IV crewmember to monitor procedures. One 

cormwnications channel per two EVA crewmembers is provided. One data 

channel will be titleshared between all EVA crewmember and the Space 

Station. High speed TV reception from a portable EVA TV camera is 

available. 

The description of the comnunications system used to support EVA is 

presented in section 4.4.3 comnuni cation and tracking subsystem 

description. 
4.4.9.3 Government-Furni shed EVA Hardware 

The NASA will develop comnon use EVA hardbare for use on Space Station. 

This hardware includes the EMU's, the MMU's, and mi scellanews EVA 

equipment such as TV cameras, tool caddies, etc. For reference, this 

hardware i s described be1 ow: 

4.4.9.3.1 EMU configuration 

The requirements for the EMU are the subject of a 1-year study effort 

issued under RFP-9BC72-72-4-36P. Current concepts indicate that the EMU 

will be regenerable, automatically reservicable, and maintainable in the 

Space Station. The EMU performance will be automatically verifiable. 

The EMU wili use a modular configuration to facil itate replacement and 

repair. There may be more than one technical approach for each modular 

function (i.e., C02 removal 1 depending on the evolution of technology, 

EVA requirements, appl ication requirements (i .e., contingency EMU's may 

allow subsystem venting). 

The mass, vol urn, and power regui red by the EMU's are presented in Table 

4.4.9-2. The consu~bles and the returnables are a1 so provided. Water 

and food are provided for an EVA crewmmber while on EVA. An approach 

for the collection of urine will be provided. 

4.4.9.3.2 EMU wares and maintenance 

The EMU will be fully maintained in the Space Station. There wi?l be at 

least one spare of each functional module. There will be at least two 

spares of all crewmember specific hardware such as arms, legs, gloves, 

and the LCVG. 

There will be provided spare piece parts to repair the EMU modules. 

Repairs will be simple and require minimal tools (i.e., replacement of a 

computer card). All repairs will be performed inside the Space Station. 

4.4.9.3.3 MU Configuration -

The MU configuration used in this reference configuration uses the same 

design concept as that used on the current Shuttle program with the 

exception that the N2 propellant capacity has been increased by a 

factor of 2.5 and that the MMU wi 11 be augnented with range-rate 

equipment the same or similar to that used at the docking ports. The !MU 

dasi gn will provi de the capabil i ty for near-in stationkeeping, 

short-di stance maneuvering, 1 ong-di stance trjverse (to 300 neters or 

more), and small object maneuvering (i .e., 600 1 bs. for 100 meters). 

Mass and volume penalties are shown in Table 4.4.9-2. A nominal ,,f two 

MU usages per week have been planned for this reference configuration. 

4.4.9.3.2.4 MMU Spares a~d Maintenance 

The MMU will be fully maintained outside the Space Station. There will 

be provided spare piece parts to repair the ::MU. 
TABLE 4 .C -9-1 - EVA RESERVICEICHECKOUT SUBSYSTEMS 

Reservicl nQ 

Recharge power supplies and power supplies checko~ 

Refill oxygen 

Regenerate heat rejection module 

Makeup cool 1 ng water 1 eakage 

Regenerate C02 removal module 

Regenerate hun;idi ty removal mdul e 

EMU drying 

Service trend data dump 

MMU Me refill 

Automati c P re-EVA Checkout 

EMU structural and leakage check 

Fan performance check 

Puap performance check 

Cmputerlcontrol ler software check 

Cautionlwarni ng software check 

Suit pressure regulators check 

Cool i ng tmperature measurement cal i bration 

Suit pressure transducer check 

C02 sensor checkout 

Cool i ng 1 oop gas separator perfomance and water f il tration 

Primary and secondary 02 supply transducer calibration 

Vent fl ow sensor performance check and software correction 

Hetab01 ic rate cal cul ation check 

Posi ti velnegati ve re1 i ef val ve checkout 

Post-EVA Performance Checkout 

Ambient pressure transducer 

Cool i ng 1 oop 1 eakage 

Purge val ve fl ow rate performance 

LCVG structural and 1 eakage 

MU systems performance check 

Automatic EVA prep operations 

Nitrogen purge and verification 

Airlock depressurization/repressurization 

IV pressure regulation 

EMU integrity checkout 

M SF'PVICE STATIOW 

Figure 4.4.9-2 

SPACE i AIRLOCK 

1 

I 

I 

I I 

1 

I 

AUTOMATIC .I a 

MMU I 

I SERVICE I 

3TAI ION I 

r------ .I! 

SPACE LTATION i 

I Loalsrlc

I ELECTRICAL ] I 

POWER 

ELECTRONICS 

CONTROL 

! 
TABLE 4.4.9-2.a - EVA SUBSYSTEM KIGHT/POUER/DIWENSIONS 

- EXERIIAL AIRLOCK 

POER PorR ACT BLOCKS 

UATTS MATTS POUEREO DIENSIONS VOLUE 

COMPONENT DAY DAY 84xx FTxFTxFT YTfLBS CU FT 

1 MU SERVICE STATION (2 EMU'S SERVICED) 

ARLK VENT LOW 100 100 17,-,25 

ARLK COOL LOOP - 30 30 17,23,24,25 

HIGH PRESS 02 LOOP 100 0 17,19,22,23 

24,25,27 

HELMET DIS L COlT 25 25 25 

WDD 

C02 REMOVAL REGEN 75 75 24,25 

HUMIDITY REMOVAL & 200 200 24,25 

SUIT DRYER 

POS PRESS RELIEF 80 90 25 

YLY CHECKOUT 

TOTAL 

2 MMU SERVICE STATION (2 WU'S SERVICED) 

~IEFI PRL~S ~2 em -x2xl 40 2 

SUPPLY 24,25,27 

PUR CND/ELECTRONICS 210 210 20,21 lxlxl 50 1 

SPAES (10%) - - - lxlx.2 17 0.2 

N2 STORAGE TANK - - - 3.5 938 7 1 

ICRYC! 

Np CRYOGZ!!!CS - - - - 1840 - 

3 EMU 

-0- 170 25 .93x.75x6 4m.9 17.05 

EMU SERVICE STAND 1 20 20 25 4xX. 5 30 18 

EMU UNIT 2 170 170 25 .95x.75x6 440-9 17.05 

EMU SERVICE STAND 2 10 10 25 4x3x1.5 30 18 

LCVG 1 - - 1.4x1.4x.4 8 1.2 

LCVG 2 - - - 1.4x1.4x.4 8 1.2 

SPARES - - - - 549.7 19.7 

4 MMU 

mi 1 - - .. 3x4~2 451.1 31 

MU SERVICE STAND 1 10 10 21 3.2x4.2xi.0 5G - 

~3 ug!~ 2- - 3x4~2 451.1 31 

MMU SERVICE STAND 2 10 10 21 - - 3-2~4.2~1 .u 3u - 

SPARES - - - 2x2~2 100 6 

5 EVA TOOLS 

PWR TOOLS (BATTEFI 50 50 24 lxlxl 80 1 

PUR TOOLS (IITILITY) 50 50 19 lxlxl 80 1 

WON-POWERED TOOLS - - - 1~1x1 40 1 

Er1JIP REPAIR TOOLS - - - 1xlx.5 70 0.5 
TABLE 4.4.9-2. b - EVA SUBSYSTEM UEIGHT/POKR/DIENSIONS - 

POWER POWER ACT BLOCKS 

WATTS YATTS POWERED OIlENSIONS VOLUE 

COMPONENT DAY DAY 84xx FTxFTxFT UTILBS CU FT 

6 MISC EVA EQUIP 

- - - - 130 3.29 

EQUIP 'STOWAGE LOCKER - - - 3x3~1 10 9 

COM MOD EQU!P FRAE - ' - - - 0 

EVA SYSTW TOTALS PEAK PUR (MATTS) = 1180 NIGHT/1550 DAY 6914.4 354.78 

AI RLoCK M IGHTIVOL~IPOUER PENALTY 

K 

- - - 6.7x8.3 1316 200 

EQUIP ATTACHENT 25 25 17,18,19,22 

LIGHTS 23 

GAS RECOVERY 1400 1400 18 

EQUIPCENT AIIZLOCK - - - 2x3 90 9.5 

LIGHTS 3 3 5x DURING EVA 

GAS RECOVERY 100 100 5x DURING EVA 

HYPERBARIC EQUIP - lxlxl 50 1 

- AIRLOCK TOTALS PEAK PUR = 1425 WS (6 MINUTES)* 1471 219.5 

fTIE SEQUENCED YITH EVA 

INTEGRATED PEAK PUR = 1425 NIGHT/1550 DAY 
TABLE 4.4.9-2 .C - SECONDARY EVA SYSTEMS JYER/YOLUME/HASS 

POUER POUER ACT BLOCKS 

WATTS WATTS POUERED DIMENSIONS VOLUK 

CWPO)(ENT DAY DAY 84xx FTxFTxFT UTILBS CU FT 

1 SECONDARY EMU -- - 

EMU UNIT 1 170 170 - .95x.75~2 440.9 17.05 

EMU SERVICE STAND 1 '20 20 - 4x3x1.5 30 30 

EMU UNIT 2 170 170 - .95x.75x6 440.9 17.05 

EMU SERVICE STAND 2 10 10 - 4x3x1.5 30 30 

2 SECONDARY -- EYU SERVICE STATION 

ARLK VENT LW2 100 100 - 

ARLK COOL LOOP 30 30 - 

HIGH PRESS 02 LOOP 100 0 - 

PUR CNDIELECTRONICS 410 410 - 

HELMET DIS L CONT 25 25 - 

CO2 RFMOVAL REGEN 75 75 - 

HUMIDITY REMOVAL 8 200 200 - 

SUIT DRYER 

PO$ PRESS RELIEF 80 W - 

VLV CHECKOIJT 

TOTAL 547 63 

3 SECONDARY AIRLOCK - - - 6.7~8.3 1316 200 -- - 

EQUIPKNT ATTACHIENT 15 

LIGHTS 25 25 - 

GAS RECOVERY 1400 1400 - 

HYPERBARIC EQUIP - - ---- - lxlxl 50 1 

SECONDARY EVA TOTALS BACKUP SYSTEM F'0UF.R = 1: 

-. 2869.8 358.1 
TABLE 4.4.9-2. d - EVA - SUBSYSTEM WE IGHT/POUER/DIKilSIONS 

COMPONENT CCNSUMABLES RETURNABLES LOCAT ION 

1 EMU SERVICE STATION (2 EMU'S SERVICED) 

ARLKVENT 

ARLK COOL LOOP 

- 

1.25# 02/RCHG/MU 

HIGH PRESS 02 LOOP 

PYR CNDIELECTRONICS 

HELMET OIS & CONT MOD 

CO2 REMOVAL REGEN .4# N212 YEEKS 1.61 C021EMU MAX 

HUMID REMOVAL A SUIT .05# H21RCHG 3.2# H20/EMU MAX 

DRYER 750 CALORI ES/MAN/DAY 5# URINEIEMU PEAK 

POS PRESS RCLIEF VL'J 2.5% H201EMU MAX 2# URINEIEMU NM 

CHECKOUT 1.3# H20/EMU NOH 

SPARES (10%) ARlK 

IN SS 

FRONTAL AREA = 60 SQ. FT. 

2 MMU SERVICE STATION (2 HU'S SERVICED) 

NIGH PRTS S N2/RcHSm 

PWR/CND/ELECTRONICS 

s 

ARLK 

SPARES (10%) IN SS 

N2 STOR. TANK (CRY01 32 MMU RCHGSISO DAYS LOG OR 

N2 CRYOGENICS EXT 

3 EYU 

l7nrmI1T : 10W/HR Cm 8 4 ARLK - 5 

EMU SERVICE STAND i ARLK 

EN! UNIT 2 10O#/HR COOLING @ 45 ARLK 

EMU SERVICE STAHD 2 ARLK 

LCVG 1 1 WASH12 EVA'S IN SS 

LCVG 2 IN SS 

SPARES IN SS 

u 

4 MMU 

OuT'mLK 

)IMU SERVICE STAND 1 OUT ARLK 

MMU UNIT 2 3UT ARK 

MU SERVICE STANG 3 GST ARLK 

SPARES IN SS 

5 EVA TOOLS 

15R'R TOOLS (BAm) BA-LY OUT ARLF 

PWR TOOLS (UTILITY) OUT ARLK 

NON-POWERED TOOLS OUT ARLK 

EQUIP REPAIR TOOLS IN SS 
TABLE 4.4.9-2. e - SECONDARY EVA SYSTEMS POWER/VDSJHE/MASS - 

COMPONENT CONSUMABLES RETURNABLES LOCATION 

1 SECONDARY EMU'S 

1 l0Wm Cm 45 AlUx 

EMU SERVICE STAND 1 ARLK 

EMU UNIT 2 10WIHR COOLING @ 45 AWK 

EMU SERVICE STAND 2 ARLK - 2 SECONDARY EMU SERVICE ARLK 

2 

ARLK COOL Loop iZ :27LEF 

HIGH PRESS 02 LOOP 

PWR CNDIELECTRON ICS 

HELMET DIS & CONT MOD 

C02 REMOVAL REGEN .4f N212 WEEKS 1.6# C021EMU MAX 

HUMID REMOVAL 8 SUIT .05# HzlRCHG 3.2 H2IEMU MAX I 

DRYER 750 CALORIESIMANIDAY 5# URINEIEMU PEAK 

POS PRESS RELIEF VLV 2.M H20IEMU MAX 2# URINEIEMU NM 

CHECKOUT 1.M H201EMU NOM I 

3 SECONDARY AIRLOCK HABl HAB2 

LIGHTS 

GAS RECOVERY 

HYPERBARIC EQUIP - 

SECONDARY EVA TOTALS 
TABLE 4.4.9-2. f - EVA SUBSYSTW YEIGHT/POMER/DIMEMSIONS 

COMPONENT - CONSUHABLES RETURNABLES LOCATION 

6 MISC EVA EQUIPKNT 

IP ---mKmr 

EOUIP 'STOWAGE LOCKER OUT KK 

COM MOD EQUIP FM (NEEDED FOR INTERNAL AUK ONLY) 

EYA SYSTEMS TOTALS 

AIRLOCK 

FYAArRLOCK 1 HAB 1 ha 2 

EQUIP ATTACt!MENT 

LIGHTS 

GAS RECOVERY 3# A1 R/DEPRESS 

EQUIPENT AIRLOCK TBD 

LIGHTS 

GAS RECOVERY .65Y A1 R/DEPRESS 

HYPERBARIC EQUIP - ARK -- - 

AIRLOCK TOTALS 
TABLE 4.4.9-2-9 - EVA ACTIVITY BLOCK DESCRIPTION 

NON-EVA 

A.B. A.8. EVA DAY MU DAY DAY 

NO. NM DESCRIPTION TIE (HRS) TIE (HRS) TIHE (HRS) 

PREPEV 

DEPRES 

EVA 

WSRV 

WCHK 

REPRES 

POSTEV 

EMUCoirJ 

CHECK 

UTLPNR 

READY 

EVA PREP-IV IN AIaOCK 

EVkAIRLOCK DEPRESS 

EVA 

WU SERVICE 

mu CHECKOUT 

EVA-AIXOCK REPRESS 

POST EVA-IN AIRLOCK 

EMU RECHARE 

MU CHECKOUT 

EVA UTILITY PI@ 

EVA EQUIP READY 
5.0 UNMANNED PLATFORMS 

l'ht: unmanned Platforms operate as complementary extensions of the Space 

Station Manned Core capabil i ties. They will be of a size appropriate to 

be useful to most potential c~rstmers. The payloads they carry will be 

canprised of (1) a scientific instrument or a compatible set of 

instruments that have a similar field of investigation, (2) a technology 

development mission, or (3) comne~cial production units. 

As scientific and programatic constraints and requirements have matured, 

the design of Platfonns as c Space Station Program Element has ~volved 

through a number of stages. In its initial form, the unmanned Platform 

was envisioned as a large structure drawn directly from the Space Station 

structural configuration which could accommodate a large complement sf 

many varied instruments. Variations ori this approach to Platform design 

included use of a resource module dupl icated from the Space Station, a 

scaled-down resource module from the Space Station, and then finally use 

of subsystem elements fran the Space Station. It is expected that the 

basic configuration design would be comnon to both polar and co-orbiting 

platforms. As the design, cost, the user requirements have matured, the 

concept has evolved into a relatively small Platform design. This 

small er Platform would accomnodate different pay1 oad packages made up of 

complementary sets of instruments. One or more of these smaller 

Platforms could be placed into polar orbit andlor into co-orbit with the 

Space Station. 

A key feature of this most recent concept is the use of a Platform 

support services core. This core provides a common Pldtform to which all 

payloads can be attached. It is replicated as funding and instrument 

payloads are available, and the capability of the core tan be increased 

to accmdate larger numbers and/or more demanding payloads in the 

growth period. The basic core is designed to meet the requirements of 

near-term (IOC) missions. A growth core, which has expanded payload 

accomnodation features beyond those of the IOC module, has a1 so been 

defined. The growth conf igur .tion can be reached either by the 

fabrication of a new core spacecraft or by expansion of the payload 

accomnodation capabi 1 i ties by subsystem enhancement vhi 1 e the core 

remains in erbit. 

The Platform will have a near-hemispheric fielC of view in one 

direction. The field of view may be either inertial, for solar or 

astrophysics studies, or Earth oriented. Payload pointing control would 

be provided by the platform ,ore itself, without payload gimbals, at an 

accuracy level suitable for most customers. The core wi 11 accept fine 

error signals from the instruments if required to meet their objectives. 

The interface to the payload instruments attached to the Platform will be 

standardized to be compatible with the instrument pqyload interface on 

the Space Station. This will allow easy interchange ~f payloads on the 

Platform. It will a1 so permit convenient exchange of instrument payloads 

between the Space Station kanned Core and the Platfonns. 
The Platform design will use Space Station Manned Core elements to the 

maximum extent that is cost effective. The basic elements of the 

Platform Core will be modular and will incorporate standard interfaces to 

allow in-orbit servicing, repair, and upgrading. This modularity will 

permit growth of the l'?i,tform capabilities while on-orbit to meet the 

increasing needs of future customers. 

The near-term core design is sized to require only one STS launch to 

place it in orbit with payload. This consideration limits the total 

Platform launch weight into polar orbit to about 30,300 lbs. The 

Flatform Core will have initial capability to provide 5 kW of electrical 

power for customers. This is sufficient for most scientific 

investigations. The growth capabil ity of each IOC-type Platform Core 

will be to 20 kW which is needed by some post IOC missions. 

5.1 DESIGN REQUIREMENTS 

The design objectives given in section 5.0 are subject to trade studies 

to determine the best approach for implementation. For the purpose of 

defining a reference configuration, tbe set of design requirements stated 

below has been derived frm. these objectives based on the needs of 

potential users a~d the requirements for Platforms that are stated in the 

RFP . 

5.1.1 General Design Requirements 

The primary design objective for the Platfonns is to provide a versatile, 

gtwwing, permanent, umanned facil i ty ir, space with the capabil ity to 

accmpl i sh significant advances in space science, techno1 ogy , and 

cocnnerce. The Platform design features are primarily driven by payload 

accommodation. The Platform must operate in several mlssion dependent 

lllodcs: 

a. Nadir orientation. 

b. Inertial orientation; instruments Sbn pointed. 

c. Inertial orientation; instruments celestial pointed. 

Design requirements for the reference Platform configuration have been 

derived from these objectives, both for an initial (near-term) and a 

growth capabi 1 i ty . These are sumarized below. 

a. The IOC Platform will be capable of being placed in orbit by a 

single STS launch with a complement of payloads. 

b. Servicing interval will typically be 2 years with a 1-year 

contingency. Hardware elements wi 11 be serviceable using Orbital 

Replaceable Units (ORU's) or an equivalent approach. Servicing will be 

prmided by the STS, the Space Station, or via Orbital Maneuvering 

Vehicle (OMV) in combination with one of these elements. 
c. A single Platform Core will be used to support a1 1 Platform 

missions. Naximum use will be made of Space Station hardware and 

software elements. Post IOC missions may require growth of the Platform 

Core. 

d. Any single failure will not preclude continurl normal 

operation of the Platform. Any subsequent failure wi 11 not preclude 

operation in a degraded mode until service can be provided. 

e. Platform will have a minimum operating life of 10 years, with 

maintenance, and a gor,l of 15 years. 

f. In polar orbit, the Platform will be capable of a nadir 

orientation and will provide an unobstructed view of the Earth including 

the 1 imbs of the atmosphere. 

g. In co-orbit the Platform will be capable of either nadir or 

inertial orientations. (Co-orbi t mans that the Platform's inclination 

wi 11 be the same as the Space Station Manned Core. ) 

h. The configuration of the Platform will accommodate selected 

sets of instruments listed in the current Space Station Mission Data 

Base. The reference configuration herein uses the data base as of July 

1984. However, the Platfarm Core is a1 so designed to accomnodate future 

undefined missions without basic changes. 

i . Capabil i ty wi 11 be provided both for the enhancement of 

payload accommodation through growth on-orbi t during the 1 ifetime of an 

initial Platform, as well as for the expansion of payload accommodations 

for future dssions through a growth Platfom design. 

j. The primary comnunication link for the Platform will be 

through the TDRSS. 

k. The Platform will be capable of operating autonomously wIth a 

minimum of ground control. Telemetry status of all systems will be 

provided for monitoring and for maintaining an abdi t trzil . 1. On-orbit assembly as well as instrument changes or upgrades 

may be used as required. 

m. For cost estimating purposes, it was assumed that the Space 

Station Program will provide the Platform Core subsystem hardware and the 

payload support structure for the Platform. Instruments will be provided 

by the customer. 

n. If greater pointing accuracy is required than that provided by 

the Platform Core, then fine error sf gnals will be provided by the 

customer instruments. 
o. The Platform will be designed so that normal functions do not 

interfere with the performance of the instrument sensors or sensitive 

surfaces. The instrulnents will be designed and grouped such that they 

will not contaminate neighboring instruments. Instruwnts that are 

especially sensitive to contamination will provide door mechanisms to be 

closed just prior to a propulsive event or the beginning of another 

contamination situation. 

p. The design must meet instrument viewing requirements and 

provide an adequate view to cold space for instrument detector active 

cooler radiators. 

q. Solar array area wi 11 be determined by overall platform power requi rements, but the 1 ocation, aspect ratio and deployment schemes nurst 

meet the foll wing requirements: 

(1 ) Solar and aerodyr~~~nic toques kept within the control 

capabi 1 i ty of the GNgC system. 

(2) Payload fields of vied requirements including reflected 

1 ight, thermal radiation and radiator view fie1 ds. 

(3) C~pabil ity to fit within the Shuttle cargo bay. 

(4) Minimum shadowing of array by platform structure. 

5.1.2 Pqyloads and Missions 

A larage, representative group of instrument has been identified for both 

initial and growth Platform concepts. These instruments will require 

both co-orbi ting and polar orbits. It is not suggested that this entire 

-epresentative group of instruments be present on-orbi t at any one time. 

.,ather, through consideration of the myriad of requirements raised by 

this group of instrunrents, an understanding of the sizing of both the 

initial and growth Platform Core modules may best be real ized. To this 

end, the representi-tive group of instruments has been partitioned into a 

nunber of mission payload sets on the basis of viewing objectives and 

power requirements. The following tuo sections describe these mission 

payload sets and their accompanying mission objectives. 

5.1.2.1 Polar Pqyloads and Missions 

The pqyloads identified for polar orbit support a strategy for obtaining 

integrated Earth science measurements from 1 ow Earth orbit. Synergistic 

groupings of instruments, ca led mi th near-sinul taneous measuremnts, 

are directed toward a caraprehensi ve mu1 ti di scipl inary appmach to 

understanding the Earth as a system. This is the basis behind the Earth 

Observation System (EOS), which nqy ultimately provide a facility 

capabil i ty for additional users such as NOAA and commercial customers. 
The EOS payloads are comprised of a wide range ~f instruments spanning 

the spectral region from the UY to the submill imeter microwave. The 

instruments cover appl icntions in oceanic, atmospheric, solar, and land 

resources sciences. The instruments ha\.e been partitioned lnto four 

mission payload sets, and the composition of each Platform pdyloid is 

addressed below. 

The initial mission parameter: of a Sun synchronous orbit with a 2 p.m 

descending node were chosen on the basis of global coverage and 

consistency of illumination. The 2 p.m. nodal crossing was chosen to 

avoid the Sun glint over water for orbit crossings near noon and to 

provide maximum contr8st with the 2 a.m. ascending node for the thermal 

and soi 1 mi sture measurements. 

The EOS science requires good "WO-dqvU coverage. In considering orbit 

altitudes for good "two-dqy" coverage, further trades were made by 

evaluating the spacecraft environments of radiation, contamination and 

drag together with amspheric path length for measurements at the 

extreme scan posjtions and the resulting pixel distortion compared to the 

nadir view for the wide swdth instruments. 

Based on these considerations, the a1 ti tude and incl ination required for 

the polar instruments are 380 mi and 98.2", respectively. This orbit 

has a 16-day repeat cycle. 

a. Polar mission payload A is considered an early mission and 

carries a collection of all the land and ocean instruments which require 

consistent, global coverage (two-day repeat cycl el. The instruments are 

(1 ) The Moderate Resolution Imaging Spectrometer (m)3IS) 

(2 ) The High Resol ution Mu1 ti frequency Microwave Radiometer 

( HWR 

(3) The LIDAR Atwspheric Sounder for Aerosol Measurements 

( LASA-A) 

(4) The Radar A1 timeter (ALT) 

(5) The Scatterometer (SCAT) and 

(S) The Autanated Data Collection and Location System (ADCLS) 

b. Polar mission payload B is considered an early rni5sion and 

comprises high spatial resolution, small swath-width, target-selective 

instruments for land and some ocean applications as well as their 

supportive ~nstrumentation. The payload incl udes: 

!1) The High Resolution Imaging Spectrometer (HIRIS) 

(2) The Thermal IR Mapping Spectrometer. (TIMS) 
(3) The LlDAR Atmospheric Sounder for Aerosol Measurements 

( LASA-A) 

(4) The LIDAR Alti&ter (LA%-B) 

c. Polar micaion pqyload Synthetic Aperture Radar (SAR) is 

assigned essential ly its own Platform to maximize the available power and 

data capability. Salar and magnetospheric monitors will also be included 

in this Platform. This is also considered an early mission. 

d. Polar mission payload C is considered to be a later mission 

set. It includes the following atmospheric instrumnts to obtaln 

simult~~~..?ous coverage of lower and upper amspheric processes and 

cmposi ti on. 

( 1) Correlation Radiometer (CORAD ) 

(2) IR Interferometer Spectrometer (MINS) 

(3) Diffe:.ential Absorption LIDAR (DIAL) 

(4) Doppler LIDAk (WPL13) 

(5) IR Radiometer (IRAD) 

(6) Tilting Etalon Spectrometer (TES) 

(7) Submi 1; ineter Spectrometer (SUMTS) 

(8) Microwade Limb Scannor JNLS) 

(9) 1IVIVi si bl e Spectrometer (UYY S) 

(10) Fabry-Perot Interferometer (FAPIN) 

(11) Michelson Interferowt~r (MIKIN) 

Table 5.1.2-1 summarizes the pwlo~C characteristics f~r the po'l ar 

mi ssi on pay 1 oads . 

!::insidered within payload requirements for polar missions A, B, and C are 

addf tional re ai rements for growth. These include 8C3 watts power and 

0.2 %Is data rate for instruwnts ur experiments. Ueight margin is not 

explicitly shown because with in-orbit integration of new !nstruments. 

weight margin is very large. 
TABLE 5.1.2-1- POLAR MISSION PAYLOAD CflARACTER1STICS 

AVERAGE 

ORBITAL DATA POINTING 

EIGHT OPERATING RATE CONTROL, ALTITUOE 

L B .- ZUR, Y Mbps ARC SEC - NM I - LAUNClf -- 

POLAR A 5,456 4,500 5.4 360 380 NLAR TERM 

POLAR B b,1+0 4,000 188.9 360 380 NEAR TERM 

POLAR SAR 2,508 2,800 300.0 360 38G NEAR TERM 

QOLAR C 10,037 7,400 2.4 60 380 LATER 

The numbers listed here are representative of these classes of pqyload 

instruments. The Space Station Mission Data Base should be consulted for 

speci flc values. "hear-term" configurations are candi dates for the IOC 

capability; "1 ater" configurations are growth. 

5.1 .?.i Co-drbi ting Mission Pqyloads 

PqyloaGs : flown as edrly missions in co-orbit are a collection of 

solar and ils~ropkysical ins,truments The 1 arge masses, nominal a1 ti tude 

selection, 2nd gointing requirements of the instrlrnrents suggest that most 

of the l~struments b? considered separna,e mi ssio? payloads. The 

instruments to fly in co-orbit are as follows: 

Co-orb i t pqy 1 oad S I RTF .Space Infrared Telescope Facil i ty 

Co-orbi t pay1 oad Star1 ab S tar1 ab 

Co-orbi t pay 1 oacl p/OF Pinhole/Occul t6 on Faci , Co-crbi t pqyload AS(; Advanced Solar Observatory 

Co-orbi t pay1 ond Hip High Throughput Experiment 

The a1 titude and inclination require~tlents of these instruments are 270 

lmi and 28.S0, -.espectiv?ly, except for SIRTF whfch is 378 to 540 nmi and 

28.5". The 378-540 nmi ,iltitude for SIRTF was selected to minimize 

optical contam:'nation. Note that co-orbi ting means that the Platfonn 

will be flown at the same inclinatiun as the manned Space Station. 

Table 5.1.2-2 sunmarizes the instmnent characteristics for cach of the 

co-orbi ti sg Pl atfoms. 

An exanple of a growth mission for co-orbi tin9 Platforms would be a 

matwials piocessing production unit which could re~uire 20 kY of power. 
TABLE 5. i .2-2.- CO-ORBITING PLATFORM PAYLOAD CHARACTERISTICS 

AV E 2.SE 

ORB1 TAL DATA POINTING 

E:!?IT OPERATING Z4TE CONTkOL, ALTITUDE 

LB F2.W Mbps ARCSEC -. - MI LAUNCH 

SIRTF 8,82<1 1,000 1.0 0.15 378-540 NEAR TERU 

STARLAB 7,956 2,400 16.0 2 2 70 NEAR TERn 

P/OF 7,938 700 1.4 10 270 NEAR TERM 

AS0 19,625 4.500 42.0 1 270 LATER 

HTP 22,050 2,400 0.3 3 270 LATER 

The fol lowing section describes the Platfom reference configuration. The 

reference Platform Core configuration is shown in 3-view planfom in Figure 

5.2-1 and in perspective in Figure 5.2-2. Three Platfom configurations 

equipped with various instrument payloads are shown in Figures 5.2-3, 5.2-4, 

and 5'2-5. 

The reference description provided here is based on studies which maximized 

the use of Space Station Core hardware in Platforns. Equal importance was 

placed on maximizing utility to the users by meting instrumnt requirements 

as defined to date. This reference configuration description is provided 

only as a departure point for Phase B studies and is not considered to 

represent an optimized desi gii. 

5.2.1 General Arrangement 

A Platform consists of a Platfom core and an instruinent payload. These 

will normally be launched as an integrated unit. However, some pqyloads may 

be better launched independently and integrated with the Platform Core in 

orbit. 

A Ylatfonn Core provides support services to the pqyload. It provides 

pointing and accepts fine error signals for improved pointing for one 

instruarent. Individual instruments provide finer pointins, if required. 

Attitude control is provided by mmentum excharge with propillsirre 5ackup. 

Ttte Platfom Core provides its own thermal control using passive techniques 

(i.e., no mechanical pumps) but does not provide heat rejection capabl lity 

for pqyloads munted to the end of the Core, such as the Star1 ab pqyload. 

Electrical power is provided by solar arrays with batteries for energy 

storage. 

8' TORSS ANTENNA 

ANTENNA MAST 

DEPLOYABLE 

OPiNlNGS FOR 

STARTRACKERS (6) 

FIXED RADIATORp 

HEAT EXCHANCER 

FLATBED ' , 

STRUCTURE */ 

2' FND OF FLAT 

BED RESERVED FOR + -- 

AN INSTRY 

MODULE 

MDDULES rJOTH SIDES // A 

OF FLAT RE0 

- POWER 

- C&Dd 

- RF 

-- GN&C 

Figure 5.2-2 - CORE PLATFORM 
Flgure 5.2-3 - STARLAL AVLOAD PLATFORM 


The Platfom Core provides data management and data links for all 

instruments. Engineering data on each subsystem's status will be 

included in the data stream. 

The Platform Core is illustrated in Figures 5.2-1 and 5.2-2. The 

backbone of the Core is a flat bed structure. It provides the 

mechanical. electrical, and thermal interfaces tn all Core subsystems. 

These subsystems are contained in orbital mplaceabl e unit (ORU) 

modules. P. bulkhead is rnounted to one ?nd of the flat bed. It provides 

mountings for the propulsion systen fuel tanks and als~ carries the two 

Shuttle rill trunnions that handle + x and + z launch loads. Two 

additional sil l trunnions that carry only +-z launch loads are mounted 

near the other end of the flat bed. The kSel pin that carries all the t - 

y lallnch loads is mounted on the bulkhead (not visible in the figure). 

On the side of the bulkhead away from the flat bed, a truss ctructure is 

deployed ii orbit. This truss is built of the same elements as the 

manned Space Station. This structure carries the solar arrays and the 

propul sion thrusters. 

The solar arrays are identical to the Space Station's arrays except for 

the initial configuratian, where each half would be 29 feet long. A full 

80-foot-long array could be used in the later, high power growth 

configuration Platfoms. 

The reference co8rf isbration uses a sing1 e dcgree-of- freedom for arrz,y 

articulatinc. lhis resu! ts in up to approximately a 30-degree cosine 

loss for the polar Platform arrays. The stellar and solar platforms do 

not scffer this loss. 

For the polar Platforms, the array location was chosen to be 27 feet out 

on the deployed truss. This location gives (1) unobstructed instrument 

viewing of the Earth's linb, (2) minimum shadowing of the array by the 

body of the platfonn, (3) maximu~r view to cold space by the radiators, 

and (4) clear field-of-view for the startrackers. 

For the stellar and solar viewing Platforms, the array locdtion was 

selected to provide minimum aerodjnamic and gravity gradient torques. 

This is illustrated in Figure 5.2-3 showing the Starlab Platforn. One 

solar array is located 9 feet out on the deployed truss behind the 

bulkhead. i-he other array is deployed on a second truss 9 fest forward 

of the Platform Core and to one side. This arrangement does not 

interfere with the field-af-view of the star trackers or of the stell 

or solar instruments. Shadowing of the arrays by the Platform is not a 

problem because tk.e .,un is always noma; to the array. The view to space 

by the radiator^ ' ;lightly less but this is ampensated for by the fact 

that the Sun nevt .irectiy shines on any radiat~r surface. For the 

stellar and solar mirsions only, balancing weights and an adai tional 

ba?ancing boom are required to tnaintair! gravity gradient torque within 

allowable 1 imits . The born is cot illustrated in Figurs 5.2-3, but it 

wo~ld bs located on the redr truss directly opposite the 
solar array. It would be made of the same deployable mast as used fcthe solar array and be the same length. The balance weights used f~ 

analysis were: 1000 lbs on the end of the balance boom, 275 lbs addeb to 

the end of the solar array opposi te the balance boom, and 1262 Ibs added 

to the end of the boom that carries the TDRSS antenna. 

The last 2 feet of the Platform Core flat bed is reserved for instrument 

mu :ting. This is adequate for several of the single large instrument 

platfunns. Figure 5.2-4 shows the polar SAR instrument ~rounted in this 

location. 

Multi-instrument payloads will require addi ti~nal mounting space. This 

will be provided by util izi,;? additional extension flat bed structures 

attached to the end of the Core flat bed structure. A1 1 flat bed 

structures will incorporate standard instrument10RU mounting pr~visions 

so that payloads can be replaced and serviced as well as transferred 

between Platforms and tnc Space Station mounting points with minimum 

difficulty. A maxim~n growth-stzed Platform is envisioned to have three 

flat bed structures attached tr, the Core flat bed for a total of four. 

Figure 5.2.-5 illustrates Polar Platform with pqyload B including the 

growth DIAL instrument. This Platform would be launched with two 

additional flat beds and with instruments in all locations excrpt the 

last 5 feet of the additional flat beds. For better launch 'Isiids 

distribution, the 2 + z sill trunnions and the keel pin would be 

relocated from theirnormal positions on the Core fl~,tbed to locctions 

on tho fnstr~ment flat beds. The DIAL 'rioii:d be added to the Platfom on 

the available 5 feet of flat bed on a later Shuttle servicing mission. 

Another flat bed structure and more instruments could be added on a later 

Shuttle fl ight. 

Each instrument module and ORU module will have built-in heat pipes 

coupled to a flat bed radiator system. It is anticipated that the sam 

standard module design used by the Core ORU modules would also be used by 

the instruments. In order to standardize the flat bed mounting hard 

poizts, these modules would cone in 1-foot width increment, with a 

minimum width of 2 feet. All modu:es will be removable by a Remote 

Ienipulator System. The figdres show standard RMS grapple fittings on 

each module to illustrate this feature. The actual handling fittings 

will probably be much smal ler and 1 ighter. 

5.2.2 Weight and Power - 

The properties of the reference Platfoms which can accommodate the nine 

pqyloads discussed in Section 5.1.2 are sumarized in Tables 5.2.2-1 and 

5.2.2-2 

Table 5.2.2-1 shows the estimated weight of the reference design 

Platforms. A comnon basic Core Platfonn was assumed for the design of 

all mission Platforms. The weight differences shown in Table 5.2.2-1 

reflect far example, the different truss/sol ar arrqy configuration for 

the polar zr?d crr-orbiting rpis;ion Plii~f~i3s, the use of balancing weights 
on the Starlab Platform, and additional power capability required for the 

Polar SAR and Polar C: mission payloads. Each near term mission pqyload 

can be accolmrodated in one 1 aunch. The AS0 pqyload may have to be 

launched as two missions because of instrument size. The Polar C mission 

payload exceeds the estimated STS capability for polar launch. Both 

require further stu*. 

Table 5.2.2-2 shows the estimated power requireaents and the 

characteristics of the array and energy storage devices required to 

provide the rquired power. The Core Platform orbltal averdge power 

estimate is 2896 U. 

TABLE 5.2.2-1 - REFERENCE PLATFORH WE lGHT 

CORE 

PLATFO W 

MISSION L t! 

NEAR TERM: 

SIRTF 11,902* 

STARLAB 14,439 

FLAT BEQ 

INSTRUKMT TOTAL 

PROPELLAW INSTRUENTS STRUCTURE P LATFO RM 

L B LB L B LB 

POLAR B 11,795 5,700 8,140 4,600 30,235 

POLARSAR 11,995 5,700 2,508 0 20,203 

- LATER: 

AS0 11,902* 2,900 19,625 0 34,427 

HTP 11,902* 2,900 22,050 0 36,852 

POLAR C 12,395 5,700 10,037 4,600 32,732 

* Required inertia balancing weights for these missions have not 

been identified. 
TAELE 5.2.2-2 - REFERENCE PLATFOM POUER 

ORBIT& MINIMUM MINIMUM 

PEAK PLATFORM Nd ERAG E ARRAY BATTERY 

HISSION BUS POWER, w POWER, Y - AREA, SQ FT - EIGHT, LB 

- NEAR TERM: 

SIRTF 4,895 3,921 882 1,074 

STARLAF! 6,796 5,308 1,194 1,341 

POLAR A 7,391 7,391 1,722 1,629 

POLAR 8 9,946 6,673 1,582 1,788 

POLAR SAR 19,596 5,818 1,420 2,079 

- LATER 

POLAR C 10,381 10,292 2,399 2,265 

5.2.3 Flight Modes 

5.2.3.1 Polar Platforn 

In nonnal operations, the polar Platform will be in a Sun-synchronous 

orbit with a 2 p.m. descending node and a 2 a.m. axen~ing node. The + z 

axis will be aligned with the local vertical (3 .e., Earth-viewing), the y 

axis will be normal to the orbit plane (POP), and the + x axis kill be 

alignec' along the direction of flight as shown in Figure 5.2.3-1. 

Reaction wheels are the prime Platfom actuators to control all external 

disturbance toques. Magnetic torquers will te used for reaction wheel 

desaturation as resulting from the accrual of momentum due to secular 

di sturbance toques. 

The solar arrqy is mintdined such that the nonnal is aligned at a nearly 

constant 300 angle with the Sun vector. 

All Platforms will be launched in a fully configured mode (i.e., core 

with instruments). Therefore, the initial flight mode will remain 

unchanged throughout the Platform's 1 ifetime, notwithstanding the later 

enlargement of a Platform by the addition of new Instruments. The only 

occasion for an interruption in the nomal flight mode will be during a 

servicing operation when the polar Platform will be deboosted to a lower 

altitude for rendezvous with the STS. 

5.2.3.2 Co-Orbiting Platform 

In normal operation, the co-orbiting Platfom will be flown at the same 

inclination as the manned Space Station, WSth only a few exceptions 

(e.g., SlRTF), the flight a1 titude will be the same as the manned Space 

Station i i .e., 270 mr~i). The orientation will be in one of two inertial 

modes, ste? lar or solar (Figures 5.2.3-2 and 5.2.3-3). The + x axis will 

point inertial ly to the observation object in space. The y and z axes 

will be oriented, in conjunction with solar array rotation, such that the 

Sun vector is normal to the solar arrqy. As a result, the + y axis faces 

of a co-oryiting Platform, which are radiating direct'ons 07 the fixed 

radiators, wii 1 have no direct solar input. 

As explained for the polar Platform, the initial flight mode will remain 

unchanged except when interrupted by a servicing operation. For the 

co-orbiting Platform, one servicing operation mode will entail transfc~ 

to, and berthing wf th, the manned Space Statian. 
TDRS WST 

171 " W. LONQ. 

c. rn 

4 a- 

SUN 

Figure 5.2.3-2 - STARLAB CB-ORBITER PAYLOAD PLATF6RM 
TORS WE81 

Flgura 5.2.3-3 - ADVANCED SOLAR OeSERVA'rORY CO-ORBITER PAYLOAD PLA
5.2.4 Launch and Serving Operations 

5.2.4.1 Launch Operations 

Co-orbiting Platforms will be launched from the Eastern Test Range (ETR) 

and polar Platforms fm the Western Test Range (UTR). All reference 

polar Platfonns and the reference co-orbiting Platform at 378 to 540 nmi 

(SIRTF) require orbit transfer from a lower initial 1 aunch altitude. 

.+ssuned STS payload launch capability in the early 1990's is 

ETR - 28.S0, 270 n~i = 58,400 lb 

378 to 540 mi = Not attainable (pqyload transfer from low 

orbit required) 

150 nmi = 72,000 1b 

WTR - 98.Z0, 3% mi = Not attainable (payload transfer from low 

orbi t requi rc 1 

150 mi = 30,300 lb 

If the above assuaed payload launch capabilities are determined to be too 

high, the heavier mission payload weights can be reduced by removing 

selected instruments. 

NOTE: The projected STS performance capabi 1 i ties (1 aunch and landing 

3qyload weights) [nay be in conflict with the above numbers. The baseline 

STS perfownc.e infomation to be used during the Definition Phase studies 

wil: be provided via the associated SSP/STS Pqyload Integration Plan which 

will be a WP applicable document. 

A transfer propu!sion subysten is i,icluded in the reference design, with 

the OMV providing an alternate approach to orbit transfer. For those 

Pl~tfoms which operate at altitudes exceeding the STS operational 

capability, the transfer propulsion subsystem will be used to lower the 

Platform for rendezvous with the STS for servicing. ,4fter servicing and 

refuel ing, the transfer propulsion subsystem will re-boost the Platfom to 

operational orbit until the next servicing operation is required. The 

transfer propulsion subsystem will contain orbit adjust propellant for a 

minfrnuin of 2 years between service intervals plus 1 year reserve !total 3 

years). 

When a growth Platform has exceeded the total propulsion capabf 1 i ty of the 

Core tanks, several a1 ternati ves are possf bl e: (1 ) an additional 

propellant tank(s) could be added, (2) the OMU could provide part of the 

delta velocity, and/or (3) the Shuttle could go to a higher altitude if 

the servicing weight was less than skawn in Table 5.2.2-1. 

5.2.4.2 Sarvicing Operations 

There are several optians available for Platfonn servicing. The polar 

Platforms may be serviced in-situ by the STS or the OMV. The co-orbitf ng 

Platforms nqy be serviced in-si tu by the OHV or thp STS, on the Station 

after berthing, or within the proximity operational zone of the Station 
Figure 6.2.4-1 - PLATFORM CONFIOURATION 

by EVA crew. Each option has operational advantages and disadvantages. 

The Platfom configuration evaluated is shom in Figure 5.2.4.-1. 

5.2.4.2.1 Servicing by STS/OUV 

5.2.4.2.1.1 Polar Platform 

In-situ servicing by the STS/OW will have ta be developed to support the 

polar Platfom. Routine servicing of the polar Platfonn could be performed 

in its orbit by an O)(V brought up by the STS. For this option the OMI 

must have basic servicing capabil i ties, such as consuables rep1 enishment, 

payload changeout, and possibly limited pqyload repair capability. For 

service operations beyond the W robotics capabil ity, it i '1 1 be 

necessary to retrieve tile polar Platform and berth it in the STS for 

EVAIRMS repair/service. The Platform would be debaosted to a lower omit 

(either by use of its own propulsion system or by use of the OW) for an 

STS rendezvous. In general, the Platfom propulsion system will be used 

to &boost to a lower orbit prior to STS 1 aunch, so that the STS may 

rendezvous with the Platform. If the OMV must be used for deboost, then a 

double rendezvous is required. The STS would be launched in a cmpatible 

orbit, the OW would rendezvous with and retrieve the Platforn, and then 

the STS wwld re-rendezyous with the W/Platfom. 

Of sajor concern in servicing the Platfom fraa the Orbiter is capturing 

it with the RMS. The deployed appendages on the Platfom are a definite 

hindrance to capture wi th the MS. If these appendages, the TORSS antenna 

and mast in particular, cannot be retracted, then grapple fixture location 

and track and capture procedures become nqjor drivers in servicing. This 

is also true for capturing the Platfom at the station with the KRMS. Once 

the Platforr is captured and placed in a surrogate pqyload bay fixture 

above the pqyload bay as shown in Figure 5.2.4-2, the RMS/MR have 

adequate access to all parts of the Platforn. 

5.2.4.2.1.2 Co-Orbiting Platfonn 

Two in-situ servicing optior,, exist for the ca-orbiting Platfom. The 

first using the STS, and the second the OMV. Although STS schedule impact 

is the main disadvantage, using the STS for co-orbiting Platfom rervicing 

is a viable option. It is conceivable th3t a servicing operation could be 

combined with an STS sate1 1 ite deployment mission if the orbits are 

compatible. It is a1 so conce?vable that a Platforn servicing mission 

could be caabined with a station resupply visit, except that the STS will 

contain a logistics module on both the up and down legs of the mission, 

plus the extra crew aboard for the crew rotation. 

The second in-situ servicing option for the co-orbiting Platfom consists 

of sending the Space Station based OMV to the Platfom location. The OMV 

will have certain capabilities developed for polar Platform servicing 

which could be applied for co-orbiting Platfom servicing. Even though 

the Platfom is initially co-planar, differential drag and nodal 

regression between the Station and Pl atfonn may require considerable 
1 

Platform pr~pulsive capability to maintain the Platfonn location relative 

to the Station to meet the servicing interval requirements. Despite these 

co~straints, the Platforn should be accessible to the OMV. 

If the Plotfom is placed at 28.50 inclination at a1 titudes other than 

the Space Station, then differential nodal regression will force the 

servicing intervals to be periodic, since the Platform fuel supply is not 

sufficient for large nodal changes. For instance, a 108 mi altitude 

difference will force the servicing intervals to be no more often than 21 

mnths. 

5.2.4.2.2 Servicing by Space Station 

Utilization of the Space Station for Platform servicing offers two 

candidate scheaes based on either direct on-Station or Station proximity 

zone operations. 

5.2.4.2.2.1 On-Station operations 

Several options for Platfonn servicing on the Space Station have been 

identified. If the Platforn is designed such that its solar arrays and 

antenna boom are retracted when the Platform arrives at the Station, the 

existing satellite servicing bays may be used for Platform servicing if no 

more than %o additional flat beds or their equivalent are added to the 

Core configuration. Retention in the bay can be accomplished either by 

surrogate payload by a structure grasping the trunnions used to launch in 

the ST$ or by an FSS with an adapter attached to the Platform. 

If the boon and/or arrays cannot be retracted, the servicing bays can be 

used only if the Platform is rotated away from the Str ion keel. 

Provision for a servicing position on the -x side of ii~c keel would permit 

boom and arrays to remain deployed. 

Satell ites to be serviced will be captured by the MRMS at the lower end of 

the Station and transported to the servicing bay or other location. A 

surrogate payload bay system is mounted on the MRMS in such a way as not 

to interfere with MRMS operations, as indicated in Figure 5.2.4-3. 

Because of the limited area available on the base for such attachments, it 

is possible that an FSS or equivalent, possibly mounted on stand-offs, 

will be required. Further study will be needed to evaluate these options. 

Several locations on the reference Platform configuration for mounting an 

FSS adapter have been evaluated. The first location is at the end zf the 

flat bed. This has the advantage of similarity to other satellites to be 

serviced, so that few special procedures will be required. The 

manipulator would have ample clearance for mating and demating the 

Platfonn. In this option, adding flat bed or other payloads to the 

Platfonn is not possible without removal of the adapter from the 

Platform. However, the servicing bays will have surrogate payload bay 

structure to handle satellites that do not incorporate an FSS adapter. 

Use of this capability to retdin the Platform in the bay would permit 

addition of flat beds to the Platform. 
1 RUNN I ONS 7 r SURROGATE PAYLOAD BAY 

GRAPPLE FIXTURE 

I x' - LOCKED SOLAR ARRAYS 

Flguro 5.2.4-3 - PLATFORM ON MRMS 

K
The second FSS adapter location studied is on the end of the Platfon near 

the RCS. This location avo1 ds the problem of FSS removal for flat bed 

extensions and can be designed for conpatibil ity with the FS. 

As the Platfom configuration becomes better defined, a lore detailed 

analysis of grapple fixture locations for RHS/IIIWS hand1 ing will be 

necessary. Detailed investigation of accessibility to specific areas of 

the Platfom using both IMSNFR and MRnS/IIFR will a1 so be required. 

5.2.4.2.2.2. Station proximity zone operations 

The option of bringing the Platforn into the proximity operations zone of 

the Station for servicing by EVA crew has been considered. The proximity 

operations (PROX OPS) zone is a 1 km sphere centered at the Station. The 

Platforn propulsion system or OMV would be used to bring the Platform to 

the rendezvous zone. For safely, the OW would be used to finally bring 

the Platfom into the PROX OPS zone, since the OW will be designed with 

redundant syte~s, possibly auto rendezvous, and man-in-the-loop control 

capabil ity. The Platfom propulsion or the OMV must have the capability 

to retrieve rnd bring the Platform to the proximity of the Station, and 

the Platforn must be miintined in an orbital position relative to :;he 

Station that meets the servicing interval requiremnts. The main 

advantage of this option is the accessibility of EVA crew to the 

Platfon. The (JCIV capability designed for polar Platform servicing could 

also be used. The operation wwld be performed at ir stand-off position of 

several hun?:-ed neters because the safety risks associated with close I 

quarter statfan-keeping (say a few tens of meters), possibly wi tR EVA crew !' 

in the vicinity. Even at several hundred meters, the Platform must 

station-keep with the Station. Platfom attitude can be control led with 

the onboard reaction wheels, but translation maneuvers still must be made 

using the Platfom propulsion system. EVA crew should not be in the : 

vicinity cf any thruster firings. In order to minimize orbital mechanics 1 

effects, the Platfom should be placed on the V-bar in the PROX OPS zone. 

This night allow shutting-down the Platfom thrusters for periods of free -3 

drift during which the EVA crew cauld perform servicing. The period of 

free drift is dependent on many variables, and studies would have to be 

perforated to see if it is even feasible. In addition, lighting must be 

provided for the EVA operations. Finally, EVA ti= is limited, so that 

servicing operations must be -areful ly scheduled. If an operation should 

take longer than EVA time constraints, the Flatfam could be mved to a 

stand-of f posi ti on of several ni 1 es for re-render vous 1 ater . Consi dering 

the safety imp1 ications, the flight mechanics problems, and the scheduling 

cql ications, proximity servicing does not look prmisirig. 
5.2.5 Dynamics and Control 

A 1 imited armunt of analysis of the Platform has been performed. Work has 

been focused on three areas: rigid body control requirements, reboost 

control by the RCS, and uncontrolled flexible response. Two cmbi nations 

of the Platfonn Core with payloads have been analyzed. The results 

presented are intended to provide insight as to where further efforts 

should be concentrated. 

5.2.5.1 Disturbance Forces and Torques 

A 1 i st of applied forces and torques is given in Secti~n 4.3-3.1 for rhe 

IOC Planned Core Station. Obviously, not all the items listed are 

applicable to the Platfonn. Of the natural enviro.ments, gravity qradient 

torques and aerodynamic forces and torques were considered. Both are 

characterized in Section 4.3.3.1. Induced enviromnts studied were RCS 

reboost and berthing which are characterized in the fol1 owing sections. 

5.2.5.2 Reaction Control System 

The RCS control system used t~ maintain attitude control during the 

reboost maneuv2r is described in section 4.3.3.2.2. Using the Polar B 

configuration and the autopilot data given in Table 5.2.5-1, the autopilot 

hysteresis (h) was parameterized to deternine its effect on 1 imit cycle 

frequency. Using a value of 0.5 degrees for hysteresis, Figure 5.2.5-1 

shows the torque acting on the Platform as a function of time. Figure 

5.2.5-2 illustrates the Platform 1 imit cycle. The torque history was used 

to drive a structural dynamics response nodel of the Platform, reported on 

and discussed in Section 5.2.5.2. 

TABLE 5.2.5-1. - AUTOPILOT DATA VALUES 

KD - - 1.0 degrees/ (degrees/sec) 

DB - - 1.0 degree 

h = 0.0 to 1.0 degrees 

T2 - - 583.0 ft-lbs 

TL - 

- 

184.0 ft-lbs 

I - 225876 sl ug- f t2 

5.2.5.3 Structural Dynamics 

A NASTRAN finite element model was consbc~cted (Fig. 5.2.5-3) for the 

Platform configuration of Figure 5.2-5. The Core body was modeled as a 

conmon element of the Space Station central keel. This structure is a box 

beam of 9-foot-square cross-section consisting of grsphite/epoxy tube 

material. The beam properties for this section were extracted directly 

from the Space Station ffnite element model. The instrument section of 

the core body was modeled as a 1 umped mass and inertia charateristics 

HYSTERESIS = 8.5 DEGREES 

PLATFORM REORBIT BURN TIME HISTORY 

ATTI TUDE ERROR, DEGREES 

Flgur~ 5.2.5-2 - PLATFORM PHASE PLANE 

at the appropriate center of mass. The solar crrays were also modeled as 

common Space Station elements utilizing the structural characteristics 

from the Space Station fin~te element model. The arrqy inertia 

characteristics were lumped at the array center .,; ~nass 1 ocated on the 

structural support born. The antenna boom consists of support stn~cture 

similar to the array support boom with a corxentrated mss at the top to 

simulate the antenna. The Platform structural mdel does not include a 

berthed Orbiter condition. This effect is not insignificant and should be 

included in future studies. Vibration mdes were assumed to exhibit 0.5 

percent of critical damping. Th? berthing port and reboost thrusters are 

located at the end of the Core ; Ay close to the solar arrays. This end 

of the model is the application point for the berthing and rebosst forcirrg 

f unc ti ons . 

The finite element model was used to characterize the structural dynamics 

of the Platform in terns of mode shapes and corresponding vibration 

frequencies (Figs. 5.2.5-4 through 5.2,5-10). Results of this analysi s 

indicate a low frequency torsion mode at 0.28 Hz for the Core body. The 

bendic~g of the Core bdy occurs at approximately 0.53 Hz. Significance of 

these modes comes from the di splacemnt of the so1 ar arrys. the antenna 

boom, and the instrunrent Platfom which participate in these Icu frequency . 

modes. Solar array torsion is predicted at slightly above 0.33 Hz; :rrqy 

bending with the Core body occurs at 0.63 Hz, and as a cantilevered beam 

at approximately 1 Hz. The antenna boom also participates in the 0.63 Hz 

body mode and also exhibits bending at 0.75 Hz. The tip mass of the 

antenna appears to have s,grificant impact on the dynamics of the antenna 

support structure. The effect of payloads on the dynamics of the Platform 

will depend on the individu~i mass of the pqyload, but will serve to 

reduce the frequenciez of the overall system. 

Forces acting on the Platform while on-orbit include, but are not 

restricted to, Orbl ter berthing and orbit reboost. These two condi tions 

are the forcing functions utilized irr the analysis set. Thp QrSftcr 

berthing force is assumed to be identdcal to tk Force used in Space 

Station analyses and consists of a 1-secczu squhre wave with a mgni tude 

of 500 pounds (Fig. 5.2.5-11). The reboost forcing function was generated 

by a Platform model that includes on-orbit disturbances and a flight 

corrtrol system. Various control system parameters were investigated to 

arrive a: a coi;tr~l input which seemed most realistic. Varying control 

deadbands have a significarlt impact on frequency content of the forcing 

function. For systems of 'low frequency, caution must be extrrised in the 

selection of these paramewrs such that the coctrol input is rtut resonant 

with the Platform structural dynamics. 

Orbiter berthing trar~sient response analysis (Figs. 5.2.5-12 through 

5.2.5-23) indicate5 acceleration levels of 0.016 g's at the Core body in 

the direction of tne appl ied force. Dynamic amp1 i fication is responsible 

for acieleraticn levels of 0.032 g's at the solar array tip. lhe maxinum 

response of the Platform occurs during 'be berthiny force application and 

diminishes rapidly In the Core body after force termination. The 


SPACE STATION PLATFORM MODEL 

FREQ (3) - -33 HZ 

Flgure 5.2.5-6 - MODE B 



SPACE STATION PLATFORM MODEL 

FREQ (7) - 1.04 HZ 

Flgure 5.2.5-10 - MODE 13 









1 d,

if- 



flexible appendages vibrate for a l~nger tine in a damped oscilletory 

motion. The Core body of the Platfom behaves as a rigid body whereas the 

appendages account for the structural dynamics of the system. Berthing 

loads produce 450 ft-lb at the root of the antenna support boos concurrent 

with 22 lb of shear. This represents the highest load in the system fcr 

berthing and is well within the 3950 ft-1 b mment capability of the beam. 

The solar arrqy support beam carries 310 ft-lb at the root and is not 

considered si gni ficant from a strength viewpoint. 

The reboost forces, neceszary to keep the vehicle at the proper orbit, are 

a train of square pulses generated from off modulated RCS jets driven by 

the attitude control algori th (Figs. 5.2.5.-24 and 5.2.5-25). The 

temporal character of these pulses dictates the amount of dynamic 

involvertent of the flexible components on the Platfom- Figures 5.2.5-26 

thrwgh 5.2.5-37 show the Platfom response to the reboost forces. The 

antenna support boom has.a response character similar to the ivut forces 

with a maximum antenna acceleration of 0.07 g's and a root nolent value of 

700 ft-lb. An interesting dynamic feature of this configuration is the 

vibratory comunication between the solar arrqy boom and the antenna 

support structure. The coupling betweerr these systems allows for 

vibration energy in the solar arrqys to transfer to vibration in the 

antenna boom. This transfer of energy is responsible for the damped 

response of the solar arrays; the energy in the arrqys is reduced, but the 

energy is not dissipated, it is simply transferred to the antenna boon. 

The growth character of the flexible response is of concern because of its 

similarity to a resonant condition. Reboost forcing functions of slightly 

different character may lead to a large increase in loads due to resonance 

and must be analyzed carefully. The analysis does not indicate a specific 

viol ation of structural a1 lrmables; however, it does indicate sensitivity 

to the reboost force inputs. 

5.2.5.4 Rigi d Body Control Dynamics 

The objective of this ar~alysi s was to determfne the control requirements 

for the unmanned P1atl'~r-m elements as a function of the natural on-orbi t 

dynamic envl ronment. However, only two representative P 1 atform missions 

were analyzed in detaii and will be reported on herein. They are (1) the 

polar payload B, and (2) the co-orbiting payload Starlab. The dynamic envi ronnents siwl atcd included gravi ty gradient torques, aerodynamic 

drag, and aerodynaai c torques. 

5.2.5.4.1 Simulation capabilities 

The SSDYWICS program, developed at JSC, and described in Section 

4.3.3.5.1.1 was used tc simulate the on-orbit dynamics of tk Platfom 

elements. 

5.2.5.4.2 Mass properties -

Mass properties used in the SSDYW.IC simulation for the two Platform 

studies are shown In Table 5.2.5-2. Two sets of mass propertf es are 

Flgure 5.2.5125 - PLATFORM REISOQST TRANSIENT FORCE 


OFtleMAL PAQf ' iS 

OF POOR QUALITY 

ui 

(I) 

Z 

0 

P 

(I) 

W 

a 

iCz 

W, 

8 

Z 

4 

a 

b- i 

C 

8 2 ' 

0 

0 

UI 

a 

I 

Lxd~i~AL PAQEViS 

OF' POOR QUALITY 

- -. ' 

,-.- 

' I"' - 

- . -. . .- - 
a r. 

f 0 

L. 

z L. 

9 

8 

H 

i, 

"4 ?B 

%m 

x 

9;; 

8, c.r 

C-' 

,; 

,; 

;3 

? 

'I? 

? 


QHdmk; P,$Q'-' 

OF POOR QUALITY 
3RIGWAL PAM' BS 

3F POOR QUALITY 
shown for the Starlab Platform representing the initial design (Starlab-1) 

and the third configuration iteration (Star1 ab-3) which had added b a l l a s t 

weights and had an added ballast boob opposite the cold solar arrqy . The 

laass property inertias given in Table 5.2.5 are relative to the Platform's 

center of mass and use the coordinate system shown in Figures 5.2-3 and 

5.2-5 for the Starlab and Polar B, respecti wly. The products of inertia 

are computed as the negative integral. The oriain of the coordinate 

sys- is in the geonrtric center of the Platfotm Core. 

TABLE 5.2.5-2. - PLATFORM MASS PROPERTIES 

POLAR B STARIAB- 1 STMAB-3 

WIGHT (LBS) 32,164 24,872 24,394 

Inertia (sl ug-ft2) 

IXX 40,688 41,710 136,160 

IYY 225,876 113,318 147,475 

! ZZ 230,415 124,908 138,513 

IXY 0 -20,814 -1,674 

IXZ 479 2,582 642 

IYZ 0 -1 1 -1,534 

C.G. Location (ft) 

XCG -5.11 0.28 

Y CG 0 .O C. 19 

ZCG 0.61 -0.08 

5.2.5.4.3 Flight nodes - simulated 

The two Platforms analyzed had different fl ight modes. The Polar B 

Platform flies in a Sun synchronous orbit with a LVLH flight mode at an ! 

altitude of 380 nmi and an inclination of 98.2 degrees. The repetitive I 

descending node is at 2:00 p.m. The solar arrqys and the antenna were t. 

held inertial ly fixed over the orbit. 

The Starlab Platform flies in an inertial mode for celestial observation 

at an inul ination of 28.5 degrees. The solar arrays have two degrees of 

freedom to point at the Sun: (1) an alpha positioning capability about 

the y-axis and (2) a roll of the Core Platform about its x-axis. 

5.2.5.4.4 Control requirements for the disturbance environment 

Simulation of the Polar B revealed minimum control requirements over an 

orbit. The configuration was flown with its solar arrays initially 

para1 lel to the x-axis at the subsolar point in the orbit. The peak 

aerodynamic and gravity gradient toques at the 380 nmi altitude were less 

than .005 ft-lb- (Figs. 5.2.5-38 and 5.2-5-39, respectively). The 

aerodyamic density encounted over the orbit is shown in Figure 5.2.5-40. 

The angular mmntum storage requirements are shown in Figure 5.2.5-41. I 

H JRONM,INCI =98 5,TIiP-U 0.11 EGO-54 17.KR-31R,KA=11 4 

POLAR PLATFORM PAYLOAD B 

H=3BONM.INCL=98.5.Tk1P

POLAR PLAT

Figure 5.2.5-40 - AERODYNAMIC DENSITY AS A 

FUNCTION OF ORBIT POSlTlON 

ORB11 

Figure 5.2.5-41 - TOT

IMPULSE HISTORY W

. ORBIT COOR
The secular y-axf s (pitch) momentum storage of 22.0 ft-1 b-sec over one 

orbit can be rcmved by using magnetic torquers to bias out the average 

gravity gradient and aerodynamf c pi tch pl arte torques. 

Initial sfulations of the Starlab-1 platform at 270 nmf and at an 

inclf nation of 28.5 degrees revealed a large secular mmentula buildup over 

an orbit (see Ffg. 5.2.5-42) of 440 ft-lb-sec. Tiis occurs when the 

Platform is rotated about its z body axis by 30 degrees placing the x 

principal axis 45 degrees out of the orbf t plane. The gravity gradient z 

axis toque is biased as shown in Figure 5.2.5-43. The 440 ft-lb-sec of 

secular momentum buildup f s large only fn the sense that it f s desirable 

to control this axis with a Space Telescope reaction wheel which has a 

mmentun capabfl f ty of only 200 ft-1 b-sec. To remedy this condi tion, the 

Starl ab Platform underwent several confi guratf on changes involving the 

addition of ballast weight and a ballast boon. The mass pmpcrties for 

Starlab-3 given in Table 5.2.5 show the inertia properties to be closely 

balanced. Sinulatf on of the on-orbi t fl ight characterf stics of Starl ab-3 

showed f n fsct that the momentum requirerrents were reduced to an 

acceptable level of 63.0 ft-lb-sec per orbit (Fig. 5.2.5-44). 
~=?70pJM, INCL~~~.~,THX=-ZB.~,THY=O,THZ=-~,KR~~~,K~=~~.~ 

STARLAB PAYLOAD PLATFORM 

Flgure 5.2.5-42 - TOTAL INERTIAL TORQUE 

IMPULSE HISTORY WRT AN INERTIALLY FIXED 

ORBIT COORDINATE SYSTEM 

H=~~OW,IN

Figure 5.2.5-43 

HISTORIES 
H-27W. I NCL-28.5.THX--21.5,THY -0,THZ--M.S.KRI..KA-12.4 

STARUB PAYLOAD PLATFORM. CONFlQURAT!ON NO 3 

Figure 5.2.- - TOTAL INERTIAL TORQUE 

IMPULSE HISTORY WRT AN INERTIALLY FIXED 

ORBIT COORDINATE SYSTEM 
5.3 SUBSYSTEMS 

The fol lowing sections describe the subsystem functions, imp1 ementation 

schemes, rommoq Core Space Station hardware applicability, a:id their 

reference desr gn concepts. 

5.3.1 Structure -

The structure subsystem provides support during launch and operation for 

all Ccre and payload module components. 

5.3.1.1 Subsystem Functions 

a. Provide for ORU mounting. 

h. Provide for 1 aunch support of deployable appendages. 

c. Provide Shuttle launch support. 

d. Provide support for paylo.~ds. 

e. Provide support for solar arrays. 

f. Provide grapple for RMS deployment. 

g. Provide f~r OMV docking. 

h. Accept launch loads of the ORU and instrument modules. 

i. Accept launch loads of the stowed solar arrays and antenna. 

j. Transmit launch loads cf the payload and Core Platform to the 

Shattle cargc bay keel and longerons. 

k. Provide support for the propul siofi subsystem. 

1. Prov'de the interface with the Shuttle RMS, Space Station and 

OMV . 

5.3.1.2 Potential Hardware Implementation 

The Platform Core flat bed structure ant i~strument flat bed structure 

(If required) may be a machi t~ed longeron!shear plate construction which 

is capable of accepting distributed or point loading ?t the pqyload 

attachment and ORU modcle interfaces. While aluminum was assumed in the 

reference configuration, composites may be used if performance vs. cost 

is acceptable. 
5.3.1.3 Space Station Reference Hardware Applicability 

The Platform structure subsystm in a large part will be totally unique 

from Space Station. Potential ccaonality exists with modular ORU 

structure and interfaces. The deployable truss will be of the Space 

Station design. 

5.3.1.4 Platfom Reference Design 

The co-orbiting and polar Platforms are configured from cmn hardware 

elenrents that can be assembled in a modular fashion to arcaimdate 

specific mission requirements. An overall description of the family of 

IOC Platform configurations is contained in Paragraph 5.2.1. The basic 

structural elements corisist of a Core fiat bed structure, instrument flat 

bed structure (if required), Platfom ORU mdules and instrument 

nodulec. Please refer to Figure -5.3.1-1 for the following di srussion. 

The Core flat ',ed will be a rectanglar box structure approximately 12 ft 

long, 10 ft wide, and 3 ft deep. At one of the 3-by 10-ft ends a 

bulkhead will be attached. This bulkhead spans the Shuttle's 15-foot 

Biaeeter and carries the two sill trunnions that react the Shuttle's 

thrust loads. Tt also carries the keel pin, Cantilevered from the 

bulkhead, away from the flat bed structure is the module that carries the 

fuel tanks. Near the other end of the flat bed structure, bvo more sf1 1 

trunnions are provided, These carry only the - + z Shuttle loads. 

Two add! tional uses for the Core Platform have been identified. In one 

case, a Core fiat bed structure with bulkhead and attached propulsion 

tanks can be flown in tie Shuttle and used to bring fuel, new 

instruments, and ORU modules to orbit for servicing a Platfom. It can 

also be used to return instruwcnts and ORU mdules to Earth. 

In addition, a Core flat bed structure could be used to point a cluster 

of instnuents that dre attached to iiic acne9 Space Station. These 

instruments might be permanently assigned to the Space Station or they 

could be "tried out" on the Space Station and later transferred to a 

Platfom ds required. This is illustrated in Figure 5.3.1-2. A 

threz-axis gimbal would wnt to the top face of the flat bed and the 

insr.wpents would be attached to the bottom face. The top face could 

cprry any ORU's required such as for the operation of the gihals. 

Balancing weights may a1 so be attached to the top face as required so 

that the center of gravity of the assdl) .dill Le near the girbal axis. 

Note that each inst~rnent is provided with heat rejection capability as 

part of the gimballed structure. Power and data handling is provided by 

the Space Station through the gsabal . After achieving orbit, a four-bay truss will be deployed iron the 

bu?khead in the same direction as the fuel tanks. Each bay is 9 ft long 

for a total of 36 ft. This truss uses the same deployable elements as 

the Space Station truss structure. The truss carries a Space Station 

type solar arrv that has been shortened to about 29 ft of actdve 
TDRSS 

ANTENNA 

DEPLOYED 

k 

TDRSS ANTENNA 

ERECTABLE RADIATOR 

FIXED HEAT- 

, EXCHANGER/ 

RADIATOR 

-YI FI PIN 

FLAT BED 

STANDARD 

MOUNTING 

- MECHANICAL 

PLATFORM EQUIPMENT 

MODULES 

Figure 5.3.1-1 - EXPLODEL VIEW OF CORE PLATFOR

length. At its far end it also carries four thruster modules and an OMV 

docking port for maneuvering operations. 

All other Platform cwnponents dre mounted to the Core flat bed structure 

in modules. These modules mount across the 10-ft width dimension of the 

flat bed and they are 2 to 4 ft wide- The last 2 feet of the 12-ft-long 

Core structure is reserved for the attachment of one or more scientific 

instruments. The large TDRSS dish antenna is deployed 36 feet above the 

Core structure using the same deployable truss design that is used to 

extend the Space Station and Platform solar arrays. Heat pipes carry 

waste heat along the 10-ft dimension of the module to a fixed heat 

exchanger/radiator mounted at the side of the module that. always faces 

away from the Sun. Erectable radiators derived from the Space Station 

plug into the fixed heat exchanger/radiator to achieve the required heat 

rejection for a given mission. The fixed heat exchangerlradi atur 

includes longitudinal heat pipes to evenly share the thermal load among 

all the radiators. 

For those Platform missions requiring additional flat bed structure for 

scientific instrument mounting, one or more 12-ft-long flat bed 

structures may be attached to ttie Core structure. For these missions, 

the +z sill trunnions may be mounted to one of the extension flat bed 

structurzs rather than to the Core structure, depending on a loads 

analysis for that configuration. Similarly, the keel pin may be 

re1 ocated if a 1 odds analysis so indicates. 

Each scientific instrument will us? a modular mounting technique 1 ike 

that used for the Platform subsystem. it will span the 10-ft-dimension 

of the flat bed structure and contain heat pipes and radiators to carry 

away waste heat. These modules will be a minimum of 2 ft wide and a 

maximum of about 8 ft wide (in 1-foot increments). 

The method of mechanically mounting any module to the flat bed structure 

so as to minimize errors in instrument pointing will be studied. This 

will also include electricdl and, if required, thermal fluid 

connections. Since any module may require replacement in orbit, the 

chosen mounting technique wi 11 careful ly consider the ease of attaching 

and detaching a ,nodule. The Platforills may be serviced by the STS and/or 

a teleoperated OMV. The co-orbi ting Platforms may also be serviced on 

the Space Station itself. Any flat bed with instruments thdt is gimbaled 

from the Space Station will also be serviced by the Space Station. Since 

a standard method of servicing all Platforns is desired, all servicing 

modes will be considered in the module interfdce design. 

5.3.1.5 Weight Table 

Table 5.3.1 shows the structure subsystem weights. 
TABLE 5.3 1 - STRUCTURE SUBSYSTEM UEIGHT 

- Item Yeight (lb) 

Core Flat Bed 

Bulkhead 

Trunnion Supports 

Total 2715 

5.3.2 Mechanisms 

The mechanisms subsystem provides latching, deployllent and daaping for 

Platform elements. 

5.3.2.1 Subsystes Functions 

a. Provide for the deploylent of Platfom truss structure. 

b. Provide for latching, alignment, and unlatching of ORU and 

instrunent modules. 

c. Release, deploy, and retract the high gain antenna. 

d. Release, deploy, and retract the solar array boors, blanket 

housings, and solar arrqy blankets. 

5.3.2.2 Potential Hardware Implementation 

Simple deploy~nts could Be accomplished with high speed, high toque 

motors, stepper rmtors, or passive stored energy &via -. Standard 

devices such as pinpullers and release nuts llay also be utilized. 

5.3.2.3 Space Station Reference Hardware Applicability 

Solar array deployment mechanisms used on Space Station will be used with 

shorter blankets for the PI atfom. Pi npul lers and re1 ease nuts qualified 

for use on the Space Station could be used in Platfom applications. 

Coronalily llsy also exist between Space Station and Platform ORU latch 

design. The deployable truss will be the same design as the Space 

Station truss. 

5.3.2.4 Platfom Reference Design 

The reference design assumes use of Core Space Station ORU latch 

mechanisms, solar array stow and deploy mhanisms, and as marly 

mi scel 1 aneous devices as feasible. Aerospace standard devices wi 11 be 

wployed where Space Station devices are not feasible with the ninirvrm 

use of special design devices. 

5.3.2.5 Yeight Table 

Table 5.3.2 shows the mechanisms subsystem weights. 
TABLE 5.3.2 - MECHANISMS SUBSYSTEM WEIGHT 

- I tm -3.- Wei ht (Ib) 

ORU Latch Mechani sm 324 

Solar Array Stow and Deployment 196 

Mi scel laneous Depl oy Devices/Dampers 6 0 

Total ?ml 

5.3.3 Thermal Control 

5.3.3.1 Overview 

The Platform reference configuration Thermal Control System (TCS) is 

composed of two-phase ammonia cold plates and heat pipe radiators which 

service Platform ORU modules located on the Core flat bed and instrument 

modules located on additional flat beds as required. Mu1 ti layer 

insulation covers all flat bed surfaces with the exception of radiators 

and sensor openings. 

The propulsion system utilizes electrical heaters, mu1 ti1 ayer insulation 

and appropriate surface coatings common to the Core Station thermal 

design (ref. section 4.4.7.4.5). Likewise, the comnunications and 

tracking hardware not located on the Flat bed have an autonomous TCS 

design common to that of the Core Station (Ref. Section 4.4.7.4.6). 

Heaters will be controlled by an on-board computer using thermal sensor 

input data. 

The GRU and the flat bed-mounted instrument modules contain cold plates 

and connections to the fixed radiator heat exchanger. The heat exchanger 

contains a heat distribution system which distributes the heat to the 

erectable radiators. This method of control allows highly variable heat 

loads whi 1 e maintaining nearly constant or isothermal temperature 

control. The thermal coup1 ing between an instmaent module or an ORU 

modale and the fixed radiator heat exchanger may be through flexible heat 

pipes if precise a1 i gnment is required. Techniques required to make and 

break these thermal connections to a1 low servicing will be the subject oi 

a trade-off study. The erectable radiator elements are commn to the 

Core Station but of shorter length. Payloads not requiring standard 

instrument modules wi 11 provide an autonomous thermal control system. 

5.3.3.2 Subsystem Functions 

The TCS will maintain the Platform within a1 lowable temperature liitl ts 

for all flight modes from launch to the end of the mission, includ~ng all 

servicing modes. The TCS is required to control Platform and payload 

interface temperatures for orbital average heat rejection loads varying 

from 4.5 kW to approximately 9.9 kW. For the 9.9 kW Platforms, the total 

instrument power is 4.5 kW, housekeeping power is 2.9 kW, and the power 

system losses are approximately 2.5 kW. The instrument module interface 

temperatures should be held to within + 5OF of the set point 

temperatures. Environmental inputs which must be accommodated follow: 
a. Those imposed on stellar or solar pointing co-orbi ting 

Platforms at 270 to 540 nni with a Beta angle of 0 to 52" 

b. Those imposed on Earth-oriented polar Platforms where the 

angle between the orbi t plane and the solar vector will be from 40 to 

25". 

5.3.3.3 Subsystem Capabilities and Sizing 

Both the Core Platfom and instrument flat beds have a cmmn thermal 

control design with a capability of rejecting 5.4 kU (Core Platfom heat 

rejection requirement) of waste heat. Instmarents have the option of 

utilizing the flat bed thermal control capability cr providing their own 

autonomous system. Also, the instrument flat bed radiator area can be 

reduced for heat rejection level requirements bel-~w 5.4 kW by removing 

the appropriate number of radiator panels from the fixed radiator heat 

exchanger. 

The radiators are mounted off the -y face of the flat bed, Figure 5.2-2, 

in the xz plane. The total radiator (including the fix d radiator heat 

exchanger radiator) is 12 ft wide by 18 ft long (216 ft 5 1. The top 9 

feet of the erectable radiator radiates from both sides. The above 

radiator was sized for a polar Platform and the following paramters: 

Coating properties - = 0.8 

ac= 0.2 

Fin efficiency = -84 

View factor to spac? .95 

Radiator temperature 5 70°F 

Under these conditions, the polar Platfonn radiators can rejec 

approximately 20 watts/ft2 1 in the -y direction and 10 watts/ft in 

the +y direction. The co-orhi ting platform can reject approximately 18 

watts/ftZ from the -y face and 15 watts/ftZ fram the +y face. 

Depending on the variations in 0RU and instmmnt module power profSles 

and the flat bed thermal balance, heaters will be used as required to 

minimize structural gradients, to provide make-up heat, for propel 1 ant 

tanks, thruster valves and 1 ines, and for non-flat bed camunfcations and 

tracking hardware which are cmn to the Core Station (Ref. Sections 

4.4.7.4.5 and 4.4.7.4.6). 

For the full 9.9 kK heat rejection load, the TCS weight requiremnt is 

estimated to be 1000 pounds for the Platform Core and payload; 600 pounds 

is the estimated TCS weight for the Platfom Core alone. 

5.3.4 Electrical Bower 

The electrical power subsystem of the Platforms provides the facilities 

for the generation, storage, control, and distribution of electrical 

energy. 
5.3.4.1 Subsystem Requirements 

a. Energy storage: The energy storage system will ,)e sized far 

the near-polar orbit described in Section 5.1.2 which has a maximum solar 

occultation period of 36 minutes and a minimum insol ation period of 63 

minutes. 

b. Distribution: Power will be managed and distributed in a 

manner which permits the sensing, i sol ation, and correction of 

non-catastrophic faults. 

c . Redundancy: The power subsystem wi 11 contain sufficient 

redundancy (block, functional , or other) to satisfy the system 1 i fetime 

requiremnt and the fault tolerance requirements of Section 5.1 .I. 

d. The near-term Platforms will provide a total of about 5.0 kW 

power i nput for customers. 

e. Provision for growth: The power subsystem will be capable of 

growth to support a continuous pwload power level of 20 kW. 

5.3.4.2 Potential Hardware Implementation 

Power genera ti on 

Roll -out or fol d-out sol ar panels are possible impiementations for power 

generation. There may be occasions when it will be necessary to retract 

the panels. Total array area will be divided into two panels and located 

so as to not block instrument and radiator fields of view while not being 

shadowed by the Platform. 

5.3.4.2.2 Energy storaqe 

Rechargeable nickel -cadmi um batteries, nickel -hydrogen batteries, or 

regenerative fuel cell s are potential choices for Platform energy 

storage. The NiCd represents the lo9est risk and longest lifetime 

alternative at this time. 

5.3.4.2.3 Power condition in^ and distribution 

Power condi tioning, ac cr dc power distribution, and the distribution 

voltage to the users must be determined as a function of Space Station 

comnonal i ty, control , efficiency, environmental effects, and fault 

detection and correction. Fine regulation may be left to the users. 

5.3.4.3 Space Station Reference Hardware Appl icabil i ty 

There are strong cost incentives for hardware and software commonality 

between the Platforms and the Space Station. Solar array technology, 

power conversion equipment, and perhaps power control and distribution 

equipment can use Space Station counterparts directly or by adaptation. 
5.3.4.3.1 Conversion, control, and distrfbution equipment 

The Platfonn power subsystem must handle high power levels, including 

growth capability to 20 kU of payload power. The Platform power 

conditioning module design should rep1 icate those of the Space Station if 

that design is small enough to accolllnodate Platfonn redundancy. 

5.3.4.3.2 Solar array 

The Platform power subsystem will have the same operating voltage as the 

Space Station to make the design and the hardware developed for the Space 

Station solar arrays appl icable. The only substantial difference will be 

in reducing the number of para1 lel strings to the nurber required for the 

shorter array used by the Platfonn. 

Energy storage 

The Space Station Regenerative Fuel Cell s (RFC) will be excessively 1 arge 

for Platfonn use if sized for 20 kU as in the Core Space Station 

reference design. Use of RFC's for Platfo~ms is therefore dependent on 

results of Phase 0 studies. 

5.3.4.4 Platform Reference Design 

The Platform reference design has a total user and housekeeping leveled 

continuous power requirement at IOC of 7.4 kV. Some instrument loads 

draw peak powers which are much larger than their average powers. For 

this reason the reference design uses the battery-share mode (load 

I eve1 ing) to hand1 e the peak power requirement, thereby decreasing the 

size of the solar array needed. 

Significant features of the major components of the reference design are 

discussed below. 

5.3.4.4.1 Power generation 

The reference design uses a conventional technology planar, silicon-cell 

solar array for power generation. A power densr ty of 20 U/lb is 

assumed. The area of the reference design array is 1720 ft2 (at 11.15 

~/ft2 BQL). Array switching is employed for coarse voltage control and 

array shorting where the full array power is not needed. Capability to 

increase array size for later missions will be provided. 

5.3.4.4.2 Energy storage 

The reference design uses conventional ntckel-cahium batteries with an 

assumed efficiency of 76 percent . If used at an average depth of 

discharge of 20 percent, a total capacity of 800 ampere-hours will be 

needed in the Core Platform. Additional battery capacity will be 

provided for especially demanding payloads (Polar C and SAR) within an 

ORU module mounted either on the Core Platfonn or on the instrument flat 
bed. The Core Flatform design includes capabil ity to connect these 

addi tionai batteries into the subsystem. 

5.3.4.4.3 Conversion and distribution equipment 

The Platform reference design incorporates distribution of power at the 

same voltage and frequency as the Space Station (high voltage, possible 

ac distribution). The overall efficiency of power conditioning is 

estimated to be 95 percent. The reference design has three 5 kW Space 

Station conversion modules, one being a spare. Djstribution equipment is 

assumed to be designed to meet Platform requirements and will use Space 

Station hardware as appropriate. 

5.3.4.5 Weight Table 

Table 5.3.4 shol~s estimated weights for items comprising the electrical 

power subsystem. 

TABLE 5.3.4 - WEIGHT TABLE 

PLATFORM ELECTRICAL POWER SUBSYSTEM 

-- ITEM 

Solar Array 964 

Batteries 

800 Ah 

Conversion Equi pment 132 

Control & Distribution 

Equipment 

The distribution equipment was assumed to use 100 W. Other power 

equipment efficiencies are as stated above. 

5.3.5 Cmunications and Tracking 

The communications and tracking subsystem provides communications between 

the Platfoiq attd TORS, GPS, and ground as well as the Space Station, STS, 

and OMV as required. 
5.3.5.1 Subsystem Functions 

a. Transfer to the ground: 

(1) Science data 

(2) Engineering data 

(3) Eiaergency data for failure evaluation 

b. Receive comnands frm the ground. 

c. Support measurement of relative range and range rate with 

respect to the TDRS/TDAS. 

d. Support orbit determination with GPS. 

e. Support rendezvous and docking operations with the STS and 

the OMV . 

f. Support transfer of data betwee3 co-orbiting Platforms and 

the Space Station. 

5.3.5-2 Potential Hardware Implementation 

The comunications and tracking system will be essentially the same for 

all Platforms, except that d co-orbiting Platform will have a direct 1 ink 

to the Space Station. The Platforms will have the capability to 

conmunicate through the TURS/TDAS systems an both S-Band and Ku-Band 

single access channels, have a GPS 1 ink as a navigational aid, be capable 

of receiving commands from and transrni tting telemetry and TV to the Space 

Station and receive comnands from and transmit telemetry and TV to an 

Orbiter, as well as interfacing with the Orbiter's rendezvous radar 

system. An S-band Mu1 ti-Access (MA) low-rate data TDRS/T9AS link for 

housekeeping purposes wi; 1 a1 so be copsidered. 

The comaunjcations and tracking subsystem consists of turn-around 

transponders, rnodul ators/demodulators , power mp1 i fiers, and appropriate 

switching at S-Band and Ku-Band for comnand, telemetry, and tracking 

through the TDRSS. In order to accomplish this throughout the orbit, a 

steerable, high-gain antenna will be required. This will be a 9-ft 

antenna for comnonal i ty with the Space Station antennas, a1 though a 

smaller antenna would suffice. For appropriate viewing of the TDRS, the 

high gain antenna will be on a boom. To reduce the requirement for 

generation of high RF power, the transponders must be at the end of this 

boom near the antenna. The antenna will have combined S-Band and Ku-Band 

feeds. Pointing will be accmplished mechanically during acquisition 

and electronically during tracking acquisition and electronically during 

tracking. lo handle the various vehicle orientations, the feed must be 

circularly polarized. A BER of less than 10-5 will be achieved without 

coding. 
Low gain S-Band antennas will also be provided for comnunications through 

the TDRS at rougtlly 1-3 kbps. Switching and sequencing is provided 

through external controls. Low gain antennas may be required for direct 

ground communications. 

Comnunications are received from the GPS through a low gain antenna. The 

antenna should be located to have the most advantageous position to view 

the GPS. 

Comnunications with the STS during rendezvm:~ and docking operations will 

be provided through a NASA Standard Transponder IS-Band) and low gain 

antennas. The antennas can be the same low gain otles used for the low 

data rate TDRS 1 ink but consideration should be given to separate ones 

located in order to support the docking opera ti^,^. This 1 ink ill 

transfer comnands and telemetry between the co-orbi ting and polar 

P;atforms and the 5TS and the OMV. In addition, TV is transmitted from 

the Platforms to the STS. 

Comnunicati ons between the co-orbi ting Platforms and the Spare Statian 

will be through a separate transponder and antenna. The antenna will be 

located to view the Space Station during a1 1 orbits. A steerable antenna 

may be requ~red even though the antenna size will be small. The link 

will support both 5 Mbps and 64 kbps channels from the co-orbiting 

Platform to the Space Station and will zlso support 48/64 kbps channels 

from the Space Station to the co-orbiting Platform. Although equipment 

for this function is included in the Pldtform reference design, the 

incorporation of this capability in the payload module will be studied in 

Phase B. The use of slow scan video on the platform as a rendezvous aid 

will a1 so be evaluated. 

5.3.5.3 Space Station Reference Hardware Appl icabil i ty 

The Space Station antennas and transmi tters/recei vers are used in the 

Platform reference design. Als3 assumed were the Space Station GPS 

system and comrnli,,ications processors which are included in the 

informdtion anu data management subsystem. 

Feasibjli ty of the use of a smaller TDRS antenna and reliability 

tradeoffs between one and two antennas should be itudied in Phase B. 

5.3.5.4 Weight and Power Table 

Table 5.3.5 shows comnunications and tracking weights and power 

requirements. 
TABLE 5.3.5 - WEIGHT AND PO#R TABLES 

PLATFOIM CWW ICATIOWS WD TRACK1 MG SUBSYSTEM 

Y IW - ->TX 1 OTK 

E QUIPMNT WANT. -- EACH(LB) -- i PURIM) 

ANTENNAS: 

Gain Ku/S-Band 1 100 100 200 

TDRS Low Gain S-Band 2 2 4 - 

Med. Gain Ant. 2 10 20 50 

GPS Law Gain Antenna 2 2 4 - 

Radar Xpndr Lw Gain 2 2 4 - 

FF SYSTEMS: 

S/Ku-Band Xmi t/Rec . 2 80 160 400 

*SS Link LDR Xlai t/Rec. 2 40 80 75 

S-Band Xpndr (ORBITER) 2 4 5 90 80 

TRACKiNG: 

GPS Rec . /Proc. 

Radar ~pndr 

-- CABLE: 350 

TOT& * To be flown only on co-orbiters. 

5.3.6 InfornaatGon and Data Managemnt Subsyste~ ( IMS) 

The IUHS provides camand and data processing and storage and serves as 

the Pl atfom executive. 

5.3.6.1 Subsystem Capabilities 

The Platform ID6 will euploy a distributed architecture +hat allocates 

processing services resident in each subsystea. 

a. The combined onboard information rate is from 3OG kbps to 

300 Mbps and will be made available for transmission through TDRSS at 

rates frou 20 to 300 Ms. 

b. The uplink corand rates will be less than 300 kbps over 

S-Band with a similar rate aver Ku-Bawd. 
c. Instrument data storage requirements are from 10 to 600 

uigabits. 

d. Holrsekeeping datd may be transmitted in real time to the 

ground via an S-Band link through TURSS and/or recorded en a special 

houseke~ping recorder for 1 ater transmi ssi on. 

5.3.6.2 Subsystem Functions 

l'he data management system nil? supply time and frequeqcy standard 

signals and housekeeping/engineering subsystems data to other spacecraft 

systems as well ds a1 1 control and synchronizfng signals necessary fcr 

commanding, data collection, and system monitoring. 

The comnand procesc will include reception, val idation, expansion, 

interpretdtion, stb~ dge, di stri bution, and execution of commands in 

packet form. The comnand types will be immediate exbute, delayed 

execute, and stor, d. Commnds may also be generated internal to the 

2latform in response to on-board algorithms or i? r.he form of comnand 

suoport requests from the subsystems or pqyload. 

Dat'. col 1 ecti on dnd processing of science and engineering data packets by 

the IDMS wili include adaptive sampling i~cluding memory dum,s of the 

pay1 oad instruments and engi ncering s~b:,~stems, merging of data streams, 

formatting of the data (and encoding where reqbiyed) for storage ar~d/or 

immediate transmissiorl. The IGMS will provide s~bsystem data bases for 

inter- and intra-scbsystems access and storage. The IDMS will monitor 

the Pla.-forms health and status in support of autonomous operation. 

Instruments that generate very high datz. rates or very large masses of 

dda may be required to provide their own separate bit stredm. The IDMS 

would not alter the information content of any bit stream, bu' would 

route it to the appropriate translllitter or mass datd storage device. 

Each Plat.form typically has one (maximum of two) instrument cf this 

cl ass. 

5.3.6.3 Potential Hardware Implementation 

A di stri b~ted data col lection and command system i~plementation i s 

envisioned with J redundant seridl fiber-optic bus providing the path for 

all informatioq (commands and data) exchanges (see Fig. E.3.6), Thscience pay 1 odd instruments and engineering subsystems are assumed LU be 

proc\?ssor based with sufficient memory to process and compose data 

packets. They ill respond to, decode, and store (if necessary) higher 

order commands. Pay1 oad instruments and engineering subsystems wi 11 

interfice with the fiber-optic bus network via interface devices (ID'S). 

The ensineeri ng sl~bsystems and pay1 odd instruments may use general 

purpose subsystem data processors (SDP's) which, under software control, 

wi 11 provide subsystem pecul iar processing, computation, and storage. 
I 'MECH I 1: I - 

SUBSYST'S 0 'POWER SOP 

DATA STORAG

SATELLITES 

PAYLOAD 

1 

AND 

TRACK 

'GN&C T -* - 

*rHERM I - 

ORBITAL 

FREQUEN

I PORT

WRKS

(OMV, EM

Figure 5.3-6 - INFORMATION AND DATA MANAGEMENT SUBSYSTEM PLA
5.3.6.4 Space Station Reference Hardware Applicabil i ty 

The Space Station IDFIS hardware is appropriate for us: on the Platforms. 

This assumes that those fu~ictions related to workstation control ( i .e., 

human control ) can be deactivated. The Space Station mass storage 

devices, which are as yet not well defined, may not be appl icable due to 

the higher data rates and higher storage requirements expected on the 

Platforms. 

5.3.6.5 Platform Reference resign 

The reference IDMS configuration i s a di strihuted archi tec ture using 

functional elements comnon with \.he Space Station. The core data 

distribution petwork will be fiber-optic. The Platfgnn reference tape 

recorders dre assumed to be larger than the Space Station reference mass 

storage devices. Two large (new development) recorders are included in 

the Platform reference design. 

5.3.0.6 Weight and Power Tab1 e 

The IUMS, for weight and power purposes, consists of a core system which 

will be the same for each Platform. Bookkept separately is the weight and 

power for two mass storage devices whicll will be chosen for each Platform 

according to data requirements. The Platform reference design includes 

Go large di3i tal tape recorders (DTR). This information is contained in 

Table 5.3.6. 
TABLE 5.3.6 - ID@ WEIGHT ANb POWER 

EIGHT TOTAL POWER POMR 

ELEeNTS - MI. - EACH(LB) UEIGHT(LB) EACH(U)TOTAL(Y) 

Opt Net/Cabl 'n 1 50 50 - - 

Data Storage 3 70 21 0 50 150 

Time Freq 3 2 0 60 50 150 

Facil MgntICtl 3 20 60 50 150 

Payload ID's 

DATA ACQ _ /TM/CCI)S: 

3 Processor/Cont. 1 40 40 7 5 7 5 

Low Rate Mux. 3 10 30 25 7 5 

High Rate Mux. 1 70 70 120 120 

SIGNAL PROCESSING: 

TDdS=a-e s so r 2 20 40 25 50 

SS Link LDR Signal 2 10 20 2 5 50 

Processor 

SS Link HDR Signal 1 10 10 20 20 

Processor 

Digital TV Processor 2 6 - 12 50 - 100 

TOTAL 841 1,790 

5.3.7 Gui dance. Navi ~ati on. and Control (GNK 1 

The GNM: subsystem provides PI atfon base-body control , sol ar array 

pointing, antenna pointing corwnds, GPS data processing, and supports 

instrunrent pointing functions. 

5.3.7.1 Subsystem Characteristics 

The GNM: subsystem, in conjunction with ttw propulsion subsystem, 

conwnications and tracking subsystems, and the information and data 

management subsystem, provides the basic functions of guidance, 

navigation, attitude control , orbit maintenance, and passive proxinri ty 

operations. Noml operation of the GN&C will not require interaction or 

assistance from ground station systems; however, the GNM: will be capable 

of accepting ground backup comnands. 
a. The GN&C subsystem will provide the following Platform 

performance: 

Accuracy (3 signal without 

fine error sensor (FES) 

Accuracy with FES 

Knowl edge of pointing (without FES) 

Knowl edge of poi n ti ng (wi th f ES) 

.03 deg 

FES dependent 

0.01 degree 

FES depevient 

Stability 0.002 deg!10 mi n 

Jitter 0.0003 deg/l sec 

b. The GNBC subsystem will have the capabil i ty to store and dump 

the secu! ar momentum expected to result from atmospheric drag, solar 

pressure, gravity gradient, and Platform rotation, and must do so for all 

initially planned configurations. 

c. The propulsion subsystem will provide a fail-safe backup 

attitude control capability for a contingency period of 12 months. 

d. Platfonn position will be derived from the Global Positioning 

System (GPS) to an accuracy of 330 ft using the coarse code. Platcorm 

position will be control led in polar orbit to maintain a ground track 

repeat every 16 days (233 orbits) with ground traces within - + 5.4 nmi 

(cross track). 

e. Solar array pointing control accuracy will be maintained to a 

suitable level. Array drive rotation rates will consider acquisition 

time, Earth orbit rate, and instrument slew requirements. 

f. Programmed yaw capabil i ty will be provided for a1 ignment of 

instrument ground tracks. 

5.3.7.2 Subsystem Functions 

The GN&C Subsystem will provide the following functions: 

a. Three-axis attitude determination, a~ 'sition, and control 

for nadir orientations end celestial or Sun-ref: .?d inertial 

orientations. 

b. Solar array pointing a~d control, arrd high gain antenna 

pointing and control commands. 

c. Platform position determination and control cmands. 

d. Delra-V control commands. 

e. Relate Platform atti tilde and attitude rate to that of ~aYl0ads 

via an attitude transfer system. 
f. Momentum management. 

g. Fault detection, isolation, and reconfiguration within the 

GNbC system. 

h. State vector prediction, propagation, and transfer to 

pqyl oads. 

i. Control and safing for proximity operations associated with 

berthing, depl oyment, and recovery. 

3. Establish and maintain a survival rode which has adequate 

power col 1 ection and control capabil i ty to ensure spacecraft survi val 

until the next servici ng opportunity. 

k. On-board a1 igmnt and calibration of sensors. 

1. Support for the functioning of instrument pointing systems if 

present and the ability to accept data from fnstnrmnt resident sensors 

to be used to supplellent basebody attf tude control sensors. 

m. Control of reaction control thrusters contained in the 

propel sion subsystem. 

n. Coordinate transformation capabil ity . o. Pqyload articulation devices, such as momentum compensated 

gidals, as needed to meet payload pointing requireaents. 

5.3.7.3 Potential Hardware Implementation 

The basic Platfom hardware complement is assumed to consist of the 

foll owing components: 

a. Actuators: 

(1) Momentum exchange devices (FED) such as reaction wheels 

or smll CMG's. 

(2) Manenturn desaturation control actuators such as magnetic 

torque rods, or devices for acti ve gravi ty gradient control. 

(3) Solar panel drive actuators and control electronics. 

b. Sensors: 

(1) Absolute attitude reference devices such as star 

trackers. 

(2) Relative attitude sensors to function in an attitude 

transfer system. 

(3) Acqui si ti on Sun sensors. 
(4) Target reference devices such as fine guidance sensors 

or high accllracy star trackers. 

(5) Inertial reference devices such as strapdown gyro 

assembl ies, rate integrating gyros, and accelerometers. 

(6) Magnetic field sensors such as flu~gate magnetometers. 

c. Processors: GNLC processors sui tab1 e for navigation and 

control functions ana supporting interface devices. 

The foll owing services are assumed to be provided by other Platform 

subsystems : 

(1) GPS data from the comnunications and tracking subsystem. 

!2) High gain antenna drive asselnbly as part of the 

comnunicati ons and tracking subsystem. 

(3) Reaction control , translation, drag makeup, and control 

torque during maneuvers as part of the propulsion subsystem. 

(4) Fine guidance sensing by the instruments as qeeded. 

( 5) Rendezvous and docking support from the communications 

and tracking su tsysten, 

5.3.7.4 Space Station Reference Hardware Appl icabil i ty 

The foll owing Space Station hardware components should be considered for 

use in the Platforms: 

a. Magnetic torquers. 

b. Star tracker triad. 

c. Inertial reference units. 

d. GNBC processors and associated ID'S. 

e. Attitude transfer system. 

Other Space Station components shou; d be examined for appl icabil i ty. 

5.3.7.5 Platform Reference Design 

The Table in Sstion 5.3.7.6 gives the system components which are 

anticipaied for lrse in the Platforms. The number of units, their 

heritage, and the relation to the Space Station are also shown. This 

section provides reference design rationale regarding the selection of 

components 1 i stcd. 
a. Reaction wheels - four are chosen for redundancy. Space 

Telescope heritage is used for reference. 

b. Magnetic torque rods - four are chosen far redundancy. Space 

Telescope heritage is assumed for consistency with the Space Station 

reference configuration. 

c. Star tracker triad - twn are assumed for redundancy. Space 

Station reference configuration corponent used. 

d. Two-axis Sun sensors - two are assumed. Heritage not 

identified but NASA standard should be considered. 

e. Wagnetaeters - two are assumed for use in the momentum 

management system. Heritage not assumed. 

f. Inertial Sensor Assembly (ISA) - one strapdom hexad unit 

assumed to have adequate redundancy. Space Station cmnal i ty assumed. 

Ring laser gyros and filament gyros aqy be used. 

g. GNLC processor and associated ID - tno assumed for 

redundancy. Space Station conrnonal i ty anticipated. 

h. Solar panel drive and electronics - representai!ve nuwbers 

shown, new design. Hust handle inertial loads and power transfer 

(including potential for growth to 23 kU total continuous power). 

i. Attitude transfer system. Arr attitude transfer or measuresent 

system will be provided on the co-orbiting and polar Platforms. This 

system will be capable of measuring the instantaneous alignment between 

the GNM: navigation base and the payload aroufitfng surfaces. Relative 

alignments shall be detenained within 5 arc 3%. Utilization of the 

attitude transfer system outputs in the guidance and control algoriths 

will be a control option, enabl ing overall Platfom control to be driven 

by a single instrument. 

5.3.7.5 Weight and Power Tdble 

Table 5.3.7 shows GNM: weights and power requirements. 

The propulsion subsystem provides translation for orbit transfer and 

orbit adjust as well as backup control torque capability. 

5.3.8.1 Characteristics 

The propulsion subsystem wi 1 1 provide the fol 1 owing perforuance 

capabil I ty for Platforms: 

a. Transfer from the baseline orbit of the Space Station or the 

STS Orbiter to tire Platfom ~npttrstfoilal orbit. 
TABLE 5.3.7 - REFERElJCE PLATFORM GN&C WEIGHT 

AND POWER ESTIMATES 

- QN WEIGHT (LB) 

Reaction Yheels 4 42 1 

(202 ft-lb-sec, Space Telescope) 

Ma netic Torq e Rods I 4 434 

?39,274A-f t , Space Tel escope ) 

Star Tracker Triad 2 40 

Sun Sensor (two axf s) 2 13 

Magnetometers 2 12 

Hexad Strapdown Gryo 1 50 

GN&C Drocessors 2 40 

Attitude Transfer System 1 2 U 

Processor BI 'I 2 10 

Solar Panel Drive 8 Electronics 2 - 375 

GN&C TOTALS: 1,415 

POWER (W) 

22 O/ 1600 

(260 avg) 
b. Maintenance of the Platform operational orbit for a period 

of nbt less than 24 consecutive months plus 12 nonths reserve in aqy set 

of calendar years. 

c. Return from the operational orbit for rendezvous in the 

proximity of the STS Orbiter or the Space Station. 

d. Provide toque capability about three orthogonal body axes 

for backup reaction control for the GNK subsystem. 

5.3.8.2 Functions 

The propulsion subsystem provides the fol lowing functions: 

a. Translational forces and three axis body torques from 

nu9 tiple, appropriately placed thrusters. 

b. Propellant transfer to the thrusters through a system of 

tanks, fluid lines, and valves. 

c. Capability for resupply fluid transfer from the Space 

Station, the STS Orbitpr or, if required, the OW. 

5.3.8.3 Potential Hardwdre I~plersentation 

Platform propulsion requirements can be mt with state-of-the-art 

technology systems. Earth storable monopropel 1 ant or bipropell ant hot 

gas systems are applicable to the basic requirements. In addition, a 

cold gas system may be required to provide a low temperature, 

non-contaminating exhaust gas environment for Space Station or STS 

Orbiter rendezvous. 

Propellant quantity requirenerrts are sufficiently high that a multitank 

supply system will be required. Multiple propellant supply tanks ~ffer 

the possibility of use of an existing design or tool ing, and adaptibil ily 

to Platform configuration volume limi tations. In addition, there is the 

possibility of reduced propellant slosh influence on the Platfom and tk 

possibility of some control of the propellant center of mass. 

5.3.8.4 Subsystea Performance 

The propulsion system selected ;n the reference configuration uses the 

hdrazine components used in the manned Space Station. It uses eight, 

40-inch-diameter spherical tanks and will have a capacity of 6300 pounds 

of hydrazine. These tanlrs contain an elastomric diaphra~l to separate 

the nitrogen pressurant gas from the hydrazine. The blowdown ratio is 3 

4x1 1. Tanks will be insulated and heated as required to prevent 

freezing. Propulsion system requirements are based on providing for the 

necessary AV for the worst case polar platform (Polar 8) but can be 

offloaded for less energetic missions. The subsystem must be capable of 

raising aqy polar Platfom which can be launched from UTR by the STS fn 

the STS parking orbit altitude of 148 nmi to the Platform's operating 
altitude of 380 nmi. It then will do stationkeeping for a period of up 

to 3 years and return the Platform to Shuttle altitude for refueling and 

servicing. With the projected STS capabil i ty of approximately 30,300 

lbs, the maximum dry weight for the polar Platform and payload may be as 

1 arge as 24,000 i bs. 

The thrusters are located at least 50 feet from the pqyload to minimize 

payload contamination and are oriented along the velocity vector when the 

Platform is in the nadi r-orielted mode to minimize reorientation 

operations for drag make-up and orbit transfer burns. The Platform 

center of gravity, propellant center of gravity , and thrust vector. are 

aligned to minimize the effect of center of gravity shift due to 

propel 1 ant usage. Addi t i o n a l l y , the propulsion system is modular for 

ease of servicing and design flexibility. 

5.3.8.5 Thruster Clusters 

Four triple redundant thruster clusters developed for the Space Station 

are used. The 25- to 75-pound size of the longitudinally oriented 

thrusters are appropriate for Platform propulsion. Smal ler thrusters for 

attitude control may be required for the Platforms. (The Space Statir~n 

reference configuration uses the same size thruster in all 

orientations). The thruster clusters will be insulated and he3ced to 

prevent hydrazine freezing. 

5.3.8.6 Hydrazine Lines 

Redundant hydrazine 1 ines will connect the thruster clusters to the 

propellant supply tanks. This line will either deploy with the 

deployable truss or will be installed by an astronaut. In either case, 

the lines will be dry at launch and will not contain fuel until the 

PTatform is at a safe distance from the Shuttle or Space Station. The 

1 ines will be insulated and heated as necessary to prevent freezing. 

5.3.8.7 Redundancy 

All systems are redundant to the point that no failure in a thruster or 

pair of thrusters will result in the failure of a function. 

5.3.8.8 Cold Gas System 

The cold gas system size provides for 4.9 ft/sec translational velocity 

capability or 25,515 ft-1 b-sec control momentum capabil i ty. The system 

is assumed to be required only for attitude hold in the final berthing 

operations. The cold gas system can provide a minimum of 4 dqys attitude 

hold mode capability. 

5.3.8.9 deight and Power Tables 

Table 5.3.8-1 shows overal l propulsl on subsystem weights. The weight 

estimate includes fluids, tankage, fluid distribution system, thrusters, 

thermal condi tioning, instrumentation, and integrating structure. 
TABLE 5.3.8-1 - PLATFOW PROPULSION REFEREKE DESIGN WEIGHT 

EIGHT - LB 

N2H4 Systea 

Fully Loaded System 7,500 

Naxium Expendable Propel lant 6,300 

Inert Weight 1,200 

Gu systetl 

fully Loaded System 300 

Maximum Expendable Propel lant 64 

Inert Yeight 236 

Table 5.3.8-2 shows the power estimate for propulsion maintenance and 

operations. 

TABLE 5.3.8-2 - PLATFORA PROPULSION HEFEREKE DESIGN POUER 

POER - Y 

N2"4 

Maintenance Orbital Average 175 

Firing Maximum 760 

GN2 System 

flaintenance OFbi tal Average Negl igi bl e 

Firing Maximn 300 
5.4 SAFETY 

5.4.1 Introduction 

The safety considerations involved in the desisn and operations of a 

low-incl ination Flztform and a polar Platform are approximately the same. 

In addition, the low-incl ination Platform hl so involves proximity 

operations with the Space Station, while the polar Platfom involves 

similar operations with the STS Orbiter. Overall safety considerations 

a1 so include interfaces and operations involved with EVA in conjunction 

with Platform( s) . 

5.4.2 Protection/Control of High-pressure or Hazardous Fl uf d Tanks 

The design of pressure vessels or tanks containing hazardous fluids will 

comply with the standard des iyn prdctice of 1 eak-before- rupture 

imp1 emented through a fracture co~itrol program. In addition, tanks/ 

pressure vessels wil i either t~e protected in such manner to prevent 

propagation of failure to other tanks or located akay from critical 

systems/subsystems or docCing/berthing areas for the Orbiter andlor Space 

Station. 

5.4.3 - Propulsion Systems 

Platfom propulsion systems will be designed to similar safety considerations as those for both the Orbiter and the Space Station. In 

addition, the proximity operations of Platforms in the vicinity crf the 

Space Station and duping EVA'S will be the subject of additional safety 

studicr dnd analyses. 

5.4.4 Proximity Operations ------ 

Safety considerations with regard to the operation of Platforms in the 

proximity of the Space Station will be the subject of additional safety 

studies. Such studies will develop safety design and operational 

requirements for the Platforms with regard to proximity operations. 

5.4.5 EVA O~grd~~ons 

EVA activities involved in the operations and servicing of the Platform( s) 

will be the subject of additional safety studies and analyses. 

5.4.6 Space Station BerthingCapabil i ties 

Sufficient berthing capabil i ty is provided by the Platform( s) and are 

adequately accommodate4 by the berthing port arrangement of the reference 

configuration. 
5.4.7 Othzr Safety Conslderatlons 

Other areas of safety concern wlll k the subject of add1 tlonal safety 

studies and analyses. Thcse areas lncltrde but are not resklcted to: 

(1) servlclng operatlons Involvlng the I--lncllnatlon P1atf"rr at or 

near the Spree Station and the pclar Platform In conjunction with the 

Orblter, and (2) assembly and repair operations associat+d tilth the 

Platfom(s). 
The Reference Configuratlon Group (RCG) of the "Skunk Works" SE&I team 

was comprised of technical speclal lsts from various NASA Centers. The 

fol low~ng 1 ist acknowledges the indl vl dual s with primary responsi bil i ty 

for various report sectiot,~ and contributing personnel. A1 Louviere 

served at head of the SE61 team wit11 Bass Redd serving as his deputy. 

Mark Craig led the Reference Configuration Group. 

REPCRT SECT ION - LEAD CONTRIBUTING PERSONNEL 

1.0 Introduction S. Nassf ff :JSC 1 

2.0 Background S. tiassiff (JSC) 

3.0 Sumry J. Butler (MSFC) D. Kaplan (JPL) 

and Desdgn Req~,'ts D. Greenshields (JSC) W. Schneider (JSC) 

B. Holder (JSC') 

G. Robitison (MSFC) 

J. Lewis (JSC) 

P. Nicaise (MSFC) 

J . Dunning ( LeRC) 

C. Sutt (JSC) 

T. Taylor (JSC) 

L. Livingston (JSC) 

D. Tomlin (MSFC) 

4.0 Space Station J. Butler (MSFC) Same as Section 3.0 

Maned Core D. Greenshieldc (JSC) 

4.1 5esqgn Reqm'ts D. Greenshiel ds (JSC) Sa:e as 5ection 3.0 

4.2 Configuration L. Livingston (JSC) J. Butler (MSFC) 

Description D. Greenshields (JSC) 

4.2.1 General 

Arrangement 

4.2.2 Flight 

Modes 

0. Meredi th (MSFC) 

S. Avans (MSFC) 

B. Ondraic (MSFC ) 

S. Croomes (MSFC) 

J . Stokes (MSFC) 

D. Deryder ( LaRC ) 

0. Garrett ( LaRC 1 

E . kempke ( LeRC ) 

J . D~lnni ng ( LeRC) 

M. Mikulas iLsRC) 

M. Dalton ;JSC) 

T. Sasseen (KSC) 

K. Yinson (JSCI 

L. Purver :', :tTS) 
REPORT SECTION 

4.2.3 Elements 

4.2.4 Mass 

Properties 

4.3 Configuration 

Eva1 uation 

4.3.1 Payload 

Accarnodation 

- LEUl 

D. Greenshields (JSC) 

D. Hows (JSC) 

B. Ondrak (USFC) 

D. Greenshields (3%) 

J. Butler (WC) 

D. Kaplan (JPL) 

4.3.2 Crew J. Lewis (JSC) 

kcorrodati ons 

4.3.3 Dynaairs 0. Holder (JX) 

8 Control 

4,3.4 Power 

Capabi 1 i ty Ih 

Util ization 

4.3-5 Them1 T. Taylor (JSC) 

Capabi! i ty 

CONTR ;BUT ING PERSONNEL 

Same as Section 4.2 

S~RE as Section 4.2 

Same as Section 3.3 

B. Ondrak (HSFC) 

J. Butler (MSFC) 

R. Duncan (MSFC) 

K. Morrison (JSC) 

C. Robinson [JSC) 

D. Suddeth IrSFC) 

J. Sos (GSFC 3 

J. Stokes (MSFC) 

G. Robinson (MSFC) 

J. Hall (LaRC) 

C. Robinson (JX) 

H. Daltoc (JSC) 

A. Benrend (JSC) 

M. Myers (JSC) 

J. Queller (Hqs) 

C. Moore (WC) 

D. Tomlin (MSFC) 

P. Nicaise (MSFC) 

J. Sch1iesi~- (JSC! 

0 7eryder C! 

D . Deryder ( L~HC ) 

L. Perez (JX) 

J. Famr (YSFC) 

K. Grady (GSFC) 

M. Lasaron (JX) 

J. Sunkel IJSC) 

J . Housncr (LaRC ) 

J. Graves (WC) 

J. Dunning (LeRC) 

E. Kqke :LeRC) 

R. Rice 'JSC) 

W. Simon (JSC) 

r. Gott (3x1 

D. Seymour (MSFC) 

G. Robfnsol IMSFC) 

J. Hall (LaKC.1 

W. Morris (JX) 

A. Fitzkee (WC) 
REPORT SECTION - LEAD 

4.3.6 Conm. S. Lenett (JX) 

8 Trackicg 

4.3.7 Launch, W. Schneider (JSC) 

Assembly, and 0. Meredith (MSFC) 

Growth 

4.3.8 Operations C. Gott (JSC) 

4.3.9 Safety H. Williams (JSC) 

4.4.1 Electrical L. Crabtree (MSFC) 

Power 

4.4.2 Guidance, P. Nicaise (MSFC) 

Nav! gati on, 

S Control 

COATRI BUTING PERSONNEL 

J. Clubb (MSFC) 

E. Bracalente (LaRC) 

M. Kapell (JSC) 

D. Rhoades (JSC) 

R. Knight (LeRC) 

0. Lercy (LeRC) 

0. Ramage (MSFC) 

A. Crurnbley (MSFC) 

S. Croomes (MSFC) 

C. Gott (JSC) 

R. Cain (JSC) 

M. Mikulas (LaRC) 

0. Ramage (MSFC) 

A. Crtrrnbley (MSFC) 

K. Morrison (JX) 

S. Croomes (MSFC) 

R. Becker (JSC) 

T. White (JSC) 

T. Saseen (KSC) 

C. Anderson (JSC 

R. Kain (JSC) 

D. Yonan (JSC! 

b. Hudson (JSC) 

A. D~punt (JSC 1 

0. t4ader (JSC) 

5, Mitchei: !JSC) 

D. aaillie (JSC) 

S. Truelock (JSC) 

D. Berlad (JSC) 

J . Graves (MSFC) 

R. Rice 'JSC) 

W. Simon JJSC) 

J. Duvning (LeRC) 

2. Kempke (LeRC) 

W. Chandler (JSC) 

D. Howes (JSCj 

D. Tom1 in (MSFC) 

J . Farmer (MSGC ) 

C. Wcgiley (LahC) 

T. Barry (JX: 

K. Grady (GSkC) 

C. Thomas (JSC) 
REPORT SECT lON - LEAD 

4.4.3 Car. 4 S. Lenett (JSC) 

Tracking 

4.4.4 Inf~mation M. Hollw (JSC) 

and Data 

tlanay-:nt 

4.4.5 Propulsion D. Yestrope (WC) 

4.4.6 Environ- A. Behrend (JSC) 

.ental Control and 

Life Support Sys. 

4.4.7 Thermal Id. Morris (JSC) 

4.43 Structures Id. Schneider (JSC) 

and hhanisms K. Hinson (JSC) 

4.4.9 EVA Systems C. Kissinger (JX) 

CONTRIBUTING PERSONNEL 

J. Clubb (MSFC) 

E. Bracalente (LaRC) 

M. Kapell (JSC) 

O. Rhoades (JSC) 

R. Knight (LeRC) 

8. Leroy (LeRC) 

A. Newconb (JSC) 

S. Anderson (JSC) 

J. Clubb (MSFC) 

1. Barry [JSC 1 

A. Dula (KSC) 

C. Herman (JX) 

C. Aukerrnun (LeRC) 

R. Donovan (LeRC) 

D. Kendrick (JSC) 

D. Anderson (KSC) 

C. Robinson (JSC) 

K. Upshaw (MSFC) 

D. Seymoor (MSFC) 

M. hers (3x1 

D. Deryder (LaRC) 

J. Hall (LaRC) 

G. Robinson (M'FC) 

C. Robinson (JSC) 

K. Upshaw (MSFC 

D. Sqmour (MSFC) 

R. Parcish (JX) 

A. F i tzkee ( GSC) 

T. Taylor (JSC) 

J. Hall (LaRC) 

G. Robinson (MSFC) 

S. Avans (MSFC) 

S. Croo~les (MSFC) 

M. Mikulas (LaRC) 

J, Hoosner (LaRC) 

L. Purves (GSFC) 

K. Hinson (JSC) 

C. Messelski (JSC) 

J. Lewis (JSC) 

J. Queller (Hqs) 

G. Robinson (MSFC) 

J. Stokes (MSFC) 
REPORT SECTION - LW CONTRIBUTItiG PERSQNNEC 

5.0 Umnned D. Suddeth (GSFC) W. Muller (GSFC) 

Platforms R. P-frey (JPL) J. Graf (JPL) 

5.1 Design Reqtn'ts S. Bergeson-Yillis (GSFC) R. Huller (GSFC! J. Graf (JFL) 

D. Vane (JPL) 

5.2 Configuration R. Huller (WC) J. Young (GSFC 1 

Description J. Sunkel (JSC J. Bell (JSC) 

B. Holder (JSC) L. Livingstop (JX) 

T. Pmck (GSFC) J. Schliesing IJSC) 

R. Berka (JX) 

5.3 Subsystems D. Greenshields (JSC J. Ywng (GSFC) A. Fitzkee (GSFC) 

T. Taylor (JSC) 

S. Lenett (JSC) 

M. Kapell !JSC) 

D. Rhoades (JSC) 

E. Bracalente (LaRC) 

K. Grady (GSFC) 

D. Suddeth (GSFC 

D. Kendrick (JSC) 

R. Klementson (JPL) 

J. Rakewicz (JPL) 

C. Gilchriest (JPL) 

C. Hermann (JSC) 

A. ilula (KSC) 

5. Anderson (JX) 

R. Hartenstrin (GSFC) 

C. Thomas (JSC) 

C. Woolley (LaRC) 

0. Weary (JSC) 

C. Bell (JPL) 

M. Mangano (JPL) 

K. Rowley (JPLI 

R. Hiyake (JPL) 

W. Wuest (JPL) 

M. Holley (JSC) 

T. Berry (JSC) 

D. Tomlin (MSFC) 

5.4 Safety H. Williams (JSC) S. Truelock (JSC) 
REFERENCES 

1. JSC-19521, Conceptional Design and Eva1 uation of selected Space 

Station Concepts, D~cember 1983. 

2. JSC-19371, Space Station Vol . 4-preliminary power profiles for !OC 

Conf iguration-MPAD, February 1984. 

3. Leondis, Alex: "Large Advanced Space System Computer Aided Design 

and Analysis Program." NASA CR 159191-1, July 1980. 

4. Space Station Preliminary Design Report, NASA JSC Report 18555, 

September 1982. 

5. Jacchia, L. G.: Thermospheric Temperature, Density, and 

Composition: New Models. SAO Speciai Report No. 375, March 15, 

1977. 

6. Hedin, A. E. : A Revised Thermospheric Model Based on Mass 

Spectroraeter and Incoherent Scatter Data. Journal of Geophysics 

Research, Vol . 88, December 1983. 

7. JSC 07/00, Shuttle Payload Accommodations Handbook, Vor . 14, 

Attdchment 1. 

8. McCormick, Caleb; editor: t6C NASTRAN Manual. 

9. Mikulas, Martin, et al: Space Station Truss Structures and 

Construction Considerations. Space Station Project Off-ice White 

Paper, August 1984. 

10. Noor, A. K.; Anderson, M. S.; Greene, W. H.: Continuous Models 

for Beam and Platelike Lattice Structures. AIAA Journal, Yol. 16. 

ho. 12, December 1978. 

11. Noor, A. K.; Anderson, C.M.: Analysis of Beam-Like Lattice Trusses. 

~onyuter ~ethods in ~p~l ied ~echan-ics and Engineering, Vol . 20, 

1979. 

12. Leondi s, A1 ex: Large Advanced Space Systems Computer-Aided Design 

and Analysis Program. NASA CR-159191, 1980. 

13. Garrett, L. Ber~ard: "Interactive Design and Analysis of future 

Large Spacecraft Coficepts ." NASA TP-1937, 1981. 

14. Heck, Michael L.: kigid body Controi Dyt~amics: Test Case Resu! ts 

and Documentation - AMA 4-9 Report Ng. 83-8, Analytical Mechanics 

Associates, Iw., April 1983. 

15. Eeck, Michael L.: Final Report Rigid Body Control Dynamics Part 

Ii. AMA Report No. 83-22, Analytical Mechanics Associates, Inc., 

April 1983. 

16. Garrett, L. B. and Ferebee, M. L., Jr.: "Comparative Analysis of 
Large Antenna Spacecraft Using the IDEAS System.' AIM Paper 

83-0798-CP Mily 2-4, 1983. 

17. Dorsey, J. T.; and Bush, H. G.: Dynamic Characteric,tics of a Space 

Station Solar Wing Array. NASA TM-85780, 1984. 

18. Load SpecificatSon - SRMS Manipulator Ann. SPAR-SG. 409 SPAR 

Aerospace, Toronto, Ontario, Canada, September 1.978. 

19. Mikulas, Martin, M., Jr.; Bllsh, Harold G.; Uallsom, Richard E.; 

Dorsey, John T. ; and Rhodes, Marvin D. : A Manned-Machine Space 

Station Construction Concept, NASA 3 85762, February 1984. 

20. Heard, Ualter L.,Jr.; Bush, Harold G.; Uallsm, Richard E.; and 

Jensen, J. kemit: A Mobile Uork Station Concept for Mechanically 

Aid Astrcnaut Assedly of Large Space Trusses. NASA TP 2108, 

March 1983. 

21. Large Spac~ Systems Technology - 1979 Technology Review, Hampton, 

Virginia. NASA CP 2118, November 1979. 

22. Large Space Antenna Systems Technology - 1982, Parts 1 and 2. USA 

CP 2269, November 30-December 3, 1982, pp. 257-283. 

23. Greenberg, H. S. : 'Development of Deployable Structures for Large 

Space Pl atf vms'. NASA CR-170689, Decenber 1982. 

24. Sox, R. L.; Nelson, 7. A.: 'Development of Deployable Structures 

for large Space Platforms." NA3A CR-170690, December 1982. 

25. Heard, Walter L., Jr.; Bush, Harold G.; Wallsom, Rich-rd E.; and 

Jenson, J. Kemit: 'A Mobile Work Station Concept for Mechanically 

Aided Astronaut Assdly of Large Space Trusses.' NASA TP 2108 

March 1983. 

26. Mikulas, Martin M., Jr. and Bush, Harold G.: "Advances fn 

Structural Concepts.' Large Space Antenna Systems Technology-1982, 

Part 1, NASA CP 2269, November 30-December 3, 1982, pp. 257-283. 

27. Dorsey, John T.: "Structural Performance cf Orthogo~al Tetrahedral Truss Space Station Configurations." NASA TM 86260, July 

1984. 

28. Bush, Harold G.; Mikulas, Martin M., Jr.; Wallsom, Richard E.; and 

Jensen, J. "hemit: "Conceptual Design of a Mobile Remote 

Manipulator Systm." NASA TI4 86262, July 1984. 

29. Mikulas, Martin M., Jr.; Bush, Harola G.; Wallsm, Richard E.; 

Dorsey, John T. ; and Rhodes, Marvin D.: "A Manned-Machine Space 

Station Construction Concept," NASA TM 85?52, February 1964. 

30. Mikulas, Martin M., Jr.; Bush, Harold, G.; and Card, Michael F.: 

"Structural Stiffness, Srrength and Dyr~anlic Characteri sties of Large 

Tetrahedral Space Truss Structures," NASA TM X-74001, March 1977. 
31. Amstrong, U. H.; Skoural , 0. E.; and Straqyer, J . U.: 'Large 

Space Erectable Structures,' Contract MS9-14914, April 1977. 

32. Heard, U. L., Jr.; Bush, H. G. ; Ualz, J. E.; and Rehder, J. J.: 

Structural sizf ng Considerations for Large Space Pl atfoms. Journal 

of Spacecraft and Rockets, Vol . 18, No. 6, pp. 556-564, NoveaberOecember 1981. 

33. Bluck, R. M.; Johnson, R. ii.: 'Fabrication of Slender Struts for 

Deployable Antennas. ' MSA CR-172 164, April , 1983. 

34. Kennel, NASA ri4-82390, Steering Law for Para1 lel Mounted Double 

Gimbaled W's--Revision A. 

SPACE STATION Yi'BSYSTEM VH IT€ PAPERS (JX-20054) 

o Space Station Truss Structures and Construction Considerations 

o Solar Panel Deployaent khanisms 

o Mule and Pqyload Truss Attac&nt 

o Module to Module Interface, Docki ng/Berrhi ng 

o Prel iminary Structural Design and Analysis of a Shuttle Launched 

Space !tation Wanned Habi tab1 e Module 

o Concept for a Space Station Solar Arrw Rotary Joint 

o Depl~y~nt Verification 

o Conceptual Design of a Mobile Remote Wanipulator System 

o Implementing Space Station Autozioaty/Automatf on 

o S;~.ce Station Large Area Magnetic Torquer 

o Space Station Fluidic Momentua Controller 

c Solar Dynamic Energy Conversion Systems for a Manned Space Station 

o A Power Managewnt and Distribution System for a Hiinned Space Station 

o Module Pattern 
APPENDIX A 

ASSEMBLY OPERATIONS CREW TIMELlNES EVALUATION 
. 

C 

0 CI 

.r U) 

r-QC, L wao 0 CUaJ C, 

a OX E m a#- aJ.r 

OW0 L r- C, s .u 0 0 C rnL Cr-rvv- acl 

.rLL - 

-0 aJ 0.4 aauw~ 

L c aw~ CC 

- 

F L oao 

C,oYmz -%I a 

!it!=* L C 

'I 0 0 mucy cnu 

roa 

w n~ 
a.' 

In 

rO 

9 

C 

0 

0 

C 

.F 

4 

L 

c, L m 0 

w r L lll 

P 

z J 0 n 

E Is = 

4 4 - E 

a%- 

m o 

4 

9 0 v V) 

aJ a &a 

0 

-.cu, c t 

Q3a 0 o 

a SF .,- OLO C, +' 
ASSEMBLY TASK EVALUATION 

- FUNCTION/TASK - -- -- ----. TIME (MIN) METHOD FEASIBILITY CO

FIRST FLIGHT (continued) 

o Plug WWS into recharie panel 5 EVA 1 

o Checkout syste;;ls No data 

lack of 

o Re1 ease 38' struct,ure 

o Stow RMS 

TOTAL 

-- 617 MIN 

13.28 HR 

Re1 ease 

latches 

backs aw

1 
ASSEMBLY TASL EVALUATION 

FUNCTION/TASK TIME (MINI METHOD FEASIBILITY 

SECOND FLIGHT 

o Dock with structure 

o Checkout MRMS 

Se

C-

30 E'l A 

o Install lower keel 60 EVA-RMS 

o Grapple lower keel (5 1 RMS 

o Release lower keel from holding (5) EVA 

fixture 

o Maneuver lower keel to installation (15) RMS 

location 

w 

m 

o EVA crewmeinbers get in position for (15) EVA 

Q, attachi ng 1 over keel 

o Attach lower keel to structure (23) EVA-RMS 

o Connect Utilities 10 EVA 

o Deploy rail extensions 15 AUTO 

o Deploy lower keel 92 AUTO 

o Orploy radiator booms (2) 5 AUTO 

o Deploy radiztor heat exchanger boom (4) 5 AUTO - 

o Erect bays (4) next to radiator boas 5 2 EVA 

3 D

at

No 

ha

1 Se

3 Se

(1

1 
ASSEMBLY TASK EVALUATIOt4 

-- FUNCTIOK/TASK -------I_.__ TIME !MINI METHOD - FEASIBILITY COM

SECOHO FLIGHT (continued) 

o Grapple rddlator package (5) KYS 1 

o.Release radiator launch restraint (5 ) EVA 1 

o Maneuver radiator pdckaye to MRMS (30) EIA-RMS 2 

and attdch 

o MRi4S transldte to position for (2U) l.lMS 

rad'a tor siowage 

o Remove radidtors (42) f rola l-1Ri4S (50) EVA 

and stow on sidds of lower keel 

2 See Assum

u o MRMS translate to position (5) f4RMS 

Q, 

1 

v radiator 6) installation 

o EV2 ingress MHMS i.lFi4 (5) EVA 1 

o EV2/W-R retrieve a rddiator from (601 MRMS/MFR 4 Questionab

the bundle, EVl maneuvers EV2/MFR siz? of ra

to radiator Installation position. tolerances

EV2 Installs radiator & tightens person ins

it down. (20 lninirad. x 3 rad.) because on

to operate 

o Move M&4S to other radiator site (5 1 b1RMS 1 

o Same as t~o steps above (3 radiatot-s) (60) t42MS/13FR 4 

o Instal 1 keel extensions (2) 2 19 MRMS-EVA-AUTO 2 

ASSEMBLY TASK EVALUATIOtI 

FUNCTION/TASK TIME (MINI METHOD FEPASIBILLTY COMM

SECOND FLIGHT ( conti nued) 

o MRMS translate to keel extension (25) 

o Erect a bay between the end of the ( 20 

lower keel and each keel extension 

o MRMS translate down keel extension (5) 

o Erect three bays between the keel (30) extensions 

o Install MRMS recharge panel on lower 15 

keel 

-A 

Q, o Plug FORMS into recharge panel rD 

5 

o EV crewmembers translate back to 15 

Orbiter 

TOTAL 

- 911 MIN 

15,2 HR 

MWS 

EVA 

MRMS 

EVA 

EVA 

EVA 

EVA 
A Cn Pi 

I * 

5 t' W 4 i n E aJaJP: Y Cn cc LLZ a s 

0 ==I- 4 OE msSF 

I c,WV I C, 

o uux 4 a t E aze ru u t' 3 LU C 

I Y LLCW ,-- C, e la 0 

C,*JI/)L E 

3 I L V) V) 3 n o a, C, 0 V) 0 

3 mi, L . C, C, 

0 CCUI m 1 .r - ?--+L 0 w t'ww Z%5cnFE t' ,-- 

3 V) 0 w 4 I I 5 C, 110 s a - u 3 3OXUIU 3 C, 0 V) 

c, C aJ EOrn u SF 

U =we CEp:.rz 0 E .P a~ 3 a 3 - a aa~wer 5 ~c 4 u Lrn* F.->CdT 0 .,- ~ cnp 3 -3 a A - m - 3 Q) CQ) w c, u aJ .- Z+":~~O g2s m-re r~ t' I r I 4 4 v. 3 13 *C,aJ-RIQ en C aJ X0.r CC, ctr>V)> L,-- u F V)F ~CI uu C, 

c 9 V)Q) .n 4 c=s C, IU u 

"~~gfP~+ 4 c, 4 LO IUU Q-uu ,rcO WL t'~ = a * I- < hC U) = IS < "9 I-E Em- <V) c C C 0 

0 

UQtYOOb uQOOuO 0 V 

I- 0 0 0 0 0 
ASSEMBLY TASK EVALUATION 

FUNCT ION/TASK l IME (MINI METHOD FEASIBILITY C

THIRD FLIGHT (continued) 

o Retrieve pigtails from keel extension (15) EVA 1 

o Connect connectors to two side panels (20) EVA 1 

o Connect connectors to end of module (15) EVA 1 

o instal 1 Airlock 1 20 MRMS 1 

o Grapple Air1 ock 1, re1 ease (10) MRMS 1 

o MRMS maneuver ASrlock 1 to HA0 1 (5 ) MRMS 1 

berthing location 

V 

V 

o Berth Alrlock 1 to HAB 1 c. 

o Temp stcd Alrlock 2 

(5) MIMS 1 

20 MRMS 1 See Ass

0 Grapple Airlock 2, release launch (10) HRMS 1 

restrainis 

o MRMS maneuver Alrlock 2 to HAB 1( 5 MRMS 1 

berthing location 

o Berth Airlock 2 to HAB 1 (5) MRMS 

o MRMS translate to recharge panel 10 MRMS 

o Plug MRMS Into recharge panel 5 EVA 

o EV cremnenbers translate back 5 EVA 

to Orbiter 

ASSEMBLY TASK EVALUATION 

FUNCT IOWTASK ---- -- ---- TI MI METHOD FEASIBILITY - COMMENTS 

FOURTH FLIGHT 

o Dock with HAB 1 moduie 

o MRMS activation 15 

o Instal 1 module mounting structure 5 9 

o Grap+le mounting structure, release (15) 

launch restraint 

o Lift mounting structure out of (10) 

cargo bay 

o MRMS grapple mounting structure dud (5) 

RMS re1 ease mounting structure 

o MRMS maneuver mounting structure (10) 

to attachment location 

o Attach mounting structure (10) 

o Berth/attdch HA0 2 module to HAB 1 7 5 

module 

o MRMS grapple CIA0 2 moduli! (5) 

o MRMS maneuver HA8 2 nodule and (40 

berth it to the HAB 1 module 

o Attach HAB 2 module to mounting (30) 

structure 

EVA 1 

MRMS-RMS-EVA 3 See Assumption

RMS-EVA 1 

MIU4S- RMS 1 

EVA 3 No data on att

hardware 

MRIJS 1 

MRMS 1 

EVA 3 No data on att

h -dware 
ASSEMBLY TASK EVALUATION 

- FUNCTIONITASK TIME (MIN) METHOD FEASIBILITY 

FOURTH FLIGHT (continued) 

o Connect side panel utilities 25 EVA 1 

o Remove Airlock 2 from HAB 1 and 20 MRMS 1 

berth it to HAD 2 

o Install upper keel 130 MRMS - RMS - 3 

EVA-AUTO 1 

o Grapple upper keel , re1 ease 1 aunch (15 ) RMS-EVA 1 

restraints and 1 ift out of cargo bay 

o MRMS grapple upper keel and RMS (5 ) MRMS-RMS 1 

release upper keel 

o MRMS translate to upper keel (30) 1 

instal 1 ation location 

o Unfold deployment rai 1 s (5 1 EVA 1 

o Attach upper keel to transverse boom (30) MRMS-EVA 3 No d

hardw

o Deploy upper keel and upper boom (110) AUTO 3 See A

o MIMS translate to recharge location, 3 5 MRMS-EVA 1 

plug MRMS in for recharge 

o EV crewmembers translate back to 5 EVA 1 

o Checkout HAB 2 

TOTAL 355 MIN 

5.9 HR 

No d

lack 
I ASSFMBLY TASK EVALUATION 

------ FtJNCTION/TASK TIME (MINI METHOU FEASIBILITY ---.-- - - --...---. ... - - - COMMENTS 

FIFTH FLIGHT 

o Dock with HAB 1 module 

I o MRl4S activation 15 EVA 1 

I o Berth logistics module Lo HAB 2 module 60 MRMS-Rl4S 2 

I o RMS grapple logistics aod~!lz (5) RMS 1 

I a Lift LOG module out of cargo bay (10) RMS 1 

o MRMS grapple LOG lnodule and Ri4S (5) MIU~S-RMS 1 

release LOG rmdule 

o MRMS maneuver LOG module and berth (25 1 t4RMS 2 

it to the HAB 2 module 

4 

V 

VI o Connect external tl~ermal connector (15 EVA 1 

o Install port soldr drray addition pkg. 172 MRl~lS-RYS- 3 

EVA-AUTO 

o Grapple port solar array package (5) RMS 

and release launch restraints 

l o Lift array package out of cargo bay (10 1 RMS 1 

o MRMS grapple array package and (5) 14R14S-RMS l. 

&#IS release array package 

o l.lfU4S trans1 ate to port end of (40) MRMS 1 

transverse boom 

o Unfold deployment rails (5) EVA 1 
ASSEMBLY TASK EVALUATION 

FUNCTION/TASK --.---------- TIME (MINI METHOD FEASIBILITY _ __------_I_-- -- ---- -- 

FIFTH FLIGHT (continued) 

o Attach port solar drray package to (35) MRMS-EVA 3 No ddt

end of transverse boom dnd mate hardwa

connectors 

o Ueploy port outboard transverse (32 AUTO 3 See As

boan structure 

o MHMS translate back to cargo bay ared (40) r4RI4S 1 

o Install starbodrd solar array (132 t4RMS- RMSaddition package EV A-AUTO 

o Grapple starboard soldr array (5 RMS 1 

addition p3 kage and re1 ease 

launch res~< dints 

V 

w 

Q, 

o Lift array package out of cargo bay (10) RMS 1 

o t4RMS grapple array package and (5) I4RflS- RMS 1 

RMS release array package 

a I.1RMS translate to starboard end (40) 14R14S 1 

of transverse boom 

o Unfold deploy~lent rai 1 s (5) EVA 1 

o Attach starboard solar array (35 MRMS-EVA 3 No dat

package to end of transverse boorn hardwa

and mate connectors 

o Deploy starboard outboard transverse (32) AUTO 3 See As

boom structure 
ASSEWLY TASK EV ALUATIOW 

FUNCTION/TASK TIME (MIN) METHOD FEASIBILITY C

FIFTH FLIGHT (continued) 

o Deploy outboard solar arrays (4) 60 AUTO 1 

o MRMS translate to recharge location, 45 MRMS-EVA 1 

plug MRHS in for recharge 

o EV cremnembers translate back to 5 EVA 1 

Orbiter 

o Module activation for inhabi tance 

TOTAL 

- 489 MIN 

8.15 HR 
ASSEMBLY TASK EVALUATION 

---- FUNZTION/TASK TIME (MINI METHOD FEASIBILITY ---------- ---- 

SIXTH FLIGHT 

o Dock with HXE 1 module 

o MRMS activation 15 

o Install module mounting structure 50 

o Grapple mounti ny structure, r2l ease (15) 

launch restraint 

o Lift mounting structure out of (10) 

cargo bay 

o MRWS grappl2 nount ti ng structure k (5 

RMS re1 ease mounting structure 

V 

u o MRMS maneuver mounting structure to (10) a~ attachment location 

o Attach mounting structure (10) 

o Berth/attach LAB 2 uodule 7 5 

o Grappls LAB 2 module, release (5) 

launch restrdint 

o Lift LAB 2 out of caygo bay (10) 

o l4RMS grapple LAB 2 module and (5) 

release LAB 2 module 

o MRMS maneuver LAB 2 and berth it (25) 

to the HA0 2 rnodule 

EVA 1 

ATV-RMS-EVA 3 See As

IWS-EVA 1 

RMS 1 

EVA 3 140 dat

hardwa

t~lRMS-KMS-EvA 3 

RI-1 S 1 

MKMS-RMS 1 

CdRIIS 3 Close 

berth 

both re
ASSEMBLY TASK EVALUATION 

FUNCT ION/TASX -------- TIME (MIIJ) METHOU FEF;IBILITY -.--------- COMFJN'r" 

-- 

SIXTH FLIGHT 

o Dock with YQB 1 wdule 

I o WS activation 15 EVA 1 

I o Instal 1 module tnoun ti ng structure 50 ATV-RMS-EVA 3 See Assumpti on

o Grapple nount ti ?ly structure, re1 ease (15 RMS-EVA 1 

launch restraint 

o Lift ~nounting structure out of (10) RMS 

cargo bay 

o MRMS grdpple mounting structure & (5) I4RMS-RMS 1 

R14S release mounti ny structure 

o MRMS maneuver nwunting structurs to 

V (10) MRMS 1 

(O attachr~ient 1 ocation 

o Attach mounting structure (10) EVA 3 

o Ber'cn/attach LAB 2 rnodule 7 5 MRMS-RMS-EVA 3 

o 6rapplz LAB 2 nodule, release (5 ) 7.24 S 1 

; aunch restraint 

o Lift LAB 2 out of cargo bay (10) HMS 1 

o MWS grapple LAB 2 module and (5) l.NMS-RMS 1 

re1 ease LAB 2 module 

o iW4S maneuver LAB 2 and berth It (25 ) MN4S 3 

to the HAB 2 nodule 

No data on att

hardware 

Close toleranc

berth and att?r

both required 
ASSEMBLY TASK EVALUATION 

FUNCTION/TASK TIME (MINI METHOD FEASIBILITY 

SIXTH FLIGHT ( contf nued) 

o Attach LAB 2 module to mounting (30 ) EVA 

structure 

o Connect side panel utilities 25 EVA 

3 No d

hardw

o MRMS translate to recharge location, 10 MRMS-EVA 1 

plug MRMS in for recharge 

o EV cremembers translate back to 5 EVA 1 

Orb1 ter 

TOTAL 

- 180 MIN 

3 HR 
ASSEMBLY TASK EVALUATION 

FUNCT IONITASK - TIM (MINI METHOD FEASIBILITY ---- - -- COMMN

SEVENTH FLIGHT 

3 Dock with LAB 1 module 

I o MWMS actlvatlon 15 EVA 1 

I o Install nodule mouqtlng structure 50 MRMS-RMS-EVA 3 See Assmptio

o Grapple mounting structure, (15 ) RI?S-EVA 1 

re1 ease 1 aunch restraint 

o Lift mounting structure out of (10) RMS 1 

cargo bay 

o MRMS grapple mounting structure (51 I4RMS-Rt-!S 1 

6 RMS release mount1 ng structure 

I 2 o MRMS maneuver mounting structure (10i IWqlS 1 C, to attachment location 

o Attach mounting structure (10) EVA 3 No data on a

hardware 

I o Serth LAB 1 module 8 0 MRHS-RMS-EVA 3 

a Grapple LAB 1 i~dule, release (5 1 RMS 1 

ladnch restralnt 

I o L!ft LAB 1 out of cargo bay (10: RMS 1 

o MRMS grapple LAB 1 module and RMS (5 1 MRMS-RMS 

release LAB 1 module 

o MRMS manet er LAB 1 I berth it to (30 1 MRMS 3 Difficult ta

HA0 1 and LAB 2 modules Brtf Wse~ i~ 
ASSEMBLY TASK EVALUATION 

- FUNCTION, TASK -- TIME (MINL METHOD FEASIBI1;l'Y -- 

Sf VENTH FLIGHT (continued) 

o Attach LAB 1 module to mounting (30) EVA 

structcre 

o Connect side pans1 utilities 25 EVA 

3 No d

hardw

o MRMS translate to recharge location, 10 MRMS-EVA 1 

plug MRMS in for recharge 

o EV cremmbers translate back to 5 EVA 

Orb i ter 

TOTAL 

-- 185 MIN 

3.1 HR 
SPACE STATION ASSEMBLY ASSUMPTIONS 

The following assumptions are made for assembling the Space Station 

concept. 

A. GENERAL 

1. Two EV creMerrbers are located in the cargo bay, +,cte IW )IZS been 

activated and checked out, and one WU is available for use on each 

mission. 

8. FLIGHT ONE - Basic Structure 

1. ( Instal 1 radiators) 

a. Assume that the radiator heater exchanger boons are not 

covered by the launch stowage of the radiator panels. 

b. Assuw some type of crew-operated quick release hand1 ing aid 

for the solar array panels. It must leave free access to one radiator 

edge, in order to permit instal lation beside the precr io~rsly instal led 

radiator panel . c. Assume approximately one inch clearance between installed 

radiator panel s. 

d. Assume radiator assenbly bars are parallel to 38-foot package 

long axis. 

e. Assume Go equipment restrai~types required, a launch 

restraint and separate panel temporary restraint. 

2. (Deploy Solar Array blanket boxes) 

a. Assume that deployment of solar arrays is automatic. 

b. Assume para1 1 el booms are deployed sinul taneousl y. 

3. (Deploy rail extensions) 

a. Assume automatic extension of two rails at a time. 

4. (Deploy boas) 

a. Assume automatic boom deploylnent of 3 minuteslbqy , plus a 

2-minute inspection of eight jointslbqy, plus a preload tPme of 2 

minutes. 

5. ( Instal 1 ATVIManipul ator) 
a. Assume ATV/manipulator has proven high re1 iabil i ty. This 

will represent new techno1 ogy. 

b. Assume a battery rechargelmaintenance station for the ATV, 

located on the Spdce S'iiiticifi structure. 

c. Assume mechanical and electrical connection guides provided 

for manipulator instal 1 ation of the ATV. 

6. (Deploy solar array blankets) 

a. Assume arrays are deploy€< sequentially. 

b. Assume crew invol vement is 1 imi te ' to observation only. 

7. (Erect baylattach docking device) 

a. Assume two-crewmember operation, with crew in foot 

restraints. 

b. Assume struts and nodes are restrained at assembly worksite. 

c. Assume that the berthing device is already attached to four 

assembly struts. 

C. FLIGHT TWO - Additional StructureiAppendages 

lC (Dock with structure) 

a. A.ssume there is no plume impingement problem. 

b. Assunis Orbiter control system is disabled during assembly and 

control is maintained by the Space Station CMG's. 

2. (Connect utilities) 

a. Assume utilities are located inside the cells. 

b. Assume crew will manual ly cOLilect electrical , instrumentation, and plumbing 1 ines to panels on side of structure out of 

MRMS path. 

3. (Deploy lower keel ) 

a. Assumed a deployment rate of 3 minuteslbay, plus a 2-mSnute 

,inspection of each bay, plus a preload time of 2 minutes for 18 bays. 

b. Assume a feasibility level of 3 due to the state-of-the-art 

level and number of bays. 
c. Assume EV crew participation limited to observation for 

noainal operations and assistance in contingency operations. 

4. ( Instal 1 /stow radiators) 

a. Assume two-EVA-crew-attach radiator panel restraint bracketry 

on both sides of lovor keel. 

b. Assurne these radiator panels will be installed later when the 

Space Station is inhabited. 

c. Assume six installei radiators are sufficient at this point. 

5. [Install keel extensions) 

a. Assme that docking/berthing ring is already attached to the 

end of one of the two keel extensions. 

b. Assume one crewember operating MU6 and the other one doing 

the attachment. 

c. Assume a deployment rate of 3 minutes/bay , plus a preload 

time of 2 minutes, plus 2 minute/bay inspection of eight joints/bay for 

10 bays. 

6. (Install lower boom and ancillary struc+~re) 

a. Assulle the two lower boom packages and 48 erectable struts 

and associated conr~ectors are attached to a pal let. 

b. Assume a deployment rate of 3 minutes/b?y, plus a preload 

time of 2 minutes, plus 2 minutes per bay inspection of eight joints/bqy. 

0. FLIGHT THREE - HA0 1 Module 

1. Same assumption as assumption C-1 (Flight t;rJo) 

2. (Temp stow airlock 2) 

Airlock 2 will be stowed temporarily to HAB 1 until HAB 2 is in 

place, which is where it will ncrmally be located. 

E. FLIGHT FOUR - HAB 2 Module 

1. (Install nodule mounting structure) 

a. Assume that the package is pre-assenrbled arld crew installed. 

2. (Install upper keel) 
a. ASS- a deployment rate of 3 minutes/bqy, plus a preload 

time of 2 ainutes, plus 2 mlnutes per bay inspection of eight joints/bay. 

b. Assume a feasibility level of 3 due to the state-of-the-art 

level and number of bays. 

c. Asswe EV crew participation limited to observation for 

nominal operations and assi stance In contingency operations. 

F. FLIGHT FIVE - Logistics Module 

1. ( Install port solar array addition package) 

a. Assume a deploylrent rate of 3 einutes/bay, plus a preload 

time of 2 minutes, plus 2 minutes per bay inspection of eight joints/bay. 

b. Assuned a feasibility level of 3 due to the state-of-the~art1 eve1 . 

c. Assuae EV crew participation limited to observation for 

nminal operations and assistance in contingency operations. 

6. FLIGHT SIX - LAB 2 Module 

1. (Install module nounti ng structure) 

a. Assume that the package is pre-assembled and crew installed. 

H. FLIGHT SEVEN - LAB 1 ;4odule 

1. (Install module mounting structure) 

4. ASSURE that the package is pre-assembled and crew installed.