NASA Technical Memorandum 102675

A Study of Concept Options for the Evolution of Space Station Freedom

Herbert R. Kowitz:

Karen D. Brender

William M. Cirillo

Lisa D. Collier

George G. Ganoe

Marston J. Gould

Martin J. Kaszubowski

George F. Lawrence

Charles P. Llewellyn

Ray Reaux

Washito Sasamoto

Patrick A. Troutman

Laura Waters

(NASA-[_-107o71,)

FqR [_4_ _V;JLUTI_()N

(_ASA) 77 n

May 1990

A STONY OF CONCFPT nPTI_NS

UF $PACF _[ATTuN F_EEnjM

C_CL 22A

GJ/l_

t_90-Z5154

Nahonal Aeronautics and

Space Admtn_strabon

Langley Research Center

Hampton, Virginia 23665-52.25

PREFACE

Space Station Freedom will be a versatile, multifunctional facility that will provide for

continuous human presence in low Earth orbit. Freedom will be capable of evolving over its

extended operational lifetime to be responsive to changing user needs and long-term national

goals. The Evolutionary Definition Office (EDO) at the Langley Research Center (LaRC) has

the responsibility for the definition of evolution paths/options beyond the baseline system to

meet future needs. In addition, the EDO is responsible for defining requirements on the

baseline station design, e.g., hooks and scars, to assure viable growth options.

Studies have been conducted to assess the impact of future missions on the station

configuration, systems, and operations. These studies have considered two major growth

options (paths). One option emphasized the development of a multidiscipline research and

development (R&D) facility to accommodate growth in science, technology, and commercial

missions. The second "option emphasized the role of the station as a transportation node to

support the proposecl solar system exploration missions. This option also allows enhancement

of research accommodations and provides the means to add functions, such as assembly,

servicing, and repair of space transfer vehicles and satellites.

This document presents descriptions for each configuration option and provides details

of the analyses performed to assess the viability of each. Much of the work described in this

document will be included in the Space Station Freedom Evolution Reference Databook. The

reference configuration is a combination of station growth paths for SEI support and R&D

emphasis.

i

CONTENTS

1.0

1.1

1.2

1.3

1.4

1.5

2.0

2.1

2.2

2.3

3.0

INTRODUCIION

PHASED PROGRAM

PURPOSE OF THE STUDY

BASELINE CONFIGURATION

EVOLUTION CONFIGURATION CONCEPTS

DESCRIPTION OF ANALYSES

3.1

3.2

3.3

4.0

4.1

4.2

4.3

R&D FACILITY CONFIGURATION CONCEPT

FUNCTIONAL REQUIREMENTS

CONFIGURATION DESCRIPTION

R&D CONFIGURATION CONCEPT ANALYSES

2.3.1 Mas_; Properties

2.3.2 Attitude Control Analysis

2.3.3 Mictogravity Envelope Determination

2.3.4 Orbit Lifetime and Reboost Analysis

2.3.5 Stru,-tural Dynamics Analysis

TRANSPORTATION NODE CONFIGURATION CONCEPT

FUNCTIONAL REQUIREMENTS

CONFIGURATION DESCRIPTION

TRANSPORTATION NODE CONFIGURATION

CONCEPT ANALYSES

3.3.1 Mass Properties

3.3.2 Attitude Control Analysis

3.3.3 Microgravity Envelope Determination

3.3.4 Orb:it Lifetime and Reboost Analysis

3.3.5 Structural Dynamics Analysis

PHYSICAL CHARACTERISTICS OF ELEMENTS/SYSTEMS

REPLICATED BASELINE ELEMENTS

GROWTH ELEMENTS

4.2.1 Solar Dynamic (SD) Power Systems

4.2.2 Truss Structure

4.2.2.1 Booms and Keels

4.2.2.2 SD Truss Extension

4.2.3 Extended Resource Node

4.2.4 Pocket Laboratory

4.2.5 Space Transfer Vehicle Servicing Facility

4.2.6 Lunar Vehicle Assembly/Servicing Facility

4.2.7 Mars Transfer Vehicle Processing Facility

DISTRIBUTI_D SYSTEMS

4.3.1 Electrical Power System

4.3.]..1 Basehne

4.3.1.2 Growth

4.3.2 Thermal Control System

4.3.2.1 Baseline

4.3.2.2 Growth

4.3.3 Data Management System

1-1

1-1

1-1

1-1

1-4

1-8

2-1

2-1

2-2

2-5

2-5

2-5

2-8

2-8

2-14

3-1

3-1

3-3

3-3

3-3

3-8

3-12

3-12

3-19

4-1

4-1

4-2

4-2

4-3

4-3

4-3

4-3

4-5

4-8

4-8

4-11

4-11

4-11

4-11

4-12

4-12

4-12

4-13

4-13

iii

PRECEDING PAGE BLANK NOT FILMED
5.0

5.1

4.3.4

4.3.5

4.3.6

4.3.7

4.3.8

4.3.9

4.3.10

4.3.3.1 Baseline

4.3.3.2 Growth

Communications and Tracking

4.3.4.1 Baseline

4.3.4.2 Growth

Guidance, Navigation, and Control

4.3.5.1 Baseline

4.3.5.2 Growth

Extravehicular Activity System

4.3.6.1 Baseline

4.3.6.2 Growth

Man Systems

4.3.7.1 Baseline

4.3.7.2 Growth

Environmental Control & Life Support System

4.3.8.1 Baseline

4.3.8.2 Growth

Fluid Management System

4.3.9.1 Baseline

4.3.9.2 Growth

Propulsion System

4.3.10.1 Baseline

4.3.10.2 Growth

GROWTH DELTA ANALYSIS

ANALYSIS METHODOLOGY

5.1.1 Attitude Control Analysis

5.1.2 Microgravity Environment Determination

5.1.3 Orbit Lifetime and Reboost Analysis

5.1.4 Structural Dynamics Analysis

R&D DELTA CONFIGURATIONCONCEPT

ANALYSES SUMMARY

TRANSPORTATION NODE DELTA CONFIGURATION

CONCEPT ANALYSES SUMMARY

4-13

4-15

4-15

4-15

4-16

4-17

4-17

4-18

4-18

4-18

4-19

4-20

4-20

4-21

4-21

4-21

4-22

4 -23

4-23

4-24

4-24

4-24

4-25

5-1

5-1

5-1

5-2

5-3

5-3

5-4

5-6

iv
1.0 INTRODUCTION

1.1 PHASED PROGRAM

Space Station Freedom will evolve from the baseline or Assembly Complete

(AC) configuration lo an enhanced capability configuration as determined by user

needs and U.S. Space Policy. This evolution will take place over several years in

incremental steps. The Langley Space Station Freedom Office is responsible for the

process of defining the long-term requirements, the establishment of evolutionary

configuration concepts that meet future national goals, and the identification of hooks

and scars to the baseline station that will assure that these goals can be met.

1.2 PURPOSE OF THE STUDY

The evolution configurations described in this document served as a starting

point for the further development of an evolution reference configuration to guide the

"design for evolution" process within the Space Station Freedom Program. In addition,

this document contains a system-by-system description of the features which currently

exist in the baseline design.

This study prc vides, a database for evolution configuration concepts. Functional

and physical characteristics, as well as supporting analyses, are provided for two

manned-base configurations, a Research and Development (R&D) configuration and a

Transportation Node configuration. The Transportation Node configuration is

comprised of a further break-down into a lunar operations configuration for near-term

exploration missions to the moon and a lunar/Mars configuration for far-term missions

to the moon and Mars. This document provades reference material for the evaluation

of the baseline station's ability to accommodate growth. It is intended to assist in the

assessment of evolution options and to facilitate definition and evaluation of software

hooks and hardware scars required in the baseline station. The analyses presented

here are unique to the configuration for which they were performed. The data are

intended solely for reference and evaluation purposes.

Section 1 pro_rides an overview of this document and describes the analyses that

were performed. Sections 2 and 3 provide the station characteristics, including mass,

center of gravity, moments of inertia, and ballistic coefficient, as well as analysis results,

for the R&D station and the Transportation Node, respectively.

Section 4.0 provides, for each element and distributed system in the growth

categories, a functional description and an enumeration of physical characteristics such

as dimensions, mass, and volume.

Section 5.0 provides a detailed explanation of the analyses performed, the

assumptions used, and the results for the growth R&D and Transportation Node

configurations.

1.3 BASELINE CONFIGURATION

The baseline space station configuration is illustrated in Figure 1.3-1, and its

physical characteristics are summarized in Table 1.3-1. The baseline configuration will

consist of a truss assembl),, called the transverse boom, on which elements, such as the

laboratory (lab) and habitation (hab) modules, and distributed systems, e.g., electrical

1-1
7

I',O

Figure 1.3-1 " Baseline Space Station Freedom Configuration
Characteristics

Mass (kg)

Average @ag area

(m z )

Average ballistic

coefficient (kg/m 2)

Inertias (kg,m 2)

with respect to the

center of mass

Center of mass (m)

from center of

transverse boom

Maximum lengths

(m)

Average flight

attitude (degrees)

from LVLH

Attitude control

momentum (Nom,s)

Value

266,000

2,310

50.2

IXX= 1.6E8

IYY= 2.2E7

IZZ= 1.7E8

IXY= 5.7E5

IXZ = -1.6E6

IYZ = -9.3E8

X = -1.27

Y= 0.21

Z= 3.51

X= 60.0

Y =155.0

Z = 80.0

Roll (X) = 0.34

Pitch (Y) = - 8.00

Yaw (Z) = 0.25

to -11.62

4,000

Table 1.3-1: Summary of baseline configuration

physical characteristics

1-3
power system, will be attached. The truss will be built up in 5 m (16.4 ft) cubes or bays

using 5.08 cm (2 in) diameter truss tubes. A utility tray that runs through the length of

the truss provides cabling and plumbing for resource distribution.

The baseline station will have four pressurized modules. Three of these

modules, one each from the U.S., Europe, and Japan, will be dedicated laboratories.

The fourth pressurized module (hab module) will provide an area for rest, recreation,

and health services for eight crew members. The pressurized modules will be

interconnected by four resource nodes which also serve as the control centers for space

station command, control, and operations. Two of the resource nodes will have an

NSTS docking mast and a cupola for direct viewing of proximity operations. The

baseline station will also have a hyperbaric airlock and attached pressurized logistics

carriers.

External to the pressurized volume will be the Mobile Servicing Center (MSC),

a telerobotic servicing platform that can traverse up and down the truss structure using

the Mobile Transporter, and the MSC Maintenance Depot (MMD). Also attached to

the transverse boom are attached payload accommodation equipment (APAE), pallets,

unpressurized logistics carriers, and an NSTS docking ring.

Attached to the transverse boom will be four photovoltaic modules that provide

the station with 75 kW of power. Also attached to the truss will be four thruster

modules, a resistojet module, and thermal radiators.

1.4 EVOLUTION CONFIGURATION CONCEPTS

Space Station Freedom must be capable of evolving over its extended

operational lifetime in order to respond to changing user needs and U.S. space policy.

r_.volution is defined as the process of increasing the capacity and capability of Freedom

after the completion of the approved baseline program. While the pace and direction

of evolution beyond the baseline is impossible to predict with any certainty, advanced

studies have been conducted in the program to assess the impact of future missions on

station configuration, systems, and operations. The general approach has been to

develop design requirements for Freedom which will insure that viable evolution

options remain open for the future.

Two evolution configuration options were studied. The first is a research and

development (R&D) oriented station that can provide extended laboratory and

servacmg capabilities for science and technology experimentation. The R&D

configuration concept is shown in Figure 1.4-1. It is a platform from which

astronomical and Earth observations can be made and an on-orbit research and

development facility on which life sciences, microgravity, and on-orbit manufacturing

can be pursued. Itis capable of performing on-orbit servicing of free-flyers as well as

station attached payloads.

The second configuration concept, shown in Figure 1.4-2, is a lunar/Mars

Transportation Node designed to support the manned Space Exploration Initiative

(SEI). The Transportation Node for the near term is a Lunar Operations

Transportation Node (shown in Figure 1.4-3). This configuration is required for lunar

missions and to support life science research and technology demonstrations for

manned Mars missions. It differs from the lunar/Mars Transportation Node concept

only by the later addition of a facility for assembly, checkout, and refurbishment of

Mars transfer vehicles.

1-4
I

k.,n

Figure 1.4-1 • R&D Configuration Concept
I

O_

Figure 1.4-2 • Transportation Node Configuration
!

Figure 1.4-3 • Lunar Operations Transportation Node Configuration
Both the R&D and the Transportation Node configuration concepts include the

addition of dual keels, solar dynamic power modules, pressurized modules, and

resource nodes to the baseline station. Each concept represents a specialized response

to perceived long-range national space program goals. Further physical descriptions of

the R&D and the Transportation Node configurations are provided in sections 2.2 and

3.2, respectively.

This document contains both functional and physical descriptions of the

elements for each configuration. There are two categories of elements: baseline

replicated elements and growth elements. The growth elements can be further divided

into those common to both concepts and those that are unique to a particular concept.

Replicated elements are those that are identical to baseline elements and will be

duplicated in the evolution station. Examples include structural components and lab

and hab modules. Elements common to both evolution configurations, but not to the

baseline station, include extended resource nodes, solar dynamic power modules, and

pocket labs. Examples of configuration-unique elements include exploration vehicle

assembly and servicing facilities.

1.5 DESCRIPTION OF ANALYSES

Most of the analyses presented here were performed using the IDEAS 2

integrated space station analysis software. This software package contains the

multidisciplinary analysis capabilities required for system-level space station

evaluations. These include:

rigid-body controls anal)_ses and flight mode characterization

evaluation of micrograwty environments

orbital lifetime and reboost analyses

determination of payload viewing envelopes

structural dynamics analyses

An important feature of the IDEAS 2 software is that a single geometric model

forms the basis for many of the various analysis functions, thus promoting consistency

and eligainating duplication of effort in generating separate models for the analyses.

IDEAS" Space Station Freedom models were created for each growth delta (shown in

section 5.0)using an enhanced version of the GEOMOD solids modeling program. The

geometric data was then processed by MODGEN which creates analytical models

compatible with the analysis modules.

The analysis programs in IDEAS 2 are integrated through a project relational

database which contains both input and output program data. Utilities are available for

reviewing and modifying input data stored in the database and for generating plots of

analysis results.

Detailed definitions of the analyses and the assumptions used in performing

them are provided in Section 5.0. The results from the analyses for the R&D and the

Transportation Node configuration concepts are provided in Sections 2.3 and 3.3,

respectively.

1-8
2.0 R&D FACILITY CONFIGURATION CONCEPT

The baseline station could be enhanced to provide greater research and

development capabilities. For the R&D configuration, these enhancements consist of

extended facilities for such research as earth and astronomical observation, life science,

microgravity, materials research, technologies demonstrations, and satellite servicing.

To provide these capabilities, the crew, power, pressurized volume, and external attach

points are increased. Additional functions, such as Space Transfer Vehicle (STV) and

satellite servicing, ate also implemented. The R&D configuration analyzed in this

report emphasizes microgravity research and materials production, although differences

in resource or pressurized volume requirements for other research emphases are noted.

2.1 FUNCTIONAL REQUIREMENTS

A comprehensive analysis of the growth of the Space Station Freedom as a

multidiscipline R&D facility has been conducted and the results are presented in

"Growth Requirements for Multidiscipline Research and Development on the

Evolution Space Station" (NASA TM 101497). A summary of this analysis is presented

below and was the basis for the functional requirements discussed in this secUon.

Facility and resource requirements for a multidiscipline space station are driven

primarily by the types of research and development being conducted and the level of

Earth-to-orbit transportation available. In the referenced study, mission sets were

developed for four specific utilization emphases on the station: microgravity research,

microgravity resear,:h and materials production, life sciences research, and

observational science. Multiple growth scenarios were then postulated from the

combination of each emphasis with three mixed-fleet transportation models

(conservative to aggressive lift capability). The analysis process generated growth

requirements for the resources: electrical power, crew, pressurized volume, attach

points, truss structure, and logistics. As an example, the growth power requirement

varied from a minimum of 125 kW for scenarios constrained to the conservative lift

capability to 325 kW for the microgravity and materials production with aggressive

transportation support.

The R&D configuration presented in this report best suited the microgravity and

materials production emphas_s. The growth resource requirements for this

configuration are sununarized in Table 2.1-1.

Power requirements for the R&D configuration will be accommodated by the

addition of solar dynamic units attached to the transverse boom. The additional

pressurized volume necessary to support the experiments and crew will be provided

through the addition of common modules, pocket laboratories, and extended resource

nodes. Additional payload attach points and servicing facilities will be accommodated

by the addition of booms and keels.

A detailed description, including the ph),sical characteristics, of the growth

elements and the distributed systems is contained m Section 4.0.

2-1
Table 2.1-1: MULTIDISCIPLINARY R&D GROWTH REQUIREMENTS

Baseline R&D Configuration

Average Power 75 kW 275 kW

Crew 8 24

Pressurized Modules

US Habitation 1 3

US Laboratory 1 3

ESA Laboratory 1 1

JEM Laboratory . 1 1

Pocket Laboratory N/A 3

Structures

Space Transfer Vehicle

Servicing Facility

N/A

APAE 2 18

Airlocks 1

S

Alternately referred to as "User Supplied Pressurized Modules"

2

2.2 CONFIGURATION DESCRIPTION

Figure 2.2-1 shows front and side views of the overall R&D reference

configuration. Attached to the lengthened transverse boom will be two full-size lab

modules (in addition to the three lab modules of the baseline configuration) and three

pocket labs. A pocket lab is a short, single-ended module attached to a resource node

port. The front view (Figure 2.2-1) shows the pocket labs attached to extended resource

nodes, which are standard station nodes with an extra cylindrical rack section to extend

their length. The pocket labs willprovide space for experiments requiring pressurized

volume. To accommodate the additional crew, the single hab module of the baseline

configuration will be supplemented by two additional hab modules. Figure 2.2-2

illustrates the top view of the module pattern. The additional lab and hab modules will

be connected to the existing modules by the extended resource nodes.

Vertical keels and upper and lower booms will be added to provide additional

locations for attached payloads and to accommodate the STV Servicing Facility. Four

sets of solar arrays and eight solar dynamic units, located on the transverse boom

outside the alpha joints, will provide up to 275 kW of power. As shown in Figure 2.2-1,

2-2
l-o

I

L,,o

STELLAR VIEWING

PAYLOADS

bU

POWER UNITS

H202 RCS

STV SERVICING

,. 255 m

VIEWING

PAYLOADS

Y (POP)

J

"" _ FRONT VIEW

Z (NADIR)

Figure 2.2-1 " Front and Side Views of R&D Configuration

RADIATORS

x (v)

Z (NADIR)

110 m

m

30 m

SIDE VIEW
hO

I

"OPTIONAL" EXTENDED

RESOURCE NODE

POCKET

(BOTTOM)

STARBOARD

KEEL

GROWTH MODULE _

(HAB 2)

I

GROWTH MODULE

(LAB 3)

POCKET LAB

(ON TOP)

POCKET LAB

(BOTTOM)

PORT

KEEL

GROWTH MODULE

(LAB 2)

Figure 2.2-2 •

R&D Configuration Module Pattern

TOP VIEW
the STV hangar for orbital transfer (geosynchronous) and vehicle storage and

maintenance will be h)cated on the lower boom.

2.3 R&D CONFIGURATION CONCEPT ANALYSES

The results of the analyses performed for the R&D configuration concept are

summarized in this section. Analyses and supporting data for the intermediate

configurations are given in section 5.0.

2.3.1 Mass Properti_

Analyses were performed on several cases of the R&D configuration to assess

the impacts resulting from changes in fuel storage and vehicle accommodations. Table

2.3.1-1 shows the mass properties for three R&D configuration cases involving potential

STV and shuttle arrangements. The origin of the geometrical coordinate system, shown

in Figure 2.2-1, is located in the middle of the center truss bay of the transverse t_oom.

The first row in Table. 2.3.1-1 reflects the configuration at its steady state heaviest with

full STV tank farms and an STV in the servicing hangar. The second row contains the

mass properties of the configuration assuming no fuel in the STV tank farms and no

STV in the hangar. The third row reflects the same configuration as the first row with

the addition of a berthed shuttle.

The center of gravity varies from 1.5 meters above the module pattern to two or

three meters below, indicating that good steady-state microgravity can be achieved.

The inertias yield gravity-gradient stability in pitch (IXX>IZZ) and small body-toprincipal axis rotations, i.e., flight attitudes will be near the local vertical/local

horizontal (LVLH).

2.3.2 Attitude Control Analysis

Steady-state attitude flight mode characterization an d control system sizing were

performed using Attitude Predict (ATTPRED) of IDEAS z'. The control momentum

gyroscope (CMG) controls analysis performed assumes only steady-state environmental

disturbances, such as aerodynamic and gravity-gradient torques. Large dynamic

disturbances, such as orbiter berthing, or STV deployment, and the corresponding

interactions between the CMGs, reactmn control system (RCS), and station attitude

will be evaluated in future studies.

Table 2.3.2-1 summarizes the microgravity, flight mode, and CMG control

characteristics of the R&D reference configuration for various combinations of

atmospheres, vehicle and fuel accommodations, and control system assumptions.

Maximum solar cycle equates to a worst-case (year 2001, flux=243, geomagnetic

index = 19.6), two-sigma atmosphere at an orbital altitude of 407 kilometers (220 nmi).

Minimum solar cycle equates to a best case (year 2007, flux=86.4, geomagnetic

index-- 16.0), nominal atmosphere at the same altitude. The gains chosen for the CMG

controller were optitnized to provide minimal deviations in pitch attitude ("pinched

pitch" method), except for certain cases where the gains" were adjusted" to emphasize. ,,

pitch momentum reduction at the cost of increased pitch deviations ("relaxed p_tch

method). All configurations studied yielded average flight attitudes within three

degrees of LVLH for roll, pitch, and yaw. Deviations over one orbit from the average

2-5
I

C_

Space Station

Configuration

Wet Tank Farm

and STV

Dry tank farm,

no STV

Wet tank farm and

STV, NSTS docked

Mass

(kg)

687,000

631,000

795,000

Center of Mass

(m)

X Y Z

Inertlas

(kg.mom)

IXX IYY IZZ IXY IXZ

-2.3 -0.2 7.0 1.22+E9 3.94+E8

-2.3 -0.2 3.8 1.12+E9 3.11+E8

0.8 0.2 8.2 1.24+E9 4.59+E8

9.21÷E8 2.16+_5 -9 02+E6

9.10+E8 2.10+E5 -9 61+E6

9.69+E8 7.41+E8 8.77+E6

IYZ

-3.77+E6

-4.29+E6

-1.01+E6

Table 2.3.1-1 • Mass Properties for R&D Configurations
I

Space Station

Configuration

Wet Tank Farm

and STV, with

,_,_v e_l_r ryrlP

Micro-G at

Lab Center

1.2

Wet tank farm and

STV, with max solar

cycle, relaxed pitch

Average Flight

Altitude (degrees)

Wet tank farm

and STV, with

min solar cycle

Dry tank farm,

no STV, with

max solar cycle

Dry tank farm and

STV, with max solar

cycle, relaxed pitch

Wet tank farm and

STV, STS attached,

with max solar cycle,

Wet tank farm and

STV, STS attached,

with max solar cycle

and relaxed

Yaw

(Z)

< 0.1

1.2 < 0.1

0.9 < 0.1

1.1 -0.2

1.1 -0.2

1.5 0.3

1.5 0.3

Pitch Roll

(Y) (X)

2.4 0.4

2.4

1.8

2.8

2.8

-1.0

-1.0

0.4

0.4

0.4

0.4

0.1

0.1

Altitude Drill

Pilch (Y)

per Orbil (degrees)

<0.1

1.4

< 0.10

0.2

1.5

0.2

1.2

Peak Momentum

(N.m-s)

10,000

4,000

8,000

9,000

4,000

10,000

5,000

Table 2.3.2-1 • Micro-G and Control Properties for R&D Configurations
flight attitude were less than +/- 0.1 degrees in roll and yaw and less than +/- 0.2

degrees in pitch for the pinched pitch control gains and less than +/- 1.5 degrees in

pitch for the relaxed pitch controlgains. Attitude error rates for all configurations and

all assumptions were below 0.005 degrees/second. The CMGpeak momentum control

requirements for maintaining attitude were all below 10,000 Newton-meter-seconds

(Nms) for the pinched pitch method and below 5,000 Nms for the relaxed pitch method.

All momentum requirements are well within the capability available from the six 4750

Nms CMGs on the baseline configuration. Figure 2.3.2-1 represents the flight attitude

history over five orbits for the R&D configuration with wet tank farms and an STV for:

1) a maximum solar cycle, pinched pitch case, 2) a minimum solar cycle, pinched pitch

case, and 3) a maximum solar cycle, relaxed pitch case. Figure 2.3.2-2 represents the

corresponding momentum control requirements over the same five orbit period. The

roll and yaw values shown on the figures are nearly identical for all cases with the only

differences noticeable in the pitch channel.

2.3.3 Microgravity Environment Dctcrminal;i0n

The steady-state microgravity profile shifts as the center of mass changes on the

station. This change in mass properues is primarily due to the presence or absence of

the STV and the amount of propellant in the STV tank farms. Fl_gures 2.3.3-1 and 2.3.3-

2 show the steady-state on and two micro-g regions for the dry R&D configuration (no

STV, with empty STV tank farms) and the wet R&D configuration (STV, with full STV

tank farms) in a worst-case 2-sigrna atmosphere (flux -- 243, geomagnetic index = 19.6)

at an altitude of 220 nmi. The highest center of mass location is obtained in the dry

R&D configuration, which provides the U.S. lab module a portion of the one micro-g

region. In this case, less than two micro-g's are sensed in the entire module pattern, as

seen in Figure 2.3.3-2. The size of these regions will increase if a less severe

atmosphere exists or a higher altitude is obtained. The lowest center of mass location is

found in the wet R&D configuration. This case shifts the one micro-g ellipse to the

lower portion of the U.S. lab module and maintains a less than two micro-g region for

the entire module pattern. During normal operations of the R&D station, the center of

mass will pass through the center of the module pattern as the STV tank farms go from

dry to wet and the STV goes on and off the station. Since the microgravity environment

is optimal at the station center of mass, it can be assumed that on v._gLgg¢ the sensed

microgravity environment in the modules will be better than shown in both Figures

2.3.3-1 and 2.3.3-2.

2.3.4 Orbit Lifetime and Rebo0st Analysis

Table 2.3.4-1 shows the reboost/orbit lifetime characteristics of the R&D

configuration. The two cases examined, with STV and fuel and without STV and fuel,

represent the potential maximum and minimum masses associated with the

configuration. Although one case is more massive than the other, there is no significant

difference in fuel requirements, because the more massive case has a higher ballistic

coefficient. However, almost an order of magnitude more fuel is required for altitude

maintenance during the solar cycle maximum, as compared to the solar cycle minimum

for the same altitude.

2-8
1"-3

I

_O

Euler Angle

(Deg.)

5.0

0.0

I I I I I I I I I I I I I I I I I I I I I I I I

Nominal Atmosphere

2-Sigma Atmosphere

2-Sigma Atmosphere (Pitch Attitude Relaxed)

Pitch

Roll

Yaw

-5.0

0

Figure 2.3.2-1 •

2 3 4

Orbit Number

R&D Configuration Attitude History Comparison

(Assumes Wet Tank Farms and Fueled STV)
cO

f_

0

a:

E

c_E

o z

2-10
/

/y' /

m/

/

/

/

i /I

//

-,3

/

Front View

Side View

Figure 2.3.3-1 • R&D Configuration Steady State Microgravity Profile

(Assumes Dry Tank Farms and No STV)

2-11
/

7

/J

/

/

/

/ Front View

/

J

/

/

"//

/

/

,/'I

Side View

ligure 2.3.3-2 ' R&D Configuration Steady State Microgravity Profile

(Assumes Wet Tank Farms a_ t Fueled STV)

2-12
I

Space Station

Configuration

Wet tank farm

and STV,

minimum solar cycle*

Dry tank farm

and no STV,

minimum solar cycle*

Ballistic

Coefficient

(kg/m 2 )

60.1

54.1

90 days

reboost

fuel (Ibm)

1,380

1,344

9,759

Wet tank farm

and STV, 60.1

maximum solar cycle**

Dry tank farm

and no STV, 54.1

m_ximum solar cycle**

*Minimum solar cycle (Flux=86.4, AP=16.0)

11,305

Reboost

Altitude

(nmi)

223

223

241

246

Lifetime

from

220 nmi

635

601

117

104

**Maximum solar cycle (Flux=243, AP=19.6)

Table 2.3.4-1 • Orbit Lifetime and Reboost, and Ballistic Properties for R&D Configurations
2.3.5 Structural _Dynamics Analysis

The typical approach for structural analysis of Space Station Freedom growth

configurations is outhned in section 5.0. No analysis has been performed on the specific

R&D configuration described in this document, but a few general comments may be

made based on previous analyses of similar configurations.

The major areas of concern for the R&D configuration are the deflection of the

transverse boom and loads on the alpha joint due to the large masses of the solar

dynamic modules, and the potential for buckling of members at the interface of the

boom and the upper and lower keels. For example, a study of a concept which had a

nearly identical lower keel configuration and similar module cluster mass showed that

during a typical reboost critical loads could be developed in truss members near where

the upper and lower keels met the transverse boom. It can be expected, since the

material properties, load paths, and mass distributions are similar on the R&D

configuration, that the potential exists for critical loads at similar locations. In general,

this is not an issue which would cause outright reiection of a conceptual design, because

slight modification to the diameter or the matenal properties of tlae struts or inclusion

of support structure can eliminate the potential failure. It is important during

conceptual design to identify those areas of concern to better plan the detail design

stage.

The transverse boom of the R&D configuration includes four solar dynamic

modules on each of its ends, with corresponding extensions to the structure. It has been

shown that the dynamic behavior of the boom during intense disturbances, such as

reboost or docking, is dominated by its overall length, the amount of mass at the ends,

and the stiffness of the alpha joint. Other studies, which investigated concepts with up

to two SD modules on each end, have consistently shown a fundamental boom bending

mode at approximately 0.07 Hz. With the additional mass and length, the R&D

station's boom configuration can be expected to have a corresponding mode at 0.06 Hz

or lower. The implication is that, as this fundamental mode frequency decreases, it

approaches the RCS jet..pulsing frequencies for reboost and the deadband frequencies

ot the control system. lhe potential control/structures interaction can result in buildup of member loads and potential failure of tlae structure. Thus, while four SD modules

on each end of the boom may be accommodated safely, analysts and designers should

be aware that structures of this sort will eventually reach a limit beyond which the

structure and control system cannot operate. In the same vein, previous work has

shown that pointing of the SD modules is sensitive to the fundamental frequency of the

transverse boom, such that, as the boom is lengthened and made more masswe near the

ends (i.e. as SD modules are added), the pointing capabilities are degraded somewhat.

Again, results indicate that four SD modules should not pose a significant problem, but,

if growth beyond this R&D concept should require additional power, alternative

approaches may be necessary.

2-14
3.0 TRANSPORTATION NODE CONFIGURATION CONCEPT

The baseline :_tation design can be augmented through evolution to provide the

scientific and technological foundation necessary to accommodate the Space

Exploration Initiative (SEI). These enhancements include additional pressurized

volume, crew, power, truss structure, and specialized facilities. For the transportation

node function, the growth station will support advanced technology development,

demonstration, and 'verification programs; exploration systems development; and the

assembly, servicing, and processing of exploration vehicles. It will also serve as a

laboratory to determine acceptable limits and/or countermeasures to long-term human

exposure to low grawty and space radiation environments and will serve as a testbed for

developing long-life, self-sufficient life support systems.

3.1 FUNCTION,Zd_, REQUIREMENTS

The facility and resource requirements necessary to support human exploration

missions are driven by a myriad of functions ranging from orbital technology

demonstration and verification of advanced space systems and precursor life sciences

research to the in-space assembly, servicing, processing, and refurbishment of the lunar

vehicles (LVs) and Mars Transfer Vehicles (MTVs).

Comprehensive analyses of the growth of the station as a transportation node

have been conducted during the past two years and are reported in the "Exploration

Studies Technical Report, FY 1988 Status, Volume 1: Technical Summary" and the

"Exploration Studies Technical Report, FY 1989 Status, Volumes I through VI".

Following President Bush's July 20, 1989, speech, a task team was created by the NASA

Administrator to conduct a study of the main elements of a Space Exploration

Initiative. The results of this study are reported in the "Report of the 90-Da), Study on

Human Exploration of the Moon and Mars". A summary of several key sections of the

90-Day Study Report is presented below and served as the basis for the functional

configuration requirements discussed in this section.

In the refere:aced study, station requirements were developed for two crucial

functions in the SEI. First, the station serves as an orbiting laboratory for conducting

the basic research and technology development necessary to implement the Initiative.

Research in the life. sciences will determine acceptable long-term microgravity and

space radiation limits and/or countermeasures andsupport the development of longhfe, self-sufficient life support systems. In the technology area, the station will serve as

a test-bed for the on.orbit verification and testing of umque lunar and Mars systems, as

well as for technology developments needed for in-space assembly and operations.

Second, the station serves as a transportation node for the assembly, test, processing,

servicing, recovery, and refurbishment of the lunar and Mars vehicles.

In addition, the station must support the R&D users, including NASA

experimenters, corarnercial users, umverslty investigators, and international

participants. In de,_eloping the Transportation Node evolution configurations, these

activities were included in the design to insure that the. multidisciplinary character of

the growth station would be .preserved to the maxamum extent possible. The

Transportation Node configurations presented in this section are responsive to these

stated requirements.

3-1
The Transportation Node configuration concept has two major phases:a lunar

operations configuration to support early lunar missions and a lunar/Mars

configuration. The lunar operationsconfiguration will also support the life scienceand

technology researchand development required for manned Mars missions. The lunar

operations configuration evolves to the lunar/Mars reference configuration by the

addition of a Mars Transfer Vehicle processingfacility attachedto the upper keels and

boom.

The growth resourcerequirementsof theseTransportation Node configurations

are summarized in Table 3.1-1. A detailed description, including the physical

characteristics, of the growth elements and the distributed systemsis contained in

Section4.0.

Table 3.1-1: TRANSPORTATION NODE GROWTH REQUIREMENTS

Baseline

Average Power 75 kW

Crew 8

Pressurized Modules

Transportation Node

Lunar Lunar/

Operations Mars

225 kW 225 kW

18 plus 23 plus

4 transient 4 transient

US Habitation 1 3 3

US Laboratory 1 1 1

ESA Laboratory 1 1 1

JEM Laboratory , 1 1 1

Pocket Laboratory N/A 3 3

Structures

Lunar Vehicle

Assembly/Servicing N/A

Facility

1 1

Mars Transfer Vehicle

Processing Facility

N/A N/A 1

APAE 2 TBD TBD

Airlocks 1 2 2

t

Alternately referred to as "User Supplied Pressurized Modules"

3-2
3.2 CONFIGURA'HON DESCRIPTION

As mentioned previously, two configurations were developed for the

Transportation Node function in support of the SEI. Figure 3.2-1 shows the front and

side views of the lunar operations configuration. The module cluster at the center of

the transverse boom will be expanded from the baseline configuration to include 2

additional habitation modules, 3 pocket labs, and 3 additional resource nodes. This

extra interior volume will house a permanent crew of 18, as well as 4 transient lunar

mission crew member:_, and will provide space for manned Mars precursor life sciences

experimentation and materials processing activities. Figure 3.2-2 illustrates the top

view of the module cluster.

Upper and lower booms and keels are added to the transverse boom to

accommodate the Lunar Vehicle Assembly/Servicing Facility and to provide additional

attached payload locations to support precursor and concurrent science activities. The

lunar facility, located on the lower keels and boom, is described in Section 4.0.

Capability to produce a total of 225 KW of power will be obtained by adding six 25 kW

solar dynamic (SD) power modules to the baseline 75 KW of photovoltaic (PV) power.

Additional structure will be included on each end of the transverse boom to

accommodate the SD modules, and corresponding levels of thermal rejection capacity

will be provided to keep pace with the increase in power.

Figure 3.2-3 shows the front and side views of the lunar/Mars node

configuration. The module pattern is unchanged from the lunar operations

configuration (Figure 3.2-2) and can accommodate the additional station crew reguirea

for lunar/Mars operations. The Mars Transfer Vehicle (MTV) Processing Facility is

attached to the upper keels and boom, and no provisions are made for

debris/meteoroid and thermal protection, since it is assumed that such protection will

be integral to the Mars vehicle.

3.3 TRANSPORTATION NODE CONFIGURATION CONCEPT ANALYSES

The results of the analyses performed for the Transportation Node

configurations are provided in this section.

3.3.1 Mass Propertk_

The mass properties of the Transportation Node configurations depend on the

presence and operational state (i.e., fueled or unfueled) of the lunar and Mars mission

vehicles. The Mars vehicle used in this analysis is the same as one of the vehicles

derived during the 90-day study (138,000 kg including 62,000 kg of fuel). The lunar

vehicles used in this analysis are also taken from the 90-day study, with a partiallyfueled lunar vehicle stack having a mass of 32,000 kg and a fully-fueled lunar vehicle

stack with a mass of 190,000 kg. Table 3.3.1-1 shows the mass properties of the

Transportation Node for several possible lunar and Mars vehicle configurations, as well

as one case with a shuttle attached. The origin of the geometrical coordinate system,

shown in Figure 3.2-1, is located in the middle of the center truss bay of the transverse

boom. The first row in Table 3.3.1-1 is the lunar operations configuration. The second

3-3
STELLAR VIEWING

PAYLOADS

SD

POWER UNITS

OMV ACCOMODATIONS

PV

----POWER UNITS

120 m

I

H202 RCS

CARGO BERTHING

ACCOMMODATIONS

Y (POP)

J

20 ml

- 255 m

LV SERVICING

EARTH VIEWING

PAYLOADS

RADIATORS

x (v)

FRONT VIEW

Z (NADIR)

Z (NADIR)

Figure 3.2-1 • Front and Side Views of Lunar Operations Transportation Node Configuration

SIDE VIEW
L.,J

I

POCKET LAB

STARBOARD_

KEEL _ ?

lw

I

POCKET LAB

(ON TOP)

POCKET LAB

I (BOTTOM)

"_ _l PORT

KEEL

GROWTH MODULE (HAB 3)

w-,,i

Figure 3.2-2 " Transportation Node Configuration Module Cluster

TOP VIEW _r_
I

STELLAR VIEWING

PAYLOADS MTV SERVICING

SD

POWER UNITS

OMV ACCOMODATIONS

PV

POWER UNITS

20 ml

H202

CARGO BERTHING

ACCOMMODATIONS

Y (POP)

z (NADIR)

255 m

FRONT VIEW

LV SERVICING

EARTH VIEWING

PAYLOADS

RADIATORS

x (v)

z (NADIR)

Figure 3.2-3 " Front and Side Views of Transportation Node Configuration

L30m

SIDEVIEW

120 m

m
I

Space Station

Configuration

Lunar operations

node, fueled LV

Fueled MTV,

part fueled LV

Fueled MTV,

fueled LV

Fueled MTV,

no LV

No vehicles

Mass

(kg)

734, """ U U II,.J

71 5,000

873,000

683,000

544,000

Fueled MTV,

part fueled LV, 823,000

and docked NSTS

Center of Mass

(m)

-1.8

-1.5

-1.8

-1.1

1.1

0.1 ).4

0.1 8.0

-0.1 2.6

-0.1 4.E

0.3 1 .;

Inertias

(kg-m-m)

5.52+E8 I 7.0/+tt_ I I.uL+co I -2 27+E6 I -2.36+E6

1.15+E9

1.44+E9

1.08+E9

9.37+E8

1.19+E9

5.21+E8

8.08+E8

4.47+E8

3.02+E8

6.02+E8

7.06+E8

7.11+E8

7.05+E8

7.01+E8

7.53+E8

1.45+E5

1.67+E5

1 46+E5

1.86++E5

6.94+E6

7.95+E6

1.64+E7

785+E6

-5.12+E6

4.07+E7

-330+E6

-2.83+E6

-3.43+E6

-3.01+E6

1.66+E6

Table 3.3.1-1: MASS PROPERTIES FOR THE

TRANSPORTATION NODE CONFIGURATIONS
row reflects the Transportation Node in the lunar/Mars configuration with a fueled

Mars. vehicle and a partially-fueled lunar vehicle. The third row represents the most

massive case, assuming a fueled Mars vehicle and a fully-fueled lunar vehicle stack.

Case four (fourth row) has the center of mass farthest above the modules, while case

one has its center of mass farthest below the modules. The fifth row corresoonds to a

case with no lunar or Mars vehicles, resulting in the lightest configuration. "rile last row

represents the mass properties of the lunar/Mars Transportation Node with a berthed

oroiter. It should be noted that the mass properties of the Transportation Node

configurations are highly dependent on the vehicle definitions derived during the 90-day

study. The center of gravity varies from 2.6 meters above the module pattern to 15

meters below. The inertias yield gravity-gradient stability in pitch (IXX>IZZ) and

small body-to-principal-axis rotations, i.e., flight attitudes will be near to LVLH.

3.3.2 Attitude Control Analysis

Steady-state attitude flight mode characterization and ct_ntrol system sizing were

performed using Attitude Predict (ATTPRED) of IDEAS". The CMG controls

analysis performed assumes only steady-state environmental disturbances, such as

aerodynamic and gravity-gradient torques. Large dynamic disturbances, such as the

orbiter berthing or lunar vehicle deployment, and the corresponding interactions

between the CMGs, RCS, and station attitude will be evaluated in future studies.

Table 3.3.2-1 summarizes the microgravity, flight mode, and CMG control

characteristics of the Transportation Node configurations for various combinations of

atmospheres, vehicle and fuel accommodations, and control system assumptions.

Maximum solar cycle equates to a worst case (year 2001, flux=243, geomagnetic

index-- 19.6), two-sigma atmosphere at an orbital altitude of 407 kilometers (220 nmi).

Minimum solar cycle equates to a best case (year 2007, flux=86.4, geomagnetic

index = 16.0), nominal atmosphere at the same altitude. The gains chosen for the CMG

controller were optimized to provide minimal deviations in pitch attitude ("pinched

i h" .... p tc method), except for certain cases where the gmns were adjusted to emphasize

pitch momentum reduction at the cost of increased pitch deviations ("relaxed pitch"

method). All cases studied yielded average flight attitudes within two degrees of LVLH

for roll, pitch, and yaw, except for the berthed orbiter case (pitch = -5.3 degrees).

Deviations over one orbit from the average flight attitude were less than +/-0.2 degrees

in roll and yaw and less than +f-0.2 degrees in pitch for the pinched pitch control gains

and less than +/- 1.1 degrees tor the relaxed pitch control gains. Attitude error rates

for all cases and all assumptions were below 0.005 degrees/second. The CMG peak

momentum control requirements for maintaining attitude were all below 15,000 Nms

for the pinched pitch method and below 4,500 Nms for the relaxed pitch method. All

_,ntum requirements are well within the capability available from the six 4,750 Nms

t.is on the baseline station. Figure 3.3.2-1 represents the flight attitude history over

five orbits for the lunar/Mars Transportation Node with a fueled Mars vehicle and

partially fueled lunar vehicle stack for: 1) a maximum solar cycle, pinched pitch case, 2)

a minimum solar cycle, pinched pitch case, and 3) a maximum solar cycle, relaxed pitch

case. Figure 3.3.2-2 represents the corresponding momentum control requirements

over the same five orbit period. The roll and yaw values displayed in the figures are

nearly identical for all cases, with the only major differences noticeable in the pitch

channel.

3-8
Space Station

Configuration

Lunar operations node,

fueled LV, max

solar cycle, relaxed pitch

Lunar operations node,

fueled LV,

max solar cycle

==..._.,-.====.=

Fueled MTV, part fueled

LV, max solar cycle

Fueled MTV, part fueled

LV, rain solar cycle

Fueled MTV, partfueled

LV, max solar cycle

& relaxed pitch

Fueled MTV, full LV,

max solar cycle

Fueled MTV, full LV, max

solar cycle, relaxed pitch

Fueled MTV, no LV,

max solar cycle

No vehicles,

max solar cycle

Fueled MTV, part fueled

LV, docked NSTS, max

solar cycle

Micro-G at

Lab Center

4.3

4.3

2.6

2.5

2.6

1.4

1.4

3.5

1.0

Average

Altitude

Yaw

(2:)

Flight

0.0

0.0

0.0

-0.1

0.0

0.0

0.0

0.0

0.0

(degrees)

Pitch Roll

(y) (x)

1.5 0.9

1.5 0.9

-1.2 1.0

-1.1 1.0

-1.2 1.0

-0.7 -1.7

-0.7 -1.7

Attitude

per

Yaw

(z)

<0.1

<0.1

<0.1

<0.1

<0.1

<0.1

<0.1

Drift

Orbit (degrees)

Pitch Roll

(Y) (X)

1.1 0.2

0.2 0.2

< 0.1 < 0.1

< 0.1 < 0.1

1.0 < 0.1

< 0.1 0.2

0.6 0.2

-1.8

1.4

0.8 < 0.1

0.4 < 0.1

-0.9 < 0.1

<0.1

0.1

<0.1

<0.1

<0.1

<0.1

Table 3.3.2-1: MICRO-G AND CONTROL PROPERTIES FOR THE

TRANSPORTATION NODE CONFIGURATIONS

Peak Momentum

(N-m-s)

3,000

15,000

10,000

6,000

1 ,000

8,000

4,500

11,000

10,000

10,000
5.0 I I I I I I I I I I I I I I I I i I i [ i I I I

I

0

Euler Angle

(Deg.) 0.0

-5.0

I t ",

D

a

I

_J

I k • I . L t I I

'l a t ,I ,..t "1

L J I

0

Nominal Atmosphere

2-Sigma Atmosphere

.......... 2-Sigma Atmosphere (Pitch Attitude Relaxed)

I I I

I I I I I I I I I I I I I I I . I 1

1 2 3 4

1 I ,I I

5

Figure 3.3.2- l •

Orbit Number

Lunar/Mars Transportation Node Configuration Attitude History Comparison

(Assumes Fueled Mars Vehicle and Partially Fueled Lunar Vehicle)

Roll

Yaw

Pitch
000

n

0

rr

o

E

ooE°z'

3-11

0

000

u4I

>-,"5

E_

E_

e-

_>

z z__

.1:=__

_ ._ o _

L- •

°

.--
3.3.3 Microgravity Envir0nmcnt Dctermin_ion

The steady-state microgravity environment changes with each operational

configuration. Microgravity profiles have been generated for a subset of the analysis

cases in order to show the wide range of possible environments encountered with the

Transportation Node configurations. These cases include the nominal lunar/Mars

operational configuration, the optimal microgravity case, and the worst-case

microgravity configuration. Since the microgravity environment is lowest at the center

of mass, the analysis cases with .the highest and lowest center of mass were selected.

Figures 3.3.3-1 through 3.3.3-4 show the one and two micro-g regions for the various

operational cases at an altitude of 220 nmi and a worst-case, 2-sigma atmosphere (flux

= 243, geomagnetic index = 19.6). Note that a less severe atmosphere or higher

altitude will decrease the sensed acceleration by as much a 0.3 micro-g.

The nominal configuration (fueled Mars vehicle, partly-fueled lunar vehicle) in

Figure 3.3.3-1 shows that the modules are not in the one tmcro-g region, but a small

portion of the modules is in the two micro-g region. Figure 3.3.3-2 displays the optimal

microgravity configuration (no vehicles), where the U.$. lab module its almost entirely

encompassed by the one micro-g region, and the entire module pattern is within the two

micro-g region. Notice that, for this case, the center of mass is located near the center

of the module pattern, producing an ideal environment for microgravity. Figures 3.3.3-

3 and 3.3.3-4 (fueled Mars vehicle with no lunar vehicles and lunar operations node

with fueled lunar vehicles) show that the modules fall completely outside the one and

two micro-g regions. For the other analysis cases, the microgravity envelopes are

between the upper and lower bound regions shown in Figures 3.3.3-3 and 3.3.3-4.

The lunar operations case with fueled lunar vehicles provides the worst

microgravity environment to the laboratory modules (i.e., the center of mass is farthest

from the module pattern center line). It is important to note that there are

opportunities to perform microgravity-sensitive experiments on the Transportation

_odes during the quiescent periods. Figure 3.3.3-5 shows the steady-state microgravity

levels sensed at the geometric center of the U.S. lab module on the lunar operations

transportation node during the lunar vehicle processing phase. The first six months are

scheduled for preparing the lunar vehicles (LVs). In this phase, the laboratory

microgravity levels increase from under one micro-g to just over four micro-g.

However, subsequent to LV departure, the levels return to acceptable limits and

sensitive R&D experiments can be performed. At a rate of one LV flight per year,

approximately six months of "uninterrupted" low gravity research time is available to

users.

It should also be noted that for the evolution path discussed here, there are

several years between completion of the baseline configuration and the start of lunar

operations, where "undisturbed" gravity-sensitive R&D can be conducted. Also, several

years of lunar operations occur before both the lunar and Mars vehicle operations on

station begin.

3.3.4 Orbit Lifetime and Reboost Analysis

Table 3.3.4-1 shows the reboost/orbit lifetime characteristics of the

Transportation Node configurations. Four cases were examined: 1) lunar operations

3-12
Front View

3

/

=== _ /

Side View

It

\, [

\

Figure 3.3.3-1 ' Lunar/Ma:rs Transportation Node Configuration Steady State Microgravity Profile

(Assumes Fueled Mars Vehicle and _rtially Fueled Lunar Vehicle)

3-13
/

Side View =

Figure 3.3.3-2 : Lunar/Mars Transportation Node Configuration Steady State Microgravity Profile

(No Vehicles)

3-14
Front View

.r...,

T

/

J,,//

/

/

/

__Z __.___'

Side View

i

V

Figure 3.3.3-3 • Lunar/Mars Transportation Node Configuration Steady State Microgravity Profile

(Assumes Fueled Mars Vehicle and No Lunar Vehicle)

3-15
LL,

:::L

0

o

_>

_o_

0

oo

..._:_

I
5_

I

:::L

:>

L--

0

i_

0

¢n

"0

(D

03

Lunar

Vehicle

1 U

Lunar Vehicle

departs

Lunar Vehicle

irns

1 2 3 4 5 6 7 8 9 10 11 12

TIME (months)

" Izg level at geometric center of U.S. Lab; levels at potential equipment locations may vary by as much as 0.5 I_g

Figure 3.3.3-5 • Lunar Operations Transportation Node Configuration Time Phased Steady

State Mierogravity Levels
I

_o

Space Station

Configuration

Lunar operations node, fueled

LV, minimum solar cycle*

Lunar operations node, fueled

LV, maximum solar cycle**

Fueled MTV, partly fueled LV,

minimum solar cycle*

Fueled MTV, partly fueled LV

maximum solar cycle**

Fueled MTV, fueled LV

minimum solar cycle*

Fueled MTV, fueled LV

maximum solar cycle**

No vehicles,

minimum solar cycle*

Ballistic

Coefficient

(kg/m 2 )

68.8

68.8

60.7

60.7

74.2

74.2

51 o0

NO vehicles 51.0

maximum solar cycle**

*Minimum solar cycle (Flux=86.4, AP=16.0)

90 days

reboost

fuel (Ibm)

1,253

11,000

1,376

11,640

1,354

11,999

1,265

10,106

Reboost

Altitude

(nmi)

223

242

223

244

222

240

223

248

Lifetime

from

220 nmi

684

133

636

117

710

139

581

101

**Maximum solar cycle (Flux=243, AP=19.6)

Table 3.3.4-1" ORBIT LIFETIME AND REBOOST, AND BALLISTIC PROPERTIES FOR

THE TRANSPORTATION NODE CONFIGURATIONS
with fueled LVs, 2) fueled Mars vehicle and partially-fueled lunar vehicle stack, 3)

fueled Mars vehicle and fueled lunar vehicle stack, and 4) no vehicles. Both the solar

cycle maximum and solar cycle minimum cases were analyzed for each operational

configuration of the transportation Node. Almost an order of magnitude more fuel is

required for altitude maintenance during the solar cycle maximum, as compared to the

solar cycle minimum for the same altitude. Although the configurations differ in mass

in some cases by ove_" 50%, there is no significant difference in fuel requirements due to

station mass, because the more massive configurations have higher ballistic coefficients.

3.3.5 Structural D_lamics Analysis

No detailed structural analyses of the specific Transportation Node

configurations have been performed for inclusion in this document. However, a recent

study investigated the response to reboost loading of a node concept which had

essentially the same transverse boom configuration as the lunar operations concept, but

a somewhat altered mass distribution on the upper and lower keels. A second study

evaluated a concept whose only substantial differences from the lunar concept was one

less habitation module and no upper boom. These concepts exhibited 126 and 113

modes respectively below 2 Hz and showed very good correlation in both shape and

frequency for transw_rse boom and appendage modes. For example, in both cases the

fundamental boom bending mode, which was the mode that most dominated the

pointing response of the solar dynamic modules, had a frequency of approximately 0.07

Hz. Each concept raaintained solar dynamic pointing accuracy to w_thin 0.05 degrees

(the requirement is 0.1 degrees), and experienced only a few inches of motion at the

tips of the PV arrays and the end of the transverse boom. It is expected that the

Transportation Node configurations will exhibit similarly high modal density, with

comparable mode shapes, frequencies, and response levels.

The Transportation Node configurations have an additional solar dynamic

module on each tip of the transverse boom. The effects of this additional mass will be

to lower the frequencies of the transverse boom modes (but not significantly alter their

shapes) and to increase the displacements of the transverse boom tips and the tips of

the PV arrays. It may also somewhat degrade the pointing accuracy of the solar

dynamic modules.

Both of the ..studies mentioned above showed some buckling failure of truss

members near the interface of the keels and the transverse boom during reboost. This

is not surprising since models of the baseline station also indicate failure of truss

members during reboost and NSTS docking. The point is that for any such large

structure, there will be members that will have to be stiffened to avoid failure. The

transportation node concepts described here will certainly require either stronger

materials, larger diameter struts, or additional support structure at the interface of the

boom and keels. Also, depending on the effects otthe added solar dynamic mass, it is

possible that the tru,ls outside the alpha joint, or the joint itself, will require additional

stiffness. Detailed analysis of these specific concepts will determine precisely which

members are subject to buckling loads.

3-19

4.0 PHYSICAL C HARACTERISTICS OF ELEMENTS AND SYSTEMS

The evolutionary space station will be composed of elements and distributed

systems. Elements are major pieces of self-contained hardware such as the pressurized

modules. They are separated into two categories: replicated baseline elements and

growth elements. Rcphcated elements are those that are identical to baseline elements

and are duplicated ia the evolution station. Growth element categories include those

common to both evclution, I configurations, elements unique to the R&D configuration,

and elements unique to the Transportation Node confi[[u.rations. The distributed

systems will be distributed throughout the station to provide resources to the station

elements and users. This section contains physical and functional descriptions of the

growth elements and the distributed systems for the evolution configurations.

4.1 REPLICATED BASELINE ELEMENTS

For station giowth, it is desirable to replicate as much of the existing hardware

as possible to reduce costs and unique spares. Several elements that will be added to

the reference configttrations will be nearly identical to baseline elements. For example,

the R&D configuration will have two additional lab modules which are identical in

design and fabrication, but not outfitting, to the initial U.S. lab module.

II

Table 4.1-1: REPLICATED BASELINE ELEMENTS

ITEM BASELINE R&D STA.

U.S. Laboratory 1 3

Hab Module 1 3

Resource Node 4 4

Airlocks (H = Hyperbaric) 1H 1 + 1H

Cupola 2 2

NSTS Docking Mast 2 2

APAE 2 18

Mobile Transporter 1 1

MSC 1 1

MMD - MSC Maint. Depot 1 1

FTS 1 1

PLC - Press. Log. Carrier 1 1

ULC - Unpress. Log Carrier 3 TBD

Truss Structure (5m Bays) 27 107

S

Quantities are totals, not the number replicated.

TRAN$,NODE

Lunar Lunar/

Ops Mars

1 1

3 3

4 4

I+H 1+1H

2 2

2 2

TBD TBD

1 2

1 2

1 1

1 1

1 1

TBD TBD

110 118

4-1
Table 4.1-1 provides a list of the replicated baseline elements, along with the

quantity that will be incorporated into each configuration. Descriptions and physical

characteristics for these and other baseline elements can be found in the Preliminary

Space Station Freedom Technical Overview dated October 1988 (no document number

assigned and not presently under configuration control) and the Weights White Paper:

$$PQ SE&I, Weight Allocal;ion. Control Weights. and Associated Matrices with

W¢igh_ Database.'Rev. 3.0. January 1989 (SSE-EZ88-R3.0).

4.2 GROWTH ELEMENTS

Elements which are new or enhancements of existing designs are considered

growth elements. Some of the growth elements are common to all configurations, while

some are unique to individual concepts. This section presents functional and physical

descriptions of the growth elements m the following order: elements common to both

configurations, R&D-unique elements, and Transportation Node-unique elements.

Table 4.2-1 provides a list of the manned-base growth elements along with the quantity

that will be incorporated with each configuration. Evolution transportation elements,

co-orbiting platforms, and co-orbiting facilities such as a fuel depot are not addressed in

this report.

Table 4.2-1: STATION GROWTH ELEMENTS

ITEM R&D STA. TRANS,NQDE

Solar Dynamic Power Module

Resource Node - Extended

Pocket Lab

STV Servicing Facility

LV Assembly/Servicing Facility

MTV Processing Facility

Lunar Lunar/

Ops Mars

8 6 6

5 4 4

3 3 3

1

- 1 1

- - 1

4.2.1 Solar Dynamic (SD) Power Systems

An SD power module provides up to 25 kW of power by transforming

concentrated solar energy into electrical power. The collector in the SD power system

concentrates solar energy into a receiver filled with Helium-Xenon gas. The excited gas

drives a thermodynamic, Brayton-cycle turbomachinery to generate electricity. To

maximize the solar impingement on the concentrator, the SD makes use of a solartracking pointing system. The SD stores energy in thermal energy salts (TES), thus

eliminating the need for batteries during orbital eclipse. An electrical equipment

assembly controls module operation and conditions the alternator power to the

4-2
required transmissionfrequency and voltage. Both the R&D and the Transportation

Node configuration conceptswill require the addition of SD modules to the baseline

station.

An SD module is shownin Figure 4.2.1-1.The concentrator is a 18.2m (60 ft)

diameter concavemirror composedof severalone-meter,spherically-contouredmirror

facets. At the focal point of the concentrator is the receiver assemblyand the power

conversionunit. The SD assemblyis mounted on a beta-joint gimbal assembly,which in

turn is mounted on the transverseboom extension. The gimbal assemblypoints the

concentratorat the sunand transferspower to the power managementand d_stribution

system(PMAD). Aft of the concentratoris the heatrejection assembly,which is slightly

under 10m (33 ft) long and 9 m (29.5ft) wide. This assemblyradiates excessthermal

energyto spaceto maintain necessarypower levelsand operating temperatures.

4.2.2 Truss Structur,_

4.2.2.1 Booms and Keels

Booms and ke.els are added to the baseline station to provide greater stability

and additional locations for items such as attached payloads, semcing bays, and

antennas. The addition of booms and keels will also provide accommodations for

hangars and construction facilities.

Booms and keels are the same type of erectable truss bays that comprise the

baseline transverse boom. Each truss bay is a 5 m (197 in) cube with a mass of 274 kg

(603 Ibm). Running within the length of the truss bay is a utility tray which provides the

wiring and plumbing used by the resource distribution systems, e.g., power, thermal

rejection, and data communication systems. The utility trays have a 0.15 m (6 in) height

by 1.5-2.3 m (60-90 in) width cross-section.

4.2.2.2 SD Truss Extension

Additional tntss structures or extensions are added for all configurations to

support and provide proper clearance for the SD modules. The truss extensions will be

located at either end of the transverse boom. These truss extensions will house utility

trays that contain coramand and data lines and sufficient cabling for 25 kW of electrical

power per module.

Like the booms and keels, the SD truss extensions will be constructed of 5 m

(197 in) cube bays. Each SD truss extension bay will have a mass of about 189 kg (416

Ibm).

4.2.3 Extended Resource Node

The baseline interconnect for laboratory and habitation modules is the resource

node. In addition to providing a shirt-sleeve environment for travel between modules,

the resource node is :s_zed to allow accommodation of much of the space station system

hardware. All configurations will require additional resource nodes; however, they will

be different from the baseline station nodes.

The clearance between modules connected by resource nodes are determined by

the distance between axial berthing ports and the location of the radial ports along this

4-3
..L",

I

.p,.

RADIATOR q_

ASSEMBLY

._- CONCENTRATOR

ASSEMBLY

/

_ i .,' _-RECEIVER ASSEMBLY

CABLE __ _ _ __--_POWER CONVERSION UNIT ASSEMBLY

6----q___ _ ",, STRUCTURE ASSEMBLY

I _ _'-__P'_'_SD ELECTRONICS

'_-..__I___-----_L_ _ I _ _ ENCLOSURE

_ _ I "-BETA JOINT/

L. TRANSVERSE_J___,_' ROLL RING

BOOM EXTENSION

Figure 4.2.1-1 • Solar Dynamic Module Components
axis. Since the location of the berthing ports in the baseline design is not in the

geometric center of the resourcenode length, a resourcenode hasan "extended"side

and a short side. The baselinerequirement for moduleclearanceof approximately2 m

(79in) is achievedbi¢interconnectingwith two resourcenodesattachedwith extended

side to extendedside. Growth resourcenodesin the sameplane will have to attach to

the short side of a baseline resource node, resulting m a reduced clearance

(approximately1me':er)betweenbaselinemodulesand growth modules. Attaching the

short endsof two resourcenodesis not viable becausecontact betweenthe modules

would occur. The 2 m (79in) clearancerequirement will not allow the useof baseline

nodesfor srowth module interconnects. Severalsolutionshave been considered;the

most promisingis to lengthenthe cylindrical sectionof the baselineresourcenode with

the addition ofa cyli_adricalrack section. SeeFigure4.2.3-1.

The resource nodeswill be extendedversionsof the baseline resourcenodes.

The overall length will be about 6.1meters(254 in), about 1 m (42 in) longer than the

baselinenode.The extendednodewill be4.4 m (174in) in diameter andhavea massof

7432kg (16,380Ibm).

4.2.4 P0¢ket Laborato_!9__

All growth configuration concepts have pocket labs. A pocket lab is a short,

single-ended module (laboratory) that can be located at many.spare resource node

ports. Uses for pocket labs have been identified by the life sciences community for

Space Exploration hitiative support and by the microgravity and materials processing

community for commercial materials processing.

The pocket labs willprovide physical separation from the common modules in

order to promote safety andisolation of environmental contamination sources. They

can be used to house equipment not readily adaptable to the common module due to

assembly constraint.,, or configuration incompaubilities. Pocket labs are also readily

detachable for possible return to Earth. Lastly, they provide incentive to the private

sector to participate by providing a standard interface and attach mechanism to the

space stauon. Three pocket labs will be added to each configuration. The outfitting of

the pocket labs will be determined by future studies.

Pocket labs have a number of potential configurations. The most likely, of these

are ones derived from elements already being fabricated for the baseline station, such

as the common module (lab and hab), the baseline resource node, or the pressurized

logistics module. Tl_._ese concepts will reduce costs by using existing designs and tooling.

The proposed applications indicate that a pocket lab should have a volume of between

20 and 40 cubic meters and power/heat rejection of up to 25 kW. Temperature

requirements will range from 3 degrees C (minimum) to 30 degrees C. Other resources

that may be required include data links, video, waste gas removal, and crew time.

These labs will, of course, have a shirt sleeve environment for crew operations.

If a common module design is used, a pocket lab would be 4.3 m (14.2 ft) in

diameter. The estimated mass of pocket labs of differing lengths is 3,292 kg (7256 Ibm)

for the two end cones and one interface adapter, plus 3704 kg/m (2475 lbm/ft) of

cylindrical section length. A sample pocket lab configuration having a length of 3.7 m

(12.2 ft), a diameter of 4.3 m (14.2 ft), and a cylindrical section of 1.9 m (6.3 ft), would

have a mass of 10,2.33 kg (22,554 Ibm). Figure 4.2.4-1 illustrates a typical pocket lab

and gives its dimensions, mass, moments of inertia, and projected areas.

4-5
_---- 6.1 M 1

EXTENSION X

X

MASS = 7432 KG

INERTIAS:

Ixx = 45103 KG*M**2

lyy = 30295 KG*M**2

Izz = 45103 KG*M**2

PROJECTED AREAS:

PAx = 26.8M*'2

PAy = 15.2M*'2

PAz = 26.8M*'2

Figure 4.2.3-1 ' Extended Resource Node

4-6
Z

Y

4.3 M

MASS = 10233 KG

INERTIAS:

Ixx = 8139 KG*M* *2

lyy = 8164 KG*M* *2

Izz = 7722 KG*M**2

PROJECTED AREAS:

PAx = 15.9 M**2

PAy = 15.9 M**2

PAz = 14.5 M**2

Figure 4.2.4-1 " Typical Pocket Lab

/4-7
4.2.5 Space Transfer Vehicle (STV) Servicing Facility

The STV Servicing Facility is an unpressurized enclosed structure on the R&D

configuration that provides a thermally controlled and meteorite protected volume for

the STV and payload servicing, as well as for orbital replacement unit (ORU) and

orbital servicing equipment (OSE) storage. The hangar is equipped with a robotic

manipulator and EVA support equipment to accommodate STV processing, payload

mating, and hangar maintenance. Interfaces to STV propellants andSTV payload fluid

services are provided. Electrical/data umbilicals, along with thermal control and

structural support, are also available.

The STV enclosure is 15.3 m by 15.3 m by 29.3 m (50 ft by 50 ft by 97 ft). Inside

the facility is a dedicated manipulator and a service track assembly. The bay has a mass

of 19,353 kg (42,650 Ibm) (see Figure 4.2.5-1) and will accommodate storage of a spare

aeroshell.

4.2.6 Lunar Vehicle (LV) Assembly/Servicing Facility.

The LV Assembly/Servicing Facility is made up of a number of distinct pieces of

hardware which provide the capability to perform the assembly, integration, checkout,

and refurbishment of an LV. These capabilities include power and data systems,

robotic manipulators, thermal control, debris/meteoroid protection, EVA support

systems, and structural supports. The facility includes a large retractable, unpressunzed

hangar-like structure attached to the lower boom between the lower keels of the

Transportation Node. Within this hangar, the Service Track Assembly (STA) provides

vehicle interface attachments and assembly fixtures, storage areas, and guide rails for

two mobile manipulators. With the STA and the mobile robotic arms, the LV can be

accessed for assembly and checkout, ORU changeout, payload integration, and vehicle

refurbishment and refueling.

The LV Assembly/ServicingFacility enclosure (shown in Figure 4.2.6-1) is a 30 x

20 x 20 meter rectangular shield whose sides can be retracted along its centerline when

clear access is needed for manipulator arm operations or vehicle ingress or egress. The

front wall is a garage-type door which opens for smaller hardware transfers. The

enclosure includes debris protection shieldmg consisting of a main aluminum alloy wall

and a thinner alloy bumper, separated by a few centimeters of Multi-Layer Insulation

(MLI).

The STA makes up the entire floor of the enclosed area. It is 17 meters wide

and 30 meters Ion 8 when fully deployed, and is composed of LV interface and attach

fixtures, the servicing base, a strongback and utility trough, and guide rails for robotic

access. Storage lockers and support equipment are located along the top of the trough

in areas not occupied by the vehicle attachment hardware.

There are three vehicle interface and attachment fixtures provided in the facility.

These fixtures serve as gantries which disconnect and swing to one side to allow the

vehicle to enter or leave the facility. Two of the fixtures are dedicated to LV

processing, while the third is used only for aerobrake assembly and attachment. Two

duplicate mobile manipulators run on parallel tracks on each side of the LV stack.

These manipulators consist of a mobile base and a multi-segment arm derived from the

4-8
29.3 M

Z

x__l

X Z

_ 15.3 M_

15.3 M

Y

x_.l

MASS = 19353 KG

INERTIAS:

Ixx = 1.958E6 KG* M* * 2

lyy = 1.958E6 KG*M**2

Izz = 1.447E6 KG* M* * 2

PROJECTED AREAS:

PAx = 448M*'2

PAy = 448M*'2

PAz = 234M*'2

Figure 4.2.5-1 ' STV Servicing Facility

4-9
|

ENCLOSURE

INTERFACE &

ATTACH FIXTURE

IMOBILE / _

MANIPULATOR

s

o _

/

/

MOBILE

MANIPULATOR

INTERFACE &

". ATTACH FIXTURE

f"

LUNAR VEHICLE

STACK (DASHED)

AEROBRAKE

ASSEMBLY &

ATTACH

FURE

MANIPULATOR

GUIDE RAIL

MANIPULAT(

GUIDE RAIL

STRONGBACK &

UTILITY TROUGH

SERVICING

BASE

Figure 4.2.6-1 • Lunar Vehicle Assembly/Servicing Facility (ASF)
SpaceStation Freedom remote manipulator arm. Two stationary EVA work stations

are also included to assist in vehicle processing.

4.2.7 Mar_ Transfer Vehicle (MTV) Pr0¢¢ssing Facility

Like the LV Assembly/Servicing Facility, the MTV Processing Facility consists

of a number of hardware elements which make up a set of on-orbit capabilities that

enable vehicle processing. In this case, it is expected that the vehicle will provide its

own thermal control and collision protection. Therefore, there is no hangar to provide

debris/meteoroid and thermal protection. Instead, the MTV Processing Facility is

limited to power and data systems, robotic manipulators, EVA support systems, and

structural supports.

The MTV Processing Facility concept is less well defined than that of the LV

Assembly/Servicing Facility. They will both provide many of the same functions;

however, the physic_d details of the MTV Processing Facility have not been defined in

detail. Several bays of truss and support structure will be added to allow the MTV to

attach between the upper keels of the Transportation Node where robotic manipulators

can access the vehicle for assembly, checkout, and refurbishment. The size and number

of manipulators will depend on the length and diameter of the MTV and the

operational details c f vehicle processing.

4.3 DISTRIBUTED SYSTEMS

Unlike elements, distributed systems are not self-contained but are distributed

throughout the space station to provide resources to the station elements and users.

Common user resources are power, active thermal control, data, communication and

tracking, fluids, and environmental control and life support. Also classified as

distributed systems are EVA, crew support, station propulsion, and station guidance,

navigation, and control. This section describes distributed systems for the baseline and

rowth configurations. Descriptions of the growth distributed systems are presently

igh level and will be defined in more detail as the system designs mature and specific

system evolution studies are completed.

4.3.1 Electrical Power System (EPS)

The EPS provides distributed power throughout the space station for housekeeping and customer use. It consists of three major subsystems: the power generator

subsystem, the energy storage subsystem, and the power management and distribution

(PMAD) subsystem.

4.3.1.1 Baseline

The baseline EPS provides 75 kW. It uses four photovoltaic (PV) modules

located on the transverse boom, outboard of the truss element alpha gimbals, to convert

solar into electrical energy. A PV module consists of a pair of PV arrays which are

made up of two solar arrays that are 29.1 m (96 ft) long and 4.37 m (14.4 ft) wide. Each

array is attached to the truss by a rotating beta joint.

4-11
The PV arrays operate in a temperature range from 160 ° to -85 ° Fahrenheit.

Array voltage is controlled by a sequential shunt umt (SSU). Under rapidly changing

load conditions, the SSU controls the voltage by shunting a small portion of the power

and converting it to heat.

Throughout every orbit, the solar arrays must maintain a sun-facing orientation.

The main truss structure alpha joints, which can rotate 360 °, continuously adjust the

orientation of the arrays. To provide seasonal tracking of the sun as it moves from the

north sky to the south sky, the beta joint can rotate up to 55 degrees.

Since all electrical energy on the station comes from the Sun, energy must be

stored to provide power during the periods of orbital eclipse, which occur for about 30

minutes out of every 90-minute orbit. The baseline energy storage subsystem consists of

20 high-efficiency, lightweight, nickel-hydrogen batteries. The batteries supply the same

amount of power during orbital eclipse as that provided by the PV during the sunlit

portion of the orbit.

The power management and distribution (PMAD) subsystem distributes and

conditions electrical power throughout the station elements. The PMAD subsystem

monitors and controls the functions of the EPS and protects the system from damage by

failed components, short circuits, and operator error. Since the PV modules and the

energy storage subsystem provide dc power at the 160Vdc level, the PMAD subsystem

is also responsible for the dc/dc conversion to 120 Vdc for secondary distribution

throughout the station.

4.3.1.2 Growth

The addition of the SD power modules (Section 4.2.1) will constitute the primary

growth path for the EPS. Each alpha joint must either be initially sized to allow

transfer of peak power, or be designed to facilitate future upgrades. The baseline beta

joint used with the PV arrays can be replicated for use with the solar dynamic power

modules. Additional distribution capacity will be provided by replicated utility trays.

4.3.2 Thermal Control System (TCS)

By a combination of active and passive means the TCS maintains the station's

elements, systems, and payloads within allowable temperature ranges. Waste heat is

collected by the TCS active subsystems and transported to a location on the station

where it is radiated into space. The TCS also uses passive thermal control techniques

to insulate most station elements from the cyclic effects of solar inputs and radiative

losses.

4.3.2.1 Baseline

The baseline TCS is sized to accommodate an 82.2 kW heat load which is

comprised of the 75 kW input of the EPS plus 8% power conversion losses and a 1.2

kW metabolic load. The active thermal control system (ATCS) consists of an external

ATCS, an internal ATCS (ITCS), and a solar power module ATCS (PV ATCS) for the

EPS outboard of the alpha joints.

The central heat collection and transportation function is accomplished with a

redundant external pumped two-phase system that uses ammonia as the working fluid.

4-12
Two separate external thermal loops operate at 2°C (35°F) and 21°C (70°F),

respectively. Wasteheatis collectedfrom the pressurizedmodulesusinginterface heat

exchangersand cold platesand transportedby the thermal busesto two radiator panels

located on the transverse boom. Modular condensers in the radiator assemblies

transfer the heat to erectableheat-pipe radiators. The 15.2m (50 ft) long by (31 ft)

wide radiatorsaremountedon anti-Suntracking rotary fluid couplers.

The ITCS is used within the modules and nodes for temperature control of

systemcomponentsand user experiments. Single-phasewater (pumped through the

ITCS loops) acquiresthe heat, holds it during transfer, and rejects it to the external

two-phasebuses via interface heat exchangers. Each of the four deployable PV

modules has an independent active thermal control system consisting of heat

acquisitioncoldplates,redundantsingle-phaseammoniatransportloops,and a radiator.

The TCS usespassivemethodsto maintain acceptabletemperatureswhenever

heatloadsdo not dictate the needfor activecooling. Thermal control tapesand paints

are used on external surfaces to control heat absorption and rejection. Thermal

insulatorsare usedto "isolate" elements from heat changes. Supplemental heaters are

used when necess_try to provide close-tolerance temperature control of critical

subsystems or to provide survival heating.

4.3.2.2 Growth

The growth requirement for the TCS is driven by increases in power, additional

conversion losses, and increased metabolic loads. Additional fluid lines and pumps may

be necessary for growth resource nodes, modules, and pocket labs and lines may be

routed along the vertical keels to the upper and lower booms. Additional radmtor

panels may be required to dissipate the increased the increased heat load. Active

thermal control may be installed for keel and boom mounted APAEs. The SD power

modules will have local, independent, active thermal control systems similar to the

baseline PV ATCS.

4.3.3 Data Management System (DMS)

The DMS p_ovides the hardware and software resources necessary to support

the data processing and control needs of the station's systems, elements, and payloads.

It also provides a common operating environment and human-computer interface for

the command and control of systems and payloads by both the crew and ground

operators.

4.3.3.1 Baseline

The DMS is made up of five subsystems corresponding to the five major DMS

functions: human-computer interface, data acquisition and distribution, data storage

and retrieval, application programprocessing, and time generation and distribution.

The components that make up the five subsystems, and the subsystems themselves, are

physically distributed throughout the station.

The human-computer interface is provided by the workstations which are called

multipurpose application consoles (MPACs). The MPAC contains the electronic

hardware and software components necessary to provide a multipurpose window into

4-13
all operations and experiment monitoring, station command and control, training,

testing, caution and warning, and user services. To support various functional needs

throughout the station and its interfaces,the hardware comesin three configurations,

fixed,portable, andorbiter MPACs.

The data acquisition and distribution subsystemis composedof manyhardware

and softwarecomponentswhichwork togetherto provide the rapid transmissionof data

betweendistributed systems,sensors,effectors,elements,and payloads. The primary

componentswhich contribute to the data acquisition and &stribution include; the

multiplexer/demultiplexer (MDM), the standarddataprocessor(SDP), and the DMS

optical and local networks. The MDM is the DMS interface to sensors,effectors,and

unique devicesof subsystemsand pay!oads. The MDM interfaces to the rest of the

DMS through a local network connecUonto a SDP,then via the DMS optical network

to other DMS components.

The data storage and retrieval subsystemprovides the hardware and software

componentsnecessaryfor the storageand retrieval of application programsand data.

Thesecomponentsarepackagedin a standard(MIL-STD-1788) enclosureand is called

the mass storage unit (MSU). Hardware components in this subsysteminclude; a

network interface unit (NIU), a userprovided embeddeddata processor(EDP), and a

massstorage device (MSD) with at least 250 megabytesof rapid, direct access,nonvolatile memory. The EDP provides the processingcapability to managethe programs

and data stored on the MSD. The NIU enablesprogram and data read and write

requestsvia the DMS communications network with other DMS components.

The application program processin$ subsystem includes the hardware and

software required for execution of all &stributed system, element, and payload

management applications. The primary hardware component used to perform this

function is a standard data processor (SDP), which consists of; a NIU, an EDP, a

Multibus II parallel system bus (PSB), a bus interface unit (BIU), and a power supply.

The EDP, which provides the core processing function for most of the DMS

components consists of a single printed circuit card, general purpose processor which

executes the Intel 80386 microprocessor instruction set architecture, and the 80387

numeric co-processor IEEE 754 compatible floating point instructions. The EDP

includes four (4) megabytes of dynamic random access memory (DRAM) with error

detection and correction capabihty. The EDP.p.rovides external interfaces with; the

PSB, a 32 bit IBM Micro Channel, and the precision time system (PTS). In addition to

its use in DMS components, the EDP may be used by other subsystems as the core

processor embedded m an orbital replaceable unit (ORU).

The software components of the application processing subsystem include the

operating system (OS), the system management (SM) application program, and the

operations management application (OMA) program. The OS isolates the application

software from the hardware on which it is running, manages the advanced features of

the 80386 microprocessor, and provides the capability of running multiple programs

concurrently. The SM application program performs subsystem and component

management functions, including monitoring the health and status of DMS hardware,

auditing the integrity of DMS software, accommodating system growth, initializing new

hardware and software components, controlling data overloads, detecting and isolating

faults, managing redundant DMS components, and providing the appropriate automatic

and manual controls to keep the DMS functional. The OMA program provides toplevel operations control for all systems, elements, and payloads on tlae station and acts

4-14
asthe primary control interface with the ground. Its functions include maintaining and

executingthe short term plan, maintaining a log of the statusof systems,elements, and

payloads, and supporting training and simulations.

The time generation and distribution subsystem is composed of hardware and

software that generate a stable and accurate time and frequency source, interface with

the universal time broadcast by the Global Positioning System (GPS) satellites, and

distribute time information to all processors on the station. This subsystemgenerates a

stable 1 million pulse,, per second frequency source and,produces a standard time code

accurate to 10 microseconds relative to universal time. tne time is distributed over the

core and payload networks using a dedicated bus.

4.3.3.2 Growth

The overall modular design with e.xpandable networks of the baseline DMS

facilitates growth of the system. The basehne network also contains extra noaes mr

!growth additions. Although the baseline processors have spare memory, the

incorporation of optical storage media into the evolution configurations will provide a

significant increase in data storage capacity. In addition, the software will be modular

and use standard protocols for communication, facilitating ease of growth.

4.3.4 Communication_ and Tr_¢king (C&T)

The C&T system provides for the transmission, reception, multiplexing,

distribution, signal pi'ocessin_, and generation of audio, video, telemetry, command and

control, and tracking data. It also provides for the manipulation of antennae on the

station.

4.3.4.1 Baseline

The transmission of data between the station and the ground is provided

primarily via the Tracking and Data Relay Satellite System (TDRSS) Ku band (13 to 15

GHz) link. Using the Ku band link, transmission of 300 Mbps return (space toground)

and 25 Mbps forward (ground to space) can be achieved. In addition to the Ku band

link, TDRSS provides an S band (2 GHz) link. The S band provides transmission rates

of 16 kbps return and 3 kbps forward. It will be used during, early assembly flights and

as a contingency link in the event that the Ku band becomes inoperative. Some payload

packages may also p:rovide independent user provided links to ground.

Space to space or "cluster" commumcations between the station and nearby

vehicles occur within a command and control zone (CCZ) that is roughly 37 kilometers

(20 nmi) ahead, behind, above, and below the station and 9 kilometers (5 nmi) to tile

Ort and starboard. These communications are of two classes: proximity operations and

nger-range communications using tracking antennas. Proximity operations are those

which take place within one kilometer of the station and include EVA and FI'S

operations, as well as NSTS docking and berthing. Transmissions are omnidirectional

at both ends of the proximity operations links. Long-range communications, .such as

with the approaching shuttle or a Man-Tended Free Flyer, occur from one kilometer

away from the station to the boundaries of the CCZ. Long-range transmission and

reception at the station require the use of directional antennas.

4-15
The tracking subs)'stemProvides station and cooperating vehicle positioning

data via RF links and Umversal Time Coordinated (UTC) time through the use of the

Global Positioning System (GPS). This information will be sent to the guidance,

navigation, and control (GN&C) system which computes the pointing angles for the

TDRSS Ku band antennas and pointing payloads. GPS data is also used to track

vehicles during rendezvous and departure. The same information that an approaching

vehicle receives from the GPS is transmitted to the station's GN&C system, where

antenna pointing angles for the vehicle are determined. A space to space tracking

antenna is then assigned to track the vehicle. The GN&C system derives range and rate

information from the data transmitted by the vehicle.

The video subsystem receives or distributes video signals to and from cameras

and monitors throughout the station. The subsystem also provides the crew with video

processing capabilities to record, edit, and play back images. Video to be transmitted

to ground via the TDRSS link are digitized prior to transrmssion.

The audio subsystem distributes au&o communications internally, throughout the

station and to the space-to-space and the space-to-ground commumcat_on systems.

Internal audio is wireless to allow the crew to operate with minimal encumbrance.

Crew members can also record experimental andoperational observations for later

playback and analysis.

As part of the C&T management function, the C&T system monitors the health

and status of its own equipment, diagnoses faults in that equipment, and controls its

configuration and operation. This control and monitoring subsystem interfaces with the

DMS orbital management application (OMA) program to receive and execute the

C&T's portion of the short-term plan in scheduling communications resources. Caution

and warning information is also exchanged with the OMA program.

4.3.4.2 Growth

To accommodate growth, the baseline space to space subsystem patch panel is of

modular design with spare connectivity. The baseline space to ground link is also

designed to operate at TDRSS full capacity (300 Mbps). However, growth

requirements may necessitate upgrades to enable utilization of a possible advanced

TDRSS which will have a 650 Mbps downlink and a 50 Mbps uplink capability. The

baseline design does not preclude either the addition of direct hnks to groundor the

incorporation of voice recognition into the audio subsystem. The video distribution

suosystem bandwidth does accommodate high-resolution video requirements, and its

design is modular to allow incorporation of future technolo__¢ upgrades. To facilitate

growth, the audio subsystem also has spare connectivity. The station is designed to

accommodate automated support for traffic control, proximity operations, docking of

free flyers, and control of multiple free flyers. Radar and laser ranging may be added

for automated docking and berthing operations.

Possible technology upgrades for the C&T system include evolution to the Ka

band or optical for space-to-space links, upgrading signal processing to accommodate

high definition TV, and apphcation of expert system technology to the control and

monitoring subsystem.

4-16
4.3.5 Guidance. Navigation. and Control (GN&C)

The GN&C System controls the station's attitude, maintains its altitude between

reboost and rendezvous, calculates pointing angles for systems and payloads, controls

proximity traffic, and maneuvers the station to avoid possible collisions with space

vehicles, debris, or meteoroids.

The station flies in a local-vertical local-horizontal (LVLH) attitude which keeps

the station perpendicular to the Earth. In the LVLH mode, the truss is perpendicular

to the direction of flight, and the bottom of the station always point towards the Earth.

This attitude is maintained within five degrees in each axis.

4.3.5.1 Baseline

The GN&C system is divided into two subsystems: the core subsystem and the

traffic management subsystem. The core GN&C subsystem is responsible for

controlling attitude and altitude and for calculating pointing angles. Attitude control is

accomplished primarily through the control moment gyroscopes (CMGs) mounted on

the GN&C pallet. The CMGs act to absorb torques which cause disturbances to the

attitude of the statioa.

Atmospheric drag will decay the station's orbit, requiring periodic station

reboost. The RCS lhrusters and the resistojets, controlled by the GN&C, are used to

increase the velocit) of the station in the direction of flight. The resultant increase in

centripetal acceleration forces the station away from the Earth. The GN&C may also

be used to decrease :_tation altitude during rendezvous operations or collision avoidance

maneuvers.

The GN&C performs attitude calculations using data collected by star trackers

and inertial sensor assemblies (ISAs). These sensors are strategically located on the

space station to minimize viewing obstruction from added elements in the growth

configurations. Star trackers are devices which measure the positions and magnitudes of

stars. This data is compared to data stored in the star catalog to uniquely identify the

star in the star tracker's field of view. The angle from the navigation base, a rigid

structural mount for the star trackers that provides a stable reference, to several known

stars can be measured with a maximum error of 60 arc-seconds. The ISAs consists of

three ring laser gyroscopes and are used to sense small rotations with extremely high

accuracy.

The GN&C traffic management subsystem is responsible for controlling

incoming, outgoing, and stationkeeping traffic within the CCZ, controlling docking and

berthing operations, monitoring the trajectories of cooperating vehicles which intersect

the orbit of the station, performing colhsion avoidance maneuvers, and supporting flight

planning of the traffic around the station. Data is gathered from the GPS, radars, and

optical sensors on cooperating vehicles and from ground tracking for noncooperating

vehicles and debris. Noncooperating vehicles are defined as those which do not

communicate with the station. Using this data, a vehicle's position, velocity, and

attitude is calculated by the GN&C traffic management SDPs. Closed-circuit television

and direct viewing support close-proximity operations for berthing and docking.

4-17
4.3.5.2Growth

Additional capacity will be required in the areas of momentum absorption

(attitude control) and fuel for reboost. There will also be a need for automated support

of the traffic management functions to accommodate expected increased volume of

traffic within the CCZ. Other growth areas will include sottware and hardware support

for an increased number and variety of payloads requiring pointing control and for the

possible addition of tethered payloads. If operations analyses indicate the need for

additional CMGs, they will be co-located with baseline CMGs to minimize interactions

due to flexure of the station.

4.3.6 Extravehicular Activity_ (EVA) System

The EVA system (EVAS) provides the capability for pressure-suited crew

members to perform a wide variety of tasks external to the station's pressurized volume.

These tasks include assembly and servicing of the station elements, distributed systems,

and attached payloads. The system includes EVA Mobility Units (EMUs), EVAS

Service and Performance Checkout Unit (SPCU), airlock depress/repress equipment

and controls, EVA hyperbaric life support and chamber pressurization equipment and

controls, EVA translation and mobility aids, EVA external lighting, EVA support

equipment and generic tools, EVA external equipment storage, and EVA

contamination detection and decontamination unit. The EMU provides an

anthropomorphic pressure suit, portable life support, and EVA crew communications

and physiological monitoring. The SPCU provides for the servicing of EVA equipment

between excursions.

4.3.6.1 Baseline

The baseline EVAS will be station-based with the NSTS-design EMUs and will

perform onl), contingency operations. All other EVAs toperform maintenance, repair,

and inspection tasks will be scheduled only when an Orbiter is present. A sfngle

hyperbaric airlock supports the extravehicular activities.

EVA subsystems are located both internally and externally to the airlock and on

the transverse boom. The EMU consists of the space suit assembly (SSA) and the

personal life support system (LSS). It is a self-contained pressure vessel that provides

the. radiation and micrometeoroid protection necessary to work in the space

enwronment. The NSTS-design suit's operating pressure is 4.3 psia and requires

substantial pre-breathing time to allow the removal of nitrogen from the body tissue

pnor to airlock depressurization.

Thermal control is accomjalished with a non-venting, regenerable, thermoelectric/wax/radiator system. This system automatically controls the temperature of

cooling water circulated through the liquid-cooled garment (LCG) worn under the

space suit. Carbon dioxide is removed by circulating the atmosphere through a

removable, on-orbit regenerable cartridge. Two silver-zinc rechargeable batteries

provide power.

Communications for the EVA/IVA crew consist of voice communication to and

from the station, data communications to the station, and voice communications

4-18
betweenEVA crew members via the C&T system's space-to-space and internal audio

subsystems. The EYA crew communicates with the ground via the station.

The SPCU provides for on-orbit reservicing of EVAS equipment between

EVAs. The SPCU provides the interface for the EVAS equipment and the service

equipment necessar.¢ to support IVA and EVA operations during EMU donning and

transition to the vacuum environment. It also provides for performance checkout and

instrumentation verification of all EVA serviceable hardware.

Translation aids are located on the station modules and truss structure to

facilitate the transpc, rt and restraint of EVA crew members and equipment. These aids

include handholds, handrails, and restraint attachment points located along, all EVA

routes, at each EVA hatch, inside and outside the airlock, and at appropriate EVA

work sites.

The Contamination Detection and Removal System (CDRS) is located outside

the airlock and uses a quadrupole mass spectrometer to detect, identify, and measure

external contamination on the EMU. If contaminants exceed tolerance levels, heat

lamps will be used to vaporize the contaminants.

4.3.6.2 Growth

Evolution in the EVAS will provide the capability for the crew to perform

routine tasks in the unpressurized environment. In addition to the basic maintenance,

repair and inspection functions, the EVAS will support installation, assembly, and

servicing of on-orbit systems and user hardware. The EVAS will incorporate the space

station-design, 8.3 p,_ia, rear-entry EMU which will minimize the time required for prebreathing and space; suit donning/doffing. The pressurized suit assembly will allow

greater crew mobility and have an automated checkout capability.

An Extravehicular Excursion Unit (EEU) will be available to provide for freeflying proximity operations of up to 300 meters from the station. The EEU will be onorbit serviceable and will be stored outside the airlock.

The Service and Performance Checkout Subsystem (SPCS) will replace the

SPCU. The SPCS will provide for automated performance checkout and reverification

following maintenance and ORU replacement. The servicing and checkout operations

that the SPCS provides include LCG water servicing, suit leak check-out, suit drying,

carbon dioxide]hun-ddity control subsystem regeneration, heat sink regeneration, and

battery recharge. The SPCS completely regenerates the EMUs in 12 hours and can also

provide one hour tam-around for contingency EVA. A one hour oxygen recharge

provides about 80% of the oxygen normally provided by a 12-hour recharge. The EMU

thermal control sub,s'stem is fully rechargeable in one hour. The carbon dioxide and

humidity control umt and the EMU batteries can be manually changed out for a

contingency EVA.

A Crew Rescue and Equipment Retrieval Subsystem (CRERS) will be provided

to rescue an EVA clew or to retrieve equipment accidentally detached from the station.

This capability accommodates separation of a crew member from any point on the

station and provides acquisition and return to the airlock within the EMU expendables

reserve time. This capability will also be utilized for detached equipment retrieval

when deemed appropriate.

4-19
Communicationsfor the EVA crew will be augmented during growth to include

voice and data to and from the station, video to the station, and the transfer of text and

graphics from the station to the EVA crew member.

Technology upgrades can be expected in the EVAS during the 30-year life of the

station, especially in the suit s_/stems. However, accommodation of the SEI on station

will require extensive operataons in the unpressurized environment to support the

inspection, assembly, processing, and servicing of the exploration vehicles.

Advancements in the areas of autonomous and telerobotic systems and cooperative

manned and telerobotic EVA systems will be needed not only to reduce the anticipated

number of manned EVAs, but also to reduce risks and to increase productivity.

4.3.7 Man Systems

Man systems consists of equipment necessary to support the crew members while

they live and work on-board the station. Although the majority of the man systems

hardware is in the habitation modules, man systems hardware and functions are

distributed throughout the pressurized volumes and interact with many other station

systems. Design emphasis for the man systems was placed on creating a good manmachine interface (MMI) which allows the crew to be productive in work areas and to

relax in the rest areas.

4.3.7.1 Baseline

Man systems are functionally partitioned into fifteen subsystems: crew quarters,

restraints and mobility aids, crew health care, operational and personal equipment,

portable emergency provisions, galley, wardroom, personal hygiene, illumination,

stowage, housekeeping/trash management, interfacing partitions and structures, inflight maintenance, integrated workstations, and inventory management.

Restraints and mobility aids are located throughout the pressurized elements to

facilitate the microgravity movement of crew and hardware. Crew member health

needs are addressed on board the station and supplemented as necessary by

communication with ground-based physicians. Nutritional requirements are met with

an extensive galley and food management system. Maintenance of personal hygiene is

supported along with station-wide housekeeping and trash management. Personal and

crew-wide stowage accommodations are also provided.

Specific areas within the pressurizedvolume of the station are dedicated as

integrated workstations which provide communications between the crew and station

systems. There are two varieties of workstation: fixed and portable. Fixed workstations

provide distributed system and element control, monitoring, verification, configuration

management, manual override of automated functions, access to procedures, caution

and warning display and annunciation, training, word processing, electronic mail, and

inventory management. The fixed workstations are configured with high resolution

color monitors that provide multifunction capability, including image windowing, video,

graphics, and text overlay.

Portable workstations consist of a portable multipurpose application console

(MPAC), a flat panel display, and a keyboard. Crew members can select and

interconnect portable workstation components in a variety of combinations anywhere

that ports are located.

4-20
Workstations located in each cupola can be used for docking and traffic

management operations. Video, graphics and text overlay, or video windowing

generated at any fixed workstation can be switched to the cupola workstation to create

t-emplates and cross hairs to assist in judging distance and alignment. Hand controllers

are provided for precision manual vehicle control.

The maintenance workbench, located in the U.S. lab, supports intravehicular

activity (IVA) maintenance and servicing of the station and other hardware. The

workbench occupies two racks and accommodates the laboratory science maintenance

servicing functions. The workbench also provides stowage for tools, supplies, support

equipment, and test and diagnostic equipment.

4.3.7.2 Growth

Evolution in the man systems is expected to be by replication of the baseline

habitation module. Each habitation module is designed to provide crew quarters, crew

health care, galley, wardroom, housekeeping and trash management, and personal

hygiene for a crew of eight. Therefore, the addition of two habitation modules during

the evolution phase_,_ will accommodate a total permanent crew of 24, which has been

deemed adequate for all configurations. Judicious outfitting of the replicated

habitation modules could increase this crew size or accommodate transient personnel

by replacing galley, wardroom, and health care facilities with crew quarters. The

provisioning of restraints and mobility aids will continue for the evolution habitation

and laboratory modules.

Modularity and design-for-maintainability requirements will facilitate

implementation of technology up.grades as they become available. Areas of technology

upgrade include health care faclhty, galley, housekeeping/trash management, personal

hygiene, and MMI. In many cases, technology upgrades will be accomplished on orbit.

4.3.8 Environmental Control and Life Support System (ECLSS)

The ECLSS maintains a safe, shirt-sleeve environment (at sea-level atmosphere)

within the pressurized elements of the space station. It controls the temperature,

humidity, and air composition of the atmosphere. It also supplies water for drinking and

personal hylgiene, processes and stores the crew's waste material, and detects and

suppresses fires. Tl_e baseline design allows the ECLSS fluid makeup requirements to

be satisfied by the water contained in food and by a nominal resupply quantity of

nitrogen. A key feature of ECLSS is its use of oxygen and water regeneration

technology to minimize resupply.

4.3.8.1 Baseline

The majority of ECLSS hardware is located in the U.S. lab and hab modules.

Redundant temperature and humidity control, atmosphere revitalization, and water

recovery and management assemblies are packaged in identical racks or mounted at

identical end-cones of the hab modules. Each life-critical ECLSS assembly is designed

to support the enti:re eight member crew for 45 days under safe haven conditions.

ECLSS uses distribtttion hardware that interconnect tlae station s pressurized elements

4-21
with air, water, oxygen,and nitrogen. For safety,the fluid resourcelines are provided

with isolation valvesat elementbulkheads.

There aresixECLSSsubsystems:temperatureand humidity control, atmosphere

control and supply, atmosphererevitalization, water recoveryand management,waste

management,and fire detectionandsuppression.

The ECLSS temperature and humidity control (THC) subsystemcontrols the

cabin air temperature, humidity, and ventilation, the air cooling of rack mounted

equipment,and the temperature-controlledstorageof food. The cabin air temperature

andhumidity is controlled by circulating the air throush condensingheat exchangers.

Power consumingequipment are air cooledwith a variable speedfan assemblyand a

heat exchangerwhich duct air to the individual racks. Additional refrigeration is

provided in the pressurizedlogisticscarrier for storageof food.

The atmosphere control and supply (ACS) subsystemregulates the total

pressureand oxygenpartial pressurefor the pressurizedelements.It compensatesfor

atmosphericleakageandairlock lossesandprovidesfor pressureventing and relief.

The atmosphere revitalization subsystem(AR) removes carbon dioxide and

trace contaminantsfrom the air and providesoxygenregeneration. The AR subsystem

is designedto accommodatea nominal metabolicoxygendemandof 0.83kg/day (1.84

lbs/day) per crew member and carbon dioxide reduction of 1.00kg/day (2.20lbs/day).

In addition to the nominal eight member crew, the AR subsystem can also

accommodatethe demandof another 1.3"crewmember"for useby animal life sciences.

The water recovery and managementsubsystem(WRM) recovers water from

variouswastewater sources,distributesthe reclaimedwater,controls the microbacterial

and chemical composition of the water, and thermally conditions the water supply.

Potable water is reclaimed from humidity condensateand carbon dioxide reduction

using a multifiltration process.Hygiene water is reclaimed from other wastewater by

reverseosmosisprocessingthroughmembranefilters andan evaporativeprocess.

The wastemanagement(WM) subsystemseparatelyprocesseshumanmetabolic

wastesand urine. Solidwastes are compacted, allowed to biodegrade, and stored for

return to Earth. Urine is processed by the waste WRM for hygiene water. The urine

brine concentrate that is a by-product of the reclamation process is also stored and

returned to Earth.

The fire detection and suppression (FDS) subsystem consists of the hardware

required to detect and suppress fires within the pressurized elements. Ionization

sensors are installed in enclosed racks, infrared and ultraviolet sensors are used in open

areas such as aisles and nodes, while thermal sensors are used in enclosed areas

containing electrical cables. Carbon dioxide is the fire suppressant used in both the

fixed and portable fire extinguishers. Halon suppressants may also be used.

4.3.8.2 Growth

Evolution in ECLSS is expected to beprimarily through the replication of the

baseline laboratory and habitation modules. However, the implementation may differ

due to the module pattern growth and ECLSS support requirements for adjacent

resource nodes, pocket labs, and airlocks. These requirements may be equal to or less

than those of the baseline module pattern in which the U.S. hab and lab ECLSS must

support the international modules, as well as adjacent resource nodes, airlocks, logistics

4-22
module, and cupolas. A possiblehigher ECLSSload may occur if the evolution hab

module is outfitted to supportmorethan a crewof eight.

Baseline modularity requirements of the ECLSS will facilitate technology

upgrades. Although in most cases,the upgradescan be accomplishedon orbit, they

shouldbeperformeclon thegrounddue to the criticality of the ECLSSfunctions.

4.3.9 Fluid Management System (FMS)

The FMS is a central management and distribution system for fluids common to

the stations elements and systems. The baseline FMS performs three major integrated

functions: resupply and distribution of nitrogen; resupply, distribution, and disposal of

water; and collecnon and disposal of waste gases.

The FMS delivers mtrogen and water to the laboratory modules for use in

experiments. It supplies nitrogen and water to ECLSS for use m maintaining the onboard atmosphere. It also supplies nitrogen to the TCS for use in purging the ammonia

working fluid from its lines during maintenance operations. The FMS sends waste gas

to the propulsion system to fuel the resistojets.

The FMS ak,,o collects waste fluids produced by daily operations. The FMS

receives excess water from ECLSS and stores it for later use in the labs. It collects

waste water not recoverable b_ on-board systems from the labs and delivered it to the

logistics carriers for return to Larth. It gathers waste gases from the lab modules, waste

gases purged from the TCS, and excess hydrogen gas from ECLSS and delivers them to

the propulsion system.

4.3.9.1 Baseline

The baseline FMS consists of three major subsystems: the integrated nitrogen

subsystem (INS), the integrated water subsystem (IWS), and the integrated waste gas

subsystem (IWGS). These three subsystems each contain the hardware, software, and

interfaces with the other station systems needed to complete the overall FMS network.

The Unpressarized Logistics Carrier (ULC) resupplies the FMS with nitrogen

and water. Because the ULC remains with the station until the next shuttle resupply

flight, it augments the INS and IWS on-orbit storage capability. System sizing is dnven

by both the baseline, 90-da), resupply interval and the requirement for contingency

storage to accommodate a massed launch.

The main truss assembly supports the distribution lines of the FMS external to

the pressurized elements. The external distribution lines allow the FMS to interface

with users on the truss (the TCS and the propulsion system) and users internal to the

pressurized elements through an interface with internaldistribution lines located at the

nodes.

To improve safety, high-pressure storage tanks for the INS and the IWGS are

located on the truss away from the pressurized modules. A majority of the internal

distribution lines are located in the resource nodes to supply the elements with common

fluids. Water is stored in the nodes to eliminate the need for dedicated thermal

conditioning and meteoroid debris shielding. Waterproduced by the shuttle's fuel cells

while docked to station is scavenged by the IWS as additional resupply. This scavenged

water is transferred from the node docking port to the water storage tanks. The IWS

low storage pressure, approximately 20 psia, does not pose a crew safety hazard. Each

4-23
of the lab modules containsits own internal distribution systemwhich interfaceswith

the INS and IWS. Likewise,wastegasesproduced within the labs are transferred to

and collectedby the IWGS.

4.3.9.2Growth

Potential evolution of the FMS existsin threedistinct areas:expansionof station

scienceactivities,accommodationof TransportationNode requirements,and reduction

of available EVA maintenancesupport. Expansion of station scienceactivities will

require the FMS to increasethe capacityof the baselinenitrogenand water supplyand

wastegascollection servicesand be capableof supplyingadditional fluid services(Kr,

At, He, and CO9).

TransporTationNode requirements entail the addition of significant cryogenic

fluid handling services,expansionof baselinenitrogen and water supplyand wastegas

collection services,expansionof baseline nitrogen servicesto include high-pressure

nitrogen lines, and the addition of Earth-storablepropellant (hydrazine and bi-props)

storage,distribution, and transferservices.

4.3.10 Propulsion System

The space station will orbit the Earth at altitudes of 190 nautical miles and

higher. At these altitudes, the station is subject to drag forces produced by interaction

with the Earth's atmosphere. The propulsion system is designed to periodically reboost

the station to maintain the nominal orbital altitude. In addition to reboost, attitude

control, and collision avoidance, the propulsion system also provides for disposal of

certain waste fluids.

4.3.10.1 Baseline

The baseline propulsion system consists of a blowdown hydrazine system

ausmented by a resistojet system. The hydrazine system provides the primary thrust for

attitude control torques and translation maneuvers. While the primary attitude control

is accomplished by the CMGs, the station propulsion system is used to desaturate the

CMGs, as a CMG back-up, and to damp out any disturbances that exceed the CMGs'

capability. The resistojet system utilizes waste fluids to produce useful impulse to

provide a supplemental reboost capability.

At each end of the transverse boom are two propulsion locations, one each

above and below the boom. At each of the propulsion locations, there are two module

interfaces, such that two modules can be collocated next to each other. When one of

the modules is depleted of propellant, it is returned to the ground, and the adjacent

module is utilized. The propulsion system will be entirely contained in the modules,

which will have everything required to produce station thrust: propellant tanks,

isolation valves, thrusters, heaters, instrumentation, and avionics. The only interfaces

with the truss will be those required for power, data, and structure. Each module will

be dual fault tolerant, since reboost performance is a station-critical requirement.

4-24
The resistojets provide for disposal of station wastefluids and produce some

supplementalreboo:_tcapability. The resistojet assemblyconsistsof a module of six

thrusters,eachdevelopin_onepoundof thrust. To conductthe heatedfluids awayfrom

the station'soptical vaewmgareas,the thrustersare located at the end of a 30 meter

boom ¢stin_er_- Wa,.te fluids are heated and exoelled out of one of these six thrusters.

The resistojets use 500 watts of power to heat the waste products to 760 ° C (1400 U F)

before expelling them from the thrusters.

4.3.10.2 Growth

For the evalutionary reference configurations, growth requirements on

propulsion system performance dictate major changes in the system. Althoush several

candidate systems are being analyzed as replacements to the baseline, initial results

favor a water electrc.lysis/Hg-O _ bipropellant system.

The major elements-of-the proposed H_-O9 propulsion system include four

reaction control modules (RCMs), two propellanFpro-duction and storage modules, two

supplemental storage modules, and two propellant transfer and distribution subsystems.

The RCMs will be located on the upper and lower keels and the transverse

boom hydrazine modules will be removed. Each module contains six thrusters. Each

propellant production module contains both an electrolysis unit for generating the Op

andthe H9 and a graphite storage tank for the gaseous 0 2 and H 2. Additional ga_

storage w_'n be provided by the supplemental storage modules. The propellant

distribution system carries the propellant from the electrolysis units to the storage tanks

via 3000 psia high pressure lines. The propellant is then regulated to 300 psia prior to

transfer to the RCMs.

The baseline resistojet system will still be used for disposal of waste fluids and to

provide supplemental propulsion.

4-25

5.0 GROWTH DELTA ANALYSIS

The results of the analysesfor the configuration conceptswere provided and

discussedearlier in this report. This section providesa more detailed explanationof

the analysesmethodology and the supportin_ data for intermediate configurations

(deltas) attainedduring build-up from the basehnestation.

5.1 ANALYSIS METHODOLOGY

Most of the analysespresentedin this report were performed usingthe IDEAS2

integratedspacestalion analysissoftware. Theseanalysesmclude:

rigid-body controlsanal.ysesandflight mode characterization

evaluationof microgravltyenvironments

orbital lifetime and reboostanalyses

determination of payloadviewing envelopes

structuraldynamicsanalysis

To ensure consistency,a single geometric model was created for each of the

delta configuratiom using an enhancedversion of the GEOMOD solids modeling

program. 'I_hese models _¢ere used to form the basis for all of the various analysis

functions of that colffiguration. The geometric data was then processed by MODGEN

which creates analytical models compatible with the analysis modules.

All analyses used the Cartesian coordinate definitions of the baseline station,

i.e., "+ x" for line of flight and "+ z" for the nadir (toward Earth) direction.

5.1.1 Attitude Control Analysis

Attitude conl:rol analysis was used to determine the Torque Equilibrium Angle

(TEA). The TEA is the flight angle that minimizes the force/torque effects of

environmental forces. Flying the station at the TEA allows the minimization of the size

of the space station's control system. The Articular Rigid-Body Control Dynamics

(ARCD) module was used to estimate the TEA to be used for control and microgravity

analysis.

The ARCD module determines the environmental forces and torques acting on

the Space Station Freedom over an orbit. The environmental effects modeled by the

ARCD program are:

Aerociynamic drag

Gravil.-y gradient

Solar radiation pressure (absorption only)

Gyroscopic effects (i.e., due to motions of articulating parts)

Inputs to the ARCD module include:

Space Station Freedom mass and projected area properties

Orbit geometry

Environmental constants

5-1
Articular part characteristics(feathering,sun-tracking,axesof rotation)

Control systemparameters

ARCD models the Freedom Station as a rigid body. The articulating parts are

modeled as rigid free bodies constrained to the station. Using the model input, the

ARCD module computes the station's mass properties and projected areas and the

torques generated throughout an orbit, accounting for variations due to the articulating

parts and their motions.

The outputs of ARCD include time histories of the following quantities

calculated about 60 times per orbit:

Environmental forces and torques

Required control forces and torques

Angular momentum requirements

Motions of articulating parts

TEA

Propellant mass required for orbit-keeping and attitude control

Steady state attitude control system sizing was performed using the Attitude

Predict (A'ITPRED) module of IDEAS/-. The modified ATFPRED software module

allows the simulation of Control Momentum Gyroscopes (CMGs) for attitude control.

A CMG control law/momentum management algorithm, steering law, and model has

been implemented into the software. This controller generates a torque command

which provides continuous closed-loop control of both the spacecraft attitude and CMG

angular momentum via state feedback and disturbance rejection. Given a torque

command, the steering law generates appropriate gimbal rate commands, which

distribute the CMG momentum vectors such that all inner gimbal angles are equal and

all outer gimbals are equally spread out. The double-gimbaled CMGs are modeled as

error-free actuators, which deliver the gimbal rate command subject to user-defined

gimbal freedom limits and gimbal rate limits.

5.1.2 Microgravity Environment Determination

A major requirement for the station is to provide a steady-state microgravityfree environment for scientific experiments and commercial research and materials

manufacturing. Microgravity requirements for these activities range from 0.1 to 100

micro-g's. The Program Definition and Requirements Document (PDRD) for the

baseline station states that the steady-state acceleration level in the pressurized

modules should be limited to 10 micro-g's. In order to assess the capability of the

evolution options to accommodate microgravity experiments, gravity level contour plots

in the form of ellipses overlaid on the ZX and YZ planes of configuration models were

generated using ARCD module data as input. The ellipses circumscribe areas of the

space station in which the gravitational acceleration field is less than any given input

level. The ellipses are determined by the values of the following steady-state

parameters:

Gravity gradient

Centripetal acceleration gradient

Atmospheric drag

5-2
Attitude drift

Dynamic inputs, such as crew activities, Mobile Servicing Center (MSC)

translation,and orbv:erdockingandberthingwere not consideredin this analysis.

5.1.3 Orbit Lifetime. and Reboost Analysis

The assumed altitude strategy allowed the station's orbit to decay cyclically from

its nominal altitude to its lowest altitude of 220 nautical miles (nmi). At the lowest

altitude, the space station will conduct shuttle resupply rendezvous, and space station

reboost. For all evolution configurations, the additionalmass and drag from the growth

elements will affect both the rate of decay and the amount of propellant required for

reboost. The Orbit Lifetime (OL) module was used to calculate the orbital lifetimes of

the confi_gurations and to determine the post-assembly reboost propellant requirements.

Gwen the spacecraft physical characteristics (mass, cross-sectional area, etc.),

initial orbit parameters, and launch date, the OL module calculates changes in

spacecraft orbits. It considers orbit decay due to the environmental perturbations of

atmospheric drag, solar radiation pressure (including effects of the Earth's shadow),

Earth oblateness, and the gravitational attractions of the sun and the moon (both on

spacecraft and on the Earth). Only long-term changes are considered. Short-term

variations are assumed to average out over one orbit.

5.1.4 Structural Dynamics Analysis

To validate the concepts which have been developed for evolution of Space

Station Freedom, a structural dynamics and loads analysis will be performed for both

the Transportation Node and R&D evolution configurations. Detailed structural

analyses have not been completed for any of the configurations contained in this report.

However, several similar configurations have been analyzed, and the results provide

meaningful insight into the structural characteristics of these concepts. The following

discussion outhnes the analysis approach taken to assess the feasibility of large

structures such as the growth configurations. Sections 2.3.5 and 3.3.5 of this report

contain qualitative extrapolations from previous analyses to the R&D and

Transportation Node configurations.

Previous structural analyses of conceptual designs have roughly followed the

approach described below.

o Generate finite element structural model. This first step requires identification

of the mass pro?erties, stiffnesses, and orientations of all components of the

system. I'.a particular, it has been shown that the behavior of large structures

such as these is highly sensitive to the stiffness of the alpha joint and to the

amount of mass near the ends of the transverse boom and somewhat less

sensitive to the mass and stiffness of appendages (solar arrays and solar

dynamic modules) and of components near the module cluster. This will be

the case far the configurations described here.

2. Solve for normal modes of the system. Past analyses have used a graphicsbased program such as SDRC/SUPERTAB to piece together the finite

5-3
element model,but the actual calculation of system modes was performed in

MSC/NASTRAN.

° Determine appropriate forcing functions. The most intense loadings

experienced by this type of structure occur durin._ reboost and docking.

Dockin_ loads can be approximated by impulses apphed at the docking ports.

Approramate reboost functions can be determinedfrom rigid body equations

of motion and mass properties, but more precise functions are gained by

incorporating flexible body effects.

. Determine system response to forcing functions. By assemblin_ the normal

modes of the system into a modal formulation, the response (dis.placements,

accelerations, and member loads) of the system due to the gwen forcing

functions can be determined at key locations about the structure. Locations

such as the tip of solar arrays, the center of the modules, the interface

between the boom and keels, and the base of the solar dynamic pointing unit

are investigated to be certain that reasonable limits are not exceeded.

Data produced from detailed analyses of the specific R&D and Transportation

Node reference configurations will be included in future updates of this document as

the analyses are completed.

5.2 R&D DELTA CONFIGURATION CONCEPT ANALYSES SUMMARY

The R&D configuration is divided into twelve growth deltas. Table 5.2-1

provides the elements added for each delta. It should be noted that these deltas do not

necessarily correspond to ETO flights and were conducted only to establish that the

station maintains the required control characteristics during growth.

Analyses were performed for each of the R&D deltas to assess the physical

characterisucs of Freedom as it grows from its baseline configuration to the R&D

configuration. A summar_ of the results of the delta analyses is presented in Table 5:2-

2. The mass of the stauon more than doubles during the growth process, with the

biggest increase occurring at delta six where the STV servicing facility and fuel depot

are added to the lower keels.

The origin of the geometrical coordinate system used is located in the middle of

the center truss bay of the transverse boom. The coordinate system has the station X

axis parallel to velocity vector, the station Y axis aligned perpendicular to orbit plane,

and the station Z axis directed toward the center of the Earth. The center of mass

varies from zero to -2.6 meters in X, from -0.2 to 0.7 meters in Y, and from zero to 7.8

meters in Z. The moments and products of inertia increase by an order of masnitude

during station evolution, with the larger increases associated with the deltas that include

solar dynamic power units and STV equipment. Throughout the buildup, the products

of inertia are always about two orders of magnitude smaller than the moments of

inertia leadingto small body-to-principal-axis rotations once the keels are attached.

Once the keels are addedas part of the second delta, all subsequent deltas yield

average flight attitudes within three degrees of local vertical-local horizontal (LVLH)

for roll, pitch, and yaw. The small flight attitudes, coupled with a center of mass near

the module pattern, result in a one to two micro-g environment for the laboratories

depending on payload placement and STV payload/fuel status. Steady-state control

5-4
Characteristics

Mass (kg)

Center of mass (m)

from geometric

center of transverse

boom

Inertias (kg.m 2 )

with respect to the

center of mass

Average flight

attitude (degrees)

from LVLH

Control momentum

(N.m.s) peak

requirement for

steady-state

operations

Ballistic coefficient

(kg/m 2 )

Range

266,000 to 687,000

X = -2.6 to 0.0

Y = -0.2 to 0.7

z = 0.0 to 7.8

IXX =1.2E8 to 1.2E9

IYY = 2.3E7 to 3.9E8

IZZ = 1.3E8 to 9.2E8

Deltas 2 thru 12

X (roll) = 0.0 to 3.0

Y (pitch)= 0.0 to 2.4

Z (yaw) = -0.5 to 0.5

4,000 to 15,000

47 to 64

Comments

Largest mass increase is at delta 6

Where applicable, STV and full STV

tank farms were assumed.

min Xdelta 10 maxXdelta5

minYdelta 2 maxYdelta4

min Zdelta 2 maxZdelta6

Where applicable, STV and full STV

tank farms were assumed.

Products of inertia are two orders of

magnitude below the moments of inertia.

Both the assembly complete and deltas

have pitch values over -5 degrees.

min X delta 1 max X delta 7

min Y delta 5 max Y. defta 12

min Z delta 7 max Z delta '3

Values depend on atmospheric and

control assumptions. For all deltas, the

flight attitude is maintained with less

than +/- 2 degrees per orbit deviation

from the average flight attitude

using the baseline CMGs.

min delta 1 max delta 7

Table 5.2-2 • Analytical Characteristics Summary for the R&D Deltas

5-5
momentum requirements are all within the capability of the six baseline CMGs with

attitude devianon (all less than +/- two degrees per orbit) depending on atmosphere

and control law assumptions. Attitude error rates for all delta configurations and for all

assumptions are below 0.005 degrees/sec.

Table 5.2-1 R&D CONFIGURATION DELTA COMPONENTS

DELTA

DO

D1

D2

D3

D4

D5

D6

D7

D8

D9

D10

Dll

D12

COMPONENTS

Baseline Configuration

1 Solar Dynamic Module Pair,

Space-Based OMV & Accommodations

Upper/Lower Keels & Booms

1 Hab Module, 2 Resource Nodes

1 Solar Dynamic Module Pair,

Attached Payloads

1 Large Pocket Lab, 1 Lab Module,

2 Resource Nodes, Attached Payloads

STV & Hangar, Assembly Platform,

Attached Payloads

1 Hab Module, 1 Resource Node

1 Solar Dynamic Module Pair

1 Lab Module

1 Small Pocket Lab

1 Solar Dynamic Module Pair

1 Large Pocket Lab

The average drag area of Freedom increases by more than 100% as it evolves

from the baseline configuration to the R&D reference station. The corresponding

increase in mass offsets the increase in area, resulting in small,variations in the ballistic

coefficient ranging from 50 kg/nff (no STV, fuel) to 60 kg/m '_ (STV and fuel) for the

R&D_configuration. During the builduo process, the ballistic coefficient peaks at 64

k_/m z at delta 7. As compared with the'baseline configuration, the R&D configuration

will decay at a slower rate, but require more than twice as much fuel per reboost.

5.3 TRANSPORTATION NODE DELTA CONFIGURATION ANALYSES

SUMMARY

The interim configuration for the lunar/Mars Transportation Node is the Lunar

Operations Transportation Node. Analysis of the lunar concept was discussed earlier in

this document. Further deltas for station augmentation to accomplish growth to the

transportation nodes is discussed in the Station Evolution Reference Configuration

Databook.

5-6
Report Documentation Page

1. Report No.

NASA TM 102675

4. Title and Subtitte

2. Government Accession No. 3. Recipient's Catalog No.

5. Report Date

May 1990

A Study Of Concept Options For The Evolution Of

Space Station Freedom

7. Author(s)

Herbert R. Kowitz, Karen D. Brender, William M. Cirillo,

Lisa Collier, George G. Canoe, Marston J. Gould, Martin

Kaszubowski, George F. Lawrence, Charles P. Llewellyn,

Ray Reaux, Washito Sasarnoto, Patrick A. Troutman, Laura Waters

9. Performing Organization Name and Address

NASA Langley Research (;enter

Hampton, VA 23665-5225

12. Sponsoring Agency Name and Address

National Aeronautics and _;pace Administration

Washington, DC 20546-0001

6. Performing Organization Code

8. Performing Organization Report No,

10. Work Unit No.

476-14-06-01

11. Contract or Grant No,

13. Type of Report end Period Covered

Technical Memorandum

14. Sponsoring _kgency Code

15, Supplementary Notes

Langley Research Center: Herbert R. Kowitz, Kaxen D. Brender, William M. Ciriilo,

George G. Ganoe, Marston J. Gould, George F. Lawrence, Washito Sasamoto,

Patrick A.'rroutman

Computer Technology Associates: Lisa Collier, Martin Kaszubowski, Ray Reaux

Analytlcal Mechanics Associates: Charles P. Llewellyn, Laura Waters

16, Abstract

Two conceptual evolution configurations for Space Station Freedom, a Research and Development

configuration, and a Transportation Node configuration are described and analyzed. Results of pertinent

analyses of mass properties, attitude control, microgravity, orbit lifetime, and reboost requirements are

provided along with a des(;dption of these analyses. Also provided are brief descriptions of the elements

and systems that comprise these conceptual configurations.

17. Key Words (Suggested by Auth(,r(s))

Evolution Configurations, Growth

Requirements, Mass Properties, Attitude

Control, Microgravity Envelopes, Orbit

Lifetime, Reboost Analysis, Growth

Elements, Replication, Service Facilities

19. SecuriW Classif. (of this report)

Unclassified

18, Distribution Statement

Unclassified - Unlimited

Subject Category 12

20. Security Classif. (of this page)

Unclassified 21. No. 77 of pages

NASA FORM 1626 OCT 86

22. Price A05