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Produced by the NASA Center for Aerospace Information (CASI)
•

i^ Contract NASS-35499

MCR-85-1308

Phase it

Final Report February 1985

Selected Tither

Applications in

Space

(gAS9-cg-1714'4 S,,LfCiEC IEI67=, N8-J-?2463

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t,eport (Martla Mariatta CcLF-- ) 138 p

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FINAL STUDY REPORT

PHASE II STUDY OF SELECTED TETHER

APPLICATIONS IN SPACE

CONTRACT NAS8-35499

FEBRUARY 1985

PERFORMED BY

MARTIN MARIETTA AEROSPACE

DENVER DIVISION

FOR

MARSHALL SPACE FLIGHT CENTER

HUNTSVILLE, ALABAMA

v
r7.

FOREWORD

This Final Study Report is submitted in accordance with the requirements o

Contract NAS8-35499, Statement of Work paragraph 6.5, and Data Requirement

DR-4. The study was performed under the technical direction of James K

Harrison, Contracting Officers Representative.

The study was performed by the Space Systems Division of Martin Mariett

Aerospace, Denver Division under Mr. Morris H. Thorsen, and in the Spacecraf

Systems Product Area under Mr. Lester J. Lippy.

The Study Manager was Mr. William Nobles. Mr. Jack Van Pelt was responsibl

+ for the comparison analysis task.

Technical consultation was provided by Mr. Joseph Carroll, Research an

Consulting Services, u,,der subcontract to Martin Marietta.

1

Areas of the study healing with tether transportation applications benefite

from the studies performed by Professor Manual Martinez - Sanchez, Sarah An

Gavit, and Dale Stuart of tl►e Space Systems Laboratory of the Massachusett

Institute of Technology.

Insight into the behavior of tethered platforms was gained from the studies o

tethered constellations performed by Dr. Enrico Lorenzini and hi

co-investigators, Mr. David Arnold, Mr. Jack Slowey and Dr. Mario Grossi

This work was performed at the Astrophysical Observatory of the Smithsonia

Institute.

i

W

i
TABLE OF CONTENTS

PAGE

FOREWORD ... .......................................................... i

LIST OF ABaREVIATIONS ............................................... vii

a 1.0 INTRODUCfIUN AND OVERVIEW.................................... 1

1.1 Concept Development Activities ......................... 1

1.2 Concept Compatibility .................................. 3

1.3 Compatibility/Conflict Considerations

1.4 Benefits Assessment .................................... 7

1 .5 Tether Rendezvous System . ............. 0................ 7

1 .6 Tether Deployer Systems ................................ g

2.0 SELECTED CONCEPTS ANALYSES AND ASSESSMENTS ................... 9

2.1 Concept A2 - The Tethered Deoroic of Shuttle from

Space Station .......................................... 10

2.1.1 Tether Recoil Characteristics ................. 12

2.1.2 Deployment and Retrieval Energy

Considerations ................................ 13

2.1.3 System Concept ................................ 15

2.2 Concept F - The Tether Assisted Launch of an

OTVMission ............................................ is

2.2.1 Requirements Coossonality with Concept A2 ...... 115

2.2.2 Reel Size and Design Concept ................ 17

2.2.3 Payload Interface Deployment Module (PIDM) .... 17

2.3 Concept E2 - An Electrodynamic Tether Auxiliary

Po-ter System for Space Station ......................... 18

i

2.3.1 System Design Requirements .................... 19

2.3.2 Tether Power Dissipation ...................... 19

2.3.3 System Efficiency ............................. 20

2.3.4 System Power Levels ........................... 24

2.3.5 Conductor Mass ................................ 24

2.3.6 Tether Length Determination ................... 26

2.3.7 Tether Voltage Considerations ................. 28

2.3.8 Current Considerations ........................ 29

2.3.9 Electrical Potential Considerations ........... 30

2.3.10 Tether Construction ........................... 32

2.3.11 Tether Temperature ............................ 33

2.3.12 Cross Track Libration Effects ................. 34

2.3.13 Tether System Characteristics ................. 36

2.4 Tethered Platform from Space Station - Concept C ....... 36

ii
TABLE OF CONTENTS (CONTINUED)

PAGE

2.5 Tether Mediated Rendezvous of an Orbital Maneuvering

Vehicle with an Orbital Transfer Vehicle - Concept D ... 36

2.6 AXAF S3rvicing and Tether Reboost - Concept B2 ......... 37

3.0 COMPARISON WITH ALTERNATE CONCEPTS ........................... 38

3.1 Comparison Criteria .................................... 38

3.1.1 Mission Model Analysis ........................ 38

3.1.2 Tether Deployment Considerations .............. 39

3.1.3 Performance Characteristics ................... 43

3.1.3.1 STS Performance .o .................... 43

3.1.3.2 OMV and OTV Performance .............. 48

3.1.3.3 Space Station Performance ............ 56

3.2 Comparison Format ...................................... 58

3.3 Comparison Results ..................................... 58

3.3.1 Tether Assisted Deployment of 220 klb Shuttle

from the Space Station - Maximum Mass

Deployment Case - Concept A2 .................. 58

3.3.2 Tether Assisted Launch of 20,000 lc Payload

to Geosynchronous Orbit - Concept F ........... 61

3.3.3 Placement of 20,000 lb AXAF Into 320 nmi

Orbit Via Orbiter - Concept 51 ................ 64

3.3.4 Retrieval of 20,000 lb AXAF from 205 mi

Circular Orbit for Maintenance and rteboost

to 320 cmi - Concept B2 ....................... 66

3.3.5 Tethered tMiV Rendezvous and Retrieval of OTV

Returning from a Geosynchronous Mission -

Concept D ..................................... 69

3.4 Benefits Analysis ....................................... 72

4 .0 TETHER DEPLOYMENT SYSTEMS .................................... 78

4.1 Deployment System Categories ........................... 78

4.2 System Performance Requirements for a

Tether Deployment System for Space Station ............. 79

iii
TABLE OF CONTENTS (CONTINUED)

w ar

E •

PAGE

4.2.1 Reference Mass ................................ 80

4.2.2 Reference Altitude(s) ......................... 80

4.2.3 Tether Lengths/Tensions ....................... 80

4.2.4 Shuttle Interface Deployment Module (SIDM) .... 80

4.2.5 Payload Interface Deployment Module (PIDM) .... 81

4.2.6 Deployment and Retrieval Energies ............. 82

4.2.7 Tension Alignment Assemblies .................. 83

4.2.8 Tether Reel Assembly(s) ....................... 83

4.3 Design Concept for a Dual Mode Tether Deployment

System for Space Station ............................... 83.

4.4 Performance Requirements for a Tether Deployment

System for Shuttle ..................................... 105

4.4.1 General Requirements .......................... 105

4.4.2 Payload Interface Deployment Module (PIDM) .... 105

4.4.3 Extendable Deployment Boom .................... 106

4.4.4 Reel Drive System ............................. 106

4.5 Tether Deployment System for Shuttle Payloads .......... 107

4.5.1 Payload Deployment Constraints ................ 107

4.5.2 Deployer Design Concept ....................... 107

4.5.3 Operations Sequence ........................... 112

5.0 CONCEPT TECHNICAL ISSUES ..................................... 116

5.1 Concepts A2 and F ...................................... 116

5.2 Concept E2 ............................................. 116

5.3 Concept 8 .............................................. 117

5.4 Concept C .............................................. 117

5.5 Concept D .............................................. 117

6.0 CONCEPT TECHNOLOGY DEVELOPMENT NEEDS ......................... 118

6.1 Concepts A2 and F ...................................... 118

6 .2 Concept E2 ............................................. 118

6.3 Concept 8 .............................................. 119

6.4 Concept C .............................................. 119

6.5 Concept D .............................................. 119

7.0 CONCLUSIONS .................................................. 120

8.0 RECOMMENDATIONS .............................................. 122

APPENDIX A EQUATIONS AND aEFERENCES USED FOR DETERMINATION

OF PROPELLANT REQUIREMENTS ............................. A-0

iv
rAVf.

LIST OF ILLUSTRATIONS

FIGURE TITLE

2-1 Dual Mode Tether Deployment System Installed on

Space Station .................................................

2-2 Tether Deorbit of Shuttle from Space Station ..................

2-3 OTV Tether Launch 'Assist ......................................

2-4 Load Power as a Function of KT ..............................

2 -5 System Efficiency as a Function of KT and Tether Voltage ....

2-6 Reboost Propellant Requirea as a Function of I ggP ............

2-7 Mass to Power Relationship for an Aluminum TetFier .............

2-8 Tether Tension Constraint .....................................

2-9 Tether Angle Relationships ....................................

2-10 Potential Map for as Electrodynamic Tether System ............0

2-11 Tether Construction Using a Bunch Stranding ...................

2-12 Variation with Time for Cross Track Component of

Tether Force ..................................................

3-1 Tether Transportation Principles ..............................

3-2 Shuttle OHS Direct Insertion Delta V Requirements .............

3-3 Shuttle OHS Deorbit Delta V Requirements ......................

3-4 Shuttle OHS Propellant Requirements/Savings :..................

3-5 OMV Propellant Requirements for Mission Set ...................

3-6 OMV Propellant Savings from Tether Launches ...................

3-7 OTV Propellant Requirements for Mission Set ...................

3-d OTV Propellant Savings from Tether Launches ...................

3-9 Potential Space Station Quarterly Altitude Losses .............

3-10 Shuttle Quarterly Tether Deployment Requirements to.

Maintain Space Station Altitude ...............................

3-11 Potential Yearly Benefits from Space Station Tether

Operations ....................................................

14

16

16

21

21 

25

25

27

27

31

31

35

42

45

45

49

54

54

55

55

73

75

76

4-1 Space Station with Dual Mode Tether Deployment System ......... 85

4-2a Tether Reel System Installation ............................... 88

4-2b Tether Reel System Installation ............. ................. 89

4-3 Tether Splice Fitting ......................................... 90

4-4a Shuttle Tether Deployment System .............................. 92

4-4b Shuttle Tether Deployment System .............................. 93

4-5 Space Station/Shuttle Attitude During Depl-3yment .............. 94

4-6a Shuttle Interface Deployment Module (SIDM) .................... 96

4-6b Shuttle Interface Deployment Module (SIDM) .................... 97

4-7 SIDM/Shuttle kelease Mechanism ................................ 100

4-8 Shuttle Release Mechanism Sequences ........................... •101

4-9 Manual Release Sequence for Shuttle Release Mechanism ......... 102

4-10 OTV Tether Deployment System .................................. 104

4-11 Deployed Payload Mass Versus Tether Length .................... 108

4-12a Tether Deployer System for Shuttle .................0.......... 110

4-12b Tether Deployer System for Shuttle ............................ 111

4-13a Shuttle Payload Deployment (Stage 1) .......................... 114

4-13b Shuttle Payload Deployment (Stage 2) .......................... 114

4-13c Shuttle Payload Deployment (Stage 3) .......................... 115

ir, 
LIST OF TABLES

s

TABLE TITLE PA

1-1 Compatibility Matrix for Space Station

Tether Applications ........................................... 

4

3-1 Tethered OMV Candidate at Space Station ....................... 40

3-2 Tethered OTV Launch Candidates at Space Station ............... 40

3-3 Basic Tether Equations ..................................•0..... 42

3-4 Orbiter Initial Weight Determination .......................... 47

3-5 OMV and OTV Considerations .................................... 51

3-6 OMV and OTV Estimated Vehicle Characteristics ................. 52

3-7 OMV and OTV Performance Characteristics ....................... 52

3-8 Space Station Ground Rules .................................... 57

3-9 Space Station Orbit Maintenance Requirements .................. 57

3-10 Comparison of Concept A2 with Baseline Approach ............... 59

3-11 Comparison of Concept F with Baseline Approach ................ 62

3-12 Comparison of Concept Bl with Baseline Approach ............... 65

3-13 Comparison of Concept B2 with Baseline Approach ............... 67

3-14 Comparison of Concept D with Baseline Approach ................ 70

3-15 Potential Benefits from Space Station Tether Applications ..... 76

ft

^ r
LIST OF ABBREVIATIONS 

ACC - Aft Cargo Carrier

E ASE - Aeros;,ace support equipment

AXAF - Advanced X-Ray Astrophysical Facility

CUR - Contracting Officers Representative

Delta V, d V - Maneuver velocity increment

ET - External Tank

ETR _ Eastern Test Range

EVA Extra vehicular activity

fps - Feet per second

GEO - Geosynchronous equatorial ori;it

I sp - Propellant specific impulse (sec)

LH2

- Liquid Hydrogen (cryogenic fuel)

L02 , LUX - Liquid Oxygen (cryongenic fuel)

MECO - Main Engine Cutoff

MMH - Monomethyl hydrazine, (storable propellant fuel)

MPAD - Mission Planning Analysis Division O SC)

MPS - Main Propulsion System (STS)

MR2p

- Angular momentum

N 2- Nitrogen (gas)

N

2 H4

_ Hydrazine (storable monopropellant)

2.04 Nitrogen Tetroxide (storable bipropellant oxidizer)

nmi - Nautical mile (6076 feet)

OMS - Orbital Maneuvering System (Shuttle)

OMV - Orbital Maneuvering Vehicle

OTV - Orbital Transfer Vehicle

PIDM - Payload Interface Deployment Module (tether system)

psf - Pounds per square foot

Q(max) - Maximum dynamic pressure during atmospheric ascent

trajectory

RCS - Reaction Control System

SIDM - Shuttle Interface Deployment Module (tether system)

STAIS - Selected Tether Applications in Space (study)

STS - Space Transportation System

TOMV - Tethered Orbital Maneuvering Vehicle

TSS - Tethered Satellite System

Wpu - Weight of useable propellant

Vii
INTRODUCTION AND OVERVIEW

This report covers the results from study Phases I and II of 

planned five phase study of selected tether applications in space

During Phase I a large number (26+) of application concepts we

examined and five were selected for more detailed development an

evaluation. The criteria used for selection were:

- Variety of applications

- Near term applicability

- Potential benefits

- Feasibility/practicality

e

The five selected concepts were developed to identify operationa

characteristics, performance levels and limitations, operation

considerations,' safety considerations, and technical issues requiri

• further study. These developed concepts were documented in the fina

report for the Phase I study. These concepts as described -'-I th

Phase I Final Report provided a baseline for the Phase II stud

effort.

The five tether application concepts selected for study were:

A. Tether de-orbit of Shuttle from Space Station

B. Orbit insertion of a spacecraft (AXAF) from Shuttle

C. Tether deployment of an operational platform from Space Station

D. Tether effected rendezvous of an Orbital Manuevering Vehic

with a returning Orbital Transfer Vehicle.

E. An electrodynamic tether used in a dual motor/generator mode 

an energy storage method for Space Station.

1.1 Concept Development Activities

Based on the information developed and the insight gained during t

Phase I studies, Concepts A, B, and E were significantly revised a

an additional new Concept F was developed during Phase II.

Concept A2 is the designation used for the revised version of t

shuttle deorbit. This revision was made to incorporate new data 

• the projected Space Station characteristics and to remedy certa

problems with the original concept. The significant changes were 

shorter tether, due to the significantly increased mass of the Spa

Station, and a revision in the concept for the Shuttle Interfa

Deployment Module (SIAM). The original Concept A required a brid

beam to be installed in the Shuttle cargo bay to serve as the teth

attachment interface. This bridge beam had to be brought up in

orbit each time it was to be u-7-d and it interfered with return car

capabilities. Concept A2 was modified to attach to the Shuttle sil

such that the interface scar weight penality is significantly reduce

and there is no interference with return cargo. The most significa

1
aspect of the revised concept is the addition of a provision to

scavenge the propellant from the Shuttle which is no longer needed

due to the tether deorbit. This Orbit Maneuvering System (OMS)

propellant must be retained on board the Shuttle far the contingency

of an aborted tether deorbit. However, for each increment of te.ther

deorbit successfully accomplished, the amount of propellant needed is

reduced. The ratio is approximately 100 pounds of OMS bipropellant

per kilometer of tether deployment. For the full range deployment of

65 kilometers of tether, 6500 pouuda of propellants is no longer

required. This consideration led to the incorporation. of a

propellant scavenging system into the SIDM. This feature allows the

propellant to be transferred from the Shuttle as it becomes surplus

and then retrieved by the tether to the Space Station for use in

resupply of the Orbital Maneuvering Vehicle (OEV).

This propellant scavenging feature greatly increases the overall

benefits of the tether deorbit for Shuttle and is considered to be

one of the significant results of this study. A new technology

report has been filed to describe the concept in more detail.

Concept F consists of a tether assisted launch of an Orbital Transfer

Vehicle (OTV) mission to geosynchronous orbit from Space Station.

The reason for adding Concept F was the realization gained from the

analysis of Concept A that th1 tethered deorbit of the Shuttle must

be coupled with another tethes application concept which would in

turn, make use of the angular momentum transferred to the Space

Station by the Shuttle deorbit operation. Otherwise, the number of

deorbit operations which could be performed would be severely limited

by the unacceptable increase in Space Station altitude. A survey of

potential applications which could act as such a user of angular

momentum identified the OTV launch assist operation as the only

viable tether transportation candidate to serve as the angular

momentum counter balance'to the Shuttle deorbit operations.

Concept B2 is a variation on the original Concept B which used a

Shuttle mounted tether system for orbit insertion of the AXAF

observatory spacecraft into its operational orbit. The Phase

analysis indicated that the concept did not provide any significant

performance advi.ntages over the baseline concept of direct insertion

by Shuttle. B2 was developed to consider the servicing mission for

AXAF which requires rendezvous of the Shuttle with the AXAF

spacecraft at 205 nautical mile altitude, refurbishment on orbit and

then use of a tether system to perform reinsertion of P.XAF to 320

nautical miles for a second 3 year operational interval.

Concepts C and D were not significantly modified from Phase I.

Concept E2 is a significantly revised concept from the Phase I

Concept E. Concept E used the electrodynamic tether as a two-way

conversion device between orbital mechanical energy and electrical

energy. The purpose was to provide a method of energy storage to

sustain the Space Station power system during solar eclipse

intervals. The basic source of the energy was solar array power.

2
generator mode (converting orbital mechanical energy into electrica

power). The orbital mechanical energy is derived from the teth

deorbit of Shuttle (Concept A2). This concept can provide a

alternative method of using the angular momentum scavenged by t

tether deorbit of Shuttle.

These concepts are discussed n more detail in 2.0.

1.1.1 Concepts Baseline for Phase !I

The concepts baseline for the Phase II study which incorporates t

s concept development activities performed under study task I is list

below.

Concept A2 - Tether Deorbit of Shuttle from Space Station.

" (Major revision from Phase I)

Concept B - Tethered Orbit Insertion of a Spacecraft fro

Shuttle, (Same as Phase I)

Concept C - Tethered Platform Deployed from Space Station

(Same as Phase I)

Concept D - Tether Effected Rendezvous of an OMV with 

Returning OTV. (Same as Phase I)

Concept E2 - Electrodynamic Tether as an Auxiliary Pow

Source for Space Station. (New Concept)

Concept F - Tether Assisted Launch of an OTV Mission fr

Space Station. (thew Concept)

1.2 Concept Compatibility

For those concepts involving use on the Space Station an assessme

was made of the compatibility.

The results of thin assessment are shown in Table 1-1. T

categories of compatibility used on the table are:

Y (yes) - Compatible for Use on Space Station

j Y*(yes) - Compatible with Shared Use of Tether Deployer Sy

N (no) - Not Compatible

1.2.1 A2 - Shuttle Deorbit with C - Tethered Platform

These concepts are not compatible because the use of a tether

platform at the upper deployment location would conflict with the u

of either of the two identified concepts for using the angul

momentum scavenged from the Shuttle deorbit operations (i.e. E2 or F

3

-• 
4

14

. ^ a;
These concepts are compatible because of the intermittent nature 

application. They should be also compatible in the shared use of 

deployer system on the down end of the Space Station.

1.2.3 A2 - Shuttle Deorbit with E2 - Electrodynamic Power Tether

These concepts are compatible with the Shuttle deployment tether 

the down end of the Space Station and the electrodynamic tether 

the upper end.

* In addition to being compatible, these two concepts form 

complementary set in that the Shuttle deorbit imparts angul

momentum to the Space Station which the electrodynamic tether in tu

converts into electrical power.

•

1.2.4 A2 - Shuttle Deorbit with F - OTV Launch Assist

These concepts are compatible with the Shuttle deployment system 

the down end of the Space Station and the OTV deployment system 

the upper end. An approach has been developed for the two concep

to share a common reel system (See 5.0).

As with the preceding set (1.2.3) these two concepts form 

complementary set with the OTV launch assist using the angul

momentum imparted to the Space Station by the Shuttle deorbit.

1.2.5 C - Tethered Platform with D - Tethered Rendezvous

These concepts are compatible but with no obvious sharing of hardwa

systems. The platform could be on either top or bottom end of t

Space Station as could the tethered rendezvous, but they could n

share a location.

1.2.6 C - Tethered Platform with E2 - Electrodynamic Power Tether

These concepts are not compatible because the electrodynamic pow

4 

system must be coupled with the Shuttle deorbit concept for i

source of energy. Also, with the lower end of the Space Stati

allocated to the Shuttle deorbit, these two would be required 

share the top end location and this is not feasible.

At

C - Tethered Platform with F - OTV Launch Assist

These concepts are not compatible. The logic here is identical 

that for 1.2.6 above. The tethered platform and the OTV laun

assist would both require a top end location on the Space Station a

they are not compatible for colocation because of conflicti

operational requirements.

5
1.2.8 D - Tethered Rendezvous with E2 - Electrodynamic Power Tether

These concepts are compatible for use on the Space Station but not on

the same end. If the electrodynamic tether is used it must be on the

upper end (See 1.2.3) and the operational requirements are

conflicting. However, the tethered rendezvous operations could time

share the lower end deployer system.

1.2.9 D - Tethered Rendezvous with F - OTV Launch Assist

These concepts are compatible for use on the Space Station because'

they are both intermittent in operation. Further they should be

compatible with the common use of the top end deployer system.

1.2.10 E2 - Electrod ynamic Power Tether with F - OTV Launch Assist

These concepts are not compatible as they both require top end

location and they are operationally incompatible for same end

location.

This is unfortunate because these tuo concepts share the capability

to productively use the angular momentum scavenged from the Shuttle

deorbit operation. This means that one or the other must be selected

for the top end location on the Space Station.

1.3 Compatibility/Conflict Considerations

The unfortunate incompatibility between the electrodynamic power

tether and the ON launch assist concepts identified (1.2.10) mean

that the two primary methods identified to productively utilize the

angular momentum made available by the Shuttle deorbit operation

cannot both be used during the same time period. This means that

some form of choice must be made between them or a transition made

from the electrodynamic power tether to the OTV launch assist at the

time the space based OTV comes on line circa 1995.

For the purpose of this study, it has been assumed that the primary

choice would be the OTV launch assist concept and a dual mode

deployer concept developed in accord with this assumption.

If an alternate choice were to be made, either for the electrodynamic

power tether to continue in use with no OTV launch assist, or for a

transition scenario where the initial power tether would be

decommissioned and the OTV laun--h assist equipment installed; then 

the design concept approach for an integrated Space Station tether 

system installation should be reexamined.

6
1.4 Benefits Assessment

The developed concepts were analyzed to provide a comparison of their

performance with more conventional non-tether methods of

accomplishing the same missions. In many cases the alternative

methods are not clearly defined and operational performance data is

not readily available. A significant portion of effort under the

study was dedicated to developing this baseline data base on the

performance characteristics of existing and planned space operations

systems. This includes Shuttle, OMV, OTV, and Space -Station. In

addition to the baseline performance characteristics of these

systems, it was necessary to identify the impact on their operations

* characteristics du! to the use of tether systems.

Another aspect of she benefits analysis, was the insight gained on

the angular momentum balance approach for Space Station. The

significance of this was originally pointed out by Ivan Bekey of NASA

headquarters staff during the course of the Phase I study. Analysis

of the Space Station mission models for OTV and OMV operations was

required to provide input data for the angular momentum balance

analyses and the benefits assessment.

Comparison criteria were developed for comparing tether applications

with alternate methods and a comparison format was developed.

Comparisons were completed for all concepts except for C where data

was not available on the operational characteristics of the

alternative platform systems, and Concept E2 where there is no clear

alternative method to providing such an auxiliary power system.

These comparison studies and benefits assessment results are

described in 3.0.

1.5 Tether Rendezvous Systems

The Phase I study identified a concern as to the feasibility of the

type of orbital rendezvous required by Concept D. The nature of this

rendezvous operation between a free flying spacecraft and a tethered

platform must be well understood in order to develop performance and

design requirements for hardware systems required.

Difficulties were encountered in accomplishing the orbit trajectory

simulations required to better define these system performance

requirements. Some limited progress was made in defining variations

on the rendezvous procedure which could enhance the probability of

successfully accomplishing the operation. However, it was not deemed

advisable to proceed with an attempt at hardware definition until an

improved understanding of the concept is achieved. The study effort

originally allocated to the definition of tether rendezvous hardware

under this task was redirected to other tasks.

7
8

1.6

A further development bearing on this tethered rendezvous concept

area was the results from the comparison/benefits analysis task.

Based on the most recent information on the mission operations

concept for the return of the OTV from geosynchronous orbit to Space

Station, there is no performance advantage for the tethered

rendezvous concept - even assuming the rendezvous car. be successfully

accomplished. This reduces the incentive to continue study

activities directed at this approach unless some more beneficial

versions of tether mediated rendezvous operations can be identified.

Tether Deployment Systems

The combination of concepts A-2 and F were used as the basis for a

dual mode tether deployer system concept compatible with the Space

Station.

Concepts B1 and B2 were used as the basis ro develop a generalized

tether deployer system for Shuttle payloads. This system is a

derivative of the Tethered Satellite System hardware.

These systems are described in 4.0.
s

In general, the study activities under this task were directed towar

the modification of the concepts studied during Phase I. The nee;-

for these modifications was based on conclusions reached at the en

of the Phase I study, improved insight gained during the study an

the availability of new data on operational characteristics o

projected space systems.

One new concept was added - Concept F: A tether assisted launch o

OTV from Space Station. This concept was needed to act as a user o

the angular momentum transferred from the departing Shuttle to th

Space Station by the tether deorbit operations of Concept A2. Th

necessity to zorrelate these concepts, where one (Concept A) impart

angular momentum to the Space Station and the other (Concept F

extracts it for use in launching another vehicle (OTV mission), is a

example of the application of insight gained during the course of th

Phase I study.

Concept E2 is a major revision of the electrodynamic tether concep

studied in Phase I, and was also based on increased insight gained

Concept E2 is for an electrodynamic tether auxiliary yuwer system

which provides an alternative method (to Concept 0 to utilize th

angular momentum available from the shuttle deorbit operation b

converting it into electrical power for use on the Space Station

This capability to provide auxiliary power at levels up to 75 k

without impacting the baseline power system design would be 

valuable adjunct to Space Station capabilities. In addition, use o

the concept would enable an ancillary benefit from the scavenging o

Shuttle OHS propellant. Additional uses of the tether deorbi

capability which could be considered are the deorbit of Space Stati

waste containers to burn up in the atmosphere, and the deorbit o

external tanks which have been brought to the Space Station. Thes

applications have not been developed into concepts because of th

inherent limitation on tether deorbit operations set by th

capability to use the scavenged angular momentum, and the resultin

judgement call that the most beneficial candidate for tether deorbi

is the Shuttle. Nevertheless, it should be kept in mind that ther

is much more angular momentum potentially available to be scavenged

The caveat is that additional concepts must be identified which, 

turn, use the angular momentum for further purposes. Conversion o

the scavenged angular momentum into electrical power is such 

concept..

Unfortunately it does not appear to be feasible to use Concept E2 an

F concurrently since they must both be installed on the upper end o

the Space Station. While the electrodynamic tether could operate 

the lower or nadir end, this location is preempted by the tethe

deorbit application for Shuttle. The OTV deployer must be on th

zenith location, and so, there results the conflict between E2 an

F. There is schedule window of opportunity for the application o

E2 prior to the advent of the space based OTV (presently planned fo

1995). For the purpose of this study evaluation, it was assumed tha

the benefits of Concept F would outweigh those of Concept E2 and th

benefits analysis was based on this assumption.

s

9

mss. 
Concept B2 is a modification to the mission operations for the

concept studied in Phase I. The modification is to perform a

refurbishment mission for the Advanced X-ray Astrophysical Facility

(AXAF) and then to reinsert it into its operational orbit by means of

a tether deployment from Shuttle. The original concept B had been to

perform the initial mission insertion to 320 nautical miles by tether

as an alternative to a Shuttle direct insertion mission. The

justification for this B2 modification to the original concept was to

determine if the refurbishment and reboost mission offered any

performance benefits in comparison with the planned baseline approach

for accomplishing the mission. Neither this modified concept nor the

original were found to provide any significant performance benefits.

Justification for continuing to develop the Shuttle deployer concept

will depend on identification of other orbit insertion applications

that require boost to higher altitudes than can be achieved by direct

insertion from the Shuttle.

Concept C - A tether deployed operational platform from Space Station

- was not modified. An alternative approach to supplying power to

the platform by means of a conductive tether was considered as a

worthwhile modification to be investigated, however, it was not 

undertaken during this study phase.

Concept D - A tether deployed OMV used to rendezvous with a returning.

OTV - was also not modified and the Phase I concept was used for the

comparison with alternative methods performed under- Task 2. The

significant new development with respect to this concept was new

information on the baseline mission operations concept for return of

the OTV from geosynchronous orbit to the Space Station. Comparison

with this baseline approach has brought out the fact that there is no

performance benefit for the concept and, hence, no reason to continue

with the development of the concept.

2.1 Concept A2 - The Tethered Deorbit of Shuttle From S pace Station

The major new development which motivated the modification of the

original Phase I Concept A, was the definition of the reference

configuration for Space Station. This configuration information was

developed to serve as a basis for the Phase B Space Station studies

and became available on a timely basis to be used in this study. The'

new information included significantly increased mass values for the

station and configuration layouts which permitted more meaningful

concepts. as to how such a tether, deployer could be integrated into

the Space Station.

In addition to this new information some problems had been identified

with the Phase I Concept A for which design solutions were developed

in A2.

10
One of these problems was the need to improve the means of attaching

the deployment tether to the Shuttle. The original Concept A

employed a bridge beam installed into the cargo bay near the Shuttle

center-of-mass location. This meant that n,,) return cargo could be

installed at that location, thus constraining the ret..rn cargo

capability of the Shuttle. Also, the bridge beam had to be brought

up into orbit for each mission where its use was planned. The design

solution was to develop a new Shuttle Interface Deployment Module

(SIDM) design which provides clearance for the return payloads and

interfaces directly with the Shuttle sills by means of relatively

lightweight attachment/release fixtures.

An insight gained from the Phase I study was the recognition of the

significant quantity of Shuttle OMS propellant potentially available

for scavenging. The problem here is the requirement to have this

propellant available on board the Shuttle in the event of an aborted

tether deorbit operation. A method concept was developed to provide

a propellant transfer/scavenging capability in the new Shuttle

Interface Deployment Model (SIDM) design such that the propellant can

be transferred into the SIDM as it becomes surplus to the Shuttle

requirements with tether deployment. A new technology disclosure

report has been submitted describing this concept. A concept drawing

of the SIDM incorporating the propellant storage capability is shown

in the Space Station tether deployment system described in 4.2.4.

Studies were made to quantify the amount of OI.S propellant which

could be made available by this approach. The relationship is

approximately 100 pounds of propellant for each kilometer of tether

deployment of the shuttle. Using a maximum length tether of 65

kilometers allows 6500 pounds to be scavenged per each full deorbit

deployment. This is the most significant item in the benefits

assessment for this concept as discussed in 3.0.

An operational consequence of this new SIDM design concept is that

the requirement to perform the tether deployment with the payload bay

doors open is even stronger than for the original concept. This

means that the requirement to keep the post release perigee altitude

for the Shuttle above 100 nautical miles to avoid an automatic

reentry on first perigee pass is still in effect. This is what

determines the limit of 64+ kilometers on tether length for Shuttle

deorbit.

A second operational consequence is due to the fact that the SIDM no

longer fits within the cargo bay envelope, and so, could not be

retained for return co earth in event of a broken tether. This puts

an increased emphasis on fault tolerant design for the release

mechanisms to insure release capability. It also leads to the

requirement for an OMV recovery method capability for the SIDM in the

event it must be jettisoned. Eunctiot ► al requirements to support such

a recovery operation have been included for the SIDM concept.

it
A third consequence of this SIDM concept is the necessity to install

propellant transfer lines into the cargo bay area for those Shuttles

intended to be used for the tether deorbit operation. For this

i study, it has been assumed that such propellant transfer and

management technologies will have been developed to support space

systems such as OMV and will be available for this application.

This OMS propellant transfer capability provided by the SIDM would

provide some attractive ancillary benefits for the Space Station

based Shuttle missions. For those Shuttle missions which are volume

limited, some or all of the excess payload weight capacity could be

used to bring up added OMS propellant thus providing a convenient

method of achieving improved payload performance. In addition, the

system would provide a capability to off-load mission contingency

reserves of propellant as the need for these reserves is reduced near

the end of a mission. In the reverse sense, such system would

provide an on-orbit capability to replenish the Shuttle propellant in

the event of unanticipated mission requirements. The SIDM would

provide this two way propellant transfer capability.

There is an obvious requirement for the SIDM to interface with the

propellant storage depot on the Space Station in order to perform the

propellant transfer operations described above. Because of lack of

information on the location and design details of this propellant

depot, no attempt was made to define the interface requirements or

operational procedures to transfer propellant from the SIDM to the

depot or vice-versa.

A fourth consequence of the new SIDM concept is an increase in the

amount of energy required to perform the post release retrieval

operation for the tether and SIDM. The structural mass of the SIDM

is increased as well as the added mass of propellant to be

retrieved. Using an estimated mass for the SIDM of 3000 pounds, 6500

pounds of OMS propellant and 65 kilometers of tether at 163 pounds

per nautical mile (40 kg/km) gives a retrieval energy requirement of

11 kWh.

The energy generated during deployment remains approximately the same

as for the earlier concept. While the length of the tether has

decreased by about 11 km, the mass of the orbiter has been increased 

to 220,000 pounds to account for return payload. The energy

generated during the deployment is 155 kWh.

2.1.1 Tether Recoil Characteristics

One of the technical issues identified during the Phase I study was

the nature of the tether recoil process subsequent to release or to

an accidental sever of the tether by debris. Preliminary simulation

indicated a possibility that the portion of the tether adjacent to

the station would recoil up past the station on the aft end and then

fall back down under the influence of gravity gradient forces.

12
Is

t a

•

of this tether release process to give improved resolution in the

vicinity of the station. Our distributed mass tether simulation

model ( ORBNET) in its present formulation does not have the

capability to perform the needed recoil simulations.

This question of tether recoil behavior remains as one of the key

technical issues for those tether transportation applications which

involve tlie release of a tethered body under conditions of

significant tension in the tether. A simulation capability is needed

to be able t, predict with confidence the detailed nature of the

recoil dynamics and, in particular, the behavior of the recoiling

tether in that vicinity of the tether deployment platform. The

simulation should also include the capability to study effects due to

viscoelastic characteristics of the tether material and auxiliary

damping effects due to core and jacketing materials. Such a

simulation would permit the performance evaluation of alternative

tether materials such as graphite or boron fibers to determine

possible advantages of these materials.

Methods of attenuating the release process to mitigate the recoil

effects were considered, but they would still leave the problem that

the severed tether case must also be accommodated. This means that

the design solution to the tether recoil problem should be sought in

the design of the tether itself. This could result in some form of

composite construction which includes the basic tension element, a

jacketing material optimized for abrasion and environmental

resistance, and a recoil damping material.

These related areas of tether recoil simulacion and tether

construction are recommended for further study effort in order to

support the development of tether transportation applications.

2.1.2 Deployment and Retrieval Energy Considerations

The quantity of energy generated during the deployment of the shuttle

(153 kWh) and the smaller, but still significant (11 kWh), amount

required to accomplish retrieval of the tether and propellant loaded

SIDM are a fruitful area for innovative engineering solution. The

tether reel drive and braking functions are performed by means of a

motor / generator drive unit. The energy developed by the system

during the deployment process appears as electrical power. For

purpose of this study, it has been assumed the deployment generated

energy would be rejected to space as waste thermal energy by a

resistive load bank which functions as a dedicated high temperature

radiator. Subsequently, during the retrieval of the tether the

retrieval energy required would be supplied from the Space Station

power system.

In the case of Concept F for the tether assisted launch of an OTV

mission, the deployment and retrieval energies are even larger than

for the Shuttle deorbit.

13
OTV 10, PAYLOAD

MOUNTED ON

UPPER TETHER

DEPLOYMENT ASSAY

1 ^

DUAL MODE

TETHER REEL

ASSEMBLY

y

SHUTTLE SHOWN READY^y

TO BE TETHER

DEPLOYED

L+arc "-7 D ual Afode Tether Deployment System Installed on S;pace ^nation

14
In order to enhance the overall efficiency of these transportatio

applications, improved concepts are needed for the utilization and/

storage of the energies involved in the operations. This may be a

area where the flywheel energy storage concept could be applied

This is another area recommended for follow-on study.

2.1.3 System Concept

Concept A2 has been developed in a complementary relationship wit

Concept F and is treated in 2.2. The resulting deployment system i

used as the basis for a Space Station deployment system trea..ed 

4.0. The resulting dual mode system is shown installed into th

current Space Station reference configuration structure in Figur

2-1. The Shuttle is shown in position just prior to release from th

Space Station. Figure 2-2 shows the post release configuration o

the Shuttle tethered deployment operation. The SIDM is show

installed on the Shuttle with the OMS scavenging propellant tank

indicated. The tether tension alignment boom and carriage are sho

on the lower portion of the Space Station. More detail

illustrations of these hardware elements are given in 4.0.

2.2 Concept F - The Tether Assisted Launch of an OTV Mission

The development of this new concept during the Phase II study wa

motivated by the results of the angular momentum balance approach t

the transportation application area. Analysis indicated that th

outstanding candidate concept to be considered in the role of a use

mission for the scavenged angular momentum is the OTV mission launc

assist. Figure 2-3 shows the OTV with payload being deployed upward.

2.2.1 Requirements Commonality with Concept A2

Because of the correlated relationship with Concept A2, requirement

for the two concepts were examined for areas of commonality.

In considering tether tension, it was noted that a deployment leng

of 150 kilometers for the maximum weight OTV missions would result 

approximately an equivalent tension to that for the Shuttle deorbi

(Concept A2). This would allow a common design solution in the are

of tether construction, to reel drive systems, and to tensio

alignment systems. The only significant differences would be in t

amount of reel tether storage volume required, in the amount o

energy generated during deployment and the amount of energy require

for retrieval.

The deployment energy generated by the maximum weight OTV missio

stack is 366 kWh and the energy required for retrieval is 29 kWh

The OTV deployment energy generated and the length of time designate

for the deployment operation are the factors which size th

motor/generator drive system for the -reel. Using our guideline of 

hours for deployment results in an average power output from the ree

braking operation of 46 kW or 62 hp. Reserve capability woul

indicate a rating of 70 hp will be required for the drive motor.

15

---IV E..
NADIR PORTION OF

SPACE STATION

KEEL STRUCTURE

AUXILIARY DOCK I 14G

PORT

ION

IACE

1TTLE INTERFACE MODUL E

fH OMS SCAVENGING TANKS

ure 2-2 Tether Derrbit of

Tuttle from Spacg Station

10E BASED OTV/PAYL0.4D

fHER

4SION ALIGNMENT TRACK

NITH LOCATION ON SPACE

M ON KEEL STRUCTURE

F'i jure 2-3 OTV Tether

Launch Assist

ISION ALIGNMENT

4CK AND CARRIAGE

'HER

16

fl.) al !- 
deployment time than 8 hours, the reel drive braking motor will ne

to be correspondingly increased in size to accommodate the high

power levels. The high temperature radiator will also requi

uprating.

2.2.2 Reel Size and Desian Concert

Assuming a tether diameter of 0.400 inch (0.01) and a length of 1

km, the estimated reel size is 3 meters long and 2.5 meters in ou

diameter with a hub diameter of 0.3 meters. These are the insi

dimensions of the reel. Such a reel would mount in a transver

orientation in the Shuttle cargo bay for launch to orbit. T

approximate weight of the tether is 6300 kg. This includes a 

excess length over the 150 km.

• The reel is to be designed as a replaceable m,)dular assembly s

that when the tether has reached the end of its service life, t

tether loaded reel can be replaced as a unit and the worn out un

returned to the earth. An alternative possibility for disposing 

the worn out tether would be to deploy the full length (150 + km) 

the tether in a downward direction and release it at the reel end 

re-entry and burn-up in the atmosphere. It would still be necessa

to return the now empty reel for reloading on the ground a

retransport to orbit.

2.2.3 Payload Interface Deployment Module (PIDM)

The PIDM must be capable of meeting the deployment tension loads 

the grapple interface with the OTV mission stack and then releas

under remote control from the Space Station. The attachment locat

of the PIDM to the OTV is yet to be determined. For purposes 

developing this concept it has been assumed that the attachme

interface would be located in the vicinity of the UTV engine and i

line with the longitudinal thrust axis of the stack. This would Ke

the tension loads during deployment aligned in the same directi

(but opposite sign) as the propulsive thrust loads from the OT

This location for the attach interface would require that the O

mission stack rotate through a 90 • angle after release and prior 

initiating the first burn. Ideally this initial burn should 

centered on the orbit perigee to provide maximum efficiency. F

this tether case, the tether release point becomes the perigee of 

new orbit subsequent to release, and the initiation of the burn mu

be still further delayed while the stack is rotated through 9

(about a transverse axis) and then stabilized. The performan

penalties for this delay need to be quantified to determine if th

are acceptable. One alternative which could be considered it 

delay the burn until the subsequent perigee pass of the new orbit.

a

a

17
Other attachment locations on the side of the stack were considered,

however, they also have associated problems. While the OTV could bt

kept in a proper orientation for burn initiation, the variation in

payload characteristics from mission to mission would require

adjustment in the location of the grapple: fixture for each mission.

In addition, the deployment operation would present a significantly

different orientation of the tether tension induced load paths in the

s t 1c k.

This area of the attachment interface of the PIDM to the OTV stack

needs further analysis and joint consideration with the OTV

definition study teams to resolve an optimum attachment method and

'.ncation.

Other functional requirements for the PIDM are given in 5.4.2.

The tether deployment of the UTV stack from $. ► ace Station is shown in

Figure 2-3.

2.3 Concept E2 - An Electrodynamic Tether Auxiliary Power System for

Space a ion

Electrodynamic tether systems can be considered for a variety of

performance objectives. This includes ( 1) Use of the tether to

convert electrical power into thrust for orbit boosting, (2) The two

,say conversion from electrical energy to orbital mechanics,, energy

and vice-versa which provides a method to use the orbital energy as

an energy storage medium and (3) The use of the tether as an

auxiliary power system which converts orbital mechanical energy into

electrical energy in a single mode. The Concept E studied during

Phase I was the dual mode energy storage concept.

The results from the Phase I study indicated some fundamental

difficulties with the dual mode energy storage concept. The most

telling was the realization that the system would always require a

full scale back up of conventional design for contingency

situations. In addition the mode switching transition of the tether

which must occur twice per orbit will significantly reduce the

operational f raction of the orbit period that can be devoted to

either the thrusting or gc aerating mode.

The new insight gained from the Phase 1 study led to the realization

that an electrodyn mic tether auxiliary power system could provide a

method to beneficially use the angular momentum available to be

scavenged from the shuttle deorbit by converting it to electrical

energy for use on the station.

18

^1 .s
2.3.2

•

Assumptions were made as to the practical sizing of such a system a

a system design logic sequence developed. These assumptions a

logic steps are described in the following paragraphs and th

resulting tether described. It should be kept in mind that this 

an example and that selecting other design requirements such as pow

level or overall system efficiency will result in variations on th

tether design. The intent during this study has been to develop 

design approach for the tether power system, and to apply it 

develop a typical example.

System Design Requirements

The system design requirements selected for use- in developing th

concept are as follows:

1. Deliver 25 kW of conditioned power to the Space Station pow

bus on a full time duty cycle.

2. No more than 5% of system power to be dissipated in the tether.

3. Capable of operating at a reserve power level of up to 75 k

delivered to the Space Station bus. (Power loss in the teth

will increase during reserve power operating intervals)

4. Tether angles with respect to the vertical are not to exceed 0

radian at maximum (75 kw) reserve power levels.

A mass of 250,000 kg was assumed for the Space Station and an en

mass for the tether system of 500 kg.

Tether Power Dissipation

The fraction of total --;stem power dissipated in the tether i

determined by the ratio of tether resistance to the total resistanc

in the circuit. This includes the net resistance of the ionospheri

current path, the contact resistances of the tether ends to th

ionosphere, the resistance of the station power processing circuitr

and the resistance of the tether.

This ratio has been designated KT , with the defining relationship:

KT = RT/R,

Where RT is the resistance of the tether and R is the tota

resistance of the circuit.

Treating the tether power system as a direct current system, th

value for KT also defines the ratio of power in the tether to tota

system power.

KT = PT/PS

Where PT is the power dissipated in the tether and P S is th

total system power.

19

2.3.1

4

W r .

_ r:
A simplified circuit for a tether of resistance RT and induc

voltage ET in series with a load resistance RL is shown in Figu

2-4. This plot shows the changing ratio of power in the load, P

to total system power, PS , as KT is varied. At the point whe

the values of RT and R

14 

are equal, the value of K T is 0.5 a

R/RT is 2. This is the impedance match condition for maximum pow

in the load for a given tether. This is shown. by the lower curve

PL with maximum value of 0.25 at R/RT = 2. The value for P S at

this point is 0.5 of the short circuit power level. As the value of

R/RT is increased the fraction of the power dissipated in the

tether, KT , decreases. The design point used for this concept is a

KT = 0.05 or R/RT of 20. At this point the system power has

decreased to 0.1 of the value at the impedance match point (RT =

RL), but the power in the load has only decreased to about 0.2 of

the value at maximum. This plot illustrates the point that a given

tether design can be operated at an increased power level above the

selected design point. The penalty is an increased fraction of the

system power will be dissipated in the tether.

Using our design criteria of 25 kW at a KT of 0.5 and the reserve

power level of 75 kW, the plot indicates that the point on the

abscissa where the value of PL has increased by a factor of 3 is at

approximately 6.5 and the corresponding value of K T at this point

is 0.15. Operating at this reserve power level would cause 15% of

the system power to be dissipated in the tether.

In order to be able to operate at both the design and reserve power

levels, both the power conditioning circuits and the tether angle

constraints must be designed for the reserve levels of 75 W. These

considerations will be applied at the appropriate points in the

design process.

2.3.3 System Efficiency

Using a definition of system efficiency, E, where

E = PB/PS,

and PB is the power delivered to the system bus, and P S is the

total power in the system. PS is equal to the rate of decrease in

orbital mechanical energy due to the drag effect of the

electrodynamic tether. Note that PL = PB + parasitic power loss,

where the parasitic power losses include the plasma contactors at

each end of the tether, the efficiency loss in the power conditioning

circuit and loss due to ionospheric heating.

20
R=RT+RL

-

FITRL

FIT = Tether Resistance

RL = Load Resistance

(including Ionosphere and

Contact Resistances)

E T KT = R T/R - PT'PS

jPS=PT+PL

_ PT = Power Dissipated in Tether

PL = Power Delivered to Load

KT

i ^ I

i

PL

i

0.9

0.8

0.7

0.6

7

0.5

Na

d 0.4

0.3

0.2

0.1

4

t

1 

1 2 3 4 5 6 7 8 9 10 20 30 50 1GO

R/R T (Also 1/KT)

Figure 2-4 Load Power as a Function of KT

i K

1 2 I 

AND K ARE POWER TO CURRENT RATIOS FOR PLASMA BRIDGES

ET= TETHER VOLTAGE (FUNCTION OF TETHER LENGTH)

LIMIT VOLTAGE e = DC/DC CONVERTER EFFICIENCY

R,= RESISTANCE OF IONOSPHERE

100

4 kV (30 km ) R = IOTAL SYSTEM RESISTANCE 10

7o

90 2kV(15b) 1

1 r —^--r t sti 9

80 — 8

1 kV (7 km)

/ -t 

i I i ac

10 ,-- ° 1

u

60 - 00 6

^

II

u

50 u 

-t — I -^ - I t 5

w j

40 i - 4

4

LL

30 W - 3

w _ POWER TO USER BUS

20 V

E TOTAL SYSTEM POWER

2

R I KI+e K2

10 yr f- E= e(1 -KT R) E

4 1

, KT

0.1 0.2 0.3 0.4 0.5

Figure 2-5 System Efficiency as a Function of KT and Tether Voltage

21

4ddiir. .. Jy ,} y
The plot of these relationships against the K T value is shown in

Figure 2-5. in this figure efficiency of the power conditioning

circuit is designated as "e", the resistance of the ionosphere plus

contact resistance as R 1 , and the induced voltage of the tether as

ET. the terms K 1 and K2 are the respective power loss terms

for the plasma contactors at the upper and lower ends of the tether.

The tether is assumed to be deployed upward from the Space Station

and that plasma contactors are used to provide a low resistance

contact to the ionosphere. These contactors have been assumed to

require power to operate as a linear function of the current. in the

system. As the tether voltage increases with a longer tether, the

current will decrease for a given power level. This gives the family

of lines for various voltage levels of the tether, ET.

To produce this figure values were assumed:

K = 50 W/electron ampere, 

K2 = 150 W/ion ampere,

e = 94%

The approximate tether lengths corresponding to the plotted values of

ET are based on an induced .-oltage of 133 volts per kilometer of

tether.

Figure 2-5 also shows a plot of the increase in conductor r.,ass as a

factor times the conductor mass required at a KT of 0.5 (or the

impedance match condition where RT = Rd . As can be seen the

tether mass required at KT = 0.05 is 10 times the mass at K T =

0.5.

A reasonable target range for the desigir power fcr this system is

assumed to be between the two inner (heavy line) curves for E T ('_

to 4 W. At a KT of 0.05 the resulting system efficiency E is in

the range from 79 to 83 percent. At the reserve power level where

KT = 0.15 the resulting system efficiency range is 69 to 74 percent.

The important consideration to keep in mind here is that the

significance of this system efficiency level is related to how the

mechanical orbit energy is to be replaced. if it were to be replaced

by a conventional propulsion system on the station, the specific

impulse of the propellant used and the costs of transporting it to

orbit drive the concept to keeping the system efficiency as high as

practicable. On the other hand, if the orbital mechanical energy is

replenished by a tethered de-orbit of the Shuttle from Space Station,

then the incentive to operate the system at high efficiency is

significantly reduced because of the amount of mechanical orbital

energy available from the Shuttle deorbit operations.

22
If the system were to be operated in the mode where the orbital

mechanical energy is replaced by a propulsion system, then the system

efficiency relates to the amcunt of propellant required to replenish

the orbital energy. For the reference Space Station this

relationship is shown in Figure 2-6. The ordinate shows the mass of

propellant required per kWh of energy delivered to the Space Station

bus. The family of curves are drawn for a range of system

efficiencies at 50%, 60%, 70% and 80%. The vertical bars indicate

typical values of specific impulse, Isp, for 3 types of propellant

that have been considered for use on Space Station;

Another comparison which can b0--Made is with an open cycle fuel cell

where a typical conversion factor is 0.45 kg of oxygen/hydrogen fuel

per kWh of electrical energy generated. The Figure 2-6 plot shows

that an equivalent fuel to energy ratio for the tether system is 0.13

kg/kWh for an 80% efficient tether system using hydrogen/oxygen as

the propulsion system. This is improvement in the ratio of fuel to

energy produced by a factor of 3.5. While this is an impressive

factor, it is not really relevant as open cycle fuel cells are not

under consideration for Space Station.

A more meaningful comparison would be to compare the amount of drag

makeup propellant required to maintain the orbit altitude of a solar

array sized to produce an average bus power of 25 kW. Using the fact

that the baseline power system for the Space Station is sized at 100

kW, and that the integrated drag of the resulting articulated solar

arrays and radiators contribute about 56% of the total Space Station

drag, we can estimate that increasing the baseline power system by

25% to 125 kW would cause a resulting increase in the drag of about

14%. Next, using an estimated average value of 5500 pounds per year

for orbit maintenance propellant, this 14% increase in drag would

require an additional 770 pounds of propellant per year. This

translates to 0.0035 kg of propellant per kWh of solar array

electrical energy.

One additional piece of information is the conversion efficiency of

orbital energy (measured in terms of orbit altitude) into electrical

♦ 

energy. This relationship is given on Figure 2-6. A tether system

operating at 80% efficiency and a power delivered to the bus of 25 kW

will cause the altitude of the Space Station to decrease by 2.6 km

per day. This in turn translates into about 29 days of operation at

this 25 kW power level for each full length (64 km) tether deorbit of

the Shuttle from Space Station. The corresponding time at 75 kW is

8.4 days.

The point to be made from the preceding paragraphs is that the

justification for an electrodynamic tether auxiliary power system

rests on the fact that the angular momentum to be converted into

electrical energy is derived from the tether deorbit of Shuttle. If

the angular momentum were to be furnished by a propulsive reboost of

the Space Station, it would require about 40 times the propellant

required for an equivalent solar array power system.

23
24

2.3.4 System Power Levels

The system power levels (P S ) are given by dividing the bus power by

the system efficiency at the selected operating point. Using a

system efficiency of 80% at the design power level ( 25 kW) and an

efficiency of 70% at the reserve power level ( 75 kW) gives the

corresponding system power levels:

PS (Design Level) = 31.25 kW

PS (Reserve Level) 107.14 kW

These are the relevant power 1' evels that will be required for the

subsequent steps of system design.

2.3.5 Conductor Mass

The expression for conductor mass as a function of system power is:

M/ PS = P a / KT KB,

where P resistivity (Ohm meters)

a = density (kg/m3)

KB = volts/unit length (m) of tether

Evaluating the product of P6 over a wide range of temperature for

aluminum and for copper it is found that the value of the product for

aluminum is slightly over one half that for copper over the

temperature range. This indicates that for a given power level an

aluminum tether will weigh only a little over half one made of

copper. Based on this information aluminum was selected as the

perferred conductor material for this design concept.

The term KB is a measure of the electrodynamically induced voltage

generated per unit length of the tether. It depends primarily on the

orbit altitude and inclination. The value for K B varies

significantly over an orbit depending on the position in orbit and on

the position of the plane of the orbit with respect to the earth's

tilted magnetic dipole field. The value of KB ranges from a

minimum of 120 volts/km to a maximum of 219 volts/km. A value of 136 

volts/km was selected for the plot in Figure 2-5. While not at the

extreme low end of the range, this value will insure system operation

at 80% efficiency or better most of the time. For the intervals when

KS drops below 136 volts/Km. The KT will increase to 0.08 and

the efficiency will drop to about 76%.
25

4

0.8

0.6

0.4

0.3

d

0.2

0.1

100 200 300 400 500 600

Ifp

Figure 2-F Reboost Propellant Required as a Function of Isp

ALUMINUM CONDUCTOR

4

3

- - 2720

E 3 2715

200 0 2

p E

2710 Y

x

-

1 2705

100 -

—^

so Xk 50 i0 30 20 10 0° +70 +20

60

Y

ao

_

—

C

--

_

0° C a

:5

50° C

M P6

20 ^` KTKB:

I

K8-0.136

V/m

8 - - - --

0.05 0.1

KT

0.2 0.4 0.5

Figure 2-7 Mass to Power Relationship for an Alwninum Tether

MP/ w 48.32

` IfP rt—

Note:

A 25 kW (e ^ 0.8 ► E.D. power tether

will puss an orbit dewy rate of

2.8 km/d.y.

I

0.6

0.6

.7

.8

p f V

^

= p O

N

Z

N

=

J
the ratio of conductor mass (for aluminum) per kWh as a

function of KT and for two temperature values is shown in Figure

2-7. Also shown in this same figure is a plot of the values of

resistivity and density as a function of temperature for the selected

aluminum conductor material.

Using this relationship we can estimate the mass of tether required

to operate at a K T of 0.05 and a design power level P s of 31.25

W. Using an estimated operating temperature of 10°C gives a value

of 82 kg/kW or a reference tether conductor mass of 82 x 31.25 = 2562

kg. This is called a reference mass because it is the mass that

would be required if the tether were a solid conductor. To obtain

the estimated total mass of the tether we must add an allowance

factor for the helical wind of an actual stranded cable and another

for the insulation required. Using a factor of 1.07 for the increase

due to helical winding effects and a factor of 1.03 for the

insulation gives an estimated actual tether mass of 2824 kg. 

This value of 2824 kg will be used to calculate tether length in the

following sections. It should be kept in mind that this is an

approximation based on estimated allowance factors. Any final design

for the tether will require recalculation using more precisely

determined values for these allowance factors.

2.3.6 Tether Length Determination

Once the mass of the tether and the outboard tethered satellite are

known the minimum length of tether required to maintain the tether

angles within the specified range (less than 0.1 radian) can be

determined. The need for this angle constraint is illustrated by

Figure 2-8. Here two cases are shown. Both cases are operated at

the same system power level which produces an electrodynamic drag

force, FD . This value for FD is the same for both cases. In

case 1 a constraint has Leen applied to keep the tether angle 8I

less than O.lr. This means that the tension force T i in the tether

must be greater than lOFD. This in turn will reduce the off -set

dimension of a tension alignment boom required to keep the tension

force aligned with the Space Station center-of-mass. For Case 1 this

alignment offset is shown as 30 feet.

For Case 2 the tether angle constraint has been relaxed to 0.4r. Now

the tension force, T 2 , need only be 2.5 FD. However, in order to

keep the tension force aligned to the Space Station center-of-mass an

off-set of 126 feet is required. The final selection of an angle

criteria for the tether needs to be determined by more detailed trade

studies. For the purpose of developing this design concept an angle

criteria of O.lr was chosen as a reasonable criteria to be used.

26
To

To

Case 1 1 Case 2

(B, < 0.1R) 1 IB ; 0.4R)

T,

i

B,

30 ft 4--

126 ft ----e•+

T T, - 10 FD

Z ^1

Space

Station

Keel

CM

FD

7

Figure 2-8 Tether Tension Constraint

End

1 ^,

/ T, iR

/ T,-2.6F0

CM

:I FD

•

a

• Total Mass M - Me + M t + Ms

• In Equilibrium Condition

the Drag Force Fd Must Act on Each Element

in Proportion to its Fraction of Total Mass.

Fd - Fe+Ft+Fs P Using the Relation Fd - V

Fe - Me P ; Ft . MtP ; Fs- M' P M V M V M V

• Note That the Drag Forces on the

End Mass and Station are Applied

by the Horizontal Component of Drag Force the Tether Tension "T".

Fd Fe = T Sins He and Fs - T Since Hs

(4 9 - Sine •' F1

Hs - Sine •' F^

O Substituting for Fe and Fs, and Using T

Small Angle Limits

ye -Me P —MA FA

M VT or M T

r Mi P—ML a

M VT °f M T

• Solving to T

V M= T- M PP

T^ V9S

Station

Figure 2-9 Tether Angle Relationships

27
The relationship of the tether tension, T; the mass of the station,

M S ; the total mass of the system, M; the system power level, P; the

angle of the tether at the station, 9 S ; and the orbital velocity,

V; is shown in Figure 2-9.

Using the values:

Ms = 250,000 kg,

M = (250,000 + 2824 + 500) kg = 253,124 kg,

P - 107 kW,

8g - 0.1r, and

V =s 7613 m/sec.

The resulting value for the tension is:

T - 139 Newtons

Using this value for the tension the tether length can be

calculated. The relationship is given by:

T (3j.1/ Rp) (Ml M2/Ml + M2) L

where:

T - tether tension (139 N),

μ - 3.992 x 10 14 N M2/kg.,

Ro n Orbit radius (6.871 x 10 6 m),_

M2

mass of Space Station (250,000 kg),

Ml

sum of end mass + tether mass/2,

= 500 + 2824/2 - 1912 kg, and

L Minimum tether length required (m).

Solving for L 19.8 km as a minimum required length. For the design

concept a length of 20 km has been selected.

2.3.7 Tether Voltage Considerations

The electromotively induced voltage in the tether, E T , is given by

the vector scalar product:

ET-$xV•1,

where

38 geomagnetic field vector (Tesla)

V - orbit velocity - earth rotation component

(m/sec )

L - tether length (m)

28

"^ ' irk1"-- - 
A useful value is the voltage per unit length of tether which i

given by f'T/L. This is the same as the term K B used in Figur

2-7. The variation of E T/L is a compli;:ated function of altitude

inclination, position in orbit, and the position of the orbit plane

with respect to the geomagnetic field. For purpose of cae concept

development, the values have been calculated at extremes to define

the maximum to minimum range of induced voltages. Tnese calculation

were based on a Lilted offset dipole model of the earth's field.

E T /L 113 V/km (minimum)

= 207 V/km (maximum)

49 

These values apply when the tether is in a vertical orientation an

straight. In the actual case as power is generated by the tether, i

will assume a bowed configuration (See Figure 2-9) which will reduc

the projected vertical length of the tether and, thus, the voltage

Effectively, it will appear as though the field strength is decreased

with increasing power levels.

In order to make allowance for this tether bowing effect, the voltag

range will be derated by a factor of 0.95. This is an estimated

derating value and is used here primarily as a place holder in the

system design logic.

Using these derated values, the minimum, median and maximum voltage

for the 20 km tether are:

ET (min) 2147 volts,

ET (median) 3040 volts.

ET (max) - 3933 volts,

2.3.8 Current Considerations

In order to operate the system at a constant power level, the inpu

impedance of the power processing circuit will need to be varied t

maintain the system current in a reciprocal relationship to th

changing voltage of the tether.

For the specified levels of design power (25 kW) and reserve powe

t (75 kW), the corresponding levels of system power are 31.25 kW an

107.14 kW (see 2.3.4).

The current range for design level is then:

I (max) 14.6 Amperes,

I (median) =

10.3 Amperes,

I (min) 8.0 Amperes.

The current range for reserve level is:

I (max) - 49.9 Amperes,

I (median)- 35.2 Amperes,

I (min) 27.2 Amperes.

29
It should be noted that if the system is operated at a constant power

level, the drag force on the tether caused by power generation is

constant and the dynamics of the tether should stabilize into an

equilibrium configuration for each selected level of power operation.

Unfortunately this situation is complicated by the presence of forces

on the tether which are out of the orbit plane. This will be

discusses in more detail in 2.3.12.

An approach to the design of the power processing circuits was

developed for the electrodynamic tether concept studied in Phase I.

This concept seems viable for use on this concept az well and will

not be ropeated here. The concept uses solid state circuitry based

on a series resonant inverter topology currently unde • - development.

The system is built up of modular units to allow sizing to any

particular requirement and for overall system reliability. The

requirement to operate at the reserve hawer level for the system

means that the, power capacity of the processing -.ircuitry must be

proportionately increased over the design level. This would be

accomctodated by an increased number of .nodules available to be bought

on line when the system power level is to be increased above the

design level of 25 kW.

The operating efficiency of these power processing units has been

estimated at 96% and wp a the basis for u3e of this efficiency value

in the calculation of the overall system efficiency (See 2.3.3).

2.3.9 Electrical Potential Considerations

In order to design the necessary electrical insulation for the

system, the electrical potential levels at various locations must be

understood. These potential levels are mapped over the tether system

in Figure 2-10.

The tether is indicated as a series of distributed voltages and

resiftances which sum to the values VT and RT. The next circuit

element is the Space Station power processing system with a voltage

drop VS and impedance RS . Next is the Space Station plasma

contactor with voltage drop V C2 and impedance RC2 . Next is the

ionospheric circuit path with voltage drop V I and impedance RI,

Finally, to complete the circuit path, is the tether end plasma

contactor with voltage drop V C1 and impedance RC1. The node

points in the circuit are at the ends of the tether and the interface

between the station power processor and the plasma contactor. The

ionosphere plasma potential is identified as V. and is shown as the

reference potential on the potential plot in the lower portion of

Figure 2-10. Referring back to 2.3.2 and Figure 2-4, the load

resistance RL used reappears in this circuit where

R = RS + RC2 + R I + Rcl

30
lonos[hPrn IRI, VII

PC - Plasma Contactor

1 Vo

Space

VC

RC TJPIi haf—^—_

` -

-- Statlnn 7

_^ 1 1 1 1 ^^ I

MT. VT) (RS. VS)

I RC1' VC1 I VT • VS • VC2 V1 ♦ VC1 • C I RC2 • VC2

I I

RL - RS • R

C2 . RI s RCI I I I

I I 2 ^ ; ; I

I I I

I I — I

I I I

I I

I I

VT

I

I I 3 I I I

I I ^ I I

I I

VI I I I VI

2 1 I I 4 I I

I. VC1 Vo I

i II I _ 1 - 

^ I _L

V

Case No KT RL KT • Tethef Power -- --

C2

1 0 RS » — Total System Power

2 0.1 9RT

3 0.5 RT

0.95 RI • R C1 ' R C2 (lCl MS • 0)

Figure ?.-10 Potential leap or , zn Elv, • trnd-m(i-r?, • 70thrr r!jj!trm

Al Foil Tape, Spiral Wrap

(Polished or Silver Plate)

0.$3 OHM/1000 ft (at 27' C) (Coated Both Sides with

Heat Sealable FEP Teflon)

2 Spiral Wrap Layer ,, with

Opposite Lead

2 Mils/layer

Breakdown Voltage 6 kV (3 kV/Layer )

Figure "-11 Tether Constr-ction 11sing a dunr_h Stminding

31

.!^
The voltages around the circuit must sum to sero.

VT ♦ Vg+dC2+VT* VCla0

The potential plot is shown for 4 cases of .system operation.

Case 1 is with the power converter open circuited. Here the Space

Station will rise to almost the full tether potential with respect to

the ionosphere. This is based on the assumption that the plasma

contactor at the top end of the tether will keep the tether potential

close to Vo at the node point.

Case 2 is the beat approximation to the design case. The potential

at the station end drops by the anount IR T and the voltage drop

across the power conditioning circuit is near maximum. The point

here is that for low values of KT (i.e. 0.05) the potential

difference across the tether insulation at the station node will

essentially be at the full tether potential.

Case 3 snows the situation where KT - 0.5, or the load resistance,

RL , is equal to RT. This is the impedance match condition where

one-half of the system power is dissipated in the tether. Here the

maximum potential drop across the insulation is a little less ti^an

half the full tether potential because of the IR drop in the tether

itself.

Case 4 is shown for completeness and illustrates the condition that

would exist if the power processing circuit were completely shorted.

The conclusion here is that the system must be insulated to withstand

the maximum voltage generated by the tether. This includes the Space

Station installation and the station end of the tether. The

potential drop across the tether insulation will decrease linearly

with distance from the station. This would indicate that the tether

insulation thickness could be varied along the length of the tether

with the greatest thickness on the inboard end.

2.3.10 Tether Construction

A proposed approach to construction of the tether is shown in Figure

2-11. A bunch stranding c3neept is used to permit the required

resolution in adjusting the conductor cross section to the desired

value and to provide flexibility in the cable.

An inner wrap of polished foil is shown as a reflective surface to

provide low absorptivity to radiation. A metallic foil was used to

avoid any differential voltages between this layer and the tether

conductor. Further thermal analyses may indicate that this layer is

not required.

32
The outer insulating portion is made up of multiple wraps of a Kap

E' film tape which is coated on both sides with a heat sealable 

Teflon. This construction method will permit the long continu

t application of insulation by a wrapping _rrocess during manufactu

and a graded thickness capability by the number of wraps appli

The insulation will be heat fused after application. The Teflon a

provides an outer surface with good resistance to erosion fr

residual atmosphere effects (e.g. atomic oxygen).

Using a rated breakdown voltage of 3kV per layer, it has been assum

that 2 layers of the tape wrap could be adequate over most of 

tether length and with the inboard region going to three or f

• layers.

This construction method would also be compatible with on orb

repair in event the insulation were damaged by handling or 

s

micrometeorite impact.

2.3.11 Tether Temperature

Using values of absorptivity of 0.15 and emissivity of 0.85 for t

tether surfaces gives equilibrium temperatures for the eclipse 

sunside temperatures for the design power (25 kW) -f:

Eclipse = - 80°C,

Sun Side = - 40°C.

And for the reserve level power (75 kW):

Eclipse = -12°C

Sun Side = + 8°C

The actual operating temperatures should range between these extrem

during an orbit. No analysis has been performed on the effects 

this temperature cycling on the proposed construction method.

Further study may indicate that higher operating temperatures wou

be preferable for materials considerations. This should 

achievable by designing for increased absorptivity and decrea

emissivity values for the tether surface.

Recall that for the original estimate of tether conductor mas

temperature of +10°C was used. With the temperature range for 

design level power identified above (-40 to -80°C), the tether 

over designed and the KT actually achieved should be less t

0.05. However, for the reserve power case (+8 to -12 °0, the des

should be near optimum.

33
2.3.12 Cross Track Libration Effects

Due to the 28.5 degree inclination of the Space Station orbit, the

tether cuts across the magnetic field lines at other than the optimum

90 degree angle when crossing the equator. This effect is further

accentuated by the 11 degree tilt of the earth's dipole field with

respect to the earth's axis. The angle between the orbital velocity

vector and the horizontal component of the magnetic field vector

ranges over 90 + 28.5 + 11

degrees. The 28.5 degree component varies thru a cycle per orbit and

the 11 degree component thru a cycle per earth rotation.

A time plot of the effects of this angle variation on the magnitude

of the cross track forces on the tether is shown in Figure 2-12.

In order to generate a system power of 31.25 kW (25 kW design level),

the in-plane component of the electromagnetic drag force on the 

tether must be 4.1 Newtons. The out-of-plane or cross track

component will vary as the plot shown on Figure 2-12. The effect of

this forcing function will be to drive the cross track libration of

the tether.

The natural frequency for tether libration in the cross track mode is

2 times the orbit frequency while the forcing function varies at the

orbit frequency. The resulting dynamics of the tether have not been

analyzed. The initial assumption has been made that the angular

displacement of the tether from vertical would be in phase with the

forcing function which would indicate that the tether would be

vertical at the high latitude portion of the orbit. If this

assumption is confirmed by further analysis, this is a fortuitous

development since the cross track libration angles in combination

with the magnetic field dip angle have an exaggerated effect on the

tether voltage developed at high geomagnetic latitudes. For the most

extreme case when the earth dipole is tilted 11 degrees toward the

orbit, the geomagnetic dip angle is almost 60° down from the

horizontal. Under these extreme dip angle conditions, r cross track

libration angle of 6° would cause an induced voltage variation,

factor of 1.15 to 0.81.

This area of cross track libration is recommended as an area for

further study and simulation. The dynamic behavior'of the tether is

a result of interaction of the geomagnetic field, the tether dynamics

and the power generated. High fidelity computer simulations are a

necessary tool to scope the integrated effects of these forces on the

tether. A capability for extended run times, which integrate the

,+

effects over multiple earth rotations, will probably be required.

34
FO - 4.1 N - Out-OUP4" Forcing Function at

Orbital Frequency Ii21

- Out-of-PW* Libration Frequwv - 2 Q

4

3

2

i

i

35

"A,A 4:.

•

^ t

I Hill

^

LL -1

-2

-3

_ 3 j

4 2-* hour Interval

180° Long ( 180° Long Ascending Node W Long

Figure 2-12 Variation With Time for the Cross Track Component of Tether Force

Blank Space

-

r
2.3.13 Tether System Characteristics

The tether characteristics resulti.n from the concept design

development are summarized below:

Conductor material: Bunch stranded aluminum wire cable

Insulation: Multiple layers of heat sealable Type F Kapton

tape

Length: 20 km

Conductor mass: 2741 kg.

Total tether mass: 2824 kg.

Mass/Length: 141.2 kg/km

Conductor diameter: 0.78 cm

Tether diameter: 0.83 cm

Tether voltage range: 2147 to 3933 V

Current range at design power: 3.0 to 14.6 A.

Current range at reserve power: 27.2 to 49.9 A.

Temperature range at design power: -40 to -80°C

Temperature range at reserve poorer: +8 to -12°C

Tension at Space Station end! 139N-2l lbs)

g level induced on Statiou: 6 x 10g

Tether end mass: 500 kg,

Space Station mass: 250,'1 30 kg.

Tether reel size: Leiagt' , 7 m

Core diameter 0.5 m

Outside diameter 1.1 m

Tension alignment system: Similar to Concept F (Figure 4-10)

2.4 Tethered Platform from Space Station - Concept C

No significant changes were made to this concept during the Phase II

study.

The baseline definition for the concept is the same as was developed

in the Phase I final briefing report.

The platform mass is 31,100 lb (14,136 kg) deployed out to a tether

length of 5.4 nmi (10 km). At full deployment the tether tension is

110 lbf (49UN). The resultingacceleration levels are 2.7 x 10-3g

on the platform and 2.0 x 10 g on the Space Station. Platform

power is supplied by an autonomous power system on the platform.

2.5 Tether Mediated Rendezvous of an Orbital Maneuvering Vehicle with an

Orbital Transfer Vehicle - Concept D

No significant changes were made to this concept during the Phase II

study and the baseline for the concept remains as described in the

Phase I final briefing report.

36
The OMV is tether deployed downward from Space Station to a tether

length of 7 nmi (13 km). The returning OTV is in an elliptical

transfer orbit of 220 x 263 nmi. At apogee of the orbit, the

relative velocities between the tethered OMV and the OTV will go to

zero. The rendezvous and capture must be effected during the short

interval of low relative velocity.

2.6 AXAF Servicing and Tether Reboost - Conce pt B2

This concept was developed to determine if the servicing and re-boost

mission provided any increased performance benefits over the original

Concept B for the initial insertion of AXAF into its operational

orbit.

The resulting revisions to the concept are in the mission operations

area which are described in 3.3.3 of the report.

Some significant findings with respect to the modified mission

operations for the revised concept are the following:

1. There is a requirement to perform a Shuttle OMS burn subsequent

to the rendezvous with the AXAF. Either the AXAF must be

mounted into the cargo bay or the burn must be performed with

the AXAF mounted onto a servicing fixture and projecting out of

the cargo bay. The acceptability of either of these is

questionable.

2. The 205 nmi orbit for the rendezvous reduces the length of

tether which can be used and, hence, the benefit to be gained

from the tether deployment operation.

Another new item of information relating to both the B and B2

versions of this concept is the sensitivity of the AXAF solar array

systems to acceleration levels. This sensitivity will probably

prevent any extensive operational checkout of the AXAF prior to

release from the tether. During the Phase I study this had been

considered a potential benefit of the concept.

37
3.0

3.1

3.1.1

COMPARISON WITH ALTERNATE CONCEPTS

At the end of Phase I of the current contract 5 concepts were

selected for further study and comparison with non-tethered

(baseline) approaches. Because of the large amount of angular

momentum imparted to the Space Station by the tethered Shuttle

deployment ( Concept A) a momentum balance technique was adopted.

The Space Station momentum balance technique considers the various

activities affecting Station altitude ( tethered Shuttle deployment,

tethered OMV launches, tethered OTV launches, Space Station drag

decay of the orbit, etc.) and keeps the Space Station within altitude

limits desired so that Shuttle payload delivery capability or other

mission constraints are not compromised. Concept F (Tethered OTV

Launch) was added to the previously selected concepts since it is the

key element for utilization of excess Space Station angular momentum.

For the analysis approach used in the study the initial effort

consisted of a study of appropriate mission models to have a basis

for OMV and OTV launch requirements for both the momentum analysis

and concept comparisons. Techniques for the estimation of Shuttle

performance, OMV performance and OTV performance were identified and

assumptions were made to determine both the baseline performance and

the tethered concepts performance capability for comparison purposes.

Comparison Criteria

The following sub-section details the approaches and displays the

results of studying the Shuttle, OMV, and OTV performance

parametrically at various tether lengths and payload weights.

Rationale for the selection of the mission candidates for tether

launch from the Space Station and also the effects of Space Station

orbit-de=ay are discussed. Tether deployment considerations, tether

transportation principles, and basic equations used are also included.

Throughout the comparison it was found that the common currency of

tether transportation applications benefits is propellant savings.

This has been found to be true for the Shuttle, the OMV, the OTV, and

the Space Station itself as will be shown throughout the following

discussions.

Mission Model Analysis

Various mission models were investigated including previous models

used for OMV, OTV, and Space Station architecture studies at Martin

Marietta Denver over the past 1-2 years. The recent mission models

associated with the Space Station for the 1991-2000 time period were

found to be the most useful in determining candidates for tether

Launches from the Space Station. Consultations were also made with

current Denver OMV, OTV, and Space Station study teams as needed to

further identify mission characteristics.

38

y
OMV mission candidates for tethered launch from Space Station are

summarized in Table 3-1 and were selected from the NASA "Nominal

Mission Model (FY 1983-2000) Rev. 7 (SS), Space Station Advocacy",

MSFC, July 1984. Potential candidates shown include placement,

retrieval, and servicing missions in the altitude range from 270 nmi

to 378 nmi altitude. Estimated payload weights vary from 2750 lb to

almost 50,000 lb. The selected missions are basically science

applications.

Since the Space Station is nominally at 270 nmi the 'retrieval

missions shown at 270 nmi were not recommended since negligible

angular momentum transfer effects would be realized. Of the 24

tether candidate missions listed for the 1991-2000 time period, only

18 were recommended as OMV tethered launch candidates for having a

potential for propellant savings and Space Station momentum balance.

Since a total of 300 Space Station OMV missions were identified in

the mission model, the recommended tether candidate missions are only

about 62 of the total. The prime reason for this is that most of the

OMV missions are conducted in close proximity to the Space Station

and are not good candidates for tether launches.

OTV mission candidates for tethered launch from the Space Station are

summarized in Table 3-2 and were selected from the NASA "Space

Station Mission Requirements Report", KSC, MRWG001, May 1984.

Selected candidates shown start in 1995 (first Space Station OTV

Launches) and are all delivery or servicing missions to

geosynchronous orbit. Estimated payload weights vary from 1500 lb to

20,000 lb, including multiple payload launches, as indicated. The

selected missions include science and applications, and communication

satellite missions.

There are 72 OTV missions listed for the 1995-2000 time period. Of

these 69 were recommended as tether launch candidates. The

recommended tether launch assist missions are 95% of the total OTV

Space Station missions planned.

3.1.2 Tether Deployment Considerations

Throughout the Phase I and Phase II contract studies certain

assumptions and ground rules have evolved relative to tether

applications for both the Space Station and Shuttle applications. In

general, these are engineering judgement factors concerning design

practice, operational procedures, safety considerations, and economic

factors. The tether assumptions and ground rules currently in use

are as follows:

(1) All releases of deployed masses assume a static (non-swinging)

condition for release.

(2) Tether deployment or retrieval should each be accomplished

within an 8 hr. shift.

39
CANDIDATE

OMV MISSIONS*

SOLAR CORONAL DIAG. MISS.

SPACE TELESCOPE SERVICING

COSMIC DEPLOY. OPT. SYSTEM

AXAF

FAR UV SPECTROS.EXPL. MISS

LARGE DEPLOY. REFL. MISS.

IR INTERFER. DEPLOYMENT

SOLAR SEISMOLOGY

YEAR (1) CIRCULAR

ESTIMATED

ORBIT

WEIGHT 92 93 94 95 96 97 98 1 99 00 ALTITUDE

^NM1-28.50 ) (LB)

1P 1R 1 360 2750

R/1 R/1 R/1 320(270R) 25500

iP 320 28800

R/1P R/If R4P 320(270R) 22640

1P W I F 1R 360 3000

1P is 378 25000(P)

3000(S)

2P 378 49610

15 360 3000(est.

*MISSIONS SELECTED FROM "NOMINAL MISSION MODEL (FY 1983-2000) REV. 7 (SS), SPACE STATION ADVOCACY"

JULY 1984, AS POTENTIAL CANDIDATES FOR TETHER LAUNCHES OF OMV/PAYLOAD FROM THE SPACE STATION.

NOTES: (1) P - PLACEMENTS; R - RETRIEVAL; S - SERVICING

(2) LAUNCH CANDIDATES ARE UNDERLINED (REPRESENTS APPROXIMATELY 6% OF TOTAL OMV SPACE STATION

MISSIONS, 1991-2000).

Table 3-1 Tethered OMV Mission Candidates at Space Station

CANDIDATE YEAR

ORBIT

ESTIMATED

OTV MISSIONS* 95 96 97 96 99 00

WEIGHT

(LB)

EXPERIMENTAL GEO. PLATFORM 1 2 1 1 1 2 GEOSYNCHRONOU 12000

PAM-A CLASS SATS (3/1) 2 4 3 4 3 4 2500 ea.

PAM-D CLASS SATS (3/1) 2 1 5 4 3 2 1800 ea.-

IUS CLASS SATS (211) 3 2 5 5 4 " 6000 ea.

CENTAUR CLASS SATS 1 13000

CENTAUR CLASS/SAT. SERVICING 1

7000 UP

4000 DOWN

GEOSYNCH. PLATFORM 1 20000

IUS CLASS/SAT. SERVICING 1 " 3000

PAM-A CLASS/SAT. SERVICING 1 1500

TT 5 1 10 11 16 1 14 113 169 TnTAL

TOTAL RECOMMENDED MISSIONS 5 1 10 11 16 1 14 1 13 169 TOTAL 1

*MISSIONS SELECTED FROM "SPACE STATION MISSION REQUIREMENTS REPORT", KSC, MRWG001, MAY 1984.

NOTE: (1) REPRESENTS OVER 95% OF TOTAL OTV SPACE STATION MISSIONS, 1995-2000.

Table 3-2 Tethered OTV Launch Candidates at Space Station

40
(3) The tether is assumed to be untapered with a maximum deployed

length of 150 km.

(4) A minimum structural design factor of safety of 2.0 is required,

and the tether should be designed for acceptable recoil

characteristics.

(5) The tether should be designed for multiple reuse for all planned

tether applications at the Space Station. The design goal is

100 or more reuses.

+ (6) The baseline tether material selected is Kevlar (low density,

high tensile strength) jacketed with Teflon ( for abrasion and

erosion resistance).

(7) In all applications, an emergency tether release (e.g.

guillotine) at both ends of the tether is required for safety.

Figure 3-1 illustrates the basic tether transportation principles

used in the transfer of angular momentum between tethered masses, as

applied in this study. In this chart, the following chart, and in

other references to tethered masses, M1 will always be referred to

as the lower tethered mass, with M 2 as the upper tethered mass

(regardless of which one is the deployer).

The left side of Figure 3-1 shows the general relationship between

the masses for a static condition prior to tether release, for a

given tether length, L. The distance that M l is below the original

orbit altitude ( h) is labeled L 1 and is proportional to the

ratio of M2 over gi l + M2 , as indicated. From this relationship

the pre-release altitudes of two masses can be obtained. 

The right side of Figure 3-1 indicates the instant of tether release

(above) with the resulting post-release orbits of the two masses

after release. Prior to release the center of sass is moving along

the circular orbit pa-,n indicated by the dashed line. This path

corresponds to the original orbit before deployment was initiated (to

first order). Before the release the tether continues to point to

the center of the earth due to the gravity gradient effect and Ml

moves slower than circular orbital velocity at its altitude (loses

momentum) while M2 moves faster than circular orbital velocity for

its higher altitude ( gains momentum).

After the release Ml descends to a lower perigee with its apogee

corresponding to the release altitude while M2 ascends to a higher

apogee altitude with its perigee corresponding to the original

release altitude of M 2 . Note that for a given tether length (L),

the approximate separation distance between the two post -release

orbits can reach approximately 7L for a static release (reduces some

from this, depending upon the ratio of the two masses and the mass of

the tether, as discussed following).

41
.I

Initial Orbit = Circular

M2

n2^

L

L2

h 0

-Mass :Original

Center Orbit

Path - h0)

LI

hl

rI

LI = L( M 2 L2 = L - LI

hl = h0- Li, h2=h0 +L2

initial Separation

Distance (L)

M2

^ M1

i

i

i

i

I

1

Maximum Separation

Distance (Approximate)

Original Orbit

Orbit I

Orbi t 2

1

I

I

M I Perigee

M2 Apogee

PRE-RELEASE "ATHS I PAST-RELEASE PITHS

Figure 3-1 Tether Transportation Pri.nciples

DEFINITIONS

MI = LOWER MASS (14CLUDING ONE-HALF OF DEPLOYED TETHER MASS)

M 2 = TIPPER MASS (INCLUDING ONE-HALF OF DEPLOYED TETHER MASS)

MI 

= LOWER MASS (EXCLUDING DEPLOYED TETHER MASS)

MI = UPFER MASS (EXCLUDING DEPLOYED TETHER MASS)

(rM =

MASS DENSITY OF TETHER (MASS PER UNIT LENGTH)

/Z - ORBITAL RATE OF TETHERED SYSTEM (RADIANS PER UNIT TIME)

M = EFFECTIVE MASS DEPLOYED = (M I M 2 ) / (MI + M2)

L = DEPLOYED TETHER LENGTH

R = RADIUS OF POINT OF INTEREST FROM EARTH CENTER

A= L/Ro : P \(V/(1+Y))

'V = M 2 /M 1

: Q = >'/(1 +V)

SUBSCRIPTS: 0 CONDI T ION BEFORE TETHER DEPLOYMENT

1 = REFERS TO LOWER MASS CONDITION

2 REFERS TO UPPER MASS CONDITION

A,P REFER TO APOGEE AND PERIGEE, RESPECTIVELY

RELAT I ONSHIPS

o R AI = R O (1-P) o R P2 = RO(1+Q) 0 h (altitude) ' R(NMI)-3444

o Rp I = RO (1-4P)/(1+3P) o RA2 = k0(I+4Q)/(1-30)

o T (Tethe r tension) = 3 AA-2 L

o E (Energy Developed/required) - 3f\2(KIL2/2+K2L3/3)

where K1 = MlI M2/(MI+M2 ) AND K 2 =CrMIM1- M2I12(MI+M2)

Table 3-3 Basic Tether Fquations

42
The basic tether equations (Table 3-3) used for computing th

pre-release and post-release orbit characteristics for circula

orbits were obtained from the paper by Prof. Manuel Martinez-Sanche

and Sarah A. Gavit of MIT entitled "Transportation Applications in

Space" which was delivered at the AiAA Symposium at Costa Mesa

California, June 5-7, 1984, with additional derivational wor

completed at Martin-Denver. Also included is ti:i standard equatio

for calculating tether tension, considering t?the y mass and 

procedure for obtaining energy developed during deployment or energ

required during retrieval (considering tether mass). The equation

listed are adequately accurate for this t y pe of analysis and ar

generally self-explanatory, with the following guidelines included.

The way the equations are written M l is always the lower mass an

M 2 the upper mass. If, for example, a payload is fully deploye

above the Space Station M 1 will be the Space Station mass includin

one-half of the deployed tether mass and M 2 will be the payloa

mass (plus the PIDM) and including one-half of the deployed tethe

mass. If a payload is deployed downward from the Space Station, the

the Space Station will be designated M 2 , etc. In all equation

units of mass must be used to obtain proper units (e.g. tension wil

be in units of force). Also note that in the definition of K

(used in the energy equation) the absolute value IMi - M2^ mus

be used. Energy is normally converted to electrical units (kWh) fo

tether deployer sizing purposes. Equations applying to elliptica

orbits may also be found in the above referenc

(Mart inez-Sanchez/Gavit).

3. 1.3 Performance Characteristics

This subsection discusses the performance characteristics of th

Shu[ r .P. OMV, OTV, and Space Station, including equations an

procedures used and data generated for the baseline approach and th

tethered approach in parametric form. Assumptions and ground rule

used for the parametric analyses are included for completeness. Th

data and procedure presented allc-i determination of propellan

savings for a general tether deployment mission from the Spac

Station or the Shuttle, given payload weights, mission altitudes, and

tether length requirements. Equation derivations are included i

Appendix A.

3.1.3.1 STS Performance

The basic Shuttle assumptions used for this study are:

(1) Because of the payload benefits at higher altitudes, direc

insertion launches are assumed for all cases (i.e., th

insertion apogee is at the desired circular orbit or at th

apogee of the desired elliptical orbit).

(2) OMS propellant savings is the Shuttle parameter that compare

tethered missions with baseline approaches.

43
(3) Adequate OMS propellant is always loaded to allow accomplishment

of an abort return with the cargo.

(4) Typical Orbiter OV-103 weight data is used, as specified.

(5) Shuttle deorbits from elliptical orbits are from apogee direct

(preferred) or can be via a 100 nmi circular orbit (if required).

(6) All SF. ► ttle tether releases are static and impulsive delta

velocity calculations are used throughout.

To dete-:mine Shuttle OMS propellant requirements and savings it is

necessary to determine delta velocity requirements for initial orbit

insertion and also for the deorbit sequence to re-entry. In addition

it is necessary to determine the weight of the Orbiter at initiation

of the first OMS burn. The following figures (Figures 3-2 through

3-4) and Table 3-3 cover the procedure for obtaining an estimate of

Shuttle payload capability and OMS propellant requirements for a

given direct orbit insertion missioa from the ETR (Eastern Test

Range).

Figure 3-2 presents the required OMS insertion velocity increment

(delta V 1 ) as a function of Shuttle direct insertion altitude for

circular orbits from 100 nmi to 350 rmi altitude and also for

elliptical orbits with perigee altitudes from 100 nmi to 350 nmi..

This data was generated using the standard orbital apogee velocity

equation given on Figure 3-2 and defines delta V 1 as the difference

between the final apogee velocity desired and the initial apogee

velocity obtained from the direct insertion conditions generated at

Main Engine Cutoff (MECO). An initial orbit perigee of 27 nmi was

assu,aed for all cases as being typical from the reference

"Performance Estimation Technique for Space Shuttle Direct Orbit

Insertion Missions - ETR Missions" NASA, JSC, October 1982. Details

of this procedure and equation derivation, as well as those that

follow in Subsection 3.1.3, are given in Appendix A.

To assist in the use of the overall procedure, 2 examples are used in

the following illustrations which also help to show how tethered and

untethered concepts are compared. Concept B (Placing the AXAF

Spacecraft into a 320 nmi orbit) is used to compare the two

approaches. Note in Figure 3-2 that a direct insertion to 320 nmi

(Ex. 1 - Baseline) with the Orbiter requires 505 fps (delta Vl),

whereas the tethered approach (Ex. 2) requires only 165 fps (delta

V1).

44
700

600

N

500

4

EY

C

400

v

O

303

N

O

200

100

100 140 180 220 260 300

Shuttle Direct Insertion Altitude, nmi

Figure 3-2 Shuttle OMS Direct Insertion

Delta V Requirements

Ex 1 505 f to 320 x 320

Circula► Orbits

—^

--^

ElliptcalOrbits

^7i

-

4;

- —

Ex 2 165 fps to 118 x290

300 nmi Perigee Orbitr

250

200 6V1 ' VA Final - VAlnitial

Where

150

VA ( Apogee ) . ( t - e) 4

 Velocity ^^

100

340

Conditions:

1. 27-nmi perigee altitude (typical) auumnd for

initial orbit to apogee in all uses.

Ref ' Performanca Estimation Technique for

Specs Shuttle Direct Orbit Insertion MissionETR Mission' , JSC, Oct 82.

2. Apogee insertion.

600

> 500

c

E4

c 400

y^

• Y

300

x

c

O 200

TI Elliptical Orbits vial

100 nmi Circular (nV2EC)

i

Ex ill2A 505 fps from T

288 x 100 77via 100 Circ

Ex 1 518 fpm from 320 Circ

Elliptical

Orbits (ZIV2E)

Ex 28 145 fps from 298 x 100 77L

100

100 140 180 220 260 300 34U

_ _ Shuttle Initial Deorbrt Altitude (NMI)

IFigure 3-3 Shuttle OMS Deorbit Delta V Requirements

AV2EC ' (VP - VC) 100 + AV2C100

Where

11+elu

VP - RIP

Circular Orbits (AV2C)

290 nmi Perigees

230 AV2E n 6V2C - (VCirc - VApogee)

Ref JSC, approximate

reentry AV requirement

170 from circular orbits vu

MacCroft. Sep, 84 (MSFC).

130

100 nmi Perigee

45
Figure i-J summarizes the required OMS deorbit velocity increment

(delta V 2 ) as a function of Shuttle initial deorbit altitude for

circular orbits from IOU nmi to 350 nmi and for elliptical orbits

with perigees from 100 nmi to 350 nmi. In both of those cases delta

V Z is made up of a singie deorbit OMS burn at apugee for ac

circular orbit altitude). For comparison purposes the delta V2

requirement for descent from elliptical orbits with a 100 nmi perigee

via a lUO nmi circular orbit (two UMS burn case) is show ► (more

conservative reentry sequence).

The basis for the calculated data is the JSC reference which gives

approximate reentry delta V requirements from circular orbits. This

is the curve made up of a seri z of straight line segments entitled

circular orbits. This data was received from Mr. Mac Crof^ of MSFC

in September 1984 and is considered a good method for first order

estimates. Delta V 2 for elliptical orbits was obtained by

calculating the difference between the apogee velocity of the

elliptical orbit and the circular orbit velocity at apogee and

subtracting the difference from delta V 2 (circular), as indicated.

In the case of the two-burn deorbit (top curve in Figure 3-3), the

initial burn is at the 100 nmi perigee of the elliptical orbit. This

portion of delta V 2 is determined uF.ue the perigee velocity

equation given and subtracting the circul,-r velocity at 100 nmi

altitude. The total delta V 2 is obtained by adding the required

delta V to deorbit from a IOU nmi circular altitude.

Continuing with the AXAF placement example, deorbit from 32U nmi

circular altitude would require 518 fps (Ex. 1 - Baseline) while

deorbit from a 288 umi by 100 nmi altitude (tethered approach) would

require 605 fps (Ex. 2A) if the 100 nmi iu..ermediate circular orbit

were used and only 145 fps (Ex. 2B) if tae ai r `ct descent trom apogee

were used.

Knowing the sum of the i-asertion delta velocity (delta V 1 ) and the

deorbit delta velocity (delta V 2 ) for a tethered approach vs. the

baseline approach allows a good initial estimate of propellant

requircments and resultiiid savings. (Second order variations caused

by differences between other on-orbit delta velocity requirements are

ignored in this analysis.) To obtain OM.' propellant required it is

necessary to obtain Shuttle payload cap-oility and the resulting

initial Shuttle weight prior to the first OMS burn.

Table 3-4 outlines the procedure to obtain initial Orbiter weight,

using the example of the AXAF placement mission (baseline approach).

46
47

REFERENCES -

o PERFORMANCE ESTIMATION TECHNIQUE FOR SPACE SHUTTLE DIRECT ORBIT INSERTION

MISSIONS, JSC, OCT. 82.

o SHUTTLE SYSTEMS WEIGHT AND PERFORMANCE

- MONTHLY STATUS REPORT, JSC, 18 MAY 1982, OV 103

(USED TO OBTAIN ORBITER EMPTY WT. AND E.T. EMPTY 'WT.)

STS GROUND RULES

o ETR, DIRECT INSERTION

0MAX = 710 PSF

FILAMENT - WOUND CASES

109% POWER, SUMMER LAUNCH

MISSION -

o CARGO Wi, APOGEE ALTITUDE, ORBIT INCLINATION

(EXAMPLE USED TO SHOW PROCEDURE:

AXAF - 20000 LB CARGO TO 320 NMI CIRCULAR AT 28.5 DEG)

LIFT CAPABILITY DETERMINATION (VALUES 'WITH ' ARE MISSION DEPENDENT)

1. NOMINAL MECO WEIGHT 351592 (LB)

2. ADJUSTMENT FOR APOGEE AND INCLINATION -3527"

3. MECO INJECTED WEIGHT 348065'

4. MPS AT MECO -13901

5.. ET NON-PROPULSIVE CONSUMABLES -423

6. ET EMPTY -69927

7. OMS LOAD -25064

8. RCS LOAD -7508

9. SSME EMPTY -20816

10. ORBITER EMPTY -142483

11. ORBITER NON-PROPULSIVE CONSUMABLES -5409

12. PERSONNEL -3614

13. STS OPERATOR -4219

14. LIFT CAPABILITY 54701 (LB)'

15. STS OPERATIONS RESERVE -3000

16. PAYLOAD REQUIREMENT -20000'

17, PAYLOAD MARGIN 31701 (LB)"

ORBITER INITIAL WEIGHT (Wo)

Wo = W(3) - W(4) - W(5) - W(6) - W(11) - W(17) = 226704"

(ABOVE WEIGHTS W(4), W(5), ETC. USED IN POSITIVE SENSE)

Table 3-4 Orbiter initial Wei g ht Determination

- --

`A '.
The table also shows the weights used for the AXAF direct insertion

baseline mission to a 320 nmi circular orbit, with an asterisk placed

by the weights that are strictly dependent upon the mission. The

nominal MECO weight of 351,592 lb for the assumed conditions is

adjusted for the specific mission altitude and results in a MECO

injected weight of 348,065 lb. All of the non-cargo weight items are

then subtracted to obtain the lift capability for this mission

(54,701 lb). The lift capability is reduced by standard STS

operations reserve (3000 lb) and the specific payload requirement

(20,000 lb AXAF) to obtain a payload margin of 31,701 lb. (for this

analysis ASE weight was not included since it has a second order

effect on the OMS propellant requirement calculation).

At this point the Orbiter initial weight (W o ) can be determined by

taking the MECO initial weight W(3) and subtracting the residue: MPS

propellants at MECO W(4), the E.T. non-propulsive consummables W(5),

the E.T. empty weight W(6), the Orbiter non-propulsive consummables

W(11), and the payload margin W07). This procedure then results in

a first order approximation of Orbiter initial weight (W o ) for use

in the OMS propellant requirement calculations.

Figure 3-4 is the final chart in the Shuttle performance series and

allows determination of OMS propellant requirements and savings for

various missions and mission approaches. OMS propellant increment

required is shown as a function of the summation of delta V .(delta

V 1 + delta V2 from Figures 3-2 and 3-3 respectively) for a

parametric range of initial Orbiter weights (W o ). Since the abort

situation is required and no weight leaves the Orbiter (other than

OMS propellant), the propellant increment is determined by the given

equation, rising the total delta velocity, OMS engine specific

impulse, and initial Orbiter weight. Note that a 7.5% margin is

added for OMS propellant reserve, in all cases.

Continuing with the AXAF placement example, the highest OMS

propellant requirement (23,500 lb) corresponds to Ex. 1 (baseline

approach.) Using the tethered approach with deorbit via the 100 rani

intermediate circular orbit (Ex. 2A) results in an 18,400 lb

propellant requirement. A eirect deorbit from apogee of the

elliptical orbit (Ex. 2B) requires the least OMS propellant (7500 lb)

and is the tether approach chosen for later comparison with the

baseline approach. (The technique of a direct deorbit from apogee of

an elliptical orbit has been discussed informally with MPAD personnel

at JSC and is considered a feasible and promising technique).

3.1.3.2 OMV and OTV Performance

Throughout the Phase 2 study continuing consultations have been

conducted with the Martin Marietta Denver OMV and OTV study teams to

obtaict reasonable estimates of representative vehicle characteristics

and capabilities to allow comparisons between baseline and tethered

approaches with these vehicles.

48
49

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The basic considerations used for this part of the analysis are

summarized in 'Cable 3-5 for both the UNV and OTV vehicles. General

considerations (applying to both vehicles) include off-loading

propellant as required for a mission and generally comparing tethered

and baseline approaches on the basis of propellant savings. All

launches are from the Space Station (270 nmi) to a higher altitude

(Space Station angular momentum consumption) and a maximum tether

length (static releases) of 150 km is considered. Candidate missions

are treated as delivery missions (or equivalent) with the empty stage

returning to the Space Station. Impulsive delta velocity

calculations are assumed.

A recent version of the Martin

obtain vehicle characteristics.

tether launches of the OMV are

stack to an aJ ee corrocnn^ .r ^, p r_.

altitude. (Eliminates first OMV

Marietta Aerospace OMV was used to

All OMV launches are in-plane. All

designed to deliver the OMV/Payload

iriinoot"I tha final desired orbit

burn.

The estimated characteristics of the Martin Marietta Aerospace Aft

Cargo Carrier (ACC) OTV were used for comparison purposes. All

tether candidate missions are launches to geosynchronous orbit (all

plane changes at geo) and an aerobraked return to the Space Station

is always used. The impulsive delta velocities used are 14,091 fps

(ascent to geo, including losses, etc.) and 6445 fps (descent from

geo, including 400 fps for phasing, etc). L► this application a

tethered launch reduces the first delta velocity requirement, as will

be shown.

Table 3-6 summarizes the estimated vehicle characteristics of both

the OMV and OTV vehicle used in the analysis. The OMV version uses

nitrogen tetroxide/monomethyl hydrazine bipropellant with a specific

impulse of 306 sec. and an initial full-fueled weight of 14,600 lb.

The burnout weight of 6,525 lb includes a fixed allowance of 5% for

RCS plus propellant margins. A usable propellant weight of 8075 lb

(maximum) was assumed. (This data was received in November 1984 and

recently the decision was made to select the tank sizes for a loaded

propellant weight of 6,700 lb, with corresponding burnout weight

changes. Since these changes do not impact the conclusions of the

study, the analysis was not revised).

Tile (ACC) OTV version uses a liquid oxygen/liquid hydrogen cryogenic

propellant combination with a specific impulse of 460 sec. The fully

loaded initial weight is 60,011 lb. with a maximum usable propellant

of 53,577 lb. The burnout weight of 6,434 lb includes the

acrobraking system and an allowance for RCS and propellant margins.

50
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51

--S w;.''
OMV(1)

w Dry WBurnout WpUseble Wlnitial

I

sp

PROPELLANT

(LB)

6100

(LB)

6525(2)

(LB)

8015

(LB)

14600

(SEC)

306 N204/MMH

(1) USES CURRENT ESTIMATED CHARACTERISTIS OF MARTIN MARIETTA AEROSPACE STORABLE

BIPROPELLANT ORBITAL MANEUVERING VEHICLE (OMV).

(2) INCLUDES 5% ALLOWANCE FOR RCS PLUS PROPELLANT MARGINS.

OTV(3)

W Dry WBurnout WP Usable Wlnitial Isp PROPELLANT

(LB)

5862

(LB)

6434(4)

(LB)

53577

(LB)

60011

(SEC)

460 L02/LH2

(3) USES APPROXIMATE CHARACTERISTICS OF MARTIN MARIETTA AEROSPACE AFT CARGO

CARRIER (ACC) OTV AT SPACE STATION.

(4) INCLUDES ALLOWANCE FOR RCS AND PROPELLANT MARGINS.

Table 3-6 OMV and OTV Estimated Vehicle Characteristics

DEFINITIONS

pVI = ASCENT & V REQUIRED WPL PAYLOAD WEIGHT

QV2 = RETURN QV REQUIRED WBO = BURNOUT WEIGHT

ME = MASS RATIO = e 

eVII01V 

Wp U = USABLE PROPELLANT WEIGHT

OMV

BASELINE (UNTETHERED) o pVI AND W2 OBTAINED FROM STANDARD ORBIT EQUATIONS FOR ORBIT DESIRED

EQN. (1) o WpU (REQUIRED) = WP'-(MI-1) + W BO (M 2M 1 -1) FOR SELECTED PAYLOADS AND

FINAL ORBIT ALTITUDES

TETHERED APPROACH o LENGTH OF TE T HER, VARIED TO OBTAIN SPECIFIC APOGEE. FIRST PERIGEE BURN

IS ELIMINATED AND APOGEE BURN IS REDUCED SINCE PERIGEE IS HIGHER

o WP11 (SEE EQN. 1) FOR VARYING TETHER LENGTHS

OTV

BASELINE (UNTETHEHED) o pVI ANDAV2 FIXED FOR r,EO MISSION

0 WpU (SEE EQN. 1) FOR VARYING PAYLOADS

TETHEREF APPROACH o LENGTH OF TETHER VAPIEO, CHANGF IN LAUNCH ALTITUDE AND VELOCITY DETERMINING ORBITAL EQUTATIONS USED TO OBTAIN NEW pV I VALUES.

o WpU (SEE EQN. 1) FOR VARYING PAYLOADS/TETHER LENGTHS

(1) SEE APPENDIX A FOR DETAILED DERIVATIONS

Table 3-7 OMV and OTV Performance Characteristics

52

.3145^:. 
OMV and OTV performance equations for determination of propellant

loading requirements for baseline and tethered launches are

summarized in Table 3-7, with detailed derivations given in Appendix

A. Delta V requirements are determined using tether transportation

relationships (Table 3-3) and standard orbit equations. Tether

launches primarily reduce (OTV) or eliminate (OMV) the perigee burn

requirement with some further reduction in the first apogee burn

requirement due to the increased perigee altitude (both of which

reduce delta V I ). Return velocity requirements (delta V 2 ) are

not affected. Propellant loading requirements (W u ) are determined

using the payload weight, stage burnout weight, and mass ratio

• relationships shown in Equation 1.

Figure 3-5 summarizes parametrically the OMV prdpellant requirements

for covering the selected mission set of candidates. Note that the

chart only extends to about 3,200 lb of usable propellant loaded,

which is only 40% of the maximum propellant load. Offloading of

propellant for missions is generally the rule, since the design

performance missions for sizing the OMV require much higher

performance but do not occur as frequently. Propellant requirements

are shown for both baseline missions (solid curve) and tethered OMV

missions (dashed curve) for altitudes from 270 nmi up to 400 nmi and

payloads from 0 to 60,000 lb.

Figure 3-6 shows the significant propellant savings for the tethered

approach over the baseline approach. Those savings approach 50% for

payloads in the 20,000 lb and above category and are at least 30% for

payloads below 20,000 lb, with appropriate tether length requirements

indicated as a function of payload and final orbit altitude. In

terms of propellant saved, the amount varies from a few hundred

pounds to over 1000 lb, depending upon the mission. Although the

savings in propellant is significant, the total number of missions

involved is small, as shown in the mission model.

^n Figure 3-7, parametric data are shown for OTV usable propellant

requirements for selected payloads (0 to 20,000 lb) as a function of

tether lengths from 0 (untethered) to 150 km (maximum considered).

As for the OMV, offloading of propellant is generally required for

those missions also, but the average propellant loading is higher

(about 606) over the mission set.

Figure 3-8 displays the propellant savings for the tethered approach

over the baseline approach. Attention should be drawn to the 150 km

length tether curve since it was selected for all OTV launches, as

the study progressed. Although the propellant savings appears small

(8Z), the amount of propellant is very significant (over 4,000 lb

saved for the 20,000 lb payload mission), and a large number of

missions are involved (69) over the 6-year period of interest

(1995-2000), as will be discussed later.

53
54

^I

=I

400 0 0 10 20 10 40 20 60 40 60 klb

Payload (klb)

380

r-T

E 360

a

L ,

a

? 340

a

O

c 320

LL Launches from

300 with OMV Burn

280

0 400 800 1200 1600 2000 2400 2800 3200

Usable Propellant Load Required, W pu , lb

Figure 3-5 OMV Propellant Requirements for Mission Set

2000

r ^-

/

lb (Payloac)

1600 4553%

Propellant

2

Savings

Ii Region

3 1200

40

m , c

800 15

oPropellant

30-45%

CL 10 Savings

400

Region

0 280 300 320 340 360 380 400

h - Final Orbit Altitude (NMi)

Figure 3-6 0MV Prot-el,tant Savings from Tether Launches

(Typical)

i

I I

Untethered OMV

Launches from i

- — i Space Station

Tethered OMV —

Space Station

to Apoyee of

Final Orbit

--

I ^

—

Space Station at 270 nmi at Apogee i

(Up to 37 km of

Tether Required)

Max. OMV Propellant

I

---- Load 8075 Ib

Space Station at 270 nmi

i

L=35 km 60k

30

_

^

25

I

20

10

L=5km

- 0'
55

60

50

b (Payload)

Y

CL

CL

30

B

J

C

yA

G 20

0

a`d

a

9

10

0 20 i0 60 80 100 120 140 160

Tether Length, L, km

Figure 3-7 OTV Propellant Requirements for Mission Set

Space Station at 270 nmi

All Launches to Geostationary Orbit

Propel;ant Savings Vary from 2% (30 km) to 8% (150 km)

l

4

I (Tether length) L 150 km

120, 3 I

Cr—

3 --

— 90

0

^ 80

0

0

U 2 4 6 8 10 12 14 16 18 20

Payload Nlivered, klb

Figur-ey 3-8 OTV Propellant Savings from Tether Launches

•

Space Station at 270 nmi

All Launches to Geostationsry Orbit

53577-Max OTV Load

i

i 20 kl

1F

--

10

7.5

5.0

2.5

0

I

0L l I I I I I

First OTV Burn Follows Tether Releess
pp

p

characteristics, tools have been developed to compare tether

applications with baseli ►► e approaches for the Shuttle, OMV, and OTV

and also to obtain the effects of tether mission applications on

Space Station angular momentum. In like manner, the Space Station

performance requirements must also be considered. In this context,

Space Station performance can be thought of in terms of the energy

(or propellant) required to keep the station at its nominal altitude.

The ground rules for all analyses involving the Space Station are

summarized in Table 3-8. The nominal mass assumed is 551,000 lb

(250,000 kg) and the nominal orbit is 270 nmi (500 km) at an orbit

inclination of 28.5 deg. The corresponding nominal Space Station

angular momentum

(MR

Z Q ) is 9.652403 x 10 15

slug-ft2/sec

(1.309055 x 10 16 kg - M2 /sec). The desired allowable orbit

altitude excursion due to tether operations is 250 to 300 nmi, but

these limits may be exceeded for short periods by scheduling

subsequent Space Station momentum balance operations. Nominal orbit

stationkeeping (drag makeup) maneuvers are assumed to be made at

least every 90 days (or quarter year). Tether transportation

operations (tether launches of OTV, etc, and tether deployment of the

Shuttle or other applications) will be conducted in a manner to

minimize orbit stationkeeping propellant usage. Momentum balance

operations will also be planned to have the Space Station near the

lower end of its altitude limits (250 nmi to 270 nmi) for Shuttle

revisits to maximize the payload weight that the Shuttle can

deliver. The latter guideiine could be made more flexible if the

Shuttle payload weight capability is not fully utilized (due to

manifesting and volume constraints) for Shuttle revisits to the Space

Station.

The Space Station orbit stationkeeping propellant requirements,

obtained through consultation with Martin Marietta Denver Space

Station study personnel, are present-id in Table 3-9. The average

quarterly propellant requirements are based on a hydrazine

monopropellant system (I . P = 230 sec) and reflect the projected air

denbity variations at 27v nmi during the solar cycle for the years

1991 thrc,zgh 2000. Quarterly average propellant requirements start

at a maximun. (near the peak of the 11-year sun spot cycle) of 2850 

lb/qti in 1991, each a minimum of 230 lb/qtr in 1997, and are again

approachi.ta the next maximur, 2455 Ib!gtr) during the year 2000.

These values can vary apprecia^ly, depen"ing upon the propellant

chusc- or the dEgree of conservatism in atmospheric model selection.

They should be thought of a. representative value,- only, and are not

critical to the conclusions of this study, 1- r er discussed.

1- he final column on Table 3-9 shows the equivalent sltitude loss of a

typical Space Station configuration, ii no orbit stationkeeping

maneuvers are made during the quarter (avrrag:-' over the ,.,articular 

year). This data is used in the benefits analysis (3 10. 

56
j .

o NOMINAL MASS 551,000 LB (250,000 KG)

o NOMINAL ORBI T 270 IJMI (500 KM)

28.5 DEG, INCLINAT!ON

o NOMINAL ANGULAR MOMENTUM 9,652403 X 10 15 SLUG-FT2/SEC.

(1.309055 X 1016 KG - !12/SEC.)

o ALLOWABLE ORBITAL EXCURSION DUE TO TETHER OPERATIONS - 250 TO 300 NMI, DESIRED'

o NOMINAL ORBIT STATIONKEEPING INTERVALS - 90 DAYS OR LESS

o CONDUCT TETHER LAUNCH, ORBITER DEPLOYMENT, AND OTHER TETHER 'TRANSPORTATION

OPERATIONS TO MINIMIZE ORBIT STATIONKEEPING PROPELLANT

o SPACE STATION ALTITUDE AT SHUTTLE REVISIT - 250 TO 270 NMI DESIRED

'THESE LIMITS MAY BE EXCEEDED ON SPECIFIC TETHER TRANSPORTATION MISSIONS, BUT

PLANNED SPACE STATION MOMENTUM BALANCE OPERATIONS WILL LIMIT THE TIME OF

EXCURSION.

Table 3-8 Space Station Ground Rules

o SPACE STATION AT 270 NMI ALTITUDE

YEAR AVE. PROPELLANT (N 2 /H 4 ) EQUIVALENT ALTITUDE

_ REQUIRED PER QUARTS° (1) LOSS_ PER QUARTER

(LB) (NMI)

1991 2890 12

1992 2750 11

1993 1880 8

1994 1080 5

1995 605 3

1996 395 2

1997 280 1

1998 330 1

1 099 1145 5

2000 (EST) 2455 10

(1) REFERENCE: INTERNAL COMMUNICATION ON SPACE STATION PROPELLANT CONSUMPTION,

G. McALLISTER, 14 DECEMBER 1984.

Table 3-9 Space Station orbit Maintenance Requirements

57

?'-!
The comparison format for evalu..ing the tethered concepts vs. the

baseline approaches (Sect. 3.3) evolved throughout the Phase 2

study. The primary purpose is to make as many direct quantitative

comparisons as possihle to eventually lead to cost comparisons in

follow-on study phases as the selection process narrows. Because of

the early state of study or development of the baseline

configurations (OMV, OTV, and Space Station) and the tethered

concepts, qu;.Iitative comparisons are also made, based on subjective

engineering assessment. In addition to quantitative and qualitative

comparisons, final remarks highlight the benefits or penalties

Incurred by the tethered approach to the particular mission.

Quantitative comparison criteria include the tethered orbit

parameters (e.g., initial, pre-release, post-release, and final

orbits), the propellant usage of the Shuttle, OMV, OTV, Space

Station, and end effector (i.e., PIDM or S1DM), as applicable, and

tether length required. Additional tether system parameters include

the maximum tether tension, the energy usage for deployment and

retrieval, and the estimated tether system weight. Space Station

angular momentum gain or loss is also evaluated, as well as the

operational time comparison with the baseline approach.

Qualitative comparison criteria include rough estimates of operations

and hardware complexity, mission success risk, technology risk,

launch (or propellant) cost factor, and hardware development cost

factor. These qualitative comparisons are useful, but subjective,

and attempt to estimate the percent change from the baseline approach.

Throughout the quantitative and qualitative comparisons, evaluations

and comments are made where applicable. The final remarks section

summarizes the main points of the comparison between the selected

tethered concept with the best estimate of the likely baseline 

approach and points out the primary advantages and disadvantages.

3.3 Comparison Results

3.3.1 Tether Assisted deployment of 220 klb Shuttle from the Space

Station-Maximum Maas DeDlovment Case - Conce p t A2

A comparison of tt.e tethered deployment of a departing shuttle

(Concept A-2) with the baseline approach of undocking and deboosting

frcm the vicinity of the Space Station is presented in Table 3-10.

Approximately 35 nmi (64.3 km) of tether is required to provide a 100

nmi perigee for the Shuttle after tether release. As the Shuttle is

deployed below, the Space Station rises from 27U nmi altitude to 280

nmi altitude, while the Shuttle is lowered to an altitude of 245

111111. During deployment, excess OMS propellant is transferred from

the integral OMS tanks ir •.te the scavenging tanks on the SIDM. After

Shuttle release, the Space Station moves into a 280 ami x 340 nmi

elliptical orbit (average altitude raised by 40 nmi) and the urbiter

moves into a 100 nmi x 245 nmi elliptical orbit. At a subsequent

return to apogee, after the Shuttle cargo bay doors are closed, the

deorbit burn is executed.

58

-'9 -I'- 
lR1TERIA

RASFIINE TETIgREO

COMPARISON EVAIuATI ON / COMMINTS

APPROACH CONCIPT

I. QUANTITATIVE MISSION CRITERIA

A. ORBITS (NMI AT 18.5 DIG. INCL)

o INITIAL ORBIT ?70x210 270x210

PRE-RELEASE OR PRE-RENDEZVOUS ORBITS

0 DEPT YER SP CF STATION N/A 280 n ?30 SPACE STATION MASS - 551 KLB

o DEPLOYED MASS (SHUTTLE) N/A 245x245 ORBITER MASS - 220 KLB

o (TETHER LENGTH - NMI) N/A (35)

POST-RELLASE OR POST-RENDEZVOUS ORBITS

n DEPLOYER N/A 280x340 AVE. SPACF STATION ALT. RAISED 40 NMI

o DEPLOYED MASS N/A 100x245

FINAL ORBITS AVE. SS ALT. KEPT BELOW 300 NMI VIA

0o D^q 270x210 280x340 OTV LAUNCHES AND DRAG DECAY

o DEPLOYED MASS 270x:70 IOOx245 ASSUMES DEORBIT FROM A P OGEE FOR TETH.

ERIP ;ASE

B. PROPELLANT USAGE (LB)

DUPING 60P.BIT ONLY. SAVES 6590 LR.OI

o SHUTTLE OP.BITLR 101CO 3600 -6500 N204/MMH FOR SS USE WITH SCAVENGING

o OMV SYSTEM

o OTV SAVES APPROX. 5500 LB. O' N2H4 ORBIT

o SPACE STATION 5500(Fst 0 •5500 STATION KEEPING PROP. ANNUALLY (AVE.!

o ENO EFFECTOR (PIDM OR SIOM) N/A <SUO - 500 SMALL COLD GAS PENALTY

C. TETHER SYSTEM

o MAXIMUM TENSION 'LB) N/A )H54 CCRRESPONOS TO 2.3 FACTOR OF SAFETY

(17144N)

ENERGY USAGE JKWHJ

o OEPLOYMf_NT — N/A 153 - 153 SS DISPOSLS 153 IiWM IN 4-6 HRS

o RETRIEVAL N/A 11 11 SS SUIPLIES 11 KWH Ua 4-6 HkS.I

o SYSTEM WEIGHT (LB) N/A 115000 -25000 ESTIMATED SS REUSABLE TETHER SYSTEM

DFPLOYER WEIGHT

D. SPACF STATION MOMENTUM 1.309060x1016 KG-M2 /SEC (REFFPFNCE)

o MOMENTUM r.AIN (KG-M2/SEC) 0 4t.99x101 • 6.99xl013 0.531 INCkEASE IN SS ANGULAR MOMENTUM

E. OPERATIONAL TIME2 (MRS.) 2.0(Est) 16.0(Est) -14 (Est) ADDROACH4 HRS. APP

REOUiRED FOR TETNEP

11. QUALITATIVE MISSION CRITERIA

A. COMPLEXITY (0-2)

o OPERATIONS 1.0 1.2 - 1,S • .2 to .5 ?01 TO 501 MORE OPS. COMPLEXITY

o HARDWARE 1.0 1.7 - 1.5 *.? to .5 ?01 TO 501 MORE HOW. COMPLEXITY

S. RISKS (0-2)

3 MISSION SUCCEss 1.0 '..0 0.0 iAMF A; BASELINE APPROACH

o TECHNOLOGY 1.0 "` 25 5, MORE TECHNOLOGY RISK

C COST FACTOR (0-2)

.l0 to .?011 0 TO 20 1. INCREASE IN DEPLOY. COST-' o LAUNCH 1.0 1.10-1.20

o HARDWARE LkILLOPMF.NT 1.0 1.20 .20 FOI MORE HOW. DEVELOPMENT COST

N01ES: 1) POWER RL'!UIRED VOR FULL IINF,TH IF "it R RETRIIVAI `64 KM) OF SHUTTII INTERFACE DEPLOYMLNI MODULE (SI(,M) 1f41

OF SCAVENGED ORS PROPELLANT 19500 LB, TOTAL)

?) TIMES ARE ESTIMATED ;ROM THE POINT WHERE THE SHUTTLE IS READY T^ MOVI AWAY FROM THE SPACE STATION,

3) DOLS NOT CONSIDER OMV AND SPACE STATION ORBIT STATIONKEEPING PROPELLANT SAVINGS

Table 3-10 Compw• i30n of Conct2pt A2 with Baseline Approac-h t

59
tether for storage of the propellant and SIDM in separate areas on

the Station for later usage. During the next 90 day (or shorter)

period, the Space Station will lose some altitude due to drag decay,

but the major reduction in altitude will be caused by one or more

tether assisted OTV launches, and/or possibly from electrodynamic

drag induced by an electrodynamic tether power generation period. By

the next Shuttle revisit the Space Station should again be in the

250-27U nmi altitude regime.

The 6500 lb of scavenged OMS propellant results from the fact that 

the baseline approach requires about 10,100 lb of OMS propellant to

achieve an untethered deorbit from the Space Str.tion, compared to

only 3600 lb for deorbit from apogee after tether release. As

indicated in 'fable 3-9, an average of a p proximately 5500 '_b of Space

Station orbit stationkeeping propellant (hydrazine) could also be

saved (annually) by tethered Shuttle deployment operations over the

10-year time period (1991-2000). A small cold gas penalty (less than

500 lb) would be incurred during tether retrieval operations with the

fully loaded SIDM (9,500 lb including scavenged propellant).

The maximum tether tension reached at full deployment is 3854 lb

(11,144 N). Although this is the maximum tether operation from the

standpoint of deployed Hass, the tension is slightly less than the

deployed OTV/payload case (Concept F), since the ON launch uses 150

km of tether (vs. 64.3 km for Shuttle deployment). Energy usage

d:jring deployment is 153 kWh (to be dispositioned on the Space

Station in 8 hrs or less), while only 11 kWh needs to be supplied by

the Space Station to the tether deployer system during retrieval.

The tether deployer system weight is estimated at 25,000 lb,

including 13,900 lb of tether mass (150 km plus 5% margin of 0.32

inch diameter Teflon jacketed Kevlar). This system is used for all

tether operations on the Space Station and will have many reisses over

the 10-year mission time period.

The tethered Shuttle deorbit operation also causes the largest

angular momentum change on the Space Station (0.53% increase), and

corresponds tc the 40 nmi average increase in altitude. The Space 

Station momentum decrease caused by a maximum payload launch of an

ON (Concept F) is almost equal and opposite to this concept (A-2).

'rot al operational time wit ti Cie tether approach would be

approximately 16 hrs, compared to 2 hrs or less for th,- baseline

approach.

60

n -- r 
and also the hardware complexity are felt to be from 20% to 50% more

complex for the tethered approach since no additional hardware is

involved for the baseline case and operations are less involved. No

additional mission success risk is anticioate3 but the technology

risk could be about 25% greater than the baseline approach. There

could be a 10 to 20% increase in deployment cost, but this does not

consider the szavenged OMS propellant savings (usable for OMV

operations) or t.^e Space Station orbit stationkeeping propellant

saved, which will be covered in 3.4. Hardware development cost is

estimated to increase by 20%.

C-.r erall, the increase in Space Station momentum allows tethered OTV

launches with further savings in OTV propellant (Concept F) and also

eliminates the need for an average snnual Space Station usage of 5500

lb of hydrazine for orbit stationkeeping to counter aerodynamic drag

effects at the nominal altitude of 270 nmi. Scavenging of 6500 lb of

OMS propellant (nitrogen tetroxide/monomethyl hydrazine) with every

full length Shuttle tether deployment reduces transportation

requirements for delivering OMV propellant (same propellant as OMS)

to the Sp ace Station.

3.3.2 Tether Assisted Launch of 20,000 lb OTV Payload to Geosynchronous

Orbit - Concept F

Concept F was added during the Phase II portion of this study and is

the prime user of the Space Station angular momentum increase

provided by the tethered Shuttle deployment concept discussed

previously in 3.3.1. The 20,000 lb payload launch was chosen since

it represents the maximum size payload in the OTV mission model and

also designs the tether system. Table 3-11 presents the comparison

of the tethered launch (above the Space Station) with the baseline

approach, which involves the use of an OMV to take the OTV away from

the Space Station for launch.

A tether length of 150 km (81 nmi) was chosen, based on the ground

rule of maximum length and a tether tension that is about the same as

the tethered Shuttle deployment case. As the OTV/payload stack is

deployed above it, the Space Station descends from 270 nmi altitude

to 261 nmi altitude, while the OTV reaches an altitude of 342 nmi.

After OTV release, the Space Station moves into a 204 nmi x 261 nmi

orbit (average altitude reduction of about 38 nmi) and the OTV moves

momentarily into a 342 nmi x 801 nmi orbit (the perigee burn for geo

transfer could commence immediately after release from the tether or

at a later perigee).

61

^ii^ wr y~ Z 
MISSION F - TETHER ASSISTFD LAUNCH OF 20000 1B OTV P AYLOAD TO

GEOSYNCHRONOUS ORBIT FPOM 111F. SPACE SIATIUN

CRITERIA BASELINE

APPROACH

TEiHEREO

CONCEIT

CnMPAPI`.ON LVAUJAIION/COMMENTS

1. QUANTITATIVE MISSION CRITERIA

A. ORBITS (NMI AT 28.5 DEG. 1NCL)

u INIIIAL URBIT 210 x 270 70 x 270

PRE-RELEASF OR PRE-RENDEZVOUS ORBITS

o DFPLOYFR SPA E STATION) N/A 261 x 261

o DEPLOYED MASS (OTV/PAYLOAD N/A 312 x 342

o (TETHER LENGTH - NMI) N/A (81)

POST-RELEASE OR POST-RENDE7VOUS ORBITS

o DEPLOYFR N/A 61 x 20ft- AVERAGE SS ALTITUDE DROP OF 37.5 NMI

o DEPLOYED MOSS N/A 342 x 801 ORBITAL ELEMENTS BEFORE PERIGEE BURN

TO CEO.

FINAL ORBITS AVE. SS ALTITUDE KEPT AT 250 NMI rig.

o DEPLOYER 270 x 270 261 x 204 APOVE VIA TETHERED ORBITER DEPLOYMENTS

o DEPLOYED MASS GEOSYNCH GFOSYNCH

B. PROPELLANT USAGE (LB)

u SHUTTLE ORBITER

o OMV 133 N/A +133 COLD GAS USAGE FOR OTV PLACEMENT/PET.

o OTV l 51158 46856 +4302 OVER H"_ OF OTV PROPELLANT SAVED

o SPACE STATION

o END EFFECTOR (PIDM OR SIDM) N/A <500 -500 SMALL COLD GAS PENALTY

C. TETHER SYSTEM

o MAXIMUMTENSION LB N/A 1970 :ETHER DESIGN CONDITION - 2.2 SAFETY

fl ? "JON) "CTOR

ENERGY USAGE (KWH)

o DEPLOYMENT NIA 366 +366 SS DISPOSES 366 KWH IN 4-6 HRS.

o RETRIEVAL N/A 29 - 29 SS S'JPPLIES 29 K.WH IN 4-6 HRS.

o SYSTEM WEIGHT (LB) N/A 25000 -25000 ESTIMATED SS REUSABLE TETHER SYSTEM

DEPLOY r RWE CHT

D. SPACE STATION MOMENTUM 1.309060x1) eKG-M2 /SEC (REFERENTE)

o MOMENTUM GAIN (KG-M 2 /SEC) 0 -6.69x101 -6.69x10 13 0.51. REDUCTION IN SS ANGULAR MOMENTUM

E. OPERATIONAL TIME (HRS.) 5.0 EST, 16.0 FST -11.0 EST

11 HR. INCREASE IN OPERATIONAL LIME

FOR TETHERED APPROACH

Ii. QUALITATIVE MISSION CRITERIA

A. COMPLEXITY (O-2)

o OPERAIIONS 1.0 1.1 - 1.2 +.1 TO .2 10. TO 20" MORE OPS. COMPLEXITY

o HARDWARE 1.0 1.1 - 1.2 +.1 TO .2 10; TO 201 MORE HOW. COMPLEXITY

B. RISKS (0-2)

o MISSION SUCCESS 1.0 1.0 - 1.05 0 TO .05 0 TO 5.,

MORE MISSION SUCCESS RISK

u TECHNOLOGY 1.0 '•.0 0 SAME TLCHNOLOGY RISK

C. COST FAC^OR (0-2)

o LAUNCH 1.0 0.9-0.9 -.1 TO-.2 101 TO 201 REDUCTION IN LAUNCH COST 

o HARDWARE DEVELOPMENT 1.0 1 1.05 1 .OS 5? INCREASE IN HOW. DEVELOPMENT COST

NOTES: 1) ASSUME ALL OF 28.5 DEG. PLANE CHANGE AT APOGEE OF TRANSFER ORBIT FOR FIRST APPROXIMATION.

2) BASED ON GETTING THE DIV FROM THE SPACE STAJION TO ITS LAUNCH POINT.

3) DOES NOT CON'd DER OTV PROPELLANT SAVINGS.

'abte 3- 1 1 Comparison of Concept F with Baaetine Approach

62
gradually retrieved for storage on the Space Station for later

reuse. In actual practice, a momentum balance will be achieved

between Shuttle deployments. OTV launches, and possible

electrodynamic tether or other tether operations, so that average

station altitudes will stay in the 250-300 nmi range, with the

constraint of keeping the Space Station near the lower part of tke

altitude range during Shuttle revisits. (A 270 nmi nominal Space

Station altitude has been used throughout the Phase 2 analysis to

allow a common reference for comparison.)

If the tether approach is used, the OMV mission (requiring about 133

lb of cold gas usage for OTV placement and return) would be

eliminated. The baseline OTV mission requires about 51,158 lb of

cryogenic propellant (LOX/hydrogen) to eo to geosynchronous orbit,

place its payload, descend and change planes, accomplish its

aerobraked reentry, and return to the Space Station vicinity. For

the tethered approach the overall operation is similar to the

baseline approach, but 4300 lb of OTV propellant can be saved by the

momentum transfer from the 150 km tether launch. A small cold gas

penalty (less than 500 lb) would be incurred by tether stabilization

operations with the PIDM.

The maximum tether tension reached is 3970 lb (17,660 N), and

corresponds to the design condition (2.2 factor of safety), which is

slightly higher than Concept A-2. Energy to be disposed during the

8-hr or less period of the 150 km tether deployment is 366 kWh, which

is the maximum achieved in any of the tether concepts. The retrieval

energy requirement (tether plus PIDM) is 29 kWh. The tethered OTV

launch causes a 0.51% reduction in Space Station angular momentum and

corresponds to the 37.5 nmi average decrease in altitude (to be made

up by a tethered Shuttle deployment later). The tethered approach

operation time is estimated to be 16 lirs (compared to 5 lirs for the

baseline approach).

The operations complexity of Concept F is expected to be 104 to 20%

higher, with the hardware complexity similarly increased. Mission

success risk shows a slight increase (to 5%), but technology risk

should be about the same as the baseline approach.

In the launch cost factor comparison, the launch operation is defined

as getting the OTV from the Space Station to its launch point, and

the cost factors do not consider OTV propellant savings, which are

covered later in the benefits analysis (3.4). A launch cost

reduction of at least 10% to 20% is expected compared to the baseline

approach due to the removal of the OMV servicing requirement.

Hardware development cost will probably increase about 5%.

Highlights of this Space Station tether application include the

elimination of the OMV service mission, the 4300 lb reduction in OTV

propellant required, and the significant reduction in Space Station

angular momentum which permits additional tethered Shuttle

deployments (Concept A-2) with further savings in scavenged OMS

propellants.

63
Concept S1 was the initial Phase II study concept selected for

comparison and requires the initial placement of the AXAF spacecraft

(Advanced X-Ray Astrophysical Facility) into a 320 nmi circular orbit

at 28.5 deg inclination. In the baseline approach (Table 3-12), the

AXAF is taken directly to a 320 nmi orbit by the Shuttle (using

direct insertion), which then returns to deorbit after deployment and

checkout of the spacecraft. In the tethered apprcach, the Shuttle

again uses the direct insertion technique, but is placed into a lower

elliptical orbit (118 nmi x 290 nmi) with insertion at apogee. A 60

km (32.5 nmi) tether is required to match the 320 nmi circular orbit

velocity conditions for an apogee release.

As the tether is fully deployed above the orbiter, the AXAF ascends

into an elliptical path (148 nmi x 320 nmi) which gives it the

appropriate velocity at apogee, while the Shuttle descends slightly

to a 116 rani x 288 nmi elliptical path. With this combination of

initial Shuttle orbit and tether length, the appropriate conditions

are met at tether release. As the AXAF goes through ite apogee it is

released and is immediately injected into its final circular orbit at

320 nmi while the Shuttle goes to the desired perigee (100 x 288 nmi

orbit). As the Shuttle moves into its lower orbit the PIDM and

tether are retrieved back to the Shuttle. At a subsequent

near-apogee passage, after cargo bay doors are closed, the Shuttle

will make its deorbit burn for reentry.

For the baseline case, most of the integral OMS propellant (23,700

lb) is required, because of the high orbit. For the tethered

approach the total propellant required is reduced by 16,200 lb if a

direct deorbit from apogee is used. In this case, the Shuttle orbit

(both before and after tether deployment) is precessing in plane by

about 11 to 12 deg per day due to earth oblateness. If the mission

takes several days, an off-apogee deorbit burn will be required to

match proper reentry conditions for the landing site location. The

off-apogee burn could require an additional several thousand pounds

of propellant but overall propellant savings are expected to be in

the 11,000 to 16,000 lb range, in any case. A small cold gas penalty

(less than 500 lb) would be incurred during the tether operations for 

deployment and retrieval.

The maximum tether tension is small for this mission (i.e., 477 lb or

2122 N) and tether energy generation during deployment is about 17

kWh (to be disposed of on the Shuttle). Energy for retrieval of the

tether and PIDM is about 1 kWh (17 kWh would be required if the AXAF

were to be retrieved in an abort situation). Shuttle tether system

64
CRITERIA BASELINE

APPROACH

TETHERED

("ON CEP I COMPARISON EVAlUAT10N/COMMENTS

1. QUANTITATIVE MISSION CRITERIA

A. ORBITS (NMI AT 28.5 DEG. INCL)

o INITIAL ORBIT n x 32( !I' t 290 LOWER STS INJECTION ENERGY FOR TETHER

CASE.

PRE-RELEASE OR PRE-RENDEZVOUS ORBITS

o DEPLOYER (SHUTTLE) N/A 11b x 288

o DEPLOYED MASS (AXA r ) N/A 148 x 320

o (TETHER LENGTH - NMI) N/A (32.5)

POST-RELEASE OR POST-REN_DEZVOIIS ORBITS

o DEPLOYER N/A 100 x 288 PERIGEE REDUCED FROM 116 TO 100 NMI.

o DEPLOYED MASS N/A 320 x 320

FINAL ORBITS

o DEPLOYER '20 x 320 100 x 288 ASSUMES DEORBIT FROM APOGEE. LOWER.

o DEPLOYED MASS 320 x 32n 320 x 320

STS DEORBIT ENERGY FOR TETHER CASE

U. PROPELLANT USAGE (LB)

o SHUTTLE ORBITER 23700 -5001 -16200 SUBSTANTIALLY REDUCED ORBITER PRO P EL0 OMV

LANT LOAD REQUIRED

o OTV

o SPACE STATION

o END FFFECTOR (PIOM OR SIDM) N/A <500 500 SMALL PENALTY

C. TETHER SYSTEM

o MAXIMUM TENSION LB N/A 477

-- t2122N)

ENERGY USAGE (KWH)

o DEPLOYMENT N/A '7 1; UPBITER DISPOSES 17 KWH IN 4-6 HRS.

o RETRIEVAL N/A ! I ORBITER SUPPLIES 1 KWH IN 4-6 HRS.

o SYSTEM WEIGHT (LB) N/A 9rnn 1000 ESTIMATED TETHER SYSTEM WEIGHT IN

CARGO BAY

D. SPACE STATION MOMENTUM

o MOMENTUM GAIN (Kr,-M 2 1SEC) N/A N'A

F. OPEPATIONAL TIME (HRS.) 1BC TBD

11. QUALITATIVE MISSION CRITERIA

A. COMPLEXITY (0-2)

o OPERATIONS 1.0 1.t - 1.5 +.2 TO .5 20; 50'^ MORE OPS. COMPLEXITY

o HARDWARE 1.0 1.I - 1.4 .l TO .4 10^ TO W MORE HOW. COMPLEXITY

8. RISKS (0-2)

o MISSION SUCCESS 1.0 1.1 1.1 10: MORE MISSION SUCCESS RISK

o TECHNOLOGY 1.0 1,05 +.05 5. MORE TECHNOLCGY RISK

C. COST FACTOR (0-2)

o LAUNCH 1.0 1.0 0 SAME COST SINCE DEDICATED LAUNCH

o HARDWARE DEVELOPMENT 1 1.0 11.C5 j -.OS 5' MORE hDW. nEVFLOPM NT COS T

!TOTES: 1) ORBITER QMS PROPELLANT RLQUIREME!+T CUULC BE INCREASED BY SEVERAL THOUSAND POUNDS TG ALLOW DLORBITS

AHEAD OR AFTER APOGEE OF FINAL ORBIT (T" HER CASE). APOGEE PRECESSES 11-12 DEG/DAY,

Table 3-12 Comparison of Concept B1 with Baseline Approach

65
payload. An operational time comparison has not been determined at

this time, but it is ex p ected that the tethered approach would take

16 hrs for deployment and retrieval with an unspecified additional

time for checkout while the AXAF is on the tether. The baseline

approach would probably take less time.

The operations complexity is expected to be substantially higher (20%

to 50% more) for the tethered approach. The hardware complexity is

also expected to be higher (10Z to 40%) than the baseline approach.

The mission success risk and the technology risk are expected to be

somewhat higher (10% and 5%, respectively) for the tethered

approach. Launch cost should be about the same for the tethered

concept while hardware development cost may be slightly higher (about

5%) .

Overall, 11,000 to 16,000 lb of Shuttle OMS propellant could be saved

with the tethered approach, but much of this would be offset by the

tether system weight of 8000 lb. The additional complexity of the

tether operations (over the baseline approach) would probably not be

desirable for this initial placement mission, particularly if it

remains a dedicated launch. The tethered approach should be

considered as a backup to the baseline approach in the event that

more pay load capability is required (about 3000 to 8000 lb

potentially available), or in the event a higher altitude orbit

becomes desirable.

3.3.4 Retrieval of 20,000 lb AXAF from 205 nmi Circular Orbit for

Maintenance and Reboost to 320 nmi - Concept B2

Since the AXAF spacecraft will gradually decay (due to drag) from its

initial altitude (in approximately 3 yrs) to an altitude of about 205

n-mi, the plan is to accomplish periodic maintenance on board the

Shuttle and then reboost the AXAF back to its original orbit. In the

baseline approach the OMV is brought up from the ground in the

Shuttle (160 nmi parking orbit) and does the actual retrieval and

reboosting of AXAF from 205 nmi down to 160 nmi and back up to 320

nmi and returns to the Shuttle. A shared payload would also be

released at 160 nmi when scheduling, cargo weight and volume limits

permit. The OMV is fully loaded with propellant to allow a return of

the AXAF from 320 nmi (after reboosting) in the event of checkout

failure. In this abort situation the AXAF would be returned to the

ground for repair and later relaunch.

A tethered approach for this AXAF servicing mission is compared with

the baseline approach in Table 3-13. In this case, the Shuttle would

maize a direct insertion to 205 nmi (vs. 160 nmi) altitude to

rendezvous with the AXAF spacecraft. The OMV would not be required,

but the Shuttle tether deployer system would be carried in the cargo

bay. After rendezvous, the AXAF spacecraft would be brought aboard

for maintenance and repair. While at the 205 nmi altitude, a shared

payload could also be released, if available.

66
CRITERIA BASELINE TETHERED

COMPARISON EVAIJATION/COMMENTS APPROACH CONCEPT

1. QUANTITATIVE MISSION CRITERIA

A. ORBITS (NMI AT 28.5 DEG. INCL)

o INITIAL ORBIT 160 x 160 205 x 205 DIRECT fNSERTICN ORBIT OF ORBITER

205 x 303 APOGEE RAISED AFTER AXAF RETRIEVED'

PRE-RELEASE OR PRE-RENDEZVOUS ORBITS

o DEPLOYER (SIP ITTLI) N/A 204 x 302

o DEPLOYED MASS (AXAF) N/A 222 x 320

o (TETHER LENGTH - NMI) N/A (18.2)

(ryQ*r RCUUCi;ON OF TETHEk 44CNGllI FROM

AXAF PLACEMENT MISSION-B1)

POST-RELEASE OR POST-RENDEZVOUS ORBITS

o DEPLOYER N/A 194 x 302 PERIGEE REDUCED FROM 204 TO 194 NMI

o DEPLOYED MASS N/A 320 x 120

FINAL ORBITS

o DEPLOYER 160 x 160 194 x 302 ASSUMES DFORBIT FROM APOGEE FOR

o DEPLOYED MASS 320 x 320 320 x 320 TFTHEREO CASL

B. PROPELLANT USAGE (LB)

o SHUTTLE ORBITER 13100 194002 -6300

INCREASED ORBITER PROPELLANT REQUIRED

FOR TETHER CASE

o OMV 6700 0 +6700 OMV 8 PROP. NOT REQ'D. FOR TETHER CASE

o OTV

o SPACE STATION

o END EFFECTOR (PIDM OR S1DM) N/A <500 - 500 _MALL PENALTY

C. TETHER SYSTEM

o MAXIMUM TENSION (LB) N/A 259

(1152N)

ENERGY USAGE (KWH)

o DEPLOYMENT N/A 6 + 6 ORBITER DISPOSES 6 KWH IN 6 HRS.

o RETRIEVAL N/A 1 - I ORBITER SUPPLIES 1 KWH IN 4-6 HRS.

(MUST SUPPLY 6 KWH FOR ABOR T

PETRIEVAL)

o SYSTEM WEIGHT (LB) 24700 22100 +2600 ASE REQUIRED TO SUPPORT AXAF REBOOST

IS LESS FOR TETHERED CASE.

D. SPACE STATION MOMENTUM

o MOMENTUM GAIN (KG-M 2

/5[C) 0 N/A

E. OPERATIONAL TIME (HRS.) 3 64.0 EST 16.0 EST 48.0 EST TETHEREC APPROACH REDUCES OPERATIONAL

TIMF BY ABOUT 48 HOURS.

11. QUALITATIVE MISSION CRITERIA

A. COMPLEXITY (0-2)

o OPERATIONS 1.0 1.0 - 1.2 0 TO .2 0, TO 20% MORE OPS. COMPLEXITY

o HARDWARE 1.0 1.1 - 1.2 +.1 TO .2 10.' TO 20: MORE HOW. COMPLEXITY

B. RISKS (0-2)

o MISSION SUCCESS 1.0 1.0 - 1.0 0 TO .05 U TO 5h MORE MISSION SUCCESS RISK

o TECHNOLOGY 1.0 1.0 0 SAME TECHNOLOGY RISK

C. COST FACTOR (0-2)

o LAUNCH 1.0 1.0 - 0.9 0 TQ -.Iu O 10 10 .

REDUCTION IN LAUNCH COST

o HARDWARE DEVELOPMENT 1.0 1 1.05 .05 5' INCREASE IN DEVELOPMENT COST

NOTES: 1) COMPLICATES OPERATION.

2) ORBITER OMS PROPELLANT REQUIREMENTS COULD BE INCREASED BY SEVERAL THOUSAND POUNDS TO AIIOW DEORBITS AHEAD

OR AFTER APOGEE OF FINAL ORBIT (TETHER CASE). APOGEE PRECESSES AT +11-12 DIG/DAY.

3) TIMES DO NOT INCLUDE AN 86 HR. PERIOD (ESTIMATED) AT THE ORBITER I"OR AXAF MAINTENANCE AND REPAIR. THIS

TIME PFRIOD IS ASSUMED TO BE THE SAME FOR BOTH VERSIONS.

Table 3-13 Comparison of Concept B2 with Baseline Approach

67

T
required co match the 320 nmi velocity conditions in this case

(substantially shorter than for Concept B1). As the tether is

deployed upward to its full length the AXAF rises into an elliptical

path (222 x 320 nmi) while the Shuttle descends slightly to a 204 x.

302 nmi elliptical path. After the AXAF is determined to be ready

for release, it will be released on the next apogee passage and

inject directly into its 320 nmi circular orbit, while the Shuttle

moves into a 194 x 302 nmi elliptical orbit. As the Shuttle moves

into ita lower orbit the PIDM and tether is retrieved for storage on

the Shuttle. At a later near-apogee passage, after cargo bay doors

are closed, the Shuttle will make its deorbit burn for reentry (in an

abort situation, the AXAF would be retrieved by tether before release

and rest,-ned to the ground aboard the Shuttle).

For the baseline case about 13,100 lb of Shuttle OMS propellant is

required (160 nmi orbit). The tethered approach requires a minimum

of 19,400 lb (6300 lb more), but this could be increased by several

thousands of pounds due to the off-apogee condition discussed in the

Concept B1 comparison (3.3.3). A small cold gas penalty (less than

500 lb) would also be incurred during the tether deployment and

retrieval operations.

The maximum tether tension is only 259 lb (1152 N) for this mission

and tether energy during deployment is about 6 kWh. Energy for

retrieval of the PIDM and tether is about 1 kWh (6 kWh would be

required for retrieval of AXAF in an abort situation). Estimated

system weight in the cargo bay is 24,700 lb for the baseline approach

(including the OMV and AXAF maintenance gear) and approximately

22,100 lb (a reduction of 2600 lb) for the tethered system (including

tether deployer system and AXAF maintenance gear). Operational times

using the OMV and tethered deployment are estimated at 64 hrs and 16

hrs, respectively. These times do not include the 86 hr period

(estimated) while the AXAF is in the cargo bay undergoing maintenance

and repair.

The operations complexity would be somewhat higher in the tethered

approach (0 to 20X). The hardware complexity is also expected to be

higher (10% to 20%) compared to the baseline approach. The mission

success risk could be somewhat higher (0 to 5%) for the tethered

approach, but the technology risk should be equivalent to the

baseline approach. Tether deployment should cost somewhat less (0 to

10%) compared with the baseline approach, using the OMV, but hardware

development costs may be up to 5% higher.

68
gpp

utilize its 4000 lb to 7000 lb payload advantage (volumewise) it

mi.-ht more than offset the launch cost advantage of the tethered

approach. A more detailed examination would be required to recommend

the best approach.

3.3.5 Tethered OMV Rendezvous and Retrieval of OTV Returning from a

Geosynchronous Mission - Concept D

Although the rendezvous and docking techniques for a tethered OMV to

retrieve a returning OTV (Concept D) have not been established,

sufficient information is now available (from recent OTV work at

Martin Marietta) to make a comparison of benefits between the

tethered approach and the baseline approach and is presented in Table

3-14. Before discussing the comparison, it is worth while to review

the overall JTV geosynchronous mission, using the baseline approach.

After the OTV and payload have been moved to the launch point (near

an equatorial crossing) the OTV main engines are fired to raise the

orbit apogee to geosynchronous altitude. At geo the apogee burn

simultaneously circularizes the orbit and changes planes (from 28.5

deg to 0 deg inclination). The OTV will then spend up to

approximately 24 hrs (time interval to the second Space Station orbit

plane crossing) to distribute up to a maximum of 4 payloads in

geosynchronous orbit. After this period, the OTV will execute a

deorbit burn (with plane change) to match the orbit plane of the

Space Station and target for an aerobraked passage through the

atmosphere (one or more mid-course corrections will be required for

accurate targeting). Some additional propulsion may be required to

control atmospheric exit conditions. The OTV will then be in an

orbit that reaches the Space Station altitude (at apogee) with the

Space Station ahead of the OTV. At apogee, the OTV will use a small

propulsion maneuver to raise its perigee above the atmosphere (e.g.,

100 nmi). The OTV will then remain in this orbit for several

circuits until final phasing occurs. At a later apogee passage a

final OTV burn is made and the OTV is then located in the Space

Station orbit (outside of proximity Zone 2) at a point about 30 nmi

behind the station. Very little phasing is required in the low orbit

since a planned fast or slow return (up to 1 hour or more variation

in return trip time from geosynchronous orbit) will eliminate any

need for long phasing times in low orbit. An OMV mission will then

be required to retrieve and return the OTV to the Space Station.

In the tethered OMV retrieval approach the first portion of the OTV

mission would be the same, but the returning OTV would be targeted to

a 220 nmi apogee subsequent to the aerobraking pass through the

atmosphere. At this point the orbit would be circularized and the

OTV would raise with the tethered OMV located below the Space

69

4j' y 
CRITERIA

rOMPARISON IVAIUATION/COMMLNTS

APPROACH CONCEPT

1. QUANIITIT IVE MISSION CRITERIA

A. ORdiTS (NMI AT 28.5 DEG. INCL)

o INITIAL. ORBIT 770 x 270 ?70x ?70

PRE-RELEASE OR PRE-RENDEZVOUS ORBITS

o OEPLOYER TSIPACF STATION) N/A 270 x 270 UPCOMING OTV IN 220 x 263 NMI ORBIT

o DEPLOYED MAY, (OMV/OTV) N/A 763 x 263 FOR RENDEZVOUS AT APOGEE.

u 1TCTHER LENGTH - NMI) N/A (7)

POST-RELEASE OR POST-RENDEZVOUS_O_RBITS

o DEPLOYER N/A 269 x 270

o DEPLOYED MASS N/A 262 x 263

FINAL ORBITS SPACE STATION ORBIT NOT SIGNIFICANTLY

0o DEPLPYCR 270 x 270 269 x 270 AFFECTED BY DIV RETRIEVAL OPERATION

o DEPLOYED MASS 270 x 270 269 x 270

WITH TETHER,

B. PROPELLANT USAGE. (LB) PROPELLAN' 11 %AGE IS COMPARED B UR THE

PERIOD AF FF OTV AERO BRAKING.

o SHUTTLE ORBITER

o OMV <500 <500 0 EST. NO P ROPELLANT SAVINGS.

o DIV <500 <50n 0 EST. NEGLIGIBLE PROPELLANT SAVINGS.

o SPACE STATION

o END EFFECTOR (PIOM OR SIDM) N/A

C. TETHER SYSTEM

o MAXI MUM TENSION (LL N/A 58

(258N)

ENERG Y USAGE (KWHI

o DEPLOYMENT N/A <1 KWH +1 KWH NEGLIGIBLE ENERGY DISPOSITIONED

o RETRIEVAL N/A <1 KWH -1 KWH NEGLIGIBLE POWER SUPPLIED

o SYSTEM WEIGHT (LB) N/A ?500 -'5000 ESTIMATED SS REUSABLE TETHER SYSTEM

- DFPLOYER WEIGHT

D. SPACE STATION MOMENTUM 1.309060 x 10 16 KG-M2 /SEC (REFERENCE)

o MOMENTUM GA114 (KG-M2 /SEC) 0 0 0 NEGLIGIBLE CHANGE 114 SS ANGULAR

MOMENTUM

E. OPIRATIONAL TIME (HRS.) 5.0 EST. 16.0 E`T - 11 TETHER APPROACH REQ. ADD'L 11 HAS.

11. QUALITATIVE MISSION CRITERIA

A. COMPLEXITY (0-2)

o OPERATIONS 1.0 1.5 - 2.0 x .50 - 1.0 502-100% MORE OPS. COMPLEXITY

o HARDWARE 1.0 1.5 - 2.0 -.50 - 1.0 SO -100: MORE HOW. COMPLEXITY

R. RISKS (0-2)

o MISSION SUCCESS 1.0 1.25-1.5 +.25-.50 252-50% MORE MISSION SUCCESS RISC

o TECHNOLOGY 1.0 1.5 - 2.O +.50-1.0 503-1003 MORE TECHNOLOGY RISK

C. COST FACTOR (0-2)

n LAUNCH 1.0 1.25 • .25 25: MORE LAUNCH COST

o HARDWARE DEVELOPMENT I.0 2.0 1 4 1.0 100% MORE HOW. DEVELOPMENT COST

NOTES: I) IIUE TO THE VRECISE NAVIGATION AND GUIDANCE ACCURACY R^QUIRLMENT FOR THE OTV 1 0 ENTER THE ATMOSPHERE, THE

OTV IS ABLE TO INJECT INTO THE SAME. ORBIT WITH THE SPACE STATION IN THE NEAR VICINITY (ABOUT 30 NMI)

BEHIND IT, THIS GREATLY SIMPLIFIES THE BASELINE OTV RETRIEVAL MISSION AND MINIMIZES PROPELLANT USAGE ON

THE OMV.

Table 3-14 Comparison of Concept D with Baseline Approach

70
rendezvous and docking would be controlled by the TUMV and tethe

deployer, zs presently envisioned.

The detailed comparison of the TOMV retrieval approach (Concept U

with the previously described baseline approach for returning the OT

to the Space Station is presented in Table 3-14. The TOMV i

att • :hed to the tether at the lower position on the Space Station (a

270 nmi altitude) and deployed with its forward (probe end) Lacin

downward. Tile tether length required is about 13 km (7 nmi) for thi

mission. Since the mass of the partially fueled OMV is quite lo

compared to the Space Station mass, the TOMV is lowered to a

altitude of 263 nmi and the Space Station essentially remains at 2i

nmi altitude. After retrieval of the OTV (rendezvous near apogee o

the 220 nmi x 263 nmi OTV final orbit), the added mass causes th

Space Station to move to a 269 x 270 nmi orbit and Zhe TOMV/OTV i

now in a 262 x 263 nmi elliptical path. After retrieval, the Spac

Station remains in an orbit of approximately 269 x 270 nmi.

The propellant usage period of interest (for comparison) is th

period after the OTV aerobraking maneuver. Because of the navigatio

and guidance requirement for the OTV to make a precise targeting o

the atmospheric entry corridor (within 5 or 10 nmi of altitude) fro

geosynchronous altitude, the OTV has the capability to do precis

orbit injection and phasing to return to close proximity to the Spac

Station. Both the OTV and the OMV will require very littl

propellant for those conditions (less than 500 lb each). With th

tethered approach, similar propellant requirements (less than 500 l

each) exist. It is believed at this time that OMV and UTV propellan

savings will be aebligible with the tethered approach.

Maximum tether tension (during retrieval of the TOMV with OTV) i

only 58 lb (258 N) and energy usage is less than 1 kWh durin

deployment or retrieval. The estimated Space Station reusable tethe

system weight is 25,000 lb (as described in Concept A-2) and Space

Station angular momentum change is negligible. Estimated operationa

time for the baseline approach is 5 hrs (OMV retrieval time) and 1

hrs for the tethered rendezvous operation (deployment and retrieval).

ne tethered rendezvous approach operations complexity and the

hardware complexity are both believed to be much greater than th

baseline approach (both are 50% to 100% more complex). Missio

success risk is projected to be 25% to 50% greater than the baseline

and the technology risk may be 50% to 100 % greater than the baselin

approach. Launch cost for the tethered approach is expected to u

about 25% greater and hardware development cost could be 100% greater

71
compared to 5), and operations and hardware complexity is

substantially greater. Because of the perceived complexity of the

tethered capture/rendezvous, both the mission success risk and the

technology risk will be significantly higher with the tethered

approach and higher launch and hardware development costs will be

incurred. 

3.4 Benefits Analvsis

A useful way of looking at the Space Station angular momentum gains

and losses is to observe the average change in Space Station altitude

as a function -)f the activities that affect altitude. Activities

considered in t',is analysis were Space Station drag decay, tethered

OMV launch ef:-ects, tethered OTV launch effects, and tethered Shuttle

deorbit effects. Sincc crew rotation is projected to be at quarterly

intervals and drag makeup will occur at least quarterly, momentum

balance was treated on a quarter year basis.

The basic approach was to take the data from the OMV and OTV mission

models (Table 3-1 and 3-2) and record the tether length requirements

and propellant savings (for later uRe) for the OMV launches (from

Figure 3-6) and the UTV launches (from Figure 3-8 for a 150 km tether

length) and calculate the average decrease in Space Station altitude

(per quarter) using the tether relationships given in Table 3-3.

Average Space Station altitude lass per quarter was taken from Table

3-9. This procedure allowed the basic activities that cause Space

Station altitude losses to be collected and summarized. Those are

presented graphically in Figure 3-9. (Note that OMV and OTV launches

cause some variat , ;n in Space Station orbit eccentricity, but this

analysis treats average altitude changes only.)

Notice that the effects of tethered OMV launches is very small (due

to the small number of OMV launches) and maximum quarterly Space

Station altitude loss is less than 4 nmi in 1998. Space station

average altitude loss due to aerodynamic drag reflects the varying

yearly atmospheric density of the sunspot cycle, with a maximum

average quarterly loss of 12 nmi in 1991. Significant altitude loss

effects begin to occur in 1995, when the first tethered OTV launches

would be initiated. Tethered OTV launch effects vary from a minimum

average quarterly Space Station altitude loss of about 29 nmi (1995)

to a maximum of 100 omi (1996). Total average quarterly altitude

losses from the three activities start at 12 nmi in 1991, reach a

minimum of 6 nmi in 1994, a maximum of 106 nmi in 1998, and remain

over 90 nmi in the year 2000. There is no intention to allow the

Space Station altitude to vary by tnose larger amounts, but to

balance altitude looses with gains from tethered Shuttle deorbit

operations.

12
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73
:altitude losses induced by the activities presented in Figure 39.

Tile data reflects the average number of Shuttle tether deployments

per quarter for each year of the mission period and is based on

average orbit altitude changes (i.e., as with OMV and OTV launches,

orbit eccentricity effects are not treated). All calculations are

based on having the Space Station at a nominal altitude of 270 nmi,

as befora.

For the first five years (1991-1995) the minimum of 1 tethered

Shuttle deorbit per quarter (4 per year) will be adequate to take

care of drag decay, the few UMV launches, and the first OTV flights

in 1995. Shuttle deorbit tether lengths will vary from a minimum of

10 km ( 1994) to a maximum of 53 km in 1995. The decreasing tether

length requirement from 1991-1994 reflects the variation in Space

Station drag aerodynamic/effects previously discussed. For this

estimate it has been assumed that no ele%,rodynamic power tether is

being used.

In general, there are always more Shuttle Flights going to the Space

Station than the momentlm balance can fully utilize, particularly in

the 1991-1995 time period. As the number of OT's launches Suild up,

the number of Shuttle missions to the Space Station also increases to

bring UTV propellant and payloads to the Space Station. The minimum

number of tethered Shuttle decrbits required varies from 4 annually

(1 per quarter) up to 11 annually (about 1 per month) in the late

90s. Up to 14 Shuttle Space Station missions per year (1999) are

planned for this time period by the "Nominal Mission Model (FY

19113-2000), Revision 7 (SS), Space Statior Advocacy," July 1984, MSFC..

The potential yearly benefits from Space Station operations are

summarized for the 1991-2000 time period in Figure 3-11. The direct

benefits are aerived from propellant savings (various types) from

using tethered launch and deployment techniques and also from the

elimination of UMV operations associate,2 with OTV launches. The

types of propellant savings shown are Shuttle OMS propellant

(nitrogen tetroxide/monemethyl hydrazine), UMV propellant (same as

Shuttle uMS), cold ,as (nitrogen) savings from the elimination of UMV

usage for OTV operations, OTV propellant (LOX/hydrogen), and Space

Station orbit stationkeeping propellant (hydrazine assumed).

74
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100 —

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16

SPACE STATION

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1991 1992 1993 1994 1995 1996 1997 1998 1999 2000

Figure 3-11 Potential Yearly Benefits from Space Station Tether Operations

i

76
77

deorbit operations. Space Station orbit maintenance is the major

user of angular momentum in the early years and allows ti:- associated

scavenging of OMS propellant and the saving of orbit maintenance

propellant. Propellant savings associated with OMV tethered launches

are not significant. Overall, propellant savings average about

59,000 lb per year over the 10 year period, with a maximum of 122,000

lb in 1998. Details of the various potential propellant savings, OMV

flights saved (69 total from 1995-2000), and minimum number of

tethered Shuttle deorbits are presented in Table 3-15.

In summary, because of the potential propellant savings from tether

operations, the Space Station propellant resupply requirements are

significantly reduced, making the Space Station a more efficient

(cost-effective) system in terms of accomplishing its mission as a

space base for launching OMVs and OTVs. The scavenged OMS propellant

can be used to support OMV operations and the OTV propellant savings

appear as reduced requirements. Since the tethered Shuttle deorbit

operations eliminate the need for orbit maintenance propellant, tnis

savings is also a reduced requirement. Further, the potential

elimination of the 69 OMV missions required for OTV launch operation

represents still another cost savings.

Other potential tether benefits include the potential for increasing

the performance envelope of the OMV (increased payload-altitude

range) and also a potential increase in payload weight to

geosynchronous orbit for the OTV. If tethered launches and Shuttle

deorbit operations are planned for the Space Station from the start

and orbit maintenance is handled without propellant, a simpler (lower

cost) Space Station propulsion system may be possible.
.

w

py

The selected cuncepts which involve tether deployment and which were

used as the basis iur defining the categories for tether deployment

systems are:

Concept A2 - Tethered Deorbit of Shuttle from Space Station,

Concept F - Tether Assisted Launch of an Orbital Transfer

Vehicle to Geosynchronous Orbit

Concepts B and B2 - Orbit Insertion of AXAF by Tether from

Shuttle.

Concept F is a companion concept to A2 which enables the beneficial

ut_lization of the angular momentum imparted to the Space Station by

the tethered deorbit of the Shuttle. Without some method of reusing

this angular momentum, the number of times the tethered deorbit of

Shuttle could accur would be limited to infrequent intervals

corresponding to the ?ngular momentum requirements for Space Station

orbit maintenance lctivity. This orbit maintenance activity would

consume only L sm-11 fraction of the angular momentum potentially

available to ue #--ritsferred to the Space Station from the Shuttle

deorbit operation-.

One other tether deployment concept from Space Station which was

considered was a tether assisted launch of the Gruital Maneuvering

Vehicle (OMV). The conclusion reached was that tha fraction of OMV

:missions which coula use tether assist was on the order of one in

twenty (Sae 3.1.1). In addition, the actual benefit in terms of

overall propellant savings due to this application was insignificant

relative to that which could be saved with the tether launch assist

to the 0TV missiunG. For these reasons the tether assisted launch of

the OMV was not considered as a basis for the design of a deployment

system. In the event it becomes desirable to perform such an OMV

launch assist, the OTV tether deployer system as described in this

section should be adequate.

The necessary teaming of Con,-ept A2 for the tether deorbit of Shuttle

at," Concept F iur the tether launch of an OTV mission to

geosynchronous orbit, leads to the consideration of a dual mode

deployer system for Space Station which could alternately perform the

downward deployment of the Shuttle and then the upward deployment of

the OTV mission. There is an obvious necessity to alternate these

upward and downward deployments to keep the angular momentum (average

orbit altitude of the Space Station) within acceptable limits.

This angular momentum balance analysis based on projected mission

models is treated in Section 3 of this report. The analysis

indicates that the controlling element in the balance is the

78
operations than can be used by the tether launch assist of the OT

missions. A conclusion is that the concept of angular momentu

balance between the upward and downward deployments co• ild accommodat

a significant increase (approximately 50%) in the OTV mission mode

traffic before reaching a limit on the amount of angular momentu

available to be scavenged from the downward deployments of th

Shuttle. As pointed out earlier, even if the Shuttle deorbit limi

were reached, other masses such as waste packages or external tank

are available for tet.her deorbit.

An alternative approach to beneficial use of this transferred angilla

momentum would be the use of an electrodynamic tether as an auxiliar

power system for the Space Station. However, concurrent use o

Concepts E2 and F are not deemed feasible due to conflicting locatio

and operational requirements for the two systems. The use of Concep

E2 until the advent of the OTV operations in 1995 would permit th

transfer of angular momentum and the scavenging of OMS propellan

from the Shuttle to begin much earlier than would be the case if OT

launch assist were the only method of using the scavenged angula

momentum. For this study, it has been assumed that the operational

and economic benefits of the tether assisted launch of the OT

r'ssions outweigh those from the use of Concept E2 to provide a

auxiliary power system. Such a planned transition :could require 

decommissioning of the auxiliary power system and institution of th

OTV launch assist capability when the space based OTV operation

begin in 1995.

For the purpose -f this study, it was decided to develop a tether

deployment system concept that would satisfy the combined

requirements of Concept A2 and Concept F. The requirements for suc

a system are given in 4.2 and the resulting design concept i

described in 4.3.

The deployment of payloads from the Shuttle into operational orbi

that are above those easily reached by the Shuttle is exemplified by

Concept B (AXAF deployment).

The requirements for this Shuttle deployment system are given in 4.

and the resulting design concept is described in 4.5.

4.2 System Performance Requirements for a Tether Deployment System for

Space Station

79
4.2.4 Shuttle Interface Deployment Module (SIDM)

The Space Station reference configuration to be used as the basis for

the Phase B studies was used for the definition of requirements and

concept development for this study.

Mass values used for cne primary elements were:

Space Station 550,000 lbm (250,000 kg)

Shuttle Deoroit 220,000 lbm (100,000 kg)

Fueled OTV with Payload 73,200 lbm ( 33,300 kg)

4.2.2 Reference Altitude(s)

4.2.3

The altitude and altitude range to be considered for the Space

Station orbit is 270 +40 -20 nmi.

The perigee altitude of the Shuttle subsequent to release from the

tether is to be constrained to be above 100 nautical miles to insure

against automatic re-entry on the first perigee pass after release

from the tether.

Tether Lengths/Tensions

The design constraint to be used for upward tether length deployment

of the OTV is to develop a roughly equivalent level of tension in the

tether to that developed by the downward Shuttle deorbit operation.

These considerations lead to a maximum tether length for the orbiter

deployment of 65 km and 150 km for the OTV. The corresponding

tension levels (rounded upward) are 3900 lbf for the Orbiter at 65 km

and 4000 lbf for the OTV at 15C km. These values include the tension

due to the tether mass at 40 kg/km and using a design factor of 2 +.

The SIDM must provide the load transfer interface between the tether

deploymen*_ system and the Shuttle. It must be capable of two fault

tolerant release capability under full tension load with the primary

release under control of the Shuttle crew and backup release

controlled by the Space Station crew. The SIDM shall not intrude

into the cargo bay payload clearance envelope so as not to constrain

the return cargo capabilities of the Shuttle.

The SIDM shall incorporate a propellant transfer interface to the

Shuttle such that OMS bipropollant can be transferred from the

Shuttle to the SIDM during the tether deployment operation. This

propellant transfer interface is to incorporate a remotely controlled

disconnect capability to be implemented prior to the release of the

SIDM from the Shuttle. The SIDM shall provide storage capacity for

OMS bipropellant in excess of 6500 pounds.

80
operations during installation and retrieval of the SIDM and for

possible recovery of the SIDM by the OMV in the event of contingency

situatious where the SIDM has been jettisoned.

The SIDM shall incorporate a cold gas propulsion and attitude contro

system to provide positive control of the attitude and location of

the SIDM during the final close-in phase of the tether retrieval

operations. The initial separations from the Space Station shall be

performed by the Shuttle RCS out to a distance such that adequate

gravity gradient tension forces have built up to complete the tether

deployment operation. This initial separation is in the range fro

0.5 to 1.0 km.

• The SIDM shall incorporate a remotely controlled disconnect or

guillotine capability to disconnect the SIDM, from the tether. This

capability is to be used in the event of a severed tether such that

the SIDM could be secured prior to being jettisoned by the Shuttle

with the intent of a subsequent retrieval operation by an OMV missio

from Space Station.

The SIDM shall incorporate a retro reflective cube surface or

equivalent system to permit location tracking of the SIDM by Space

Station systems during the tether retrieval process.

The SIDM shall incorporate a berthing interface capability with th

tension alignment carriage on the lower tether deployment assembly.

The SIDM shall incorporate a command/control data link to the Spac

Station to permit statusing and control functions to be performed for

the onboard systems. This would include the tether disconnect,

propellant transfer disconnect, the secondary release mechanisms and

the cold gas attitude control system.

4.2.5 Payload Interface Deployment Module (PIDM)

The PIDM must provide the load interface between the tether

deployment system and the OTV/payload assembly. It must be capable

of releasing from the OTV under full tension load. The releas

operation will be performed by remote control from the Spac

Station. The location of such a tether system load interface wit

the OTV has not been determined. It has been assumed for study

purposes that it is in an axial location on the main engine end of

the OTV. This location will minimize sensitivity of the interface to

variations in the mass distribution of the OTV/payload missio

stack. It may be desirable to incorporate a controlled tip off rate

release capability to allow the use of forces generated during tether

release to effect the 90' reorientation of the OTV stack prior to

initiating the orbit transfer burn. For study purposes assume a

symmetrical release with zero tip off rzte.

81
the Space Station until tether tension forces build up to adequate

levels to complete the deployment operation and to provide control of

the location and attitude of the PIDM during the final close-in phase

of the tether retrieval operation.

The PIDM shall incorporate a standard grapple fixture accessible by

the Space Station KLMS for any handling operations required on the

Space Station. The grapple fixture will also be compatible for use

by the OMV in the event of a contingency mode where retrieval of the

PIDM by the OMV is required.

The PIDM shall incorporate a remotely controlled disconnect or tether

guillotine capability to disconnect the tether from the PIDM.

The PIDM shall incorporate a retro reflective cube surface or

equivalent system to permit location tracking of the PIDM by Space

Station systems during the deployment and retrieval operations.

The PIDM shall provide a berthing interface capability with the

tension alignment carriage on the upper tether deployment assembly.

i The PIDM shall incorporate a command/control communications link to

the Space Station to permit the exercise of statusing and control

tunctions. These will include the tether disconnect, the release

function, and the cold gas attitude control system.

4.2.6 Deployment and Retrieval Energies

The energy generated during a 150 km deployment of a maximum OTV

stack is 366 kWh. This includes the energy generated by the

deployment of the PIDM and tether. The retrieval of the PIDM and

tether requires a retrieval energy to be supplied of 29 kWh.

The energy generated during a 65 km deployment of the Shuttle is 153

kWh and the energy required to be supplied for retrieval of the SIDM

(loaded with 0500 pounds of OMS propellant) and the tether is 11 kWh.

82
the tether tension force with the Space Station centerolmass to

control disturbance torques on the station. The tension alignment

system is to provide an angular range of + 0.1 radian about the local

vertical in the plane of the orbit. Cross plane angles are

considered to be negligible and no provision for cross track

alignment is required.

The tension alignment assemblies are the most outboard portion of the

station in contact with the tether. They are to include a provision

• to sever the tether at this outboard location for use in event of a

broken or severed tether.

4.2.8 Tether Reel Assembly(s)

The reel assembly shall provide the capability to deploy and retrieve

the tether under the required tension levels. This includes the

capability to modulate the deployment and retrieval rates in

accordance with specified tension control laws. The energy generated

during the deployment operation is to be converted to electrical

energy by means of the tether motor/generator drive system operating

in a generator mode. The system shall be capable of either supplying

this power as input to the Space Station power bus or of directi<ig it

to a high temperature resistive load bank such that it can be

radiated to space as waste heat.

The reel assembly shall include a friction brake capable of locking

the reel against full tether tension once full .eployment has been

attained.

During retrieval operations the tether drive system will function in

a motor mode and will be supplied with electrical power from the

Space Station power system to perform the retrieval.

4.3 Design Concept for a Dual Mode Tether Deployment System for Space

Station

• The concept must provide the capability to alternately deploy

payloads downward (Shuttle) and upward (OTV). This means that

deployment locations will be required at both the nadir and zenith

locations on the Space Station centerline. Due to the vertical

dimension of the station (400 ft) the following options were

considered:

1. Two complete systems - one nadir and one zenith.

2. One system that could be translated between the locations by

means of the traveling manipulator on the station.

83
4. Two reels colocated but with a single drive motor serving both

reels.

The weight penalties involved with the use of two separate reel

systems as suggested in items 1 and 4 and the difficulties associated

with transporting the system from one end of Space Station to the

other as noted in item 2 were the reasons for deciding on a single

centrally located reel system per item 3.

The Space Station configuration used for this study is the NASA,

reference configuration selected for the phase "B" Space Station

study effort. The future operational configuration (FOC) projected

by NASA would not materially affect the implementation of a tether

deployment system as described in this report so long as the gravity

gradient tower concept for the Space Station is retained.

Figure 4-1 shows the baseline Space Station with the addition of the

tether deployment system equipment. in this configuration, Shuttle

is shown berthed at the earth pointing or nadir end of the station

prior to release for downward deployment. The OTV with payload is

shown positioned on the top or zenith end of the Space Station prior

to release for deployment upward. Centrally located on the . Space

Station structure is the dual mode main reel system. Reaching from

this point to each end of the structure are separate leader tethers

feeding the two deployment systems (upper and lower). A tether quick

release connecting device to be discussed later is located in

proximity to the reel. As one system is in use, the leader tether

for the other will be disconnected.

Each end of the Space Station structure is equipped with a tension

alignment boom. This bcom translates fore and aft on the upper (or

lower) end of the Space Station structa re, and in plsne with the

orbit (These alignment booms are shown later in more detail in

Figures 4-4a and 4-4b for the lower system and in Figure 4-10 for the

upper). Kiding on the boom is a translating tension alignment

carriage which includes a berthing ring for the Payload Interface

Deployment Module (PIDM). As the payload is deployed, the

translating carriage will change position as required to maintain

alignment of the tether tension with the Space Station center of

mass. This allows the Space Station in-plane attitude to remain

vertical as the tether to the OTV or Shuttle payload deviates fre.a

vertical during the deployment process. The booms are designed to

accommodate up to + 7° (slightly in excess of the D.lr requirement)

of in-plane tether angle. Out of plane angular deviations for these

missions are minimal, therefore, no out of plane alignment provisions

are included.

84
7 •K y

4-10

01V Spacecraft

Payload Interface

I ..fir Deployment Module (PIDM)^1 ^ i

i ^- — leper tension Alignment Roum System

IT 

I ^^I

Flight Path I I ^IJ

NAD I P.

I

1 3262

Space Station C/G Location Band

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See Fig

Main Reel

yY LShuttle Interface 1leplovment Module (SIDM) -.

Lower len,fon Alignment Buom System

^St..*,tIe

ice. Station w, 'th Dual Mode ^'eth!?2 Deployment .System

85

k^

See Figure 4-4
1. heel drive assembly;

2. Keel with tether

3. Upper tether alignment assembly;

r. Lower tether alignment assembly.

These subsystems .+-.: brought up as preassembled unirs to be attached

to the structure by standardized fastening devices. It is

anticipated that such devices will he standardized throughout Space

Station for attachment of components. The reel module with tether

will be removed and changed out as a unit for tether replacement.

The installation of the interconnecting tether between the reel

assembly and the boom assemblies will be perfo nned by EVA.

Attached to the lower end of the Space Station structure is a

standard Shuttle berthing "aterface. This location has been selected

to position the Shuttle in an ideal location for the installation of

the SIDM. The SIDM could also be installed onto the Shuttle at other

berthing locations on the Space Station. This SIDM attachment

interface location is forward of the Shuttle center of mass. This is

to cause the Shuttle to hang on the tether in a nose up attitude so

as to avoid any danger of the tether fouling the Shuttle tail

structure during tether deployment and release operations.

Installation of the SIDM to the Shuttle will require EVA to mate the

latches and to mate the UMS propellant transfer lines.

The corresponding attachment of the OTV spacecraft to the upper PIDM

will. be somewhat different than for the installation of the SIDM onto

the Shuttle. The UTV mission stack to be deployed will be

transported from designated assembly and servicing areas to the upper

deployment area by the Space Station traveling manipulator. The

I manipulator will plane the mission stack onto the PIDM and the

retention latches closed by remote control. After the OTV mission

stack has been attached to the PIDM which in turn is held by the

alignment boom .,ystem, the manipulator can be detached and moved out

of position.

The PIDM will differ from the SIDM with respect to its spacecraft

interface configuration. Also, there is no PIDM requirement for an

OMS propellant scavenging system. The functional requirements for

on-board systems suc:i as the propulsion system required for retrieval

and communications will be similar.

86

r.
leader exchange mechanisms, tether control unit and guide pulleys.

Reel - The reel is sized to hold up to 81 nautical miles (150 km) of

Kevlar tether. The tether will be teflon jacketed to prevent long

term degradation from exposure to atomic oxygen. There is adequate

space for a larger capacity reel if needed.

Reel Motor/Generator - The reel motor sizing is dependant on poi

generated during the deployment phase -a nd the requirement to develop

adequate braking torque to halt the deployment at any stage.

Depending on the amount of time available for OTV deployment, it is

anticipated chat the power rating of the motorie.enerator drive will

be in the range of 75 to 150 horsepower.

Radiator - The radiator is a high temperature electrical resistor

bank which receives electrical power from the reel drive

motor/generator during deployment and converts it to thermal energy

that is then radiated to space as waste treat. This radiator can be

located away from the proximity of the reel system and can be placed

on the Space Station in a location selected flo ootimize its view

factor to space and to avoid illuminating sensitive areas with

thermal radiation.

Tether Splice Fitting - Figure 4-3 shows a method of connecting the

main ree tether to either the downward or upward tether leader

deployment assemblies. Each of the leader tethers and the main reel

tether are spliced to their re3pective connector fittings. The

mating fittings connect together by a hook and pin method. To mate

the two fittings, they are oriented at 90° to each other which allows

tl,e hook to fit over the pin;,. The spring plunger is then retracted

allowing the hook to fall in place over the pin. When the two

fittings are returned to an in line position they can not be

separated as the hook end will bear on the opposite fitting bearing

face. The assembly is designed to pass over the 12 inzh diameter

guide pulleys. This will occur during the first 200 feet of tether

deployment when both velocities and tensions are lcw. Guides will be

incorporated at each pulley to insure that the fitting will pass over

the pulley in a flat position.

Tether Control Unit - This unit will contain tether control equipment

fur measuring tether velocity, tension, iepluyed length and a device

to provide tension control at tl ► e reel during intervals of low

tension operation. Tension control will be required to insure smuuth

rendering of the tether onto the reel.

87
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pulleys, tether and reel assembly are omitted from drawings in this

report.

Figures 4-4a and 4-4b show a close up view of the Shuttle oertned to

the lower end of Space Station. This berthing location is ideally

located for installation of the SIDM on Shuttle from the lower tether

deployment assembly. The Shuttle interface station location shown

will always be located forward of the Shuttle center of mass

resulting in a Shuttle nose up hang angle attitude at all times. The

nose up attitude angle can be controlled by shifting the SIDM

location forward or aft. The SIDM attaches to the Shuttle sill rails

by means of the SIDM latches located on the aft end of the SIDM. The

forward end of the SIDM bears on the sill bearing pads with no

physical attachmer'_. With the Shuttle hanging on the tether in a

nose up attitude position the SIDM latches will always be in tension

and the bearing blocks in compression. When the latch pins are

pulled by the pyrotechnic pin pullers, the SIDM will be pulled away

by the tether tension. Installation of the SIDM onto the Shuttle as

shown here will require EVA.

The tension alignment boom translates on rollers attached to the

Space Station support structure. The tension alignment carriage in

turn translates on the boom tracks. Motor driven rack and pinion

drives will provide the force to translate the carriage and boom and

to maintain them in position. These boom and carriage drives will be

controlled by the Space Station attitude determination system to

control the in-plane attitude of Space Station. See Figure 4-5.

Figures 4-6a and 4-6b show an expanded view of the SIDM. The

spherical unit on top is shaped to fit the cylindrical berthing

interface on the carriage. The shape aids the final retrieval

docking of the SIDM after Shuttle release. Latches on the alignment

carriage berthing ring will secure the SIDM to the carriage for

berthing and stowage. The SIDM strong back structural section serves

to transfer tether tension loads to the release latch mechanisms at

the Shuttle sills resulting in vertical loads only at the sill

fittings.

Figures 4-6a and 4-6b also show the propellant storage tanks for the

OMS propellant scavenging system. No attempt has been made to show

the propellant transfer interface or disconnect fittings. The

propellant tanks shown are the same size tanks as those planned for

the Orbital Maneuvering Vehicle. The total capacity of the four

tanks is 6700 pounds which is an ideal sizing for this application.

91
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As the tether is deployed, ONS propellant (N 2 04 on the right side

and MMH on the left side of the Shuttle) will be transferred to the

appropriate receiving tanks on the SIDM strong back. At any time

during the deployment operation, only as mach propellant will be

transferred to the SID% scavenging tanks as has become surplus to

Shuttle deorbit requirements. In the ek , ent of tether system

malfunction or other abort situation, the SIDM could be rele9sed from

the Orbiter and adequate propellant would always be available

on-board the Shuttle to complete deorbit and reentry. Under norinal

conditions, 6500 lbs of propellant will be transferred and retrieved

to the Space Station for rise by other systems (such as the OM' 3 ).

The modification required to the Shuttle cargo bay wil'

'

consist of a

nitrogen tetroxide line from the OMS kit oxidizer transfer panel at

the aft cargo bulkhead up to the Sli)n iustallation station on the

right side of the cargo bay and a similar line for the monomethyl

hydrazine fuel on the left side of the cargo bay.' This plumbing will

be permanently installed in those Shuttles to be used in tether

deorbit operatiuna. The SIDM interface will ha-e quick disconnect

flex lines which would connect with to the y anks mounted OLI the SIDM

strong b slck beam. These same connecting lines will subsequently be

used to transfer the scavenged propellant ::to propellant storage

tanks at the Space Station.

The release latch mechanism design for the SIDM is shown in Figure

4-7. When the shuttle has reached the release point, it is essential

that release of the SIDM takes place for the sake of crew safety.

The cargo bay doors cannot be closed with the SIDM in place. The

release mechanism is designed to be three fault tolerant. The remote

control rel-are can be effected by either the Shuttle crew or the

Space Station crew. Each of them has control of a pair of pivoting

release latch fit-ings equipped with pyrotechnic pin pullers. The

SIDM side swing latch consists of an open face hook. If the Shuttle

latch pin is pulled, the Shuttle side link can rotate all.,wing the

latch to fall free of the SIDM latch hook. Likewise, if the SIDM

side pin is pulled, the SIDM latch hook rotates, falling free from

the Shuttle latch. See Figure 4-8.

It is intended that the two Shuttle swing latch pins will be the

primary release mode to be used and under control of the Shuttle

crew. This permits these Shuttle release latches to be refurbished

on the ground rather than the SIDM latches which would require refurb

at the Space Station. Release will be performed by firing the pins

on both sides of f.he Shuttle simultaneously. In the event only one

pin should fire, full release will still occur because the open face

hook on the opposite side will simply fall out after release of the

first: hook. If both should fail, the SIDM pins can be activated by

the Space Station crew. The same sequence of events applies to the

SIDM side as for the Shuttle side. The result is that four

pyrotechnic pin pullers are involved bur only any one of the four

need operate to effect release.

98
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position the fitting in an upright position while the SIDM ie being

installed.

The SIUM latch pivot point contains an eccentric shaft. The

eccentric shaft is manually activated by the "T" pin. See Figure

4-9. When the ecr^ntric shaft is rotated, the latch will pivot Lo

the release pr• .ition allowing inatallation of the SIDM to the Shuttle

sill fitt ; :Ags. After the SIDM is in position the "T" handle is

rotated to latch the open face hook into the Shuttle side latch.

This feature could be used as an additional back-up release method in

case the pin release system should fail. This emergency condition

would require that the tether first be severed by means of the SIDM

guillotine. After cutting the tether the SIDM would then be secured

prior to being jettisoned for later retrieval by an OMV mission from

Space Station.

The dynamic release reaction of the SIDM must be studied further to

insure that no possibility of collision with the Shuttle or cargo

exists. In addition, the reaction dynamics of the SIDM i ► only one

pin fires needs to be analyzed. More clearance may be required

around the Shuttle payload than is shown here.

Figure 4-10 shows the placement of the Orbital Transfer Vehicle (OTV)

on the upper tether deployment system. This assembly is the name

design concept as the lower tether deployment system except !.or

i.ncreased boom length. The additional length is required becau p e the

center of mass of Space Station is located closer to the lower end of

the Space Station. The added length permits the same range of

angular alignment for the tether angle during deployment.

The OTV will be placed on the Payload Deployment Module (PIDM) by the

Space Station traveling remote manipulator and latched in place. On

release of the latches, the OTV will separate with the aid of in-line

thrusters in the PICH. Thexk thrusters will control the separation

until tether tension reaches approximately four pounds at which time

the gravity gradient tension forces will be used to complete the

deploymer.c proceso.

An analysis was made of the loads in the Space Station truss

structure for the upper tether deployment system using a tension of

4000 pounds and with the tension alignment boom and carriage at

maximum excursijn (45 feet from center line). The resulting strain

induced in the upper end of the Space Station keel structure caused

21 inches of displacement from the unstressed :.onditicn. This is

over a length of approximately 300 feet from the Space Station

center-of-mass to the upper deployment system. This is a preliminary

estimate based on assumed characteristics of the Space Station

structure and will need reverificat:on. It has been deemed an

acce,itable effect for this study. In actuallity the maximum angular

displacement and the maximum tension levels will not occur at the

same rime and the condition analyzed is an extreme worst case.

103
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The deployment system is to be designed as a modular unit to be

mounted into the Shuttle cargo bay by means of the standard payload

trunnion fittings.

The module is to be compatible with a location either forward or aft

of the payload to be deployed.

The design goal is to minimize the mass and the length of the

assembly consistent with a reasonable range of performance capability

for the overall system.

The primary sizing consideration for the system performance is set by

' the tether reel capacity. A decision was made to use the reel size

from the Tethered Satellite System. This reel has an inner diameter

of 38.5 inches and length of 48 inches. The payload deployment

envelope of mass and tether length is to be determined by the amount

of appropriately selected tether which can be carried on this reel.

4.4.2 Payload Interface Deployment Module (PIDM)

The system shall incorporate a PIDM which interfaces with the payload

to be deployed by means of a grapple fixture mounted on the payload

in a suitable location to accept the interface loads during tether

deployment. The PIDM shall be capable of releasing from this grapple

fixture under full tension load, and without inducing significant tip

off angular rates in the payload. This release operation is to be

remotely commanded from the Shuttle. The payload is to be installed

onto the PIDM by the Shuttle RMS prior to the extension of the

extendable boom. This payload installation operation is to be

capable of being performed without requiring EVA.

The PIDM shall incorporate a cold gas attitude control and propulsion

subsystem to provide the initial separation of the payload from the

orbiter until the gravity gradient tension force builds up to

rdequate levels to complete the deployment. The system shall also

provide location and attitude control capability for the PIDM during

the final stage of retrieval operations.

The PIDM shall incorporate a tether guillotine or equivalent

disconnect method to separate the tether from the PIDM.

The PIDM shall have a standard grapple fixture accessible to the

Shuttle RMS for any required handling operations on board the Shuttle.

The PIDM shall incorporate a berthing/structural interface with a

mating support structure on the end of the extendable deployment boom.

The PIDM shall incorporate a retro reflector to provide tracking

capability by the Shuttle Ku Band tracking radar system.

I

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105

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propulsion/attitude control system.

4.4.3 Extendable Deployment Boom

The system shall incorporate an extendable deployment boom of

ade q uate length to permit the initial deployment and the final

retrieval operations to be performed away from the immediate vicinity

Of the cargo bay. The boom shall provide a structural interface to

support the PIDM during Shuttle launch, re-entry and landing

operations. It is also to be compatible with retrieval and berthing

1 operations for the PIDM. It shall provide a retention latching

capability for the PIDM which is remotely controlled by the Shuttle

crew. The boom shall incorporate a tether guillotine near the outer

1 most portion of the boom.

The boom shall be capable of reacting bending moments on the boom

caused by the tether tension.

The boom shall incorporate a high temperature resistive load bank

located near the outboard end to permit the deployment energy to be

radiated to space as waste heat.

4.4.4 Reel Drive System te

Control of the tether reel drive is to be by means of an electrical

motor/generator drive which can either provide the tether braking

forces during payload deployment or provide the reel-in tension for

retrieval of the PIDM subsequent to release of the payload.

The reel shall incorporate a friction brake lock to hold the reel

against tension forces once full tether deployment has been achieved.

The reel drive system is to be controlled by modulating the system in

accordance with specifies tension control laws designed to optimize

the deployment and retrieval operations.

The reel drive system shall incorporate level wind mechanisms and

tension control systems to maintain adequate tension levels the

reel at all times.

The drive system shall have a provision for a modular energy storage

unit which can be sized to provide the retrieval energy required by

the particular mission requirements. This ;Azing will include

consideration of whether or not a contingency capability is to be

provided to retrieve the payload in case of pre-release malfunction.

106
payloads is adapted from the Tethered Satellite System (TSS) hardware

currently under development. The major modifications are the

elimination of the spacelab pallet as the structural interface with

the Shuttle, the elimination of subsystems modules for support of

scientific payloads, increased power requirements for the reel drive

motor, and increased bending moment on the extendable deployment boom.

A decision was made to keep the .ether reel size the same as TSS and

to retain the same approach for the tether tension control, level

wind mechanisms, and for the extendable deployment boom.

4.5.1 Payload Deployment Constraints

The capability to deploy payloads is a function of the payload mass,

the length of the deployment tether and the capacity of the tether

storage reel.

Using the TSS reel and a requirement that the tension level be less

than 50% of ultimate strength for the Kevlar tether, the relationship

between payload mass and deployment distance was calculated. This

relationship is shown in Figure 4-11. This curve treats the

relationship as though any desired tether diameter were available.

In the actual case the curve must be further derated in accordance

w_ r h the size step availability of tethers. The curve should be

considered as an upper limit on the performance capability of the

system to deploy payloads. As an example, using a maximum Shuttle

cargo weight of 65,000 pounds and subtracting an estimated 5,000

pounds for the deplover system hardware, the curve indicates that the

system could deploy a 60,000 pound payload spacecraft to a tether

length of 62 km. Less massive payloads can be deployed to

correspondingly longer tether distances.

4.5.2 Deployer Design Concept

The adapted version of the TSS hardware is shown in Figures 4-12a and

4-12b. The major apparent difference is the elimination of the

r Spacelab pallet and `he provision of a direct structural interface to

the Shuttle. This rrangement reduces the length required in the

cargo bay by 18 inches.

The deployment boom is shown positioned centrally within the tubular

support structure and with the PIDM latched in position directly

above the boom stowage canister.

The reel and drive assembly, subsystem modules and the auxiliary

battery module are shown attached to the tubular truss structure.

107
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PAYLOAD MASS (klbs)

Figure 4-11 Deployed PayZoad .Mass Versus Tether Length

1

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The most significant differences are the elimination of the Spacelab

pallet and the deletion of the science support subsystems.

4.5.3 Operations Sequence

Figures 4-13a, 4-13b, and 4-13c show a typical sequence for payload

deployment using the Advanced X-Ray Astronomy Facility (AXAF) as a

-epresentative example. The tether deployment system may be located

as shown in Figure 4-13a in the aft region of the cargo bay or,

alternatively in the forward end. The Shuttle Remote Manipulator

System (RMS) will be used to position the AXAF onto the PIDM while

the release mechanism is latched onto the AXAF grapple fixture. This

is shown in Figure 4-13b. Once the payload spacecraft is mounted on

the PIDM, the boom is extended. As it extends the guy wire cable is

pulled out of the guy wire reels mounted on the cargo bay sills.

These reels are attached to the sills such that the tether tension

during deployment will be reacted through the guy wire cables rather

than through a bending moment on the extended boom. The guy cable

reels are located so as tc align the tension force with the Shuttle

center-of-mass. The relationship among these forces and the

center-of-mass location will determine the attitude angle of the

Shuttle during deployment. This attitude angle of the Shuttle is

controlled to keep the zetner well away from possible contact with

Shuttle surfaces such as the tail. Any libration of the tether away

from the vertical will be compensated by a change in the aspect angle

of the Shuttle to keep the tension force directed along the guy

cables. The guy cables will be pulled out of the stowage reels as

the boom extends and lock into position to react the tension force as

the boom reaches full extension. When the boom is retracted the guy

cables will rewind onto the stowage reel by a negator spring drive.

Figure 4-13c shows the location of the high temperature resistive

load bank to be used to radiate excess energy generated by the reel

braking motor during deployment of the payload. The load bank is

located on the boom to provide a good view factor to space and to

minimize radiant heating of cargo bay surfaces.

After the payload has been released, the tether is retrieved to where

the PIDM is in the Shuttle vicinty. The PIDM cold gas propulsion

attitude control system will be used during the final stages of

capture and berthing. Once the PIDM is relatched into the berthing

ring the boom is retracted to its stowed position and the system

secured for re-entry.

112
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115

Figure 4-I3c Shuttle Payload Deployment
Phase II baseline concepts are described in the following sections.

These issues are intended to provide focus for follow-on concept

definition efforts.

5. 1 Concepts A2 nd F

The similar nature of these application concepts means they share

most of the same technical issues. Where an issue is unique to one

only it will be noted.

Tether Recoil Dynamics - The issue here is can the tether be designed

to incorporate suffi-ient internal damping to result in acceptable

recoil characteristics in event of a severed tether while at maximum

tension.

Tether Erosion/Abrasion Susceptibility - can a tether be designed to

provide an adequate number of reuses.

Debris Collision - What is the probability of the tether being

severed or damaged.

Deployment/Retrieval Energy Management - Can methods be identified to

either use and/or store the energy generated during deployment and to

supply the energy needed for retrieval.

Induced Acceleration - Can the Space Station operations accept the

levels of acceleration induced by tether operations. This includes

structural effects on solar arrays, radiators and antennae, and

materials processing operations.

Orbital Perturbations - Can Space Station operations be modified to

accept the orbit perturbations required by the angular momentum

balance concept.

Structural Stress - Can the Space Station withstand the internal

stress due to tensioned tether runs.

Tether Attachment Interface - (peculiar to F) What is an acceptable

mode to impart tether tension loads into the AXAF mission stack.

5.2 Concept E2

Ionospheric Conductivity - Will the ionospheric conductivity behave

as assumed in studies.

Plasma Contactors - Will the plasma contactors function per concept

assumptions.

116
Space Station Potentials - Can design solutions be found to safely

accommodate the tether induced potentials on the Space Station.

5.3 Concept B

Tether Recoil - Can tether recoil be safely managed in event of a

broken or severed tether.

Induced Acceleration Levels on Payload - Can the candidate payloads

accept the induced acceleration levels and tether tension loads.

Energy Management - Can the energy management for deployment and

retrieval be accommodated on board shuttle.

I

5.4 Concept C

Tether Service Life - Can the tether be designed to provide adequate

service life under continuous usage.

Contingency Recovery - Can a platform be stabilized for subsequent

recovery in event of a severed tether.

Induced Acceleration - Are the induced acceleration levels compatible

with typical platform mission operations.

Platform Stability - Can adequate platform stability be achieved and

maintained against anticipated perturbations.

Power Supply - How best can power and other utility services be

supplied to the tethered platform.

5.5 Concept D

Rendezvous - is is feasible to accomplish a rendezvous with the

available levels of trajectory control and tracking.

Capture - What type of equipment will be required to effect

successful capture in the short time available.

Docking - Once captured how are the residual velocities and angular

rates damped out prior co hard docking the two vehicles.

Tethered UMV - Are the maneuver dynamics of the UMV controllable

while suspended on the tether.

117

^^L A.M.
the Phase lI baseline concepts are given in the following sections:

Concepts A2 and F

Tether Construction - To achie'^, e recoil damping, erosion and abrasion

resistance, continuity sensing (for broken tether warning), and

extended service life.

Tension Alignment - Tension alignment assemblies compatible with

Space Station.

Tether Reel Design - Two stage reels designed to achieve maximum

tension at different lengths. Level wind tranaition mechanisms and

tension control devices.

Reel Drives - High power (75-150 horsepower) motor/generator high

torque motors. Qualified for space operation.

Energy Management Technology - Inch ► des high density energy storage

methods and high tem —!rature radiators for rejecting energy to space.

Propellant Transfer/Management - Gauging transfer and remote control

disconnects for OMS bipropellant.

PIDM/SIDM Design - Design of the integrated systems required for the

tether end effector deployment modules.

Release Mechanisms - Remotely controlled mechanisms capable of low

tip off rate releaae.

Concept E2

Tether Construction - The tether must be designed to meet

requirements in the following areas.

- Flexibility

- Nigh conduct 4 vity to mass ratio

- Abrasion/erosion resistance

- Insulation against high voltage breakdown

- Thermal control properties

- Repairability on orbit

Plasma Contactors - The design of t y.^ plasma contactors and the

required support systems - particularly for the ouL board end of the

tether.

6.1

6.2

Tether Reel - Insulated reel design and high voltage termination of

the tether including the high voltage circuitry to the power

converter/processor circuit.

118
Simulations - Simulation codes for massive tethers with

electrodynamic dr forces.

6.3 Concept B

Deployer Systems - Design for minimum weight, tether capacity, high

tension drives.

Tether Construction - Design for recoil damping abrasion/erosion

resistance, a-,d continuity sensing.

Energy Management - Design to reject energy generated during

deployment and su 1 supply retrieval energy. 8Y• Includes high tem eratu-e_ K, P

radiators for waste heat rejection, and energy exchange interfaces

with the Shuttle power system.

PIDM - Design of required subsystems including releasa mechanisms,

cold gas propulsion/attitude control system, tether

guillotine/discor.nect, and communications, control, tracking systems.

6.4 Concept C

Tether Construction - Design for extended exposure durability .tnd to

incorporate utilities services such as fluids and power.

Reel System b Drive - Design for ease of deployment, retrieval and

yo-yo stabilization maneuvers.

Berthing/ Servicing - Design for ease of access to perform periodic

maintenance and servicing of platform while retrieval to the Space

Station. 

6.5 Concept D

Simulation Programs - For trajectory prediction and for Tethered OMV

maneuver dynamics.

Capture/Docking Hardware - Develop performance specifications and

design concepts.

119
1

r

Phase II study:

A. The application of tether technology has the potential to

significantly increase the overall performance efficiency and

capability of the integrated space operations and transportation

systems thru the decade of the 90's.

B. The primary concepts for which significant economic benefits

have been identified are dependent on the use of Space Station

as a storage device for angular momentum and as an operating

base for the tether systems.

C. These Space Station based concepts must be coupled into

operational pairs that are functionally related such that one

concept uses the angular momentum derived from the other.

D. The outstanding candidate concept for the source of angular

momentum is the tethered deorbit of Shuttle from Space Station.

An ancilliary benefit is the significant quantities of

bipropellant scavenged from the Shuttle during the tether

deorbit operation.

E. Two alternative candidates have been identified as leading

contenders for the role of angular momentum consumers. First is

a tether launch assist to an OTV mission which uses the angular

momentum to reduce propellant requirements for OTV. Second is

an electrodynamic power tether which gradually converts the

scavenged angular momentum into electrical power for use on the

Space Station.

F. Concurrent usage of the OTV launch assist and the electrodynamic

power tether concepts are not feasible. Electrod y namic power

tether usage could start with Space Station. Tether launch

assist to OTV cannot commence until the advent of the spare

based OTV circa 1995.

G. Use of these functionally related concepts require that

provision for them be incorporated into the Space Station design

concept.

H. Certain operational impacts on the Space Station are inherent

with these tether applications. The primary ones are the

acceleration levels induced by tether deployments, and the orbit

perturbations caused by tether mediated angular momentum

transfers.

120

Affie"T .
be used, stored or rejected to space. Retrieval requires

smaller but still significant quantities of energy to be

supplied to the tether retrieval system.

J. Any overall strategy for tether usage must consider the

compatibility aspects for those concepts using the Space Station

as an operations base.

K. Economic benefits to be derived from the use of tethered

platforms from Space Statior has not been analyzed during this

study phase. However, it should be noted that the ust of such

platforms would not be compatible with those concepts for which

significant benefits have been identified.

t

L. Tethered OMV rendezvous concept for retrieval of returning OTV

is not economically viable.

M. While tether insertion of AXAF from Shuttle does not present a

unique performance capability and is not economically

advantageous. Other missions do exist where the orbit insertion

capability is unique to the tether method.

11

121
y gpy 

sets.

8.1 Angular Momentum Balance Concepts (A2 with F, A2 with E2)

These recommendations relate to the following concepts:

A2 - Tether Deorbit of Shuttle from Space Station

E2 - Electrodynamic Power Tether for Space Station

F - Tether Launch Assist to OTV Mission from Space Station

A. Continue to pursue the definition of these concepts as top

priority and including:

Space Station integration requirements definition

Trade studies of relative merits of pairing A2 with E2, A2

with F, and transition from one to the other.

B. Perform an assessment of incorporating capability to transfer

OHS propellant into the Shuttle.

C. Brief Space Station and OTV program staffs on potential tether

application benefits.

D. Initiate technology development activities in key areas to

resolve technical issues identified.

8.2 Tether Deolovment from Shuttle (Concept B)

Although the Concepts B and B1 involving tether insertion of AXAF did

not identify any outstanding benefits, it should be remembered that

there are areas of performance where the tether deployment provides

unique capabilities. The AXAF operational orbit is barely within the

Shuttle direct injection capabilities limit. For missions inside

this Shuttle performance envelop the tether deployment is not

justifiable, however, for missions requiring higher orbits the tether

provides a significantly enlarged performance envelop.

Recommendations are:

E. Develop parametric performance envelop for an optimized Shuttle

tether deployer system to identify areas of unique performance

capability.

F. Continue concept definition studies to define an optimized

performance design for the deployer system. (Optimize mass,

tension limits, tether length/volume, energy management).

122
G. Develop methods to provide utilities support (power, fluids) to

the platforms via the tether.

H. Define retrieval/servicing operations

8.4 Tethered Rendezvous (Concept D)

This study has shown there is no economic benefit justification for

further study effort on this concept. Recommendation:

I. Discontinue study effort on this concept

f

8.5 Other Concepts

During the course of the study, other interesting areas have been

identified which could be of significant benefit. Preliminary

studies are recommended t( . explore the merits of the following

application areas:

J. Use of a tether deployed mass to provide a collision avoidance

maneuvering capability to Space Station. This would be an

ancillary benefit from Concept A2 for the tether deorbit of

Shuttle.

K. Tether dep'oyment of Space Station waste packages for disposal

by burn-up re-entry.

L. Utilization of a tether from Space Station to perform a residual 

propellant scavenging from an external tank with subsequent

tether deorbit of the tank.

M. Tether launch assist to expendable launch vehicles from both

Shuttle and Space Station.

r

123
APPENDIX A

EQUATIONS AND REFERENCES

USED FOR DETERMINATION

OF PROPELLANT REQUIREMENTS

I
A. SHUTTLE DIRECT INSERTiON MISSIONS

Refs: (1) Performance Estimation Technique for Space Shuttle Direct

Orbit Insertion Missions - ETR Missions, JSC, Oct. 1982.

(2) JSC, Approximate reentryOV requirement from circular orbits

(via Mac Croft of MSFC), Sept. 1984.

(3) Shuttle Systems Weight and Performance - Monthly Status

Report, JSC, 18 May 1982 (for typical OV103 orbiter weight

data and external tank weight).

1. Obtain,&V l (OMS insertionOV at apogee, summarized in Figure 3-3).

r Assumption No. 1: h perigee (typical) = 27 NMI (from Ref. 1)

a) Circular orbits

AV 1 = V CA - VA

V

1 = -j7i^-

(1)

,f'A R(ft),AV^FPS)

genera ly

where: VCA = circular orbital velocity at apogee

VA = velocity at apogee

A = GM = 1.4076452 x 10 16 ft3/sec2

Rp = radius of perigee

R A = radius of apogee

R E = radius of earth = 3444 NMI

R = R E + h(altitude)

6076.1 ft/NMI

b) Elliptical orbits

Z^ V l = VA - VA

Q V1 =j 2Rp2 J - 2Rp I jU

R A + RP2 RA + R P , (2)

RA

where: VA I = velocity of apogee of direct insertion ellipse

VA 2 = velocity of apogee of final elliptical orbit

Rp l = radius of initial Perigee

Rp 2 = radius of final perigee

RA = radius of common apogee

and VA = 2Rp /U ( 1 -e)μ

N R A RA +R P RA

A -1
V2C

( FPS)

276

.72h+183 (3)

1.30h+84

1.46h+48

A. SHUTTLE DIRECT INSERTION MISSIONS (Cont)

2. ObtainAV 2 (OMS total deorbitL V, summarized in Figure 3-4).

a) From circular orbits ^V2C)

Assumption No. 2: -- a- hcircular

(from Ref. 2) (NMI)

h (130

130(h '110

170 ^^ 230

h) 230

where: h is in NMI

andAV2C = deorbitQV from circular orbit (FPS)

b) From elliptical orbits via 100 NMI circular (elliptical

100 NMI perigee)

orbits with

AV 2EC = V P100 VC 100 + 276 (from a)

QV2 EC = 2RA AL

JRA + Rp + 276 (4)

RP

where: =

total deorbitAV(FPS) (2 burns) from apogee ,LV 2EC of elliptical

orbit to 100 NMI circular to deorbit

V

P100 = Perigee velocity (100 NMI)

V C100 = circular orbit velocity (100 NMI)

R P

= radius of perigee (100 NMI)

and Vp =

FR

R ^l =(1+e)^l.

RA+R P RP

c) from apogee of elliptical orbit (direct)

0V 2E = OV 2C - (V CA VA)

(j— R

AV 2E = OV 2C μ— A- (5)

RA

where: OV2E = direct deorbit,6 V from apogee of elliptical orbit

OV2C = deorbitOV from circular orbit at apogee altitude (from a)

VCA = circular orbital velocity at apogee

VA = velocity at apogee before deorbit burn

A-2
A. SHUTTLE DIRECT INSERTION MISSIONS (Cont)

3. Obtain (AWp) propellant required for ascent and deorbit (summarized

in Figure 3-5).

a) Obtain initial orbiter weight at initiation of first OMS burn (WO)

Use Refs. (1) and (3) from Page 1 and obtain W O a= described in Table 

b) CalculateQWp

,&Wp = 1.075WO [1 - e

--by/9IsP^

(6)

where: 1.075 = factor for OMS propellantmartin

t WO = orbiter weight at start of first OMS burn

= 2.183. . . (natural log, base)

,AV =T.AV =AV1 +0V2 (appropriateAV1 &AV 2 from A.1

and A.2)

g = reference g = 32.174FPS

I S p = 313.2 sec.

4. Compare alternatives to obtain,&Wp savings (Figure 3-5)

B. OMV DELIVERY MISSIONS

Assumptions No. 3: Use current (Nov. 84) approximate characteristics of

Martin Marietta Aerospace storable bi-prop reusable

OMV as follows (per Table 3-8):

WBO (burnout wt) = 6525 lb (1)

WpU (usable propellant) = 8075 lb (maximum)

W1 (initial stage wt) = 14600 lb (maximum)

Propellant = N204/MMH

I S p = 306 sec. (vacuum)

(1) Includes 5% allowance for RCS plus propellant

margins (held constant for all propellant loadings 

first order estimate)

Assumption No. 4: Static tether releases above the space station with

tether length limited to 150 km maximum.

Assumption No. 5: Tether launch OMV to apogee of desired orbit.

1. Procedure

o Off-load propellant as required for mission

o Investigate untethered and tethered launches

o All launches from Space Station at 270 NMI

o OMV delivers payload to orbit and returns empty to Space Station

o All launches in-plane

A-3
B. OMV DELIVERY MISSIONS (Cont)

1. Procedure (Cont)

a) ObtainjAV for ascent (&V I ), untethered for selected h A values

0 V 1 = AVp +AVA

OVI = (V P - VCp) + (VCA - VA)

(^ "Pμ

OV1 

= JR—ARAP - RA+RP

Rp FRA

RAA+Rp

- 1 1 -

RAPRp

orAvi+ (7)

Rp A

b) Obtain£,,^y for descent ^Vh), untethered for previously selected

h A values

AV2 AV1 (from a)

Note: R P = Space Station orbit radius (3714 NMI)

RA = Final orbit radius

c) Select incremental altitude of OMV/payload above mass center ^U),

tethered case (mass center at 270 NMI).

AhU = 0 for Ltet

whereAh U = Ws

WU + WS

where: WS =

W U =

L =

her = L =0

L (8)

Space Station Mass (551 KLB)

OMV/pay'.oad mass (or upper mass)

Tether length

Note: h S = h U - L

VaryOhU from 0 to maximum value required for problem, considering

payload and tether length limitations.

d) Obtain data required to calculate0V 1 &AV2 , tethered cases.

Rp = RS +AhU 0 (9)

VP(FPS) = 6076.1 Rp(NMI)e270 (10)

R 

her 9270 = .00110677 rad/sec

P P R A = 2μ- RpVp (11)

A-4
B. OMV DELIVERY MIk1SIONS (CONT)

1. Procedure (Cont)

hA = R A - 3+44 (NMI) (12)

a = (R A + Rp)/2 (13)

VA A

(14)

aVp

VCA = (15)

R n

e) ObtainAV, &AV 2 , tethered cases as a function ofAh0

AV1 = V CA - V A

(16)

from eqns. 14 and 15

OV 2 = QV P + ^VA

7777 - 2Rp

QV2 RA

+RP 

I

I _ RA+Rp —

R

RA

from eqn. 7

where: R P

 = Space Station nominal radius (3714 NMI)

Assumption No. 6: Small changes in Space Station orbit do not

significantly affectOW p calculation since

QV2 is for returning stage without payload

(first order approximation).

f) Find W PO (usable propellant) required, untethered case. iyU = 0,

summarized in Figure 3-6).

Wp L = Wpb + WBO 0 - MIM2)

M1 _ 1

(for delivering payload and returning empty stage)

or for selected payloads (WpL)

W PU(required) = W PL (M 1 -1) + WBO(M1M2-1) (17)

where: ML =eaVE/gISP (18)

QV I =AV2 from a

I S p = 306 sec.

g = 32.174 ft/sect

W

BO = 6525 lb.

A-5

y w^ ^• v 
B. OMV DELIVERY MISSIONS (CONT)

1. Procedure (Cont)

g) Find W PU b L required, tethered cases, as a function ofAhU and

W PL (summarized in Figure 3-6).

UsinV j and AV 2 as a function of©h U (determined in e).

W^ U (required) = W PL (M1-1) + WB0(M1M2-1)

from eqn. 17 for selected payloads.

U

L =^►

hU/fW^ from eqn. 8

` S

where WU = WPL + WPU (req'd) + WBO

(19)

2. Compare tethered cases to untethered cases to obtainAWp savings for

OMV (summarized in Figure 3-7).

a) Plot h orbit vs WPUrequired

for all cases at given payload values

b) Plot horbit vs L (tether)

for tethere H cases at given payload values

c) PlotAWp savings (tethered vs untethered) at given W PL values vs

final orbit altitude using Plot a) data. Plot constant I. values

using Plot b) data.

C. OTV DELIVERY MISSIONS

Assumption No. 7: Use estimated characteristics of Martin Marietta Aerospace Aft Cargo Carrier (ACC) OTV as follows:

Aero braked reusable OTV W BO (burnout wt) = 6434 lb. (1)

Wp U (usable propellant) = 53577 lb. (max.)

W1 (initial stage wt) = 60011 lb (max.)

Propellant = L02/LH2

I S p = 460 sec (vacuum)

(1) Includes allowance for RCS and propellant margins (held constant for

all propellant loadings - first order estimate).

Assume - Static tether releases above Space Station (Ref. Assumption 4)

Assume - Tether length limited to 150 km (Ref. Assumption 4)

Assumption No. 8: All plane changes at apogee (first order estimate)

(28.5 deg)

A-6
C. OTV DELIVERY MISSIONS (CONY)

Assumption Ne). 9: /\l req.irements for geosynchronous mission from

t.'0 NMI (untethered)

AVI =,&Vp +AVA + 350 (FPS) (losses and mid-course

maneuvers)

GVI = 14091 FPS

AV2 =AVA2 +AVA3 (350 FPS) (after aerobraking) +

50 (midcourse maneuvers)

AV2 = 6445 FPS (geo to 50 NMI and up to •-..2^O NMI circ)

1. Procedure

o Off-load propellant as -equired

where:

V = VAi '2 + VCA2 - 2VA1,2VCACOS28.5 Equation (20) AI 2

o Investigate untethered/tethered launches

o OTV delivers payload to geo and returns empty stage

o All launches from Space Station (270 NMI)

a) Obtain W PU (usable propellant) required, untethered cases ^&U=O),

(summarized in Figure 3-8)

M ` = AVO Isp from eqn. 18

where: VI = 14091 FPS

$V2 = 6445 FPS

for selected payloads (WPL)

and WPU WPL(MI-1) + WBO (MIM 2 -1) fro m 

eqn, 17.

b) Obtain data required to calculateQVI , EAV2, tethered cases, for

selected,&hU (as in OMV procedure)

(constant) RA1,2 = 22816 NMI (geosynchronous orbit radius)

(constant) Rp2 = 3494 NMI (50 NMI aerobrake altitude)

' RpI = 3714 NMI (mass center) + OhU (selected) (from 9)

VpT (ve l ocity of tether tip) = 6076.1 Rpl(NM^70(rad/s

(from 10)

whereG270 = .00110671 rad/sec

Vp l = 21AR A1

Rp l Rp l +R AI ) (21)

0 Vp I = Vp l - V DT (22)

A-7
^. C. OTV DELIVERY MISSIONS (CON?

VA = 2μ_Rpl

1 RA 1 ^ pR l + R 

Al ) from (2)

i

2

OVA1 - V^, 1 + VCA2 - 2VAIVCA cos28.5 from (20)

where VCA = 10077FPS (Geo)

c) ObtainOV I &AV2, tethered cases as a function of hil.

OV I =0V p 1 fQVA I + 350 (23)

^Vp1 R,LV A1 from b)

G V 2 = 6445 FPS (same as a)

Note: Approximation fo &V 2 can change for longer tether 'engths due

to larger changes in SS orbit (beyond scope here).

d) Find Wp U and L required,

(summarized in Figure

ML = e0y9ISP

ISp =

9

W BO =

tethered cases, as a function ofAh U and WPL'

3-8).

from eqn. 17 (P-10)

460 sec.

32.174 FPS

6434 lb

WPU (req'd) - 

Wp L (Ml -1) + WBO(MIM2-1)

from eqn. 17 for selected payloads

W

L =AhU/ 

WU+WS

from eqn. 8

where Wjj = W PL + 

WPU (req'd) + WBO

from eqn. 19

2. Compare tethered cases to untethe r ed cases for various payloads vs

L to find AWp savings (summarized in Fiqure 3-9)

A-8