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EL_CTi{OL_AGNZ_IC LAU_ChEi_5 Final _echnical

a_poi:t (tJattell¢ ColUmbUs.Lais,, Ohio.)

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FINAL TECHNICAL_REPORT

on

PRELIMINARY ANALYSIS

OF SPACE MISSION APPLICATIONS

FOR ELECTROMAGNETIC LAUNCHERS

tO

NATIONAL AERONAUTICS. AND SPACE ADMINISTRATION.

LEWIS RESEARCH CENTER.

(Contract Number NAS 3-23354)

August 30,. 1984

by

Lisa A. Miller

Eric E. Rice

Richard W. Earhart

Robert J. Conlon

BATTELLE

Columbus Laboratories

505. King Avenue

Coiumbu$, Ohio 43201
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FOP, EWOSO

ThiS Battelle C01umbus LabOratories study was sponsored by

NASA Lewis ReSearch Center unde_ NASA COntraCt Number NAS3-Z3354, for

the purpose Of providing a preliminary anaiysis of space mission

applications for electromagnetic launchers. Work was conducted from.

NOvember 1982 thrOugh June 1984, Battelle's assessment Involved:

(1) deveiopment of mtsston models, and requirements for electromagnetic

launchers; (2) reference system concept development; (3) preliminary

economic assessment; (4) preitminary environmental assessment; and

(S) assessment of technoiog¥ developmen_ needs, To support the system

COnCept development, NASA Lewis ResearCh Cen_er supported Dr. Henry H.

Kolm and Or.. Peter P. MongeaU, EML Research, Inc., Cambridge, I_, and

Dr. O. K. Hawardi, Collaborative Pianners,.Znc., Cleveiand Heights, OH.

InfOrmation developed during the study perlod is contained

in this.final, report. Inquiries regardln9 th_s study should be addreSSed

toL_

Hr. William R. Kerslake, COR

NASA Lewis Research Center

Mail Stop 50i-7

Space Propulsion Technology

DiviSion

Cleveland_ OH 44135

TelephOne: (216) 433-4000 X51831

FTS: 294-5183

M). Lisa A. Miller, Project Hanager

Space Systems Section

Defense and Space Systems Department

Battelle Columbus Laboratories

505 Ktng Avenue

Columbus, OH 43201.

Telephone:, (614) 424-4!46

FTS: 976-4146

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,_,ATYE*.t.E -- ¢OLUMm_JI
ACKH rs

The principal authors, LiSa A. Hiller, Eric E. Rice,. Richard •

W, Earhart, and Robert J. Conion, acknowledge the assistance of the

following Battelle staff whO. provided contributions to the technical

content of this _eport:. ..

Corinne M. Buoni

Donald J, Butz

Thomas H. Crabb

Harry H. Collts

Norman H. Fischer

Mtchaei N. Golov_n

Richard. G. Oung

Nell E. Miller

Richard F. Porter

Robert C. Reynolds

Samuel H, Smith

Albert E. Weller

Acknowledgement is aiso made for the NASA-funded COntributions

in the aeea of coaxial accelerator concepts provided by: Henry H. Kolm

and Peter P. HOngeau, EML Research, Inc., Cambridge, MA, and O. K. Mawardi,

Collaborative Planners, INC.., Cleveland Hetghts, OH, Special appreciation

is given to. William R. Kerslake, NASA/LeRC study monitor, Lynnette M.

Zana, NASA/LeRC, and Fred F. Terdan, NASA/LeRC, manager of the Advanced

Propulsion Research Section, for their interest and guidance throughout

the study,

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TABLEOFCONTE S

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

ACKNOWLEDGEMENTS t l

• • • • • • • • • • I • .- • • • l • • • • • • ., •

1• 0 EXECUTIVE SUMMARY • • • • • • • • • • • • • ., o, • • • • • • • • 1_' 1

1.1 Study Background . , .......... . ........ 1-1

1.2 Study Objectives ; ............ . ......... 1-2

1.3 Approach , ...... , ........ . ...... ;., . 1-2

1 k 4 Study Guidelines e)l_ltl'ttil¢l,il#lltl I 2

i.5 ReferenCe Concepts .......... 1-4 1.s.1 Eart.-to-Orbi ail'Lau c e ::i. ...... 1-4

1.5.2 Earth,to-Orb_t Coaxial Accelerator ......... 1-4

1.5.3 Hybrid Railgun/Rocket ............. , . . . 1-6

1.5.4 Hybrid Coaxial ACCelerator/Rocket ......... 1-6

1.6 Summary of Major Results .... , ............ 1-g

1,6,1 Alternative EML Concepts . : . : ........... i-9

1,6.2 Mission Models and Requirements ......... 1-9

1.6.3 Safety and Environmental Impact ......... ,. i-ll

1.6.4 Cost Estimates ................ 1-14

1.6.5 Technology Requ;rem;nts: .................. 1-i4

1.7 Conclusions ................................ 1-14

1 ., 8 Recommendations ..i•............,...i.i 1-17 .

2.0 MISSION MODEL OEVELOPMENT ..................... 2-1

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Z.1 Earth Orbital Mission• , ...................... 2-1

2.1.1 OTV Propellant Supply .................... 2-3

2.1.1.1 Mission MOdel s ....... 2-3

2.1,i.2 implications of'STS'Scavenging . . . . . 2,5

2.1.Z Space Station _upply ..................... 2-6

2.I.2,1 Life Support ............... 2-6

2.1.2.2 Spares . . , ....... 2-7

Makeup Propeilan1:s. : . ,-. 2-7

2.1.2.4 Fuel Cell Makeup Propellants ......... 2-8

2.1.2.5 Miscellaneous Supply Items ....... 2-8

2,1.3 Materials for Space Processing ........... 2-8

2.1.4 Traffic Models .................. 2-9

2.2 Lunar Base Supply Mission ............ 2-9

2.2.1 Lunar Base Characterizationl ........... 2-12

2.2.1. I Life Support Systems ........... 2-13

2.2.1.2 Power Systems ........... 2-14

2.2.1.3 Lunar Transportation .......... 2-14

2.2.1.4 Earth-to-Moon Transportation ........ 2-15

2.2.2 Lunar Base Supp!y Scenarios ......... 2-16

2.2.3 Lunar Base Supply MiS_.I(}..Model:..: ........ 2-18

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2.2.3.1 Crew Supply Requirements Durirtg

Transpor_t, . , .........

2.2,3.2 Lunar Life Support Consumabies .....

2.2.3.3 Equil_oent ................. 2-21

2.2.3,4 HatErials. :.... ,.. , ........... 2-21

2_2,3,5 Pr'opeilants .......... , . , . 2-21

2.2.4 T r,afftCModel, , , . , , ..... '''"'" i i 2-22

2.2,5 L.nar Base supply Advantages and Disadvantages . 2-24

2,2.6 LUnar Base Supply Issues ..... _ ........ 2-24

2.3 SOlar System Escape Mlsston,, ............ , . . 2-25

2.3.1 ESRL HtsstOn A Summary ,.... , ............ 2-25.

2.3.2 TransuPantC Waste, . , .... , .... , 2-26

2._,2.1 Reprocessing and Refabrtcatt_n _astes[ . 2-26

2,3.2.2 TRU Waste Available CoP Space Disposal , 2-2g

2.3,3 Total Htsston Summary .... , ........ , .... 2-29

2,4 Earth EsCape MtSStOn , ............... 2-32

2.5 SubOrb.ital Hission _del : :. : :, :.. .......... 2-33

2_5.1 NASA Soundlng Rocket Program ........ , .... 2-33

2.5,2 Hiss ion Summary ................... 2-33

2 . 6 Eiect_omagnettc BoOSt , Hisslon: • • • • • • • • • • • • • -• 2-36

2.6.i Hybrid EHL/ROCkel_ ................. 2-36

2_6.2. SSTO/TAV BooSter. ..... , ........... 2-37

2 7 Space-Based Missions • 2-39

• ' ' ..... D's'o'a:l p s i ........ 2 39

2.7 • 1 Satellite Kick System • • • • • • • • • • •

2.7.2 Space-Based Nuclear Waste . ; . . . . . . 2-40

3 0 EHL SYSTEHS ANALYSIS 3-1

2-19

2-19

3.1 PreliminawAnalysis of Candidate Concepts

3.1.1

3.1.2

3.1.3

3.1.4

3.I.5

3.1,6

3.1.7

3.1.8

3.1.9

• 3-!

Ratlgun Concepts ........... 3-1

Earth to Orbit EHL System ............ 3-8

Earth-to-Or_bit Hybrid Launcher ........... 3-11

Lunar Supply Launcher. , .............. 3_16

Solar SystEm Escape Launcher : ............. 3-18

Earth-Escape Launcher . . . • • . . • • • • • • • • 3-18

SSTO/TAV Booster • . . . . • . . . • i, . • • • • 3-21

Suborbital Launcher: 3-22 • • ,. • • • • • • • • • • • •

3.1.10 Space-Based Systems .... , ........... 3-22

3.1,10.i Sate!lit, Kick SyStem ........... 3-25

3.i.10.2 Space-Based Nuclear Waste DispoSal . . . 3-26

3.2 Analysis of Radiation from Orbiting EML (Nuclear Waste

DispoSal system) . . , .... 3-2g 3.3 La,nch,tndo,Analysis y;t;mE;c pe rom

Orbiting EHL Syste,i ..................... 3-29

3.3.1 Previous Work: .................. 3-29

3.3.2 Orbiting EHL System: : .............. 3-31

3.4 Launch Window Analysis for Solar System Probes . . i . . . 3-32

3.5 Analysis of Requirements for Launching Payloads to

Earth Orbit ......... 3-36

3.5.1 Apogee Hotor impuis; Requ;rements: ......... 3-36

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3.5.2 Range Safety Limitations...

3.6 Lunar 8ase Resupply Mission Requirement; for'Vertic;1

EquatOrial EML Launches ................

3.6.i Frequencyof LaunGh Opportunities. i i I ......

3.6.2 Launch Velocity Requirements., ......

3.6.3 Lunar LandinO RetroRequirements..,...

3.6.4 Launch WindOws ................., .....

3,6,5 COnclusions.., ...... 3.7 Earth-to- eos chronous OrbitLaunch eio i y'R;q;i e;, t;

3.8 Projectile ConceptS..., ..... ..........

3,8.2 Hybrid EML/Rocket ........,...

3,8.3 Lunar-Supply Projectile ...............

3,8,4 TRU Baste Projectile ........,...,..

3,8,5 GEO Projectile...,

3.9 Nose Cone Material Selection:.:............... . ..........

3_9.1 AblattonConsiderattons, ......... ....

3,9.2 EconOmic ConSiderations ...............

3.i0 Projectile Stress Analysis ................

3-41

3-41

3-41

3-43

3-4S

3-46

3-47

3-47

3-47

3-49

3-49

3-53

3-53

3-54

3-54

3-55

3-56

4

.

0 REFERENCE CONCEPTS.

_,,•••.•••••••••••••••

4

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4.1 Preliminary Evaluation of EHL COncepts ..........

4.2 Earth-tO-Space Rail Launche_ Concept ............

4.2,1

4.2.2

4-1

4-2

Concept Definition ................ 4-2

4.2.1.1 ProjectilePayload Fabrication ...._ 4-2

4.2.1.2 Surface Transport of Projectile/Payload. 4-2

4.2.1.3 Projectile/Payload Preparation at

Launch Site ............

4.2.1.4 Launch Operations_ ...........

4.2.1.5 Trajectory Nonitortng_ .........

System Element Definition ....

4.2.2.1 Projectile/Payload Characteri;tics ...

4.2,2.1.1 Payload ....... 4.2.2.1.2 Propulsio y;t;m: ....

4.2.2.1.3 P_ojectile Elementsl ....

4.2.2.2 Surface Iransportation Systems ..........

4.2.2.3 Launch Site Support Facilities ......

4,2.2.3.1 Power Plant .....

4.2.2.3.2 Projectile St;rag; and

Checkout Facilities.

4.2.2.3.3 Genera] Storage Facilities.

4.2.2.4

4.2.2.3.4

4.2.2.3.5

4.2.2.3.6

4.2.2.3.7

4.2.2.3.8

Administration and

Engineering Facilities ....

Industrial Area ....

Community Living Areal..

Liquid Gas and Water Plant.

Other Facilities ........

Launcher System ......- ......

4.2.2.4..1 Bore and Rails ........

4.2.2.4.2 Energy Storage ........

4-5

4-5

4-5

4-5

4-6

4-6

4-6

4m7

4-9

4-9

4-10

4-10

4-10

4-I0

4-10

4-10

4-10

4-12

4-12

4-12

4-16

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4.2.2.4.3 Launcher Support Structure. 4-16

4.2.2.4.4 Preaccelerator System .... 4-I6

4.2.2.4.5 Switching and Control.... 4-18

4.2.2.4.6 Temporary Projectile

Sto_age. Factlities.,.. 4-18

4.2,2.4.7 Service and AcceSS Systems: 4-i8

4.2,2,5 i_nitortng Systems ........_ .. 4-18

4226 Sp ce Des¢i[ 18

...a nation, ,, ......... 4-

4.3 Earth,to Space Coaxiai AcceieratoP Concept, .....,. 4-19

4.3,i Concept Defini_ion, ....., ......... 4-1g

4,3,2 System Element Definitton_. 4-19 4.3.2.i proJectile/Payloa 4.19

4.3.2,i,1 ....Pay|oad,, ...., ..... 4--19

4.3.2.1,2 Propulsion System ...... 4-20

4.3,2,1,3 Projectile.,,,.,... 4-20

4.3.2.2 Surface Transportation S_'stems ..... 4-20

4.3.2.3 Launch Site SupportFactlittes ..... 4-20

4.3.2.4Launcher System ....... 4-22

4.3.2,¢.1 Drive Coiis;_ i ; ..... 4-24

4.3,2,4.2 Energy Storage.. 4-24

4.3.2.4,3 Launcher Support Structure. 4-24

4.3.2.4,4 Preaccelerator Syste_ .... 4-24

-4.3.2.4.5 Switching and Control ..... 4-26

4.3.2.4.6 Temporary Projectile

S¢orage Facilities 4-26

4,3,2.4.7 Service and Access SystemZ] 4-26

4.3.2.5 MonitOring SyStems .......... ..... 4-26

4._.2.6 Space Destination .............. 4-26

44 Hybrid* Railgun/Rocket Concept iii!@1,11111._1114I 27

4.4.1 Concept Definition .............. 4-27

4.4.1,1 Projectile/Payload Fabrication .... :: 4-27

4.4.1.2 Surface Transport of Projectile/Payloa.d 4-27

4.4..i.3 Projectile Preparatlonat Launch Site.. 4-2.7

4,4.1.4 Launch Operations ................ 4-27

4,4.1.5 Trajectory Monitoring ............ 4-28

4.4.2 System Element Definition ...... 4-28

4.4.2.1 Projectile/Payload 6haracteri;tic; .... 4-28

4.4.2.1.1 PaylOad .... 4-28

4.4.2.1.2 Project"_ Elements]... 4-29

4,4.2.2 Surface Transportation Systems ..... 4-31

4.4.2.3 Launch Site Support Facilities ...... 4-31

4.4.2,3.1 Power Plant., ..... 4-31

4.4.2.3.2 Projectile Storage anJ

Checkout Facilities ..... 4-31

4.4.2.a Launcher System ............... 4-32

4.4.2.4.I Bore and Rails ....... 4-32

4.4.2.4.2 Energy Storage. 4-36

4.4,2.4,3 Launcher Support St;uct;r;: 4-36

4,4.2.4.4 PreacceleraLo_ System .... 4-36

4.4.2.4.5 Switching and Control ..... 4-36

4.4,2.4.6 Temporary Projectile

Storage F_cilities ..... 4-36

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4.4.2.4.7 Service.and /_ccess Systems. 4-36

4.4.2.5 MOnitOring SyStemS ....... , .... 4-36

4.4.2,6 Space Destination .... : ........ 4-37

4.5 Hybrid Coaxial Accelerator/Rocket Concept ......... 4*37

4.5.1 Concept Oefinttion ...... , .... ..... 4-37

4.5.2 SystemElementOeftnitton ..... ,.-...,. 4-37

4.5.2./ ProjectilePayload Characterisis,,t c . 4-38

4,5.2. i.1 Payioad.. . . , . 4-38

4.5,2.1.2 P_ojecti,e:: :: : .... 4-38

4.5.2.2 Surface Transportation Ssemsy t ..... 4-40

4.5,2.3 Launch Site Support Facilities ........ 4-40

4.5.2.4 Launcher Systems . , ........... 4-40

4.5,2.4.1 Drive Coils ...... , . -4-43

4.5.2,4.2 Energy Storage .., " 4-43

4,5.2.4_3 Launc_erSupport Structur;-: 4-43

4,5.2,4.4 Preacce]ePator System .... , 4*43

4,5,2.4.5 Switching and Control .... 4-43

4.5.2.4.6 TempOrar_ Projectile

Storage Faclllt{es . , . . ; 4-43

4.5.2,4,7 Service and Access Systems . 4*43

4.5_2.5 Monitoring Systems , . _ ...... , . 4-43

4.el5 2 6 Space Destination .O.OOIIt_.I. 4-44

5.0 SAFETY AND ENVIRONMENTALIMPACT ASSESSMENT........... 5-i

5.1 Preilminary Safety AsSessment ............ . . . . 5-I

5.2 Environmental Impact AsseSsment ........ 5-I

5.2.1 Facilities Development and C,onstruction_ ..... 5-2

5.2.2 TesttngandOperattons .... . . . , ...... 5-2 5.2.2.1 Son1c Booms............. 5- 4

5.2.2,2 Power Plant Emissions:: ........... 5-6

5.2.2.3 Chemical Effluents .............. 5-6

5.2.2.4 Solid Wastes ............... 5-6

5.2.2.5 Materials Usage... : ............ 5*6

5.2.2.6 Reentry of Hardware ........... 5-8

5.2.2.7 Orbital Debris ............. 5,8

5.3 Conclusions, .......................... 5-8

6.0 EML COST ESTIMATES, ..................... . . . . 6-1

'<l

6.1 Development and Investment Cost Estimates ......... 6-5

6.1.1 Facilttfes and Supporting Systems ........... 6-5

6.1.1.1 Land ...... ; ........ 6-5

6.1.1.2 Power'Piant: ..., , ....... , . 6-6

6..I.1.3 Pers0nne] Support'Facilities ........ 6-7

6.1.1.4 Transport Facilities . .. 6-7

6.1,1.4.1 Shipplng Docks.'Sto;age" and

, Transport Facilitles ..... 6-7

6.1.1.4.2 LandTransport Facilities... 6-7

6.1.1.5 Airfield and Hangar ......... . . 6-8

6,i.1.6 InduStrial Area ............. 6-8

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6.2

6.1.2

6.1,1.7 Adminlstration/Engiheering Bulldlngs .. 6-8

Launcher Systems .......... 6-9

6.1.2.1 Launcher Tube'Hou;ing_ ,... • , . . . . 6,9

6.i.2.2 Launcher Tubes 6-9

6,1.2.2,1 COp;;Ai;"' e 1 o LaunCher. .....

Conductors , . , ...... 6-10.

6,1.2.2.2 Electrical Insulation,, :: 6-1o

6,i,2,2,3 Radial Force Containment ,. 6-12

6,1.2.2,4 VaCuum COntainment..... , 6-12

6.1.2.3 EieCtrlCai Power Storage . . . , 6-12

6,i,2.3.i Brook, Cotl..,., i i i i 6-13 6.1.2,3.2 Homopolar Generators H G ) 6-13

6,1,2,4 Switching and EML inductors ........ 6-i6

6.I,2.5 Projectile injection Systems ....... 6-i7

6.1,2.6 Power Conversion Piant (Raiigun

Launchers) , . , .......... 6-18

6,1•2,7 water 01Stiiilng Plant : ...... ,.... 6.19

6.1.2.8 Gas Handling Facil_t!es .......... 6-19

6.1.2,8.i Liquid Nitrogen Piant and

Storage ..... ...... 6-20

6.1,2.8,2 Evacuation System for

LaunCher Tube• . . , . 6-20

6,1,2.8,3 Water Electrolysis Plant : : 6-20

6.1.2.9 Handling DeviceS/Systems , .

6.1.2.10 Control Center, Controls and Montto;ing • 6-22

Systems ........... 6*22

6.1.2.11 Tracking System: :S:s:e:" 'ytm " " 6.i.2,12 Accident Recovery s..... • . . 6-23

EML Operations cost Estimates. : ......... , .... 6-23 6.2.I Projectiles and Mission Peculiar Equipment . , : . 6-23

6.2.1.1 Hybrid EML/Rocket Projectile (Coaxial

and Railgun) ............. 6-23 6.2.1.1.1P;yio :

6.2.1.2

6.2.1.i.2

6.2.1.1.3

6.2.1.1.4

6.2.1.1.5

6.2.1.1.6

6.2.1.1.7

6.2.1.1.8

6.2.1.1.9

6,2.1.1.10

6123

Nose Cone'o; Pa;1;ad Sh;o;d: 6-26

Structure. , .... 6-26

Thermai P_otecttonSystem"

(SldeBody) ........... 6-26

Fins ...... 6-26

Sabot'o; Coil;IRings 6-26

Auxiliary Handling Equi;ment

and Supplies . . . . • 6-26

I.stru t Pack ge..... 6-27

Propulsion System... .... 6-27

Trafisportation • • . . . . . 6-28

Earth Orbital Projectile (Coaxial

Launchet_)•

6.2.1.2.1

6.2.1.2.2

6.2.i.2.3

6.2.1.2.4

6-28 l • • l l • • • •

"Payioad:: . . l l l . l l • 6-28

Nose Cone • • • l • . • • • 6-28

Structure. .... 6- 28

Thermal Protect.ion Syst;m

(Side Body) l • • .. • • • l l 6-28

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6.2.1.2.5 Ftns ......... . . . 6-29

6.2.1.2,6 Coils ...... 6-29

6.2.1.2.7 Checkou q i ment

Supplies , ......... 6-29

6.2.1,2.8 InStrument,Package ..... 6-29

6.2.1.2.9 propulsion System...... 6-29

6,2.1:2,10 Transportation .... 6-30

6.2.1.3 Earth-Orbital Projectile (Raiigun"

Launcher) ......... .. 6-30

6.2,1_3,2 Nose Cone, ....... 6-30

6,2.1.3,3 Structure. , .... 6-30

6.2.1.3,4 Ther_iialProtection SySt;m_ . 6-30

6.2.1.3.5 Fins , ............ 6-3_

6.2.1.3,6. Sabots ...... 6-31

6.2,1=3.7 Checkout _qGll_;nt and

Supp]ies . . . , , . . , . . 6-31

6.2.1.3.8 Instrument Package ...... 6-3i

6.2.1,3.9 Propulsion System ...... 6-31

6.2.i.3,10 Trahsportation ........ 6-31

622 Operations 6.31

6.3 Development Test Program ............ .6-33 ....

6.4 Economic Evaluation of EML concepts:........... 6,35.

6.4.1 Launch Systems Alternatives .............. 6-35

6.4.2 Comparison of Alternatives .......... .... 6-40

7 . 0 TECHNOLOGY ASSESSMENT .ol....eeOOOQleeoeeea 7.1

7.1 Technology Evaluation ....................... 7-1

7• 1.1 Scale*up ......... 7-1 7.1.2 EnergyDi;t i u ion'and'S i c tng:.ii. ..... 7-2

7.1.3 Projectile Development ................. 7-2

7.1.4 Energy Storage Systems ........... 7-3 7.I.s Projectile Brushesand_a_u_es : .......... 7-4

7.1.6 HPG Brush Material ................... 7-4

7.1.7 Launcher Tube Bore ....... ........... 7-5

7.1.8 Preboost Systems ............ 7-5

7.1.9 EML System Structural'Support[ . •

7.2 Supporting Research and Technology (SR&T) Recon_endations_ 7-67.6

7.2,1 EML Experimental Research. , .... ; • • 7-6

7•2.2 EML Mission Requirements Studies • : ....... 7-8

7•2.3 EML Systems Studies. , .............. 7-8

7.2.4 Special Studies .................. 7-9

8 . 0 REVIEW OF ALTERNATIVE EML CONCEPTS m • • • • m • • • • • • • • • 8-1

8.1 Coaxial Accelerator Concepts ........... 8-I

8.1.1 Frequency Controlled Coil'Drlver .... ...... 8-I

8 00 1 2 Solenoid Earth Launcher .o,..•.-.o.•,_.. 8-3

81.2 E1ectrothemal Thrusters , .............. 8-6

821 Electrothe_al Rocket: 8-6 I! Illlllllllllll

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8.2.2 ElectrO_her_al Ramjet 8-7

8.2.3 Aflnular-F1owElectrothermal Ramjet ......... 8-7

8.2.4 LauhCheP Requiremeht_ ................. 8-7

8.2.5 EnePgyTra.s_e_ ..... :: .............. 8-11

8.2.6 Suma_y, ................... .. 8,13

8.3 Electromagnetic Rocket Gun ................... , 8-14

8.4 Eiectro_agnetic Theta Gun .................. 8-16

8" 5 SOviet EleCtrOmagrfletic. Accelerators 04"0"0,..0o0, 8" 17

8.5.1 Induction Acceleration .................. 8-17

8._.2 Ratlguhs ...................... 8-17

APPItNDIX A• REFERENCES• • • • "0 • • • • • • • • • • • . . . .,. _ . . Ai 1

APPENDIX.B. ACRONYMS AND.ABBREVIATIONS ................. B-1

APPENDIX C. METRIC TO ENGLISH CONVERSION FACTORS .......... C-1

APPENDIX D. EML RESEARCH, INC., REPORT ................ D-i

APPENDIX E. COLLABORATIVE PLANNERS, INC., REPORT .......... E-I

APPENDIX F. REPORT DISTRIBUTION LIST ................ F-I

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Table 1-1•

Table 1-2a•

Table 1-2b•

Table 1-3•

Table 2-1a•

Table 2-1b.

Table 2-2a•

Table 2-2b.

Table 2-3•

Table 2-4.

Table 2-5.

Table 2-6.

Table 2-7•

Table 2-8•

Table 2-9.

Table 2-10•

Table 2-11.

Table 2-]2.

Table 2-13.

LiST OF TABLES

Mass (MT) of Earth-Orbital Payloads Per Year ........ 1.12

Number of EML Launches Per Day for Earth-Orbital Launcher. 1-13

Number of EML Launches Per Day for Hybrid EML/Rocket

LaUncher .......................... 1-13

Expected Costs ($, M, 1981) ............. .... 1-15

Projected LEO Space Station Personnel for the Years 1991

Through 2000 ......................... 2-2

Projected Space Station Personnel for the Years 2000

Through 2050 ....................... 2-2

Early OTV Flight Projections ......... •....... 2-4.

OTV Flight Projections for 2000 Through 2050 ........ 2-4

Mass (MT) of Earth-Orbital Payloads Per Year ...... . 2-10

Number of Flights Per Year for Earth-Orbital Launcher.._. 2-11

Number of EML Launches Per Day from Earth-Orbital •Launcher 2-12

Life Support Resupply Estimates Required for Lunar.Base

Evolution (kg/person/day) .................... 2-13

Alternatives to Vehicles in the Earth-Moon Transport

Regimes ................ . ............ 2-16

Annual. Lunar Base Supply Requirements (MT/yr) in 2010• . . 2-20

Annual Lunar Base Supply Requirements.(MT/yr) in 2030... 2-20

Payload Manifesting in 2010 (Number of Vehicle Flights).. 2-23

Payload Manifesting in 2030 (Number of Vehicle Flights).. 2-23

TRU Waste from Fuel Rep-ocessing ............ . . . 2-30

TRU Wastes from Fuel Refabricatio.n . . . . . ....... 2-_0

6  ̧

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Table 2-14•

Table 2-15.

Volume of U•S. Commercial Waste Available for Space

Disposal ..........................

Mass of U•S• Commercial Waste Available. for Space

DIsposal • • • • • • • • 6 • • • • Q • • 6 • • • • ,e • • •

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Table 2-16.

Table 2-17.

Table 2-18.

Table 3-1.

Table 3-2.

Table 3-3.

Table 3*4;

Table 3-5.

Iable 3-6.

Table 3-7•

Nuclear Waste Disposal in Space Traffic Model (HLW and TRU

Wastes) ............................ 2-32

Hu_ber of Flights Per Year for HybridEML/Rocket Launcher. 2-38

Number of EML Launches Per Day for Hybrid EML/Rocket

Launcher .................... ,....... 2-37

EML Payload Requirements Summary .............. 3-1

Spark Gap Lengths for Needle Point ElectrOdes ....... 3-6

Rallgun Concept Sizing Summary .............. 3-4

Coaxial Launcher Concept Sizing Summary .......... 3-9

Calcuiated Railgun Parameters for Selected Hybrid Cases.. 3-15

Radiation Dosage as a Function o_ Distance from the Waste

Form Projectile .........................

Representative Midcourse Velocity Increment Requirements .

3-30

3-46

Table 4-1.. Earth-to-Space Rail Launcher Projectile Mass Summa.ry .... 4-9

Table 4-2. Earth-to*Space Coaxial Accelerator Projectile Hass Summary 4-22

Table 4-3. Hybrid Railgun/Rocket Projectile Mass Summary ........ 4-2g

Table 4-4. Hybrid Coaxial/Rocket Projectile Mass Summary ....... 4-38

Table 5-1.. Facility Construction Materials Requirements ........ 5-3

Table 5-2.. Critical Distances for Expected Sonic Boom Overpressures . 5-5

Table 5-3.

Table 6-1.

Table 6-2.

Table 6-3.

Tabie 6-4.

Table 6-5.

Table 6-6.

Major Operational Materials Requirements for Earth-toOrbit EML Reference Concepts (Hybrid ConceRts not

Included ........................... 5-7

EML Development and Investment Work Breakdown Structure. •. 6-2

EML Development Test Program and Operations Work Breakdown

Structure .......................... 6-3

EML Investment Cost Estimates ($, M, 1981) ......... 6-4

Launcher Tube Cost Calculations (2000-m Tube) ........ 6-11

Energy Storage. Requirements ................. 6-14

Liquid Nitrogen Requirements ............... 6-21

xii

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Table 6-7. Cost Estimates for EML Projectiles ($, K, 1981) .......

Table 6-8. Personnel and Supply Cost Estimates (People and $, M,

1981) .................. ............

Table 6-9.

Tabie 6-10,

Development Test Program Cost Lsttmtes ($, M, 1981) . . .

EML Cost Estimate Summary ($ 1981) • | • • • • Q • -q • -e • • •

6-24

6-25

6-35

6-34

Table 6-11. Cost Comparis_ of Alternatives for Space StatiOn Supply

Mission(2000-2029) ($, M, 1981, Expected) ........ 6-41

Table 6-12. Cost Comparison of Alternatives fOr_ 2010-2030 Lunar Base

Supply Kiss_on ($, M, 1981, Expected) ...........

Table 7-1. Estimated Funding Requirements (1984, KS) for Recommended

EML SR&T Activities • • • • - • • • • • "o • • • • al • ,o • • •

Table 8-1. Mass and Velocity Summary for Mawardi's Earth-to-Orbit

Projectile ............................

Table 8-2. Payload MaSses as a Function of Structural Mass for

IO00-kg Projectile .................. ...

6-43

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Figure 1-1. Study Appt-Oach ..... _ • , ................. 1-3

Figure 1-2• Overview of the Earth-tO-0rbit Rat1 Launcher Concept ...... 1-5

Figure 1;3.

Figure 1-4.

Figure 1-5•

Figure 1-6•

Figure 2-1.

Figure 2-2.

Figure 2-3.

Figure 2-4.

Figure 2-5.

Figure 3-1,

Figure 3-2•

Figure 3-3•

Figure 3-4.

Overview of the Earth-tO-Orbit Coaxial Accelerator Concept 1-7

Overview of the Hybrid Railgun/Rocket Concept.. ..... 1-8

Overvtew of the Hybrid Coaxial/Rocket Concept, ....... 1-10

Comparison of Tota] Program Costs ............. . 1-16

LUnar Logistics Scenario , ............... 2-i7

Overview of "Standard" Nuclear Waste Disposal in Space

Mission ....... ..... ,......... , , .... 2-27

Overview of ESRL Nuclear waste Disposal in Space MisSion . 2-28.

U.$. Sounding. Rocket Launch History ............ 2-34

Sounding Rocket Payload HassTrends ............. 2-35

Operattona] Principle of a Railgun ............. 3-3

Mutual Inductance Gradient as a Function of Projectile

Coil Position ........................... 3-5

Discrete Coil Launcher ................... 3-6

Selected Reference Design Launcher ............ 3-6

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Ft 9!re 3-6.

Figure 3-7.

Earth-to-Orbit Options ............. , ..... 3-10.

Electromagnetic Boost Options ...... ............ 3-12

Hybrid EHk/Rocket Maximum Payload as a FunCtion of

Launched Hass ......................... 3-13

i'.°

Figure 3,8•

Figure 3-9.

Figure 3-10•

Figure 3-11.

Lunar Base Supply Options • •, • • e • • • _ • • • • • •- • • 3-17

Solar System Escape and Earth Escape Options ....... 3-I9

Overview of Nuclear Waste Disposal in Space Mission

SCenarto (Ratlgun Launch Illustrated) ........... 3-20

Railgun Launcher-for TAV or SSTO Vehicle ......... 3-23

PRECEDING PAGE BLANK NOT Fli.b¢"l)

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Ftgure 3-12,

Figure 3-i3.

Figure 3-i4.

Figure 3-15.

Figure 3-16.

Figure 3-t7.

Saddle COt] Launcher for TAV and S$TO Applications ....

Suborbital Options .......... ..........

Space-BaSed Options.. , ; . . . . , , . , ....... . .

Possible Railgun Configurations for Satellite Kick Systems

Waste Form Geometry .................. . . . .

Launch From SO0-km Orbit on BeSt Day ............

Figure 3-18. Launch From 500*kin Orbit on Worst Day ..............

Figure 3-10.

Figure 3-20.

Figure 3-21.

Figure 3-22,

Pa_ t

3-23

3-24

3-24

3-26

3_ 30

3-32

3-32

Solar System Probe Aphelion/Perlhelion Range ........ 3-34

Solar System PrObeS_-Maxlmum Distance from Sun ....... 3-35

ApOgee Motor impulse Requirements for Circular Equatorial

Orbits .................... .......

Longitude of Shield impact Relative to Equatorial Launch.

Site ..............................

Figure 3-23.. Longitude of Shield Impact Relative. to Belem, Brazil,

Launch Site ............................

Figure 3-24. Longitude of Shield Impact Relative to Quito, Equador,

Launch Site ..........................

Figure 3-25. Flight Time and Impact Velocity Versus Launch Velocit_ . .

Figure 3-26. Velocity Impulse Required to Establish Circular Parking

Orbit About Moon .............. , .........

Figure 3-27. Velocity at Lunar Surface Versus De-Orbit Impulse.....

Figure 3--28_

Figure 3-29.

Figure 3-30.

Figure 3-31.

Figure 3-32.

Figure 3-33.

Figure 3-34.

3-37

3-38

3-39

3-40

3-42

3-42

3-,44

Total Retro Impulse for Soft Lunar Landing ........ 3"44

LaunCh Velocity and Angle Requirements for 500-km Circular

Orbit .............................. 3-48

Delta-V Requirements for 500-km Circular Orbit ...... 3-48

Earth-Orbital Projectile (Railgun) ................ 3-fRO

Earth-Orbital Projectile (Coaxial) ............... 3-50

Hybrtd Raiigun/Rocket Projectile ...... . ....... 3-51

Hybrid CoaxialRocket Projectile , ........ .... 3-51

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Figure 4-1..

Figure 4-2.

Ftgure 4-3.

Ftgure 4-4.

PaylOad Mass aS a FunCtiOn of AV/I and Propulsion System

Hass FoacttOn

Preliminary Screening Of Ratlgun Systems .........

Preliminary Screening o_ Coaxial EHL Systems .... ....

Earth-Orbital Proje¢¢ile for Rail LaunCher System, , . . .

EML Launch Site

3-52

4-3

4-4

4-8

4-1t

Figure 4-5. RatlgunSystem Cross-Sectional View ...... , ..... 4-i3

Figure 4-6, Earth-to-Orbit Railgun System Side Yiew ............. 4-14

Figure 4-7. Options Chart - Earth-to-Orbit Rafl Launcher ,....... 4-15

FtguPe 4-8, Homopolar Generato_ and Inductor . . . ........... 4-17

Figure 4" 9; Proposed Preboost Concept oeeoem_e, tele••we 4" 17

Figure 4'10• Earth-to-Orbit Coaxiai Projectile ............... 4-21

Figure 4-11. Coaxial Accelerator System Cross-Sectional View ....... 4_23

Figure4-12. Options Chart - Earth-to-Orbit Coaxiat Launcher ...... 4-25

Figure4-13. Brooks Coil with. Dimensional Relationships ........ 4-26

Figure 4-14. HybrtdRailgun/Rocket Projectile . . . . . . . . . . . . . 4_30

Figure 4-15. Hybrid Railgun/Rocket System Cross-Sectional View ....... 4-33

Figure 4-16. Hybrid Railgun/Rocket.System Side View ............. 4-34

Figure 4-17. Options Chart- Hybrid Railgun/Rocket Launcher ...... 4-35

Ftgure 4-18. Hybrtd Coaxial/Rocket Projectile .............. 4-39

Ftgure 4-19. Hybrid COaxiai/Rocket System Cross-Sectional view. ..... . 4-41

Ftgure 4-20. Options Chart - Hybrid Rocket/Coaxial Launcher ....... 4-42

Figure 6-i. Effects of Discounting on Total Program Costs ....... 6-42

Figure 8-1. Frequency Controlled Coil Driver ............. 8-2

Figure 8-2. Solenoid Earth Launcher .................... 8-5

Figure 8-3, Electrother_nal Rocket Concept, . . . . , , . . . . , . . . 8-8

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Figure 8,11.

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Eiectrothe_nal Ramjet Concept . • • • . • • • • • • • • • • 8 8

Annular Flow E]ectrothemal Ramjet 6 • • • • • • • ,• • • • • •

Proftle ot' Launch Tube

8-9

8-10

Power" Proftle • - 8-11

Themal-to-Nechanical Energy Efficiency Peoftle ...... 8-12

ElectromgnettC, Rocket Gun . , ........ , ..... 8-15

Shape of Teavellng Hagnet.tc Wave . ...... , . . . , 8-15

Schemttc of Electromagnetic Theta• Gun . . . . . . . . . . 8-16

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1.0 EXECUTIVESUltRY

This Fi.nal Technlcal Report documents the findings of the

Electromagnetic LaUnChers, "Space Mission Analysts of " NASA Contract NAS3 233B4. The Executive Summary (Section 1.0) contains the study

background, objectives, approach, principal assumptions and requirements,

and "summariZes the reference concepts, major results, COnclusions, and

recommendatiOns, Sections 2.0 through 8,0 present the technical details

of the StUdy results, Appendix A provides, the references. AcronymS..

and abbrievtattons are defined in Appendix B, while Appendix C contains

metriC-to-English conversion faCtOrs. Separate- NASA-funded studies

by £1ectromagnetic Launch Research, inc., and Collaborative Planners,

Inc., are attached to this report in Appendices O an_ E, respectively.

Appendix F contains the distribution iistfor this-final report.

1,1 Study Background

In 1982, Battelle Columbus LabOratOries• conducted a feasibility

asSesSment of an electromagnetic Earth-to-space rail launcher (ESRL)

to determine the viability of developing a rail launcher System in the

2020 time frame to launch, material into space .(Rice, Miller, and Earhart,

1982). Based upon an evaluation Of potential applications, a reference

concept was Selected. The reference concept consisted of two missions;

the first, mission would launch prOjeCtiles to solar system escape

velocities to dispose of comercial high-leve'l nuclear waste, the second

would launch cargo to low-Earth orbit to resupp]y U.S. space stations.

The ESRL system would be based at an equatorial site, with two separate.

rail launcher tubes placed in underground shafts. One tube would be

inclined at 20 degrees from the horizontal for Earth-orbital missions;

the other would be vertical for solar system escape nuclear.waste disposal

missions. Each launcher tube would be 2.04 km in length, and surrounded

by I0,20& homopolar generator/inductor units to transmit power to the.

rails. The two rai.1 launchers would be powered by a common power plant.

Total projectile masses were 6500 kg for Earth-orbital missions and

2055 kg for solar system escape missions. These projectile masses

corresponded to payload masses of 650 kg and 250 kg respectively_ The

seemingiy large ratios of projectile-to-payload were due to shielding

requirements for nuclear waste and orbtt*circulartzation propulsion

system requirements for Earth-orbital payloads. Based upon preliminary

technical, environntental, and economic analyses, it was concluded that

the ESRL system appeared to be technically feasible and economically

beneficial.

The previous ESRL study investigated only electromagnetic

railgun launcherS for Earth-to-space missions. Other types of

electromagnetic launchers (EMLs) are currently being studied, the most

well known of which are the coaxial accelerators. This study was

conducted to investigate all types of EML concepts for performing a

variety of space missionS.

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1.2 Study 0b;_ecttYes

The overall, objectives of this study were: (1) to provide

NASA. Lewis Research Center (NASA/LeRC) With sufficient information

such that a c0mpaPiSon could be made between various promising EML Space

mission concepts; (2) to develop., mission models and requirements for

(MLs; (3).to. define reference system concepts; (4) to conduct preliminary

analyses Of economics and performance;, and (5) to recommend areas Of

technology development,

1.3 Approa,(:ll

The study approach emphasized the .assessment of factors which

would contribute to the COmpariSOn of the different types of electromagnetic launchers. Thase factors include mission definition and

requirements, pPelimfnary conceptualization of EML systems,.economics,

and technology status,

The analysis included four tasks: (I) Characterization of

Candidate EMLS; {2) Development of MisSion MOdelS and Requlrements_

{3) EML COnCept Analysis; and (4) Technology Assessment. Figure I-.

emphasizes the overall study approach. The specific study tasks and

their interrelationships are outlined.

Initially, the study addressed space missi-on applications

of EMLS beginning in the year 2020 and continuing through 2050. The

study was revised midterm to reflect a more near-term operational start-up

in 2000 for the Earth-to-orbit mission and in 2010 for the lunar base

supply mission.

In order to stay abreast of the rapid developments in this

field, it was necessary to maintain contacts with others working in

this area. This COntaCt included attendance at two EML conferences,

the American Defense Preparedness AssociationSeminar on Electromagnetic

Launchers in February 1983 and the SecOnd Symposium on Electromagnetic

Launch Technology in October 1983. Also, experts in coaxial

electromagnetic accelerator technology were contracted_ separately by

NASA/LeRC to provide input to the preliminary conceptualization task.

1.4 Stud_ Guidel|nes

The guidelines which were used in the performance of this

stay include:

• Battelle made maximum use of related studies and other

associated data, as appropriate.

•Battelle considered EML systems available in the open

literature.

• Peaceful space applications of EMI. sys%ems were considered.

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STUDY APPROACH

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• Reusability of potentially recoverable

was considered.

EML projectiles.

• All costs were given in 1981 dollars to COmpare With the

previous Battelle study.

• Study activity was scoped to follow allocated funding

resourceS.

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Fo_P reference concepts were selected following preliminary

conceptualizatiOn and evaluation of the seven identified missions using

railgun and coaxial acce)erator concepts The four concepts selected

for further study are.briefly described in thfs section.

1.5.1 Earth-t_-Orbit Pail LaunCher

The Earth-to,orbit rail launcher cOnCept was envisioned to

supply materials to an orbiting Space StBtion. These supply items would

include hydrogen/oxygen p_opellants in water form,, life. Support.

consumables, spares, and mlscellaneous _aterials.

Projectile systems and certain payloads would be fabricated

and transported to the launch site, where they would be placed.in storage

until launch time, The required number of p_jectiles (with integrated

payloads) would be transported daily to, the launcher. The projectile

would be placed in the preboost _ection of. the launcher and then launched.

The muzzle velocity would be 6.85 km/s, with an additional 2.1 km/s

required for orbital insertion supplied by the projectile propulsion

system at 500 km altitude.

The launcher woUld, be inclined, along a mountain side at

20 degrees to the horizOntal, to .optimize atmospheric drag losses and

orbit-insertion propulsion requirements. The iau_cher tube would be

surrounded by 3600. homopolar generator/inductor unitS providing energy

storage. The power would be provided by a nuclear power plant. The

launcher tube would be partially i_nbedded in a concrete foundation to.

provide structural support to preveat da(aage to the bore .from the bore

stresSes.

An overview of the Earth-to-orblt rail launcher system is

shown in Figure 1-2. Included in the figure are an illustration of

the launcher elevation angle, crosS-sectional and side views of the

launcher system, and a projectile concept drawing. Detailed discussion

Of this system is provided in Section 4.2 of this report ............................

i.5.2 [aPth,to-0rbtt Coaxial Accelerator

The Earth-to-orbit coaxial accelerator was also conceptualized

to deliver supply materials ._ Earth orbit. The payloads would be the

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same as those described tn Section 1.5.1. The overall mission scenario

is also the same.

The launcher would be built at a 20 degree elevation angle

on a mountain side. The launcher tube wOuld be partially imbedded in

a concrete foundation. The power would be supplied by a nuclear power

piant and stored in a single large Bt'oolcs. ¢0il inductor. The current

is suppiied to individual turns of the drive coil.

Figure i-3 provides an overview of the Earth-tO-orbit coaxial

system, A crOSS,SeCtional View of the launcher system, a projectile

concept drawing, and an illustration of the launcher elevation angle

are shown in the figure. Section 4.3 of thtS report discusses this

SyStem in mope detail..

1.5.3 _Rodket

Payloads would be launched to low-Earth orbit via a hybrid

railgun/rocket system. These. payloadS would, be identical to those defined

in the Earth-orbltal traffic model, including propellants,, life support

expendables, spareS,., materials .for spaceprocessing, and miscellaneous

items. --

The three solid-rocket stages would be manufactured and loaded

with propellant prior to delivery to the launch site storage facility.

Before launch, the motors would be stacked and the payloads integreted.

Projectiles would be transported daily to the launcher. At its scheduled

launch time, each projectile would be placed in the breech of the launcher

and after the launcher system has been fully charged, the projectile .

would be launched at 2 km/s, The projecti.le would continue along its

trajectory, through three stage burns, to its destination at low-Earth

orbit.

The launcher would be located on a mountain side at a 35,degree.

angle from the horizontal (no rocket vehicle kick angle is requlred).

Energy storage is provided by 750 homopolar generator/inductor units

lined along the 2-km long launcher tube. The power would be provided

from commercial utility power plants. The launcher tube structural

support is provided by a concrete bed surrounding half of the tube to

prevent structural damage to the rails and boredue to launch stresses.

Figure I-4 provides an overview of the hybrid railgun/rocket

system. Cross-sectional and. side views are illustrated, as are the

projectile concept and launcher elevation angle. Further discussion

of this reference concept is given in Section 4.4.

_rtd Coaxial Accelerator/Rocket

The hybrid coaxialrocket concept was envisioned to supply

the Same payloads to low-Earth orbit as the hybrid railgun/rocket. The

mission scenario would be the same as well.

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A mountain with a 35-degree elevation angle over a 2-km length

is conceptualized. Energy would be provided from a utility power plant

and stored in a single large Brooks coil inductor. The launcher tube

would be partially imbedded in a concrete foundation to provide structural

support to prevent damage to the drive coils from hoop stresses during

launch_

An overview -of the hybrid coaxial/rocke_ is ShOwn. in Figure

1-5, The- launcher, crosS-section, the projectile conceptualization,

and an illustration of the launcher elevation angle are given in the

figure. Section 4.5 provides further discussion of the reference concept

system.

_or Results

The major results of this preliminary assessment of EML space

missions are summarized in the following paragraphs.

1.6.1 A1ternati ve EPIL Concepts

A survey of EML conceptswas conducted to identify those which

could perform the selected space missions. The open literature (U.S,

and foreign) was reviewed and EML experts contacted. Fi_,': concepts

were identified and reviewed:

¢ Railgun

e: Coaxial accelerator

e Electrothermal tl_ruster

• Electromagnetic rocket gun

• ElectrOmag_etlc theta gun.

Of these concepts, two were selected for the reference concepts based

upon technical evaluation--railgun and coaxialaccelerator.

1,6,2 lltsSton Hodels and Requirements

Seven missions were identified for definition in this study

and are listed below:

e Earth-orbital launch

• Lunar base supply

• Solar system escape

• Earth escape

• Suborbital launch

e SSTO/TAV boost.

• Space-based launch.

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FOr a particular mission to merit development of an EML system, large

amounts of payload deltvery are required. The Earth-orbitel mission,

Including Space Station supply, OTV propellant delivery, and raw material

supply, had the largest material deltvery requirement and was .the basis

for th, reference concepts.

Table 1-1 presents the two Earth-orbital mission models,

indicating yearly mass deltvery requirements. The low model assumed

that there wOuld be 270 persons in orbit in the year 2050; the high

model assumed G70 people. Traffic models for the Earth-to-space EML

(G50 kg maximum payload) and for the hybrid EML/rocket (800 kg maximum

payload) are shown in Tables 1-2a and 1-2b, The latter two tables

illustrate the daily launch rate for each system between the years 2000

and 2010,

The lunar base supply mission, added tO the mission list late

in the study, also had large material delivery requirements. Although

the costs were not studied in detail, it was fel:t that an equatorial-based

EML system could be a cost-effective method of transporting material

to low-Earth orbit, geosynchr+onous orbit,, and the MOon.

1.6.:3 Safety and Environmntal Impact

A brief safety and environmental impact assessment, evaluating

various issues, was conducted. Development and construction of. the

EML systems are expected to have some environmental effects. Local

sonic boom effects are expected;, however this is not expected to be

a critical issue, Overpressures on: the order of supersonic aircraft

are expected to occur betwee,_ 355 and 700 m for the Earth-to-orblt EML

and at about 200 m fop the hybrid system.

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Major accident events for the EML systems are likely to be

projectile break-up after launch, llquid-propellant spills, or on-pad

fires. Care must be taken to pt.otect.workers and the local population

from hazardOuS exposure. The safety risk is expected to be comparable

to current space, activities if the launch site is selected to avoid

overflight of populated, areas.

A benefit of using the Earth-to-orbit EML systems may be to

reduce adverse environmental impact from the Space Shuttle by reducing

the number of Shuttle flights. Reductions would be in the areas of

effluent quantities and impact to the. ozone layer from Shuttle HCl

emissions. Although the hybrid EML/rocket systems would release HC1 .

and A120 3 into the atmosphere during first-stage burn of the three-stage

rocket, the emissions would be dispersed and pose less of a hazard than

the conventional systems which have a large ground cloud of exhaust.

Also, the hybrid systems may add to orbital debris by emitting Alp03

particulate into Earth orbit at Space Station altltude_. Further analysis

of a liquid-propulsion system may be desired for the hybrid EML/rocket

system.

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TABLE 1.-2a. NUMBEROF EML LAUNCHESPER DAY

FOR EARTH_ORBITALLAUNCHER

Launches Per Day

Year Low Hodel Htgh Hodel

2020 1.4 1.5

2005 2.8 3.9

2010 3.5 5.0

2015 4.1 6.5

2020 5.1 8.0

2025 5.2 9.8

2030 6.3 11.1

2036 6.5 _ 12.6

2040 7.5 14.4

2045 7.7 15.9

2050 8.7 17.0

F1fry-Year Average 5.3 9.6

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FOR HYBRZDENL/ROCKET LAUNCHER

Year

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Launches Per Day

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Low Hodel Hlgh Hode'l

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2000 0.2 1.4

2005 1.4 1.8

2010 2.0 2.2

20i5 2.6 3.•5

2020 2.5 4.2

2025 2.8 4.9

2030 3.2 5.7

2035 3.4 6.5

2040 3.8 7.4

2045 4.0 8.1

2050 4.4 8.7

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Ft fty-Year Average 2.8 5.0

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1-14

1. 6. 4Cost£stimtes

Preliminary cost estimates are given in Table 1-3. The expected

costs are shown for each of the four reference concepts. Expected

investment costs ranged from $1.2 B to $2.2 B, w_th annual operating

costs between $35 M and $40 M, not including projecti]e costs. SeCtiOn

6.0 diScusses the costs in detail.

At high launch rates, the prOjectile Costs dominate capital

costs. At low launch rates, the capital costs are not Spread over many

launches, So the EML System is not as cost effective. When amortized

over a 30*year period, One launch per day results in costs between $496

and $757 per kilogram. A launch rate of ten per day corresponds tO

COSts between $181 and $234 per kilogram.

Figure 1-6 compares total program costs for the current STS,

a conventional four-stage rocket (800_kg payload), and the hybrid

EML/rocket system, all fort an initial operating capability (lOt) date

of 2000. A $1.B billion development cost was assumed in this study.

With launch rates below two launches per day, the front-end.investment

causes discounting of cost streams to favor the four-stage rocket for

payload delivery. At higher launch rateS, the EML front-end investment

is recovered and the hybrid system is favored.

Coaxial accelerators have potential for lower costs than railgun

SyStems for several reasons, The projectile stresses are lower because

multiple projectile coils distribute the acceleration loads throughout

the projectile. This means that structural masses may be less. The

launch tube hoop stresses are also lower which correlate to reduced

tube structural masses. For the 2000 IOC, a .single Brooks coil was

assumed for the coaxial accelerator concepts as opposed to multiple

homopolar generators and inductOrs for the railgun systems. The- single

energy store would be less expensive than the multiple stores.

1.6.5 Tech.olo 

EML technology needs to advance to, further define the reference

concepts and to improve the cost estimates. Areas of needed technology

development include system scale-up, switching and energy distribution,

energy stoeage, brusheS, projectiles, and structural support. Battelle

has recommended that NASA conduct a supporting research and technology

program in experimental research and system studies to further evaluate

the potential benefits of the EML for space mission applications.

1.7 COaclustonS

Based upon this preliminary assessment, Battelle concludes

that electromagnetic launchers appear to be technically feasible and

economically beneficial in supplying material to space. However,

large-scaie EML development can be Justified only when large amounts

of material are launched. There appears to be no near-term (2000-2010)

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cost-effective civilian space ap_licatlon for EML Systems, bu.tprojections

O? future traffic to space stations and to lunar bases indicate that

there is-justification _or EML development beyond 2010, ' ...

1.8 RecnmeFdat|ons

Because of the potential long-term economic benefits of

electromagnetic launchers, for space applicationS, Battelle recommends.

that NASA continue investigations fn _reas related to these space missions.

a_d Continue- to track technology developments by other government

agencies, The NASA investigations should Include _ further analysis of

EHLs for lunar base applications. Detailed system and projectile designs

are eequired, aS. are Space, Station studies ,regarding EML logistics,.

such as Orbital Maneuvering Vehicle operations, automated rendezvous

and docking, and handling andstorage issues,

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.2.0 RISSIOE I()DI[L OEVELOi_IENT

This section presents the mission models and describes model.,

development for the seven identified electromagnetic launcher (EML)

space miSsiOns,. These missions are:

• Earth.Orbital

• Lunar Base.Supply

• Solar system Escape

• _Earth Escape

• Suborbital

• Electromagnetic Boost

, Space-Basod _L.

The mission mOde1_ iqc_ude projections of thetype of material

to be launched, the amount of ma_, and number of launches per year

for each of the m|ssions. The Earth-orbltal mission model spans fPom

ZOO0 to 2050, The lunar base supply mission model beglh_ in 2010. The

tlmefran_e for the other missions under conslderatJon lasts thirty.years,

beginniOg _n the year 2020.

2,1 Earth Orbit_l l_lssJon

Several science fiction writers nave considered the use of

electromagnetic launchers for launching cargo; recently, _tudtes have.

sho_ that EHLs are i,deed applicable for the cargo-launching missions.

These studies show that :he high accelerations necessary to maintain

reasonable launcher lengths preclude the use of ENLs for transportation

of personnel and sensitive equipment. However, lot bulk items, including

propellan_s,, materials for space processing, and certain food items,

EHL_ appear to. beet_r_ctive _Iternatlves..to the Space Sh,ttle and its

der_vatlve_.

The Earth-oroic_l mi,sion model assuages a significant manned.

presence in =pace-by the year 2020, Table 2-1a indicates the projected

personnel growth of an initial Space Station in Cow-Earth orbit from

199Z through the year ZOOU. An orbiting Space Station was assumed, to

be operational In Ig9_. The Initia] Space Station would be small and.

modular with up to eight persons permanenuly 1ocatea onboard (ninety-day

replace,nent cycle).. B_ adding additional modules, the Space Station

will evolve. In 1995, a_, increase of four persons was predicted, for

a iota| of twelve persons in orbit. By 2000, sixteen people are assumed

to be living and working in a U.S. Space Station. These estimates are

baseo u_on the results of the NASA Space Station Task Force Concept

Development Group defin4tion and from eight NASA contractors which

participated in the Space. Station Needs, AttributeS,. and Architectural

Options. Stud_ for NASA Headquarters in late 1982 and early 1983.

Reference_ to the reports of the contractors are 11sted as.follows:

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TABLE 2-1a.

2-2

PROJECTEDLEO•SPACESTATION PERSONNEL

FOR THE YEARS 1991 THROUGH2000

I

Number ofYear Fersonnel

i

1992 8

1993 8

1994 8

1995 12

1996 12

1997 12

1998 16

1999 i6

2000 16

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TABLE 2-1b. PROJECTEDSPACESTATION PERSONNEL.

FOR1}IE ¥EARS.2OOO]'HROUGIt2050

i f ,if i I

Year

lit r ii I i i i w.

Number.of Space Star,ion Pev'sonnel

Low Hodel High Hodei

2000. 16 16

2005 24 32

2010 40 60

2015 72 120

2020 100 220

2025 125 . 280

2030. 150 390

2035 175 450

2040 200 560

2045 225 620

2050 250 670

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e Boeing Aerospace Company (1983)

e General Dynai_tcs Convair Division (1983)

' e- Gru_an Aerospace Corporation (1983)

e Lockheed Rtssile and Space.Company (1983)

e" Hartin Marietta Aerospace(1983)

.e McDonnell Douglas Astronautics COmpany (1983)

• Rockwell International Corporation (1983)

e TRW (1983).

The projections beyond the year .2000 are categorized in low and high modeis and are shown in Table 2 lb. The low model assumes

that there will be a populatlon of 250 peop'le in Earth orbit by the

year 2050; the high model assumes a population of 670 people at this

same time, BOth models included a small military LEO station and added

a manned GEO station. The majOrity of the personnel in orbit are

projected to be in low-Earth orbit in a number of Space Stations.

2.1. It OTV Propellant Supply

Propellant transportation, is a large portion of the EML

Earth-tO-LEO mission model. The transport of orbital transfer vehicle

(OTV) propellants in their LH2 and. LO 2 forms, as well as water, was,

considered. Many more flights would be required to launch LH2 and LO2

(instead of water) because water is 2.89 times denser than an equivalent

amount of LH2/LO 2 propellants. Also, LH2 and LO2 tankage is more costly,

because cryogenic storage is required. Safety concerns also point to

Water as the propellant payload of choice. This concept assumes that

an electrolysis facility and adequate power would be located onbOard

the stations. Since the ratio of oxygen to hydrogen in water is 8:1

and most hydrogen/oxygen propulsion systems utilize an oxid.izer-to-fuel

(O/F) ratio of 6:1, there would likely be an excess of oxygen; however

O/F ratios of 8:1 are being considered for propulsion systems. Any excess

Oxygen delivered to orbit would be used for other station needs, such

as for life support or orbit;l _rag make-up.

2..X,1.1 Nissl0n ,llode|s

OTV preliminary designs and mission models are widely available

and greatly varied. Sources quoted in this section are: Boeing, 1980

and 1983; General Dynamics, 1980 and 1983; and Davis, 1982. All of

these sources studied cryogenic systems; recent storable-propellant

OTV system_ studies are not currently available.

Round-trip (LEO-to-GEO-to-LEO) OTV propellant requirements

quoted by the named references range from" 15,000 kg to 32,500 kg. In

this Study, the total propellant requirement for each OTV round trip"

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TABLE.2"2a. EARLY.OTV FLIGHT PROJECTIONS

I

Number.of OTV E11ghts Per Year

Year NASA/HSF.C. a Boe_ngb

J J i j i w

1991 8. 1

199S 7 6

2000 14. 7

2005 -- 9

2010 ....

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Notes: (a) From NASA/MSFC, 1980 (Nominal Hodel)

(b) From Boeing, 1983

(c) From HcDo_nell Douglas, 1983

(d) From Davis, 1982

(e) Projection for 1990

(f) Projection for 1999 ........

TABLE 2-25. OTV FLIGHT PROJECTIONSFOR

2000 THROUGH2050

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Year

Number of OTV Fltghts(a)

Low Hodel High Hodel

2000

2005.

2010

2015

2020

2025

2030

2035

2040

2045

2050

10

12

12

15.

15

18

18

21 l

21

24

10

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16

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23

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was asSu_ned to be 22,000_k9 (18,855 kg oxygen and 3145 kg hydrogen),

If transporting w_t_r to be electrolyzed on orbit, this corresponds

to 28,300 kg of water required per OTV fltght. Assuming losses in

transportation, handiing, and the electrolysis process, 29,500 kg was

used as the baseline for the OTV model.

Tables 2-2a and 2-2b sumarize the projected OTV flight

schedu]es. Projection3 for 1990 through 2010 Shown in Table 2-2a are

from Boeing (1983), McDonne11Douglas (1983), Davis (1982), and NASA/MSFC

(1980). Low and high models are Shown in Table 2-2b for the OTV Flight

projections for the fifty-year period from 2000 to 1050 used for this

study. Cohsideredtn these tables are the continued use of Earth-orbit

for Communicatlonsj e]|vironmental _onitoring, etc., and the likely manned

traffic between LEO and GEO and between LEO and the MoOn.

FoP an EHL propellant supply mission, with a launch directly

to GEO, propellants carried from LEO to GEO on the OTV for the return

deita*V to LEO is avoided. From Section 3.7, with a launch velocity

of 11.7 km/s, 430 kg of water cou]d be delivered to a GEO Space Station

water hOlding tafik_ The ,ater couid then be electrolyzed, liquified_

and loaded onto the OTV for Its return to LEO. The total propellant

mass required for the OTV round trip would then be reduced, because

the return-trippropellants are notcarrted as cargo from LEO to GEO.

2.1.1,2 implications of STS Scavenq_n_

Several studies are currently being funded by NASA ¢o

tnvesttgat_ the possibility of scavenging propellants from the Space

Shuttle E):ternal Tank (ET), Orbiter lines, and tankage in the payload

bay. Rockwell is studying payload bay tankage concepts and math

propulsion system (HPS) transfer systems for NASA/JSC (Rockwell, 1984),

while Harttn Marietta ts investigating for NASA/MSFC several scavenging

concepts using the ETand Aft Cargo Car_io_ (ACC_ tank (Hartin Marietta,

1984). NASA estimates cost reductions of $150_-$185 M per year through

recovery and use of the suPp|usand residual propellants remaining in

the ET and MPS (Gtl_nore,. 1984),

Preliminary results lndlcat_ that all OTV missions could be

met using two reu;able OTVs based at the Space Station with propellants

Scavenged from the ET and/or surplus tanks in the ACC. Using NASA,s

nominal Space Shuttle mission model for 1991 to 2000, scavenging from

the £T yields 2,901,00 lbs (1316 HT) of available propellants fro_ 204

STS flights. When the ACC is used in conjunction with ET scawnging,

2,276,00.1bs (1032 HT) of propellant is available over 198 flights (Martin

Hartetta, 1984). The available propell_nt maSs does not account for

losses due to botloff, transfer, and trapped propellants. The mass

of the scavenging system hardware must also be taken into account.

Nevertheless, the scavenged propellant figures should be suf6icient

to supply the required OTV propellants..

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There are, however, technical _ssues which must be resolved

before an ET propellant scavenging system is put In effect. A _reltminary

B A'r T ii L. L. E --.COL. UMBU8
2-6

technical assessment has been performed (Gilmore, 198¢), 6tlmore

indicated a number of areas where special design requirements may be

needed:

e Long,term storage

• Zero-g propellant management

• Zero-g propellant gaging

• Zero leak fluid couplings and disconnects.

Multiiayer Insulation (MLI) was proposed to insulate the storage tanks.

Although MLi has been sucCessfuily used fOr many years, Gilmore felt

that handling and reuse of the material may be a "challenge". The

Cryogenic Fluid Management Facility (CFMF) is a self-contained test

package which will be flown in a series of tests in 1988 to demonstrate

LN2 handling and transfer. Propellant management techniques using the

surface tension Of the fluids are now in use in the RCS and OMS tanks

of the Space Shuttle Orbiter (zero-g conditions, but using storable

propellantS). Critical technology areas cited by Gilmore were zero-g

propellant gaging and "foolproof" zero leak fluid couplings and

disconnects. These are areas in which: the technology has to advance

before the propellant scavenging system could be built. These are not

considered to be as technically challenging, however, as those issues

which face an EML supply system (demonstration of scale-up, etc.).

Another issue which NASA will face if a decision is made to

go ahead with a propellant scavenging system is the questiOn of who

will Day to transoort the propellants to orbit, it has been stated

that "the cost of putting these propellants in near orbit has already

boen paid" (Gilmore, 1984). However, the people paying for the Shuttle.

launch may not wish to pay to transport someone else's propellants.

When the EML supply of OTV propellants to a Space Station

is eliminated, the near-tem mission models are severely reduced. OTV

propetlants represent the primary missiOn of a 2000 IOC EML system,

and, as such, would appear to be the concept driver.

2.1.2 S_ S_tton 5u_Dly

Transportation of certain types of non-sensitive payloads

for use on Space Stations is another portion of the mission model. The

payloads might include food, oxygen and nitrogen for life support systems,

spares, and miscellsneous supply items. The requirements are discussed

in this section.

2.1.2.1 Ltfe Su_

Partlally-closedenvironmenta1 control and life support systems

(ECLSS) are envisioned to be on the _pace Station. Oxygen and nitrogen

must De supplied on a reguiar basis to support the ECLSS systems. ECLSS

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supply requirements were found tn a number of sources (primarily Guy,

1982; General Dynamics, i983; Rockwell, 1983; and Hamtlton Standard,

1983), There was a wide variation in the estimates for required ECLSS

ropellant mas.ses presented tn these documents, _angtng from 640 kg/yr

8-man StatiOn) to 40,000 kg/yr (12-man station) for partially-closed

systems. For thts study, the ECLSS mass requirement is approximated

at 125 kg/person/year between the years 2020 and 2050.

Food requirements for Space Station personnel were estimated

from Rockweli {1983) and Carlisle and Romero (1982) with masses Of 1,18

and 0.62 kg/person/day, respectively. Averaging the requirement values

from these two sources yields 0.9 kg/person/day {food only, does not

include packaging). Zt was felt that only about half of the required

food could be launched vta the high accelerations of an EHi, launch (100

to 1200 g's depending upon method of launch), which led to an estimate

of 0.45 kg/person/day for EML launch.

By summing the food and ECLSS requirements, the total life

support masS requirement ts assumed tO be 290 kg/person/year or

0.8 kg/person/day. Realistically, this number should decrease with

time, Supporting the predicted increasing number of people on orbit

would be prohibitively expensive with only a partially-closed system.

it is 'likely that, before a commitment tO orbiting a large number of

people is made, a completely closed life support System would be

developed. For the purposes of thts study, however, the ltfe support

resupply requirement (less than 1 kg/person/day), was held constant

throughout the 50-year period of study.

2.1.2.2 Spares

It was assumed that only one-third of any spares, required

for Space Station maintenence could be launched from an EHL, because

many of the spares could not withstand the high accelerations of launch.

The spares supply requirement was projected to be about 0.11

kg/person/day. This number .represents one-third of the spares mass

estimated in Rockwell, 1983.

I 2.1.2.3 ACS and Drag Makeup Prepellants

Several concepts are underStudy for use as attttude control

These and d_ag makeup thrusters on the Space Station. concepts include Earth Storable, hydrogen/Oxygen, and reStstojet propulsion systems (NASA,

1984). FO_ simplicity, Boetng's Space Operations Center (SOC) study

' to(Boeing, determine1982)propellant was used resupply, as a reference quantities. for orbit maintenance requirements

The SOC study calculated propellant requirements for orbit

maintenance of a full-up SOC {NASA, 1979), To maintain an orbit at

490 km (265 nml), 2500 ib sec/day impulse Is required. This requirement

means that 1633 kg {3600 IbS) Of monopropellant hydrazlne (Isp Of 240

to 256 secondS) are needed per year, corresponding to a daily launch

13AT'II'ELI. E -- C_l. UIVIIEUI
fl

2-B

rate of 4.47 kg/day for each SOC in orbit, if a cryogenic (LH2/LO 2)

system were. used tnstead, approximately 1000 kg/yr of these propellants

would be required, Reststoj_t Concepts are .included In ECLS$ recycling,.

so no resupply maybenecessary, (

Hydra_lne has approximately the same density (1.008 g/cc at

room temperature) as water, sc I633 kg/yr represents 5 launches per.

for the Earth-Orbital launcher. This function was left off the

m-_-_ion model) however tt could be performed if an _HL system were

available,

2.1.2.4 F.uel Cell Makeup Propellants

Should primary fuel cells be. used for Space Station power,

an EML system could be used. to transport the oxygen and hydrogen necessary

for operation.

From NASA/LeRC Space Station PIR-18 (February 1983), the

reactant consumption of a H2/O 2 fuel cell was given as 0_42 kg/kWh. The

oxygen*to-hydrogen fuel cell reactant ratio is 8.1, which is the same.

ratio as that occurring in water. _or Continuous Operation at an average

of 25 kW power, 91,975 kg of oxygen and hydrogen are required each year

(81,756 kg oxygen and 10,219 kg hydrogen). This corresponds to a daily

requirement of 252 kg/day (less.than one launch per day).

These figures however dO not assume the use of regenerative

fuel cel]s. The regen cells include an electrolyzer as part of the

system to convert the water produced during the process back to its

original LH2/LO 2 form. Using regenerative fuel cells would severe]_

reduce the amounts of reactants necessary for operation.

t

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Since solar a_rays are likely to be selected as the Space

Station primary power source, this requirement has been dropped.

2.1.2.5 Niscellaneous Supply items

The miscellaneous category inciudes such items as personnel

equipment, clothing, hygiene supplies, sh_p stores, EVA supplies, and

maintenance items. The daily requirement was estimated at about

3.22 kg/person/day from Rockwell, 1983, Three-quarters of this figure

(2.42 kg/person/day) was used. in developing the mission models, because

it was felt that some miscellaneous items may not be able to withstand

the high accelerations of an EML launch.

2.1.3 Materials forSpace Processtn_

A major function of a Space Station could oe materials

processing in microgravity. Several experiments are currently In progress

and many more are planned. The 1982 ESRL report (Rice, et. al., 1982)

projected at least one launch per day (650 kg/day) was necessary to

support the materials processing in space activity from 2020 Do 2050

! • ._,

iA1"_St.*-E -- COLUMBWB

t
U 2-g

with _ML launches or' non-sensitive materials. The same requirement

was used as the htgh model tn this s.tudy from 2010 on. Between 2000

and 2010, the requirement increased up to one launch per day. It was

felt that during this period, technology development would stt11 be

occurtng, preceding the full-scale manufacture of products. T.he .low

model assumed one-half the htgh model requirements.

_2,1.4 Traffic Nodell

Based upon the payload requirements set forth in Sections

2.1,1 thrOugh 2.1.3, EML traffic models were determined for the

Earth-to_-_ _-Earthorbtt mission. Table 2-3 presents the masses to be

launched per year for the various payloads. The estimated payload

densities are Indicated under each column heading. These densities

were used to estimate payload mass per EML launch, in Sectton 3.1,5.2

of the ESRL report (Rice, et ai, 1982), p_ioad masses aS a function

of payload density are given for the ESRL projectile (650 kg maximum

payload), Several payload densities and the CorreSpOnding masses are

shown below:

Type of Payload Density (g/cc) Payload Mass (kg),

Water 1.0 320

Life Support/Spares 1.5 440

Materials Processing 2.7 660

The masses were used to determine the number off'lights that are required

per year. The resulting traffic, models for the Earth-orbital launcher,

indicating the number of flights per year for each model category, are

shown in Table 2-4: The total daily launch requirement for the

Earth-orbital. mission is summarized in.Table 2-5_

2.2 Luna r Base Supply itisston

A lunar base represents a logical step beyond the placeme,t

of an Earth-orbiting Space Statton. Much of the technologx gained from

Space Station development would be directly applicable to the build-up

of a lunar base. The base could be used for astronautical and

astrophysical observations, life sciences and ecosystems studies, analysis

of engineering/industrial resources, and colonization. AdditiOnally,

the Moon could be used for military purposes, However, this application

was not considered in this study.. Currently, NASA/JSC is supporting

efforts to analyze the requiren_ntS and development approach of a lunar

base. The purpose of this mission analysts is nOt to Justify a lunar

base nor to specify a lunar base concept, but rather to develop an

understandlhg of typica_ _upport required and to. determine the generic

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2-12

effectiveness of an electromagnetic launcher (EML) in supplying a iunar

base. From recent NASA efforts and past Studies of irunar utilization,

a lunar base characterization, supply estimate, and traffic model are

presented. ..

TABLE 2-5. NUI6ER OF EML LAUNCHES PER DAY FROM

EARTH-ORBITAL LAUNCHER

m

Launches Per Day

Year Low Model High Model

2020 1.4 1,5

2005 2.8 3.9

2010 3,5 5.0

2015 4.1 6.5

2020 5.1 8.0

2025 5.2 9.8

2030 6.3 11.1

2035 6.5 12.6

2040 7.5 14.4

2045 7.7 15.9

2050 8.7 17.0

Fi fty-Year Average 5.3 9.6

....... m

)

C -

2.2.1 Lunar Base Characterization

Recent efforts at NASA/JSC propose a five-phase program for

lunar base development. The first phase consists of preliminary surface

explorations and lunar mapping. Phase 2 represents the initial, temporary

manned base. w_th very. limited research capability and transportation

versatility. Phase 3 consists of a permanently manned base and begins

to exploit the lunar resources (especially the extraction of lunar oxygen

for propellants). Phase 4 is an advanced, fully operational base

including a mOderate manufacturing facility, preliminary lunar

industrial_zatlon, and experimentation of biological life support and

self-sufficiency. Finally, Phase 5 consists of a self-sufflcient base

with an operational Controlled Ecological Life Support System (CELSS)

and a growing manufacturing capability. After reaching the fifth phase

the lunar base would be fully operational as a transportation node,

laboratory, and industrial complex that iS nearly independent of Earth

supply.

Major systems of the lunar base throughout all phases include

the habitat and life support system, power and thermal System,

manufacturing and research facilities, lunar rtranspOrtation and

liA'rTEI, kE --_ COL, UM_uSL

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2-13

manipulator system, and Earth-lo-Moon transportation systems. The type,

amount, and size of the equipment and capabl.litles would change as the

lunar base evolveS. Characteristics ang posslb!e alternatives of the

systems are described below.

2. Z.!, 1 Ltfe Support Systems

The initial llfe suppo.rt system (LSS) will undoubtedly be

an extension of the Space Station LSS using• mechanical and chemical

waste processing systems, Thus, the initial system will be only partially

closed, with C02 recycling and part_al water recovery. As technologies

grow, these systems will evolve Into the biological and ecologljalcycles which are mOre favorable toward self-sufflciency. Table 2 6 depicts

typical LSS supply requirements for the different phases of the lunar

base. AS the LSS matures f_om Phase 2 to PhaSe 3, water recycling is

likely to become more effiCien_ (assumed here to save 60 percent more

water than Phase 2). Further reduction of miscellaneous clothing and

materlals supply would result from increased recycling Capability. The

fourth phase lunar base was assumed to have _ater recycling 90 percent

of the initial LSS, mlscellaneous m_terlals recycling 10 percent of

Phase 2, an initial biological C6LSS to provide 50 percent of the food

nd almost total 02 recovery. A goal of the lunar LSS is the Closed

colOgtCal Life Support System (CELSS) which consists of plant grouch

providing C02/02 exchange and food. The last phase of the lunar base

claims self-sufficiency'; however, realistically, vttamlns and small

food supplements may amount to about 3 percent of the tnitlal food

resupply. In addltion to the growth of tne habitat LSS, the

exlravehlcular activity {EVA) will evolve from Shuttle and Space Station

capability to support longer duration EVAs on the lunar surface.

TABLE 2-6 LIFE SUPPORTRESUPPLYESTIMATES REQUIRED

FOR LUNARBASE EVOI._.ION (KG/PERSON/DAY)

..... m i r rl if ,

02

H20 (drink)

H20 (wash)

Food

Mtscel 1aneous

'l m

TOTAL

III I tlf • II , I I ' I L I ml H I f I k .... , , ,I

i m

Phase

1 2 3

i

-- 1. O0 O.20.

-- O.50 O.20

-- 2.60 1.04

-- 0.70 0.70

-- 0.20 0.15

i | | i

-- 4.90 2.12@

i

O. 50

O. 26

0.35

0.02

0.68

5

ll

l.ll,

II

0.02

i

0.02

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2-14

A Boeing Aerospace study has estimated that a lunar CEL_S

that contributed 50 percent of the crew diet wou)d pay off in 5.5 years.

The remaining SO percent would have to be supplied as packaged food.

A CELSS supplying 97 percent of the diet _3 percent suppiemented by

vitamins and condimentS) would beCOme economically feasible in seven

years (Spaceflight, December, 1983).

2,2.1.2 ..Power S:fStElS

The power and thermal systems must supply adequate support

for both the 14-day cycies of day and night. The temperatures may range

frnm about 400 K to 120 K for these two different environments. Power

requirements are estimated at the lO0-kW level for the tnttial base

and grow by a factor of 10 for each of the subsequent development phases

mentioned above (NASA/JSC, i984). Thus, by the ftfth phase where the

lunar baSe is Self-sufficient, up to 100 HW of power may be required

(NASA/JSC, lg84c). Major uses of po_er in the latter phases of the

lunar base tnclude lunar resource processing (especiaily lunar liqutd

oxygen_LLOX) and environmental control of large volumes (for C(LSS),

$01ar l, nuclear, and regenerative fuet cell systems have been considered.

The preferred systems for power generation may initially be

the solar array with regenerative fuel cell supplement during the night

cycle 1. AS the base evolves and power requirements increase to provide

power for lunar resource processing (such as LLOX), the higher-capacity

continuouS nuclear power generators Willl be much more efficient land

effective. In these latter pnases, a combination of both solar and

nuclear may be the best choice to provide steady and peak power

requirements, las well as system redundancy.

2,2,1,3 Lunar Transportation

Transportation requirements on the lunar surface will initially

be small as preparation of the operating base will be the primary mission.

Equipment may be limited to a short distance transporter for crew mobility

and a mechanized shovel to dig and bury habitats. However, as the

misstons expand, longer distances and longer duration travel will be

required, Rockwe)l estimated an averagelunar sortie mission of 45 days

(North American Rockwell, 1971). Thus, equipment required in latter

phases may also include a long distance transporter with large payload

capability and a transportabie, temporary habitat for remote operations.

In addition to the actual transportation equipment, the crew

would need machined assistance for manipulation of bulk supplies and

equipment (such as solar arrays), and for assembly of piece parts into

larger systems. The manipulation may be done with crane-type robotic

arms that can be adapted with end effectors for grappling, lifting,

and digging.• These manipulators and devices may be attached to vehicles

as appropriate for a specific mission.

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2_2.1, _ Transpo,-tatton

The most costly operational expense of a lunar base is for

transportation from the Earth to the Moon. The lunar base may never

become totally self-sufficient and would certainly not be close to total

self-sufficiency for many years after the inittal base, Thus, early

in the program, a convenient, inexpensive ,,_thod for iogiStiCs transport

would be beneficial, This section briefly identifies characteristics

of the Earth-Moon transportation systems.

Three major classes of cargo existz crew, equipment, and

bulk materlal. Assuming no reduction in crew requirements On the MOon,

the flow of crew would always be either predominantly from Earth to

Moon or with equal flow in both directions. The flow of equipment would

nearly always flow from Earth to Moon with little .or none returned to

Earth. The initial flow of bulk suppiy materials such as fuel, food,

and other consumables would initially be from Earth to Moons however,

after useful operation of the lunar base is achieved, the flow direction

may be equai in both directions or may become Moon-to-Earth dominated.

Supply of mateHals is deferred to discussion in the next section.

Earth-Moon transportation alternatives must operate in three

different orbit regimes, namely Earth orbit, Earth-Moon transfer orbit,

and the Lunar orbit regimes. Transportation systems exist that treat

each regime separately (segmented systems) and that combine two or more

regimes (nOn-segmented systems). Segmented transportation systems treat

these regimes with specialized vehicles. A key example of such a system

is a Shuttle-OTV-Lunar Lander alternative. The Space Shuttle operates

most efficiently between the Earth's. surface and. LEO, where lligh thrust

is required. The OTV operates most efficiently in Orbit, never having

to land on a planetary surface. Similarly, the Lunar Lander iS optimized

as a launcherlander in a 1/6-g environment. The specialization of

vehicles increases transportation efficiency and thus, the mass of payload

capability. However, these vehicles must dock and exchange payloads.

Although this type of operation seems "natural" with the operational

STS, upcoming LEO Space Station, and LEO-to-GEO OTV, three vehicle systems

and additional orbiting facilities are required. Many alternatlves

exist for each of the vehicles including use of expendable launch vehicles

(ELVs) versus reusable vehicles. Table 2-_7 identifies the major

alternative vehicles throughout these regimes.

t

The non-segmented transportation systems ideally combine all

the transportation regimes, but usually only combine two of the three

regimes. The EML concept is a primary example of a high-impulse system

that can combine the Earth-Moon transfer orbit with either an Earth

or lunar launch. Given certain launch constraints, an EML payload

launched from Earth may traverse to orbit the Moon or to impact the

Mooa's surface with some guidance propuiston. Similarly, payloads can

be launched from the lunar surface at high velocity to intersect Earth

orbit or other orbits. With sufficient onboard propulsion (approximately

3 to 4 kin/S), bulk material could be soft landed on the Moon directly

8AT.TILLE -- COL.*.JMBUI

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'1

2-16

from Earth, or could be placed in an orbit about theMooti from the Earth.

HOwever, Since very-high accelerations (thuusands of g's) occur during

launch, payloads are restricted largely to bulk supply material. Thus,

the disadvantages of the EML_type transport system are the lack of

pointing capability (this my also be overcome wtth increased onboard

prOpUlsion) and the requirement of supplemental transportation for crew

and special equipment.

t

TABLE 2-7. ALTERNATIVES TO VEHICLES IN THE EARTH-MOON

TRANSPORT REGIMES

Earth to LEO

(Vehicle)(a)

LEO-Lunar Orbit

(Prop_llant Type)

Lunar Orbit-Lunar Surface

(Propel;ant,Type)

Shuttle

ELVs

HLLV/ULV

EML

HybridEML

LHz-LOX (OTV) LH2-LOX

Storable PrOpellantS Storable Propellants

Lunar-Derived Propeilants(b) Lunar, Derived Propellants(b)

Nuclear-Ion Drive

Solar Electrtc

(a) ELV = Expendable Launch Vehicle; HLLV = Heavy-Lift Launch Vehicle;

ULV = Unmanned Launch Vehicle.

(b) Lunar-derived propellants include 02/SiH 4, 02AI powder+and 02 thermal

propulsion systems all using lunar-processed oxygen. (presented

in NASA/JSC, 1984)

2.2.2 Lunar Base Sul_enari0____!s

By the year 2010, a permanently-manned lunar base may be

operational and would require supplies from Earth such as crew, life

support consumables, propellants, equipment, and bulk materials. Of

these supplies, an EML could supply the life support consumables,

prOpellants, and bulk materialS, Crew and most equipment are too

sensitive to withstand the high accelerations of "launch (2500 to

3600 g's).

Two lunar logistics scenarios (an OTV-based scenario and an

EML-based scenario) were developed to provide insight to possible

advantages and disadvantages of the use of an EML. Figure 2-1 shows

the proposed traffic flows between the Earth (E), Earth space station (S),

a lunar space station (L) and the Moon's surface (M). The baseline

OTV lunar logistics scenario would use conventional hydrogen/oxygen

prop, lston between all transportation nodes (STS, Unmanned Launch

Vehtcle--ULV, OTV, and lunar STS--LSTS). The scenario which used an

BATT¢I. Lm -- cot. uMlauS

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equatorial-based EML to supply materials to the lunar and Earth space

sta_,ions would reduce the number of STS, ULV, and OTV flights. The

materials launched by ENL would include life support consumables,

propellantS, and other materials. As in the Earth orbit supply mission,

hydrogen and oxygen propellants would be transported in the form of

water to increase payload density, and simplify storage and handling

over cryogenic propellants. Propellant tsSues and tradeoffs are

dtScuSSSed in Section 2.6. The STS, ULV, OTV, and LSTS systems would

still deliver crew, equipment, and other payioads sensitive to high

acceleration.

in the baSeline lunar-logistics scenario, the STS would consist

of the partially-reusable Shuttle with 29.5-NT capability to. LEO and

15,4-NT return capability. The ULV was assumed to be a Shuttle-derived

vehicle using existing STS Solid Rocket Motors, an ET, reusable engineped,

and a cargo volume protected by a large ShrOud. The ULV was assumed

to transport 68 _ of payload to LEO in a_ single flight, The

Earth*orbiting space station would provide a base for OTV refurbishment

and refueling, a stOrage depot for propellants and materials, and a

_elay station for ct, ew and equipment .between the STS/ULV and OTV systems.

The OTV was assumed to be a 17-MT cryogenic bystem, with a 480-second

spectf.ic impulse and 8-MT payload capability. This assumption is

reasonable for the 2010 timeframe and is compatible with Current

space-based OTV concepts. The lunar orbiting space station may be derived

from.the Earth space station to provide similar basing, cryogenic storage,

and rendezvous capability for the OTV and the LSTS systems. The LSTS

would be optimized for lunar-surface landing and return to the lunar

orbiting station. For compatfbility with the OTV, LS.TS was assumed

to be a 15-NT cryogenic, vehicle with a specific impulse of 480 seconds

and payload capacity of 8 HT. Both the OTV and LSTS would accommodate

manned modules for crew transportation.

The EML-based lunar logistics scenario contains the same four

nodes: Earth, Earth space station, lunar space station, and a base

on the Moon's surface. However, no UL.V fltghts are Included for

Earth-to-space station (ES) transportation. The equatorial EML system

would accelerate a 500-kg payload at 3600 g's, at a 12-km/s launch

velocity to the vicinity of the lunar space station, and would accelerate

a 650-kg payload at 2500 g's to the vicinity of the LEO Space Station.

For each case, additional on-board propulsion is required for orbit

circulation at the stations. The preliminary EML concepts for this.

mission aredtscussed in Section 3.1.

2.2.3 Lunar Base Supply Mission NOd_l .....

Two lunar base timeframes were addressed for each. scenario.

The first, at year 2010, would consist of a modest base and lunar space

station with pennament habitation for 12 crew, three of the crew being

rotated every 90 days. Logistics for lunar base build-up was not •

analyzed. The life support system (LSS) was assumed to be a growth

of the Earth space station LSS so that food and a portton of consumed

mAi"'i'msm,m.mE -- cm,-uM'_um
_

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Ii)

i'i

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2-19

L,

water would be a supply requirenmnt.: Propellants for the OTV must be

supplied, to the Earth space station for the trip to the Moon (by either

th_ ULV or the EML) and to the lunar space _tation for the return trtp

(by either the 0TV or the EML), P_,opeliantS for the LSTS must also

be supplied (by either the 0TV Or the EML) to the lunar space station

in addition to equipment needed for build-up of base capability and

self-suffic_ency (by the OTV only).

A self-sufficient base with c)oSed life. support system was

assumed fo_ the year 2030. A crew of 48 has been assumed with crew

rotation of three persons every 45 days (2-year stay time per person).

OlV and LSTS propellants must be supplied, along with a Small amount,

of equipment and materials to maintain the base and supplemert lunar

resource processing.

The supply categories have been categorized aS the crew, life

SuppOrt consumables,, equipment., materials (bulk), and OTV and LSTS

propellants. Tables 2-8 and 2-9-summarize the total annual lunar base

supply requirements for each scenario in the years 2010. and 2030. The

derivation o_ these requirements Is discussed below.

' .,!

(

I '

2.2,3.1 Crew Supply Requirements O,.rtng Transport

Crew rotations must be accomplished with conventiona| space

transportation because of the high accelerations of EMLs. A five-day

trip to lunar orbit, a one-day change.out in lunar orbit, and a five-day

return_ trip were assumed for a crew.of three. The STS would not require

additional equipment for manned transport. However, a 5.1-MT manned

module for the OTV. and a 2.4-MT manned module, for the LSTS would be

required to provide life support and a habitat for three crew members

(Roberts, 1984). Each p_son was estimated to weigh-136 kg, including

clothing and peripherals. For the year 2010, four rotations were assumed

each year, giving each c_.ew.member a one-year stay on the Moon. During

the transit, life support consumables would be approximately 4,9 kg

per man per day (see Table 2-6).

For the year 2030, eight rotations of three crew members would

give each person two-year duty cycle. The same manned modules were

i assumed; however, the transport consumables would, likely be reduced

a

to approxlmately 0.7 kg/person/day (see Table 2,6).

2,2,3.2 .... Lunar.L!f _ Su_pO.rt Consumables

The life support system of the initial lunar base. would most likely be an extension of the Earth space station having a closed C02.

cycle and partially,-closed water cycle. Table 2-7 lists the estimates

of life support requirements for the various phases of the lunar base. The year 2010 capability would be best represented by Phase 3 at

2.3 kg/person/day for food, water, and miscellaneous supplies. Total

BAtTelLE -- COLUMBUS

¢ : ,..;!i., _ .
TABLE 2-8.

2-20

ANNUAL LUNAR BASE SUPPLY REQUIRENENTS (Kr/YR) ZN 2010(a)

Transport Regime( b )

Baseltne Lunar

Logistics Scenario

EML-Based Lunar

Logistics Scenario

ES 1444.7

SE 1.6

SL 304.8

LS 22.7:

LN 78.3

RL 11.3

ES'

EL'

m. , J ,in f . • rl i , ,

10.0

1.6

31.0

22.7

29.3

11.3

162.0

136.4

(a)indtvidual supply items are discussed in Section 2.2.3.

(b)Transport regime aceonyms are defined in Figure 2-1.

TABLE 2-9. ANNUAL LUNAR BASE SUPPLY REQUIRENENTS (RT/YR) IN 2030(a)

Transport Regime( b )

Baseline Lunar

Logistics Scenario

EHL-Based Lunar

Logistics Scenario

ES 2450.8 7.4

SE 3.6 3_6

SL 532.2 48.8

LS 48.8 48.8

LH 127.8 26.4

HL 26.4 26.4

ES'

EL'

309.3

241.2

(a)Indivtdual supply items are discussed in Section 2.2.3.

(b)Transport regime acronyms are defined tn Figure 2-1.

8A'r'f'|LI. | -- COLUMBUS

, ,. •

b
2-21

year-2030 capability • (Phase 5) was assumed to have a nearly closed life

support System with negligible supply requirements (see Table 2-6).

2- 2.3.3 Equipment

Equipment supplted to the lunar base would support maintenance

and growth of various systems and capabilities. NASA has estimated

that approximately 40 MT of equipment would be required to develop an

advanced base from the tnttlal, permanently-manned base (NASA/JSC, 1984c).

Assuming equipment would be transported to achieve an advanced base

in five yea,-s from 2010, the annual supply of equipment would be 8 MT/yr.

Spares must also be delivered; however, the intttal amount of spares

would ltk_ly be small and increase as equipment ages. Thus, 8 MT/yr

has been assumed to be constant throughout the 2.010 to 2030 timeframe.

BeyOnd the 2030 ttm._rame, a stmple estimate of 4,0 MT/yr waS assumed

based on the likelihood of increased automation and larger tasks on

the Moo,.

2.2.3.4 Materials

Bulk material supply to the Moon may include metals, gaSes,

polymers, and organic materials that may be processed On the Moon or

lunar space station or may be combined with lunar resources to create

useful (and posstbly marketable) products. Lunar resource materials

may become valuable in LEO so that material transfer from the Moon to

Earth orbit is a possibility. Such a production process was not seen

to be operational in: the inttial 2010 timeframe; thus, no materials

are transferred to or from the MOon. However, an advanced i unar base

would likely require material tranSportation. An annual materials supply

of about 1 MT may supplement the production of metallic alloys,

lunar-derived oxygen, and cement-type products. If materials-processing

plants become economical on the Moon, this number could easily increase.

Mass flow from the Moon to LEO was estimated at 4.0 HT/yr by 2030.

2.2.3.5 Propellants

Major uses of propellant include the OTVs, operating between

Earth orbit, and lunar orbit and the LSTS, operating between lunar orbit

and the lunar surface. Lunar surface vehicles, the lunar space station,

and regenerative fuel cells consume only a small amount of fuel, and

were considered to be negligible _n the Overall mission mOdel.

The LSTS w,Julddeliver payload to the Moon's surface including

the propellant required for the return trip. This propeilant requirement

is independent of the lunar logistics supply scenario. Tables 2-8 and

2-9 show the propeliants that must be carried as payload for the years

2010 and 2030 to supply the required crew, consumables, equipment and

materials. The LSTS is assumed to be a 15 MT vehicle (dry) capable

of delivering 8 MT with cryogenic propulsion system of 480-second specific

impulse. The propellants for the LSTS would either be supplied at the

lunar space station by the OTV or the EML system a 30 percent packing

factor was assumed for propellant delivery in the OTV and ULV/Shuttle.

|A1"1"|1.1_ | -- C(=LUMI'iUlt
2*22

The 0TV incurs about a 4-km/s veloctty increment for the trip

tO the Moon and abOut 1 kl./s with aerobraking capability on the return

_TV t s percent of trip. General the propei lant weight 60-70 the total

wight of an ; thus, a great deal of propellant Is required to and

from the Noon. For the baseline iunar 1oglstlcs scenario, the propellant

supply at the lunar space station (for OTV and LSTS use) would be

transported from the Earth Space Station to the Noon by the 0TV. These

propellants delivered would have been supplied to the Earth Space station

by the ULV.

Introducing an EML into the transportation system would relieve

the amount of OTV and SYS/ULV-deilvered propellant requlred. By

transporting water {propellant) to an Earth space station and lunar

space station {the OTV, LSTS departure nodes), and eleCtrolyZing it

into hydrogen and oxygen and llqulfylng, the number Of OTV flights needed

to support the missions is dramatically reduced. Even after taking

out the 30-percent packaging •factOr shown inTables 2-8a and 2-ga, more

than 80 percent of the baseline propellant supply to the Earth space

station would be saved using an EML. More than 30 percent of baseline

propellant delivery would be saved with EML propellant delivery to a

lunar space station.

t

li

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I

2.2.4 Trafftc Node1

Tables 2-10 and 2-!1 present the number of flights required

for both logistics scenarios (see Figure 2-1). A 100 percent load factor

for each vehicle was used, assuming the following maximum payloads:

Shuttle 2g. 5 HT

ULV 68.0 MT

OTV 8.0 MT

LSTS 8.0 HT

EML 0.5 HT

The tables show the Earth to Earth space station traffic (ES)

in Shuttle flights. The Shuttle flights would be required for manned

transfer; thus, at least four STS flights would be needed in 2010 and

at least eight in 2030. If arl Unmanned Launch Vehicle (ULV) were

available, 20 and 33 ULV flights may be used the years 2010 and. 2030,

respectively, to supplement the required manned STS flights.

iI

Introduction o.f the EML into the traf_flcmodel would eliminate

the need for a heavy tra,sport vehicle between Earth and the Earth space

station {ES). The STS would supply only those flights wIlere

acceleration-sensltive payloads must be transported. The EML would

aiso reduce the number of OTV flights by approximately 20 percent. The

entire propellant supply could be transported with 577 and I000 EML

launches per year for the years 2010 and 2030, respectively. With the

EML-supplted prope.liant reserve at the Earth and. lunar space stations,

'"I

(

8ATTIEL. L.E -- CDL. tJMIlltJB

)
I 2-23

G

G

TABLE 2-10. PAYLOADNANIFESTING, I_ 2010 (NURBER

OF VEHICLE FLIGITT$)taJ

Transp9r_

Regime_o)

Baseltne Lunar

Logistics.Scenario

EML-Based Lunar

Logistics Scenario

ES 48(1.0). 1(0.98) 1(0,29), 3(0.02)

SE 4(0.03), 45(0) 4(0.03)

SL 22(1.0), 1(0.20) 3(1.0), 1(0.9)

LS 4(0.35), 19(0) 4(0.71)

LM 3(1.0), 1(0.65) 3(1.0), 1(0.65)

RL 4(0.35) 4(0.35)

ES' .... 324(1.0)

EL' -- 253(1.0)

(a)Table asSumes vehicles can be loaded to I00 percent of capability.

Number in parentheses indicates the fractlon_f vehicle capability.

(b)Transport regime acronyms are defined in Figut_o2-I.

TABLE 2-11. PAYLOADI@3IFESTING.I_ 2030 (NIJNBER

OF VEHICLE FLIGHTS)(a)

Transport

Regime{ b)

Baseline Lunar

Logistics Scenario

EML-Based Lunar

Logistics Scenario

ES 83(1.0), I(0.07). 1(0.22), 7(0.02)

S£ 8(0.03), 72(0) 8(0.03)

SL 38(1.0), 1(0.15) 1(1.0), 1(.0.9), 6(0.7)

LS !(1.0), 1(0.9), 6(0.7), 31(0) 1(1.0), 1(0.9), 6(0.7)

LM

ML

1(0.91), 7(0.35)

1(0.85), 7(0._5)

i(0.85), 7(0.35)

1(0.85), 7(0.35)

ES' -- 61S(1.0), 1(0.6)

EL' -- 4a0(I.0),I(0.4)

i | i l " i -' ii _i i i ' ii i i m

(a)TaBle assumes vehicles can be loaded to 100 percent Of capability.

Number in parentheses indicates the fraction.of vehicle capability.

(b)Transport Pegl_ acronyms ,are defined _n F|gure 2-1.

BATYELI.| -_ P--_LU,_liUli

¢ 'ii.,::!'.....
2-24

the OTV would become more efficient as tt would not have tocarry return

propellants and tankage for that propeilant.

2.2,5 ENL Lunar Base Su 1 Advantages and Dtsadva_

Because an EML system could be_ an efficient method of

transporting water, many advantages exist in supplying water in space

for propellant production over conventional supply. However, several

disadvantages are also introduced by using the EML.

The majo£ advantage of the EML supply relates to a dedicated

propellant-supply system that lessenr the STS, ULV and OTV fliglit and

maintenance burdens. The OTV would be ltghter and operate more

efficiently in delivering the desired cargo (crew and equipment). At

the lunar station, the EML-Supplied propellant would eliminate the need

for lunar-derived prOpellantS (Such as oxygen, see Section 2.2.6). Also,

the EML projectiles may be taken to the Moon'S surface and used as

pre-processed rawmatertals.

The advantage of supplying H/O in the form of water is that

large cryogenic storage is reduced, handling is easier, and boll-off

losses from dewar storage tanks is nearly eliminated. Cryogenic

propeilant storage is kept to a minimum because only enough H2/02 to

fill the nex_ OTV or LSTS vehicle would need to be available. The

disadvantage to this is that electrolysis production equipment must

be OnbOard the space stations. The energy for this production may be

obtained from solar or nuclear power. Other d_sadvantages inherent to

the EML system are the launch constraints caused by fixed launcher tubes,

Also, She htgh accelerations do not allow delivery of crew or sensitive

equipment. Finally, the EMLS have relatively small payloads, which

create additional retrieval and handltng burdens at the space stations.

4

2.2.5 Lunar Ba_

Many Issues and trades have been identified that need to be

studied In further detail, Major issues involve the OTV payload

capabilities, space station capabilities,.type and source Of propellants,

amount of crew transportation needed, and the amount of materials delivery

to and from the Moon. In addition to these issues, EMLs may be placed

on the 1.unar surface and used to supply 02 to the transportation nodes.

The effects of a larger OTV may reduce the number of OTV

flights; however, the propellant requirement will likely remain the,

same order Of magnitude. An OTV sized for payloads larger than the

manned module could transport large amounts of bulk materiat or propellant

in addition to the manned mOdule.

The capabilities of the Earth and lunar Space stations wi.l

affect the economic savings apparent inan EML supply System, The level

of on-orblt• manpower, facilitles, and resources needed to refuel and

refurbish OTVS is an important cOnSideration.

8A'fTiRL. L.£ -- ¢OL. UMBU8

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2-25

The propeliant alternatives (especially for the LSTS) are

cryOgeniC and storable. Cryogenic propellants would l_kely be Sto_ed

in large dewarS and house re-liquiftcatton systems. This type of facility

could be costly to maintain. An alterna.tive is to use Storable propellant

systems (such as MHH/NTO), especially for the LSTS, to reduce the storage

and handling burdens. However, these propellants are not as efficient

{specific impulses at 300-350 seconds) and therefore larger masses of

propellants would be required. An alternative to cryogenic Storage

is to transport water, s_ore it and electrolyze and liquify the hydrogen

and oxygen where .needed for each flight, Although the electrolysis

and liquiflcation equipment add additional cost, the water is easily

and safely stored and the power may be provided from a solar oK nuclear

source.

An alternative to propeilant supply f_om Earth is lunar-derived

propellants. Oxygen is very plentiful on the Moon, as it is available

in oxl)es of all major metals. Hydrogen is not abundant. An estimated

5500 m_ of lunar soil must be processed to obtain I HT of hydrogen

{hydrogen is trapped in the soll from the solar wind). Hydrogen would

have to be suppll)d if large quantities of propellant are needed.

Increasing the oxidizer-to-fuel ratio from 6-to-I to 8-to-l may prove

advantageous. Alternative propellants have been suggested that use

more of the lunar materials and reduce or eliminate the requirement

for hydrogen. Propellant systems such as 02/SiH4, 02/Al powder,

02/AI/HTPB hybrid, and 02 electric systems are possible, but have lower

performance and some production problems (NASA/JSC, 1984_). All of

these propulsion alternatives have an effect on the mass of lunar supply.

An EML could be used tO supply materials tO construct the

lunar base, including the lunar space station. Piece parts of various

systems ma_ be delivered by EML and then asSembled in lunar orbit or

taken, to the surface.

2.3 Solar S_stem Escape Mission

Launching high-level nuclear waste to solar system escape

velocities was MisSion A of the Reference Concept in the Battelle 1982

ESRL Study. The study considered complementing the U.S. nuclear waste

mined geologic repositories by launching domestic high-level waste (HLW)

out of the solar system. Based upon DOE and NASA funded Studies, it

was determined that the HLW mission would probably not be a driver for

ESRL development (Rice, 1982). This disposal approach still left a

majority of the nuclear waste to be buried in the repositories. An

alternative to this concept was investigated during the current study.

By launching HLW and transuranic (TRU) wastes via solar system escape

trajectories, the need for mined geologic repositories wOuld be totally

eliminated, thus significantly reducing the cost of ground-based disposal

systems.

"k

t:

2.3..1 ESRL Mission A Summary.

The nuclear waste disposal concept lnvestlgaLed in the ESRL

study looked only at launching the high-level waste out of the solar

iATT{LLE -- taLUMBUi
2-26

system, while leavtng the lower-level and short-lived high-level (Cs

and Sr) nuclear wastes in underground burtal sites (see Figure 2-2).

This was an: effort to compare this methOd with the so-called "Standard

Space Disposal" concept, sponsored by NASA Marshall Space• Flight Center.

The standard Concept disposed of the HLW in a heliocentric orbit midway

between Earth and Venus,. at 0.85 A.U. In this Concept, two spherical

nuclear waste packages are launched to low-Earth orbit by a

Shuttle-derived vehicle where they are docked wi_h an orbit transfer

vehicle and a solar orbit insertion stage. The dual propulsion system

combination delivers the nuclear waste to its final orbit around the

Sun (see RiCe et al, 1982, fOr further details regarding the standard

space disposal concept).

The ESRL study considered launchlng the HLW using a railgun

system (see Figure 2-3). The ESRL Mission A projectl.e contained 250

kg of nuclear waste in cermet form. After reviewing the availability

of high-level nuclear waste for. space disposal (Rice, et al, 1982),

the traffic mOdel was established, with a requirement of two launches

per day to dispase of all the.HLW.

2.3.2 Transuranic Waste

Transuranic (TRU) waste is that radioactive waste which is

contaminated with transuranic alpha,emltting radlonuclides or U-233

at levels gceater than (traditionally) 10 nano-Curies pen gram (10 nCi/g

or 10 x 10-_Ci/g) of waste. In the early 1970's recognition was given

to the higher risk to human health posed by the long-lived alpha-emitting

radionuclides and the Atomic Energy Commission began segregating and

storing TRU waste for future d_sposal in a federal repository. Gradually,

the commercial shallow-land-burial disposal facilities .for radioactive

wastes ceased accepting TRU wastes.

The value of 10 nCi/g, was established on the basis that it

was similar to the concentration of naturally occurring levels of radium

in soil. Recently, promulgated regulations by the NRC will permitburial

of wastes up to !00 nCi/g, if properly packaged and buried in licensed

burial grounds, This increased activity level will permit more accurate

segregation of TRU and non-TRU and will result in some reduction in

the volume of TRU waste requiring disposal.

In commercial nuclear power activities, TRU waste is generated

only by reproceSsing spent fuel and by refabricating the recovered, fuel

(plutonium and uranium) into new reactor fuel. Transuranic elements

do not occur prior tO the fissioning of uranium in the power reactor.

Defeflse activities conducted by DOE produce large volumes of TRU wastes

which are planned to be disposed of in the Waste Isolation Pilot Plant

now under construction in New Mexico.

2,3,2,1 Reprocessing and Refabriration Wastes

During fuel reproc_ssing, spent fuel rods are chopped into

small pieces arid the. fuel is dissolved from the cladding by an acid

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solution. The uranium and plutonium are extracted from the actd solution

by an organic.solvent and subsequently separated into individual streams

for: eventual recovery. The residual acid solution contains the fission

p_oducts and remaining traTJsurantc elements and is, by definition, highlevel liquid waste. This very radioactive waste stream, although a

TRU waste, must be treated and disposed of by special regulations being

prepared by the NRC. The other TRU wastes generated in fuel reprocessing

activitieS, which are the subject of this analysis, are Contaminated

with much lower levels of radioactive materials. The TRU wastes from

reproceSsing fall into the general categories, of hulls, filters, process

trash, fa_led equipment, fluorinator ash, ion exchange resins, silica

gel, degeraded solvents, and concentrated, liquids_.

TransuraniC wastes in a fuel refabricatlon plant occur from

materials contaminated with plutonium which has_ been recovered from

spent fuel and which is mixed with urani.um to produce mixed oxlde, fuel.

The fission product contamination is very low in ,this waste and packages

will have low surface-dose rates, There are f_ve major categories of

this waste: fllte_s, process trash, failed equipment, process liquids,

and incinerator ash. The assumed methods for treatment are similar

to tllosefor the same categories of waste from reprocessing,

2.3.2.2 TRU Haste Available for Space Disposal.

The treated waste volumes and masses for each category of

TRU waste are. given in Table 2-12 expressed• per metric ton of heavy

metal of fuel repr_cessed (MTHMp) or in Table 2-13 as metric ton of

heavy metal of fuel refabricated (MTHMf). It is assumed that I MTHMf

results from 5 MTHMp.

The data are calculated from results in DOE/ET-O028, Technology

for Commercial Radioactive Waste Management, the technical basis for

the. Generic Environmental Impact Statement (GELS). The waste volumes

and masses are taken directly from the reference source which assumed

one reprocessing plant operating at a capacity of 2000MTHM/yr supplying

plutonium to one refraction plant operating at _0 MTHM/yr.

2.3.3 Total Mission Summary

The total nuclear waste disposal mission consists of the

available TRU waste added to the high-levei nuclear waste available

for space disposal. The mission summar_ is presented, in two tables.

The total annual volumes and masses o_ HLW and TRU waste for space

disposal are given in Tables. 2-14 and 2-15 assuming the spent fuel

reprocessing projection of Rice, etal, 1982. This projectiofl assumes

that fuel reprocessing and refabricating capacity and waste treatment

capacity are available to mee_ the predicted rates.

The traffic model, which indicates the required number of

launches per year to dispose of the nuclear waste, is presented in

E) AT_rELL. i -; co_uM|ue

¢
2"30 "

TABLE 2-12, TRU klAS_ FRQI( FUEL_REPROCESS[NG

Category m3/HTHHp

' I i ' ii i '1" " i

l

kg/HTHHp

I' I Jr[ r II'T i " ,IM'. lilt r

HultS 0,076 266

FtlterS 0,040 i2

Process Trash (inciuded In failed

equip_entand ash)

0.045 225

0.034 51

(2.8 included in ash)

(included in ash)

(Included tn ash)

(included tn ash)

0.094 24

O. 289 578

Failed Equipment

(and noncombustlbie waste)

Fluo_fnator Ash

Sillca Gel

Degraded Solvent

Ion Exchange Resins

Concentrated Liquid Residue

Incinerator Ash

TOTAL

ii if i lille L i[" ,-- . , -'_ r i

TABLE 2-13, TRU WASTEFROMFUEL REFABRICATION

Category m3tHTHHf kg/HTHHf

Filters

Process Trash

Failed Equipment

Ltqutd Residue

Incinerator Ash

TOTAL

Total

Total (both plants) _

r I " 'h

I

0.025 16

Combustibles to incinerator

NoncombuStibles to failed equip.

O. 05 250

0.045 89

0.022 5.5

O. 142 360

m3/HTHHp kg/HTHHp

0.03 72

O. 32 550

.... , m

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TABLE 2-14, VOLUE OF U.S. COMlqERCIAL MASTE AVAILABLE

FOR SPACE DISPOSAL

+

,m rI + _

Year Haste

Avat 1able

: '1 i

1989

1g90

1991

1992

1993

1994

1995

1996

199._

1998

1999

2000.

• rl ' f 1"jr _" it + r 'L_E .............. _ mlJ " i " ....

HLW V91un'M_ TRU VQlume Total Velum

(m3) (m3)

..... r mr ) I

43 1885 1928

13 576 589

15 672 687

18 778 796

20 886 906

23 1008 1031.

25 1107 1132

26 1206 1232

29 1267 1296+

30 1338 i368

32 1389 1421

33 1434 i467

307• 13,546 13,853

• jrl _l I _+_' rr ' • t '_ if" i ' r .

i " i- r

TABLE 2,-15. HASS OF U.S. COH_RCZAL MASTEAVAILASLE

FOR SPACE DISPOSAL

• _, . ............. , +, T i t I i i IL

Year Waste HLW Mass. TRU Mass,

Available (RT) (HT)

i| i = •

1989 279 3828

1990 85 1170

1991 100 1365

1992 115 1580

1993 131 1800

1994 149 2048

1995 164 2249

1996 166 2450

1997 188 2574

1998 198 2717

1999 206 2821

2000 _21_2

1993 27,514

Total Hass

(HT 

4107

1255

1465

1695

193i

2197

2413

2616

2762

2915

3027

3124

29,507
r "=

2-32

Table 2-16. Twenty-six launches per day are required to dispose of

the HLW and TRU wastes In the yearsY020, to 2050.

TABLE2-16. NUCLEARWL$' DISPOSALIN SPACEI ¢AFFICMODEL

(HLWANDIRU taSTES)

J ii " = " i z

Launches Per Year

Year --" HLw TRU Total

2020 730- 8760 9490

2025 730 8760 9490

2030 730 8?60 9490

2035 730 8760 9490

2040 730. 8760 9490

2045 730 8760 9490

2050 7'30 8760 9490

| r i -, • , , ,, i1 i

I '

it Should be noted that the mission model sumlarized above

isbased upon a risk assessment done for NASA in 1981-1982 (Rice, Denntng=

and. Friedlander, 1982) which assumed a nuclear powe_ capacity of 200

GWe in the year 2000. Current projections indicate that 164 GWe is

a more likely figure (U.S, Department of Energy, 1982), and wou]d reduce

the total mission projections to 0.82 (164/200) of the original estimates.

2.4 Earth Escape MtsStOn

The Earth-escape mission model was developed for planetary

exploration missions in the thirty-year period, 2020 to 2050. In 1980

NASA appointed a Solar System Exploration Committee (SSEC) to investigate.

possible planetary missions form 1988 through 2000. Theprimary purpose

of this committee was to recommend a program strategy to -evitalize

the U.S. planetary science program. Morrison and Hinners (1983) sunvflarize

the first task of the SSEC which was to develop a "core program of lowand moderate-prlced missions". The core program excludes the larger _

Viking.-type mission which many U.S. scientists feel are high-priority

mission, but which the SSEC felt were too expensive to be included in

the basic core program. Nevertheless, the SSEC program consisted of

14 missions to be launched between 1988 and 2000, which corpesponds

to a launch rate of greater than I per year. These missions include

one lunar, three Mars, two VenuS, and three Saturn (including Titan)

missions, and four to study the cometsand asteroids.

The ESRL Study estimated planetary launches to range from

one to four launches per year. The mission model for this study is

shown below.

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Mode...._l Launches, per Year

Low O. 5

Medi um 1.0

High 2.5

The low model represents a launch every two years; the _¢edium model,

a launch every year; and the high model, five launches every two years.

These launch figures are considered, to be reasonable projections,

especially if the recommendations of the SSEC are followed_ that is,

o implement the. core planetary program and supplement it with other

arger important scientific programs.

2.5 Suborbital Mission Node1

Electromagnetic launchers could perform a variety of suborbital

missions. This section presents the mission model development for the

EML suborbital launches. The missions under consideration in this study

include atmospheric, astronomy, physics, planetary, and re-entry studies.

At the present time, these studies are supported primarily by the NASA_

Suborbital Program with airborne, balloon, and sounding rocket launches.

2.5,i liASA Soundin 9 Rocket Pro(jrw

NASA has used sounding rockets to launch scientific payloads

since 1959. The program has grown in size, peaking in the late 1960's

and early 1970's. Since then, sounding rocket launches have declined

and seem to have stabilized at approximately sixty launches per year.

The decline has been caused by a number of factors, namely high inflation.

rates, limitations in NASA budgets, and increased costs.due to larger,

heavier, and more sophisticated " loads. The launch history through

1980 is presented in Figure 2-4. _ decline in sounding rocket launches

is not, however, an indication of lowered demand, as the.demand, measured

by applications from scientific investigators wanting to fly their

experiments on sounding rockets, consistently exceeds the number of

available rockets (Teeter and Reynolds, 1982).

Sounding rockets are used for those scientific packages which

do not require long viewing times or heavy masses. The average package

mass is 205 kg, with a typical viewing time of several, minutes, longer

for payloads with parachutes. Airplanes and balloons are used for heavier

payloads with longer -viewing times (up to several days). However,

sounding rocket payloads have increased in mass throughout the years,

asAs shown in Figure 2-5.

2.5.2 Mission Summar2

EMLs would likely replace only the sounding rockets for

suborbital missions for several reasons. The EML impulsive launch is

similar to that Of the rockets, in that a ballistic trajectory__ts_-

8

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vt

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0

1959

2-34

1955 1970 1975 1980

Year

r'

FIGURE 2-4. U,S. SOUNDZNGROCKETLAUNCHHISTORY

BAT?£LL.E -- COLUMBUS

!!
I

I.

t

ql

0

Q.

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0

0

q..

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V)

:P

210

200

180

160

140

120

100

8O

6O

50

1960

2-35

1965 1970

Year

FIGURE 2-5. SOUNDING ROCKET PAYLOAD MASS

IAT'f E LI.i[ -- COLI.i MiilJ|

.: . ;.:

1975.

TRENDS

1980

[
2-36

,

_o

followed. There is no Opportunity for th_helOnger viewing times of aircraft and balloons with an EML launch, . accelerations of launch

for sounding rockets and EMLs would be similar, and higher than those

experienced by aircraft and balloon payloads..

The suborbital mission model for EML launch, during the

thirty-year period of 2020 to 2050 is summarized below:

_r

_d

Model Launches _er Year

Low 50

Medium 100

High 150

The projections were considered to be representative of demand for

suborbital launches during the timeframe in question. There would of

course be some variation depending upon any new discoveries or technology

developments which may occur or upon the political cltmate of the times.

2.6 Electrom_ettc Boost Ntssion

MissiOn models are presented in this section for two types

of EML space applicatlons, where the EML provides, the initial boost

phase of launch, The first mission is an EML launch of a

chemically-propelled rocket to deliver cargo to orbit, the so-called

hybrid rocket/EML. This mission, which was originally conceived by

Henry Kolm, is a vehicle concept called the Electro-Scout. The second

application uses the EML to give an initial velocity to a mar,ned

single-stage-to-orbit vehicle (SSTO) or to a manned Transatmospheric

Vehicle (TAV), a new vehlcle concept currently under study, sponsored

by the U,S, Air Force (Rice, et ai, 1933).

2.6.1 Hybrtd ENL/Rocket

Initially, the hybrid EML/rocket was considered only to replace

the currently-available small U.S. expendable launch vehicles. These

launch vehicles, the Scout in particular, primarily launch small

satellites into low Earth orbits. The demand for small satellites is

not significant at the present time, and is not expected to increase

much in the future. However, if an inexpensive method of launching

were available, three to ten launches per year could be expecte_ in

the thirty-year period from 2020 to 2050.

Later investigation concluded that a hybrid EML/rocket system

.could launch the payloads currently studied for the Earth-orbital launcher

(e.g_, space station supply missions). Table 2-3 indicates the mass

available for launching to orbiting Space Stations. A traffic model

,j

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2-37

indicating number of launches per year fOr the available payload (compared

to the Earth-orbital projectile payload) and is shown in Table 2-17.

Table 2-18 reduces the traffic model Information to tndtCate daily !aunch

rates to suppprt the mission model.

TABLE 2-18. NUMBER OF EIL LAUNCHES PER DAY

'FOR HYBRID E]qL/ROCKET. LAUNCHER

u ._,u*uw*w,_

LaUnches Per Day

Year Low. Model High Model

2000 O. 7 1.4

2005 1.4 1.8

2010 2.0 2.2

2015 2.6 3.5

2020 2.6 4.2

2025 2.8 4.9

2030 3.2 5.7

2035 3.4 6.5

2040 3.8 7.4

2045 4.0 8. i

2050 4,4 8.7

Fi fty-Year Average 2.8 5.0

2.6.2 SSTO/TAV Booster

.m,m.m_

An EML system was also studied to give a small initial velocity

(on the order of 500 m/s) to an airbreathin 9 or rocket system. These

sytems might include stngle-stage-tooorbtt (SSTO) vehicles or

Transatmosphertc Vehicles (TAVs). The systems would be manned, and

therefore requite vew low.accelerations.

The vehicles under consideration would, under normal

circumstances, have a relatlvely low launch demand (estimated at a weekly

launch rate). Should circumstances change, the vehtcles would be required

to launch on demand. Therefore, the mission model for TAV/SSTO vehicles

was projected for the years 2020 to 2050 as follows:

Model _er Year

Low 50

High 200

In this case, the low model (an average of one launch per week) represents

the expected launch rate. However, the System .should be designed to

withstand datly launches.

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2.7 Space-Based Mlssions

Two spaCe-baSed EHL systems were cOnsidered in this study:

a satellite kick system, to transfer satellites from low-Earth Orbit.

to geosynchronous orbit, and a high-ievel nuclear waste disposal system.

The mission model development for each system is discussed in this

section.

2.7,1 Satellite Kick S_. t_m

The NASA Outside Users. Payload Model (Battelle, 1983) contains

payload schedules for non-NASA, non-DOD reimbursable payloads to be

flown .by non-Soviet-block countries in ,1983 to 1998. Detailed low and

high models are given for twelve mission categories, including

international, U.S, domestic, and. foreign regional communications; U.S,

and foreign observations (both geoSynchronous and lo_-Earth orbits);

materials prOcessing; and scientifictechnical development. Those mission

categories which may be serviced by a space-based EML satellite kick

system are the communications and GEO observation satellites. The Outside

User_Payload Model projections for the. years 1983 to 1998 are summarized

below.

Category Low Model High Model

U.S. Communications

Foreign Communications

U.S. GEO Observations

Foreign GEO Observations

Total

107 143

98 157

8, 11

16 27

229 338

The average number of payloads flown per year over the defined

sixteen-year period for these mission categories is 14.3 for the low

model and 21.1 for the high mode]. The trends indicated in the model

summaries show that communications satell, ite launch schedules dip in

the early 1990s and increase again in the late 1990s, The observations

satellite launches appear to-remain constant over the indicated timeframe.

By the years 2020 to 2050, communications satellites in GEO

should be saturated with satellites placed in modular platforms. Earth

observations missions were projected to continue at approximately the.

same launch rates as currently experienced. Therefore, the mission

model for a space-based satellite kick system was projected as follows:

Model Launches per Year

Low 12

Medium 18

High 24

BATi'EI,,t. E -- CQLU MBU, 8
2-40

2.7.2 _ed Nuclear Waste Dts___sal

A combination of and alternative to the "Standard Space

Disposal" concept described in Section 3.2.1 and the ESRI. Mission A

concept is presented here. Small billets of hlgh-!evel nuclear waste

(HLW) were launched from an orbiting electromagnetic launcher syS.tem

to solar system escape velocities. The HLW biliets are delivered in

shielded spherical containers to the space-based EML via a Shuttle-derived

vehicle.. The billets considered here are the same• as those studied

in BOeing, 1982. Each cylindrical billet {with radius of 2.926 cm and

height of5.858 cm.) has a mass of I kg. Therefore, to dispose of all

US commercial HLW (Rice, Mi'iler, and Earhart, 1982), 500 launches per

day from the space-based system are required to dispose of the. high-level

waste. Disposal of TRU wastes (Section 2.3) in this manner was not

considered economically feasible beCause of the extremely large number

of Space Shuttle flights required..

Fi

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3" 1

3.0 F.NL SYSTENS ANALYSIS

This section summarizes the results of the systems analysis

conducted during the course of this study. Section 3.1 presents the

preliminary mission scenarios and system concept sizings based upon each

mission payload model. Table 3-1 summarizes the payload requirements

for each mission derived from data in Section 2.0. AdditiOnal subsections

describe other analyses that were performed, including radiation estimates

(from nuclear waste payloads;, launch windows, flight mechanics, and

projectile considerations.

3.1 PrelimtnaryA.alysts of Candidate Concepts

Preliminary concept sizings were performed for the missions

Identified in Section 2.0 and were based upon the payload information

shown in Table 3-1. The railgun sizings were done by Battelle in the

same format aS that done in the previous ESRL study (Rice, et al, 1982).

Sizings for the coaxial accelerator concepts were done by EML Research,

Inc., under a parallel NASA Contract.

3,1.1 Ra|lgun Concepts

Raflgun concepts for space mission applications were addressed

in the 1982 Battelle-ESRL Study (Rice, et al, 1982), and are only briefly

summarized in this section. The simplest railgun consists of two

conducting rails which are shorted by a, moveable solid metal armature.

A magnetic fteld is proauced by the electrical current passing through

the rails and armature. The, Lorentz force,. F = Nqu x B = i x B, propels

a projectile down the bore of the railgun (see Figure 3-1). The solid

armature is restricted to operation below 2 or 3 ks/s, due to the

fundamental heating iimlts of metals.

A plasma armature may replace the solid armatures for velocities

beyond the 2 or 3 km/s limit. The plasma ts a good conductor, which

supports the .high pressures necessary to propel the projectile along

the rails. A problem with plasma armatures occurs during the first

1 km/s or so of travel, where the plasma causes erosion of the rails.

Currently, this problem is avoided by preboosting; that is, accelerating

the projectile to I km/s-.by other means before it enters the railgun

proper.

For high velocity launches, a single energy store system may

not be appropriate, because longer rail lengths require a very large

energy store to maintain the current throughout the launch. Also tall

resistance becomes a problem. Distributed energy storage (DES) is a

potential solution. Energy stores are distributed along the length of

the railgun; energy is switched tnto the rails as the projectile travels

down the bore. Switching becomes a much larger problem for DES railguns

than for single-stage systems, where the switching is done at the beginning

of the acceleratien, where it is the simplest to accomplish.

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Some concern has been expressed that, because of the high

voltages required for Earth-to-space railgun operation, the bore would

not be able to stand off these voltages. Table 3-2 (excerpted from CRC

Press, 1975-1976) lists the spark gap lengths for different peak voltages

for the worst case (needle point electrodes). These are the spark gap

lengths, in air. The table indicates that, when operating at a voltage

of 100 kV, the bore of a railgun must be greater than 15.5 cm to avoid

arcing between the rails. For the Earth-Orbital railgun (Section 4.2)

the peak voltage is 65 kV and bore diameter is 1.0 m, whtch is an order

of magnitude above that required to avoid arcing. It should be noted,

however, that the environment inside the railgun would be different from

that defined for the table; but with the order-of-magnitude difference,

no problem is anticipated.

Railgun sizing data foreach of the eight missions are sumarized

in Tabie 3-3. These concepts are discussed in Sections 3.1.3 through

3.1.10.

3.i_2 Coaxial Electromqnettc Accelerator Concepts

The second type of electromagnetic launcher selected for

consideration in this study is the coaxial electromagnetic accelerator.

Several coaxiai accelerator concepts which could be used to perform space

missions ace briefly discussed here and in Section 8.0. For more detailed

discussions, refer to Appendices D and E and to Mongeau, 1981.

A coaxial elect_omagnetic accelerator operates by passing a

projectile coil through a drive coil. Figure 3-2 illustrates the

relationship between projectile coil position and the mutual inductance

gradient. The thrust of the launcher is proportional to the mutual

inductance gradient and is given by the equation: |

l

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where Ip and I d are the currents in the projectile and drive co_l.s

respectively. As seen in the figure, the thrust reverses direction as

the projectile coil passes through the drive coil. One method o_

alleviating this problem is to use a stnusotdally oscillating drive coil

current. The current Would be synchronized to coincide with the travel

of the projectile cotl through the drive cotl.

I!

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TABLE 3,2. SPARK GAP LEN6]'HS FORNEEDLE POIITr ELECTRODES

Peak voltage (kV)

Length of Spark Gap

for Needle. Point,

Electrodes (cm) l  ̧ ,

10 0.85

50 5.20

100 15.5

150 26.1

200 35.7

250 45.2

300 54.7

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The major advantage of the coaxial launcher over the railgun

is that projectile stresses may be reduced by adding multiple projectile

coils. This allows the stresses to be distributed along the projectile,

so that Jrmay be much longer than a reilgun projectile.

The "mass driver" is a discrete.coil coaxial launcher which

allows the projectile to be pushed and pulled through the drive col!

region. Figure 3-3 shows a section of a discrete coil. accelerator. The

current is fed Into the projectile coil by a set of brushes. Two mass

driver prototypes have been fabricated. Mass Ortver I was built at

Massachusetts Institute of Tech.ology in 1976. This launcher accelerated

a 0.5 kg projectile to 800 m/s through a 2-m length.

Brush-commutation• wtll not work above certain Speeds, so for

higher velocities, induction is used to introduce curre,ts into the

projectile cotls. Thepulsed induction: accelerator was first demonstrated

in the USSR by Bondaletov in 1978, A Z-g projectile ring was accelerated

to 5 km/s. Pulsed induction,launchers only operate in the "push-mode"

since the drive coil itself induces the current. One advantage to the

induction methods is that no mechanical contact is necessary between,

the projectile and the drive coil barrel, so there can be effectively

no wear of the system.

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CONTINUOUS.;HELICAL BA_EL

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............................_

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FIGURE 3-4. SE_C_D _EE_NCE DESI_ _UNCHER

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Another coaxial accelerator concept is the helical railgun.

Instead of dis:fete drive coils, a continuous helix is used. Brush

commutation feeds the excitation current to the projectile coils. MIT

demonstrated the concept in 1979 when 4.5 kg gilders were accelerated

to 60 m/s.

The coaxial ENL concept selected for this study by Henry Kolm

and Peter- Mongeau (see Appendix D)+ts illustrated in Figure 3-4. The

reference coaxial launcher is a combination of the discrete-coil

pulsed-induction accelerator and the helical rallgun configurations.

As with the induction accelerator, the projectile coils are excited by

induction, not by mechanical contact; and+ like the helical railgun, the

concept uses a continuous winding instead of discrete coils, so. that

most of the commutation energy is stored in the winding rather than in

the commutation capacitors, The drive "col:Is" consist of a continuous

helical winding of rectangular copper-alloy wire, instead of discrete

coils. Each turn or set of turns is individually commutated by switches,

Such as solid state or triggered vacuum gap switches, so that an energized

segment, consisting of several turns (4 to 16 depending upon the mission

under consideration) is synchronized to travel with the projectile.

The projectile coils are solid copper or aluminum rings

distributed along the projectile at approximately one radius apart. These

coils are imbedded in a non-conducting composite, which would al.so serve

as themal protection and structure, as recommended, by Kolm and Mongeau.

The drive barrel and the projectile are in mechanical contact,

but no electrical contact occurs. The current is induced from the drive

coils to the projectile rings.

A sumary of the coaxial launcher concepts for each mission

from Appendix D is shown in Table 3-4, The preliminary sizing, was done

by EHL Research, Inc., and is documented tn Appendix O.

3,1,3 Earth-to-0rbit Ell S_sten

The Earth-to-orbit ENL system would launch bulk materials to

low-Earth orbit. Materials which were studied include. Orbit Transfer

Vehicle (OTV) propellants (In water form), Space Station supplyrequirements (life support consumables, spares, and other miscellaneous

items), and raw materials for processing in space. Concept options are

l)lustrated in Figure 3-5.

The operational scenario for this m_ssion begins at the

projectilepayload fabrication plant. The projectile, propulsion system,

and in some cases, the payload would be integrated before shipment to

the launch site. At the site, the projectiles would be stored until.

time of launch approaches. Water payloads would be supplied by the water

plant at the launch site.

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8AYTELL. E --_GOLUMBUB

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Just prior to launch, the projectile/payload would receive

final checkout and then be loaded into the electromagnetic launcher..

The projectile would be accelerated to the .6.85 km/s launch velocity

before leaving the muzzle of the launcher.

The acceleration limit of launch was set at 2500 g's; however

for the 2040-m launcher envisioned the acceleration is approximately

1225 g's.

For a payload of 650 kg, the total projectile mass would be

S900 kg for a railgun system and approximately 3000 kg for a coaxial

system. The difference in the two masses is due to the ability of the

coaxial accelerator to distribute the acceleration loads tn¢o many segments

by using multiple projectile coils.

4:

f!

The Earth-to-orbit system would have an elevation angle of

20 ° from the horizontal, The lends itself well to a launcher system

along the side of a mountain, saving construction costs of digging an

underground launcher tube 2040 m in length.

One problem for the near-tem Earth-to-orbit EML launcher is

the Space Station orbit requirement of 28.5 ° inclination. To supply

the Space Station, a launcher at 28,5 ° latitude cou]d launch once per

day to place its payload near the Station. If the system were placed

at the equator, two launchers per day could be possible; however, two

launcher tubes must be built--one pointed toward the Station on its

northward pass and one toward the sOuthward pass. if the Space Station

were in a 0° inclination Orbit, the station could be supplied on every

orbital pass by an equatorial-based EML system. For an orbital altitude

of 500 km, the periOd is approximately 1.6 hours. This corresponds to

a maximum launch rate of 16 per day to a single Space Station..

if significant savings could be shown using an EML system based

at the equator, NASA might consider a Space Station in an .equatorial.

orbit. This however is not foreseen in the near future.

i

3.1.4 Earth-to-0rbit ,H_orid Launcher

A hybrid EML/rocket. launcher was envisioned to launch cargo

into low-Earth orbit using the same mission model as the Earth-to-orbit

(all EML) launcher. The EML in this case- would replace the rocket's

first (and largest) stage by providing a velocity boost of 2 km/S from

the Earth's surface. Concept options for this and other electromagnetic

boost concepts are illustrated in Figure 3-6,

The hybrid system operational scenario is described here. At

the solid rocket manufacturing plant, the motor cases wquld be built

and loaded with propellant. The motors would then be transported Co

the launch site by rail or a combination of rail and ship, depending

upOn the location, of the site. Upon arrival, the motors would be placed

in. storage.

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PAYLOAD

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SYST£M

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• 0 • (TAVI

• 0 - 100 (SSTO)

• 0 * 30* (AIRCRAFT)

LATITCDE

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AZIMUTH

• VARIED

•' 'TI

FIGURE 3-6. ELECTROIM6NETiC BOOSTOPTIONS

BATTEi. I.E -- COL. WMStJ6
3-13

The payloads would be delivered separately to the launch site

and stored until launch time. Before launch, the motors would be stacked

and the payload integrated. The assembied vehicle/projectile wouid be

loaded into the electrOmagnetic launcher. All launch systems would be

readied, and countdown begun. When the vehicle leaves the muzzle, it

would be traveling at 2 km/s. Trajectory simulations indicate that,

for maximum performance, the ftrst stage of the rocket would fire after

15 to 20 seconds, and the vehicle would Continue alOng its trajectory

to lOw-Earth Orbit.

A mass of 15,000 kg was Initially selected for the rocket. Peter

Koren of Thfokol was consulted to preliminarily size a. three-stage rocket

(because of the better-performance cOmpared with two Stage rockets). A

total AV of 9.45 _/s is required for Earth orblt insertion, with 7.45 km/s

of this requirement supplled by the rocket and 2.0 km/s supplted by the

EHL.

The rocket was Initially sized for steel cases wlth a mass

fraction of 0.88. The mass breaJ,_own was as follows:

If

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....................................................................................... ' ......................... i

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Stage Propellant Mass (kg) Stage Mass (kg)

I i

1 7873 1074

2 3810 519

3 1016 138

i , i ...................... i ..... r ................. iii ....... i llf I

Later, a brief analysis was conducted to see if a 15,000 kg

rocket was acceptable or tf another vehic]e size would be more appropriate.

A preliminary comparison of various masses of 3-stage solid-propellant

rockets was conducted using Battelle's Launch Vehicle Performance Program.

This computer program is used for advance planning purposes to analyze

launch vehicles. Thrust, ISD or mass flow rate, burn time, and motor

jettison masses are entered for each motor, Payload and payload adapter

m_sses are input; payload Shroud mass and time of Shroud jettison are

entered as well. Coast times are assigned by the user to further define

the trajectory. The program computes two-dimensional and three-dimentional

trajectories and performance reserves. An iterative process is used

to calculate the maximum payload for a given vehicle configuration to

a defined orbit from a defined launch site.

Battelle's Launch Vehicle Performance Program was run for varying

masses at two initial boost velocities (2 and 3 km/s). Coast times between

burns were allowed to vary from vehicle to vehicle to detenlline the maximum

payload capability. The results are shown in Figure 3-7 for input

conditions of 290 sec Is_ (typical so!id propellant rocket), 88% stage

mass fraction, and final c_rcular orbit altitude of 500 km.

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Four representative points were chosen off the payload curves

to calculate railgun iaunch parameters as a method of comparison. The

four cases .are:

(1) _5,000- kg launched mass wtth

(resulting maximum payload -- 800 kg)

(2) 15,00_,_kg launched mass

(maximum payload = 1250 kg)

EML AV = 2 km/s

with EML _V = 3 km/s

(3) 20,O00-kg launched mass

(maximum payload = 1080 kg)

with EML _V = 2 kin/s

(4) 20,O00-kg launched mass

(maximum payload =1650kg)

with EML _V = 3 km/s

For all cases, lO0-g acceleration was aSSumed. The results of the brief

investigation are tabulated in Table 3_5.

TABLE 3-5, CALCULATED RAILGUN PARANETE_

FOR SELECTED HYBRID CASES

Case Number

Parameters 1 2 3 4

Minimum Length (km) 2.04 4.5 2.04 4.5

Acceleration Xime (sec) 2.04 3.0 2.04 3.03

Force (MN) 14.7 14.7 19.6 19.6

Current (_) 7.67 7.67 8.85 8.85

Launch Kinetic Etlergy (GJ). 30. 67.5 40. 90.

Peak Power (GWe) 58.8 88.2 7a.3 117.

Peak Voltage (kV) 7.67 11.5 8.85 13.3

i

One of the parameters calculated by the Performance Program

is the kick angle (the angle through which the launch vehicle "kicks

over", usually during the first stage burn). This angle was used to

define the necessary elevation angle of the mountain launch site.. For

the EML hybrid system, the minimum elevation angle was 35 degrees. This

angle is fairly high for most mountains, but it was felt tha_ there may.

be mountains with this elevation angle, over a 2-km length.

At the defined acceleration of 100 g's, the length of an EML

system is 2 km. Inside bore diameters are O.S m for the coaxial

accelerator a[zd 1.2 m for the railgun.

For-the coaxial accelerator concept, the projectile col_s could

be imbedded inside the motor cases during fabrication. For the rat!gun

conceptS, a sabot with conducting armature would be placed at ths rear

of the projectile,

§A'lr'r2{.i.| -- CaLUM_iU|
3-16

For the same launched mass, increasing the boost velocity from

2 km/s to 3 km/s more than doubled the energy requirement, while only

resulting in a @OZ gain in payload mass. The energy-to-payload,mass

ratios are approximately equal for cases with the same boost vel.ocltles;

i.e. for Cases 1 and 3 and for Cases 2-and 4. It was decided to continue

with the 15,000-_ rocket vehicle. The decision was based in part on

the lower mass (the lower the better for near.term EML development).

Also, 800 kg was felt to be an acceptable payload capability; for

comparison purposes, the Scout launch vehicle is capable of placing 235 kg

into a SO0 x 500 km orbit when its smaller 0.80-m diameter heat shield

iS used (VOught, 1980).

3.1.5 Lunar Su_.Ep..lyLauncher

The EML tunarsupply launcher was envisioned to supply necessary

life support consumables and OTV propellants to a manned lunar base, The

year 2010 was selected as the initial operating capability for such an

EML system. Figure 3,8 presents lunar supplyconcept options.

For preliminary sizing purposes, theEarth-to-orbit projectiles

were used. Maximum accele_atlon was set at 2500 g's and the launch

velocity set at 12 km/s (reference Section 3.6). These requirements

correspond to a minimum launcher lengthof 2940 m.

Ideally, the luna_ supply launcher would be located at the

equator and would be. placed verticail_. A 3-km long vertical launcher

would be built in an underground shaft similar to that used in the ESRL

concept (Rice, et al, 19,82),

The projectile would be fabricated in the United States and

transported by Ship or aircraft to the equatorial launch site. Water

payloads would be supplied at the launch site; Other payloads (food,

for example) would be transported byair,

Launch support facilities would be similar to those defined

in Rice, et ai (1982) for the ESRL system whi.ch was sized for a iO-km/s

launch velocity,

One major problem cf the lunar supply launcher is the large

amount of auxiliary propulsion which must be carried for midcourse

correction, lunar orbit insertion, and, if required, lunar landlng.

Section 3.8.3 describes the propellant requirement necessary and- shows

that. a payloaa of approximately 50_ kg would be possible for lunar-orbit

insertion when using the Earth-to-orbit projectiles.

Assuming a launch capability of One launch per hour (one on

either side of a daily launch window), 12 launches would be possible

per month (144 per year).. With. a maximum payload mass of 500 kg, this

corresponds to a maximum yearly payload delivery of 72 MT. If a launch

every half-hour were attainable, 18 launches per month could be possible.

This would correspond to a maximum yearly delivery of 108MT of payload.

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3,18

3.1.6 Solar Syste)m Escape Launcher

High, level and transuranic (TRU) nuclear wastes produced by

U.S. commercial nuclear power plants would be sent to a waste processing

facility (See Section 2.3 for discussion of TRU waste processing and

Rice, et. al., 1982 regarding high-level wastes), This facility would

be located in the United States and would perform a dual function by

processing the waste and fabricating the projectile. The waste would

be stored for a number of years before becomillg part of the projectile.

Concept opttons for the solar system escape and Earth escape missions

are presented in Ftgure3-9.

The nuclear waste is encorpora_ed In a stainless steel radiation

shield. The shield requireent is thicker for the high,level waste form

than for the TRU wastes, because of the higher levels of radiation.

Structural strength is also provided by the shield. Carbon-carbon

materials are applied to the outside of the projectile for thema!

protection. A tungsten nose cone is believed to be required _to wlthstand

the 20 km/s launch velocltleS.

The finished projectiles would be transported to the remote

island launch site by railroad, then by an ocean vessel. At the launch

site, the projectiles would be placed in a storage and checkout facility.

At launch.time, the projectile is loaded into the EML system.

and, when the launcher is ready for launch, the projectile is launched

at 20 km/s to escape the solar system. Figure 3-I0 provides an overview

of this concept scenario.

Since it was felt that the projectile could withstand high

launch accelerations, I0,000 g's was set as the acceleration limit. This

corresponds to a launcher len0th of 2.0 km. Because of the length of

the launcher,, the vertical launch requirement, and the safety

considerations, an underground launch system on a. remote location is

needed.

To dispose of all the high-level and TRU wastes produced by

U.S. commercial nuclear power plants, 26 launches per day are necessary

during the 6-hour launch window. Four launcher-tubes would be required

at a minimum, and the power requirements would need to be increased over

the ESRL requirements (Rice, et al, 1982) to provide for one launch every

15 minutes (one launch per hour per tube).

3.1.7 Earth-Escape Launcher

The Earth-escape launcher would be used to launch planetary

explorers and probes. A payload mass of 600 k9 was believed to be lar9e

enough to provide an adequate payload, yet at the same time allow a

reasonable projectile mass. The projectile would consist of the 600-kg

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payload, a small auxlliary propulsion sys.tem for mid-courSe guidance, l

strUctural support for: the payload and propulsion system,, stabilfzation

fins, and a tungsten nose cone to. withstand atFnOSpheric flight during

the 20 km/s launch velocity. Figure 3-_ illustrates the concept options.

An acceleration limlt of I0_000 g'S was set to provide reasonable

launcher lengths. At a launch veloclty of 20 km/s, the minimum length

would be 2040 m. The coaxial accelerator would appear to be more

attractive for this case, because the acceleration loads could be

distribute d along the projectile, saving a significant amou_it of structur&l

mass. Rrellminary conceptualization yielded projectile masses of 2055 kg

for a rail gun system and 750 kg for a coaxial accelerator system, with

the major difference being Structural mass.

The Earth-escape launcher would be placed vertically and

therefore would require an underground site to support the system. The

launch velocity fiiust be variable to accommodate launches to different

planets and solar system locations (12-20 km/s velocities were envisioned,

see Section 3.4 for details).

3.1.8 SSTOr AVBooster

The Tr_nsatmoSpheric Vehicie (TAV) and Single,Stage-to-Orbit

(SSTO) vehicle are concepts studied by the USAF (Battelle, 1983) and

NASA (Jackson, i983; Eldred, 1984), respectively. Most of these vehicle

concepts utilize rocket prOpulsion, although airbreathing propulsion

is being considered. The various concepts range in mass from about 500,000 .......

to 5,000,000 Kg.. Concept options are shown.in Figure 3-6.

The design goals for the TAV are to use a fully reusable,

horizontal takeoff, single-stage vehicle (with ground assist) to deliver

9100 kg (20,000 Ibs) of payload to a low polar orbit. The vehicle must

be ready, to launch within 5 minutes following an alert with turnaround

time to allow 2 launches per day. The TAV operational requirements are

summarized: manned operation; minimal ground support; horizontal takeoff

and landing; flexible basing; and adverse weather operations. During

the Phase I USAF/ASD study conducted by Battelle, 14 TAV concepts were

evaloated, of thOse, six were considered in: the class of

"single-stage-to-orbit" type systems and three used ground-assist sleds

of various types. These systems would be excellent, candidates for EML

use for the.initial boost.

The civilian SSTO goals and requirements are similar but less

severe. Ground support and quick turnaround time are: not as critical

and the use of dedicated, vulnerable launch facilities are not critical

issues.

The projectile (vehicle) mass was set at 900,000 kg for this

study. Since the vehicle will be manned, a 5-g acceleration limit was

set. It was felt that an initial velcity boost of 500 m/s was sufficient

aAT,TIELL, a -- COI-UMStJ6

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tO launch the vehicle_ This led to EML system lengths which are of the

order of existing aircraft runways (2.55 to 3,83 kin),

The railgun concept presented here for a TAViSSTO launch is

illustrated in Figure 3-il. The projectile/vehicle sits on top of a

solid sliding armature. Tbe vehicle for the coaxtal accelerator concept

rests upon support structure containing a saddle-shaped projectile

coil (shown _n Figure 3-12 and on page 92 of Appendix D), The support

StrUCture slides alongthe length of the drive coils.

3.1.9 Suborbtt.al Lauacker_ •

For most of the missions deSCribed in Section 2.5 (atmospheric

soundings, astronomy experiments, and reentry tests), a vertical or near

vertical launch is preferred over a 20-35 ° launch. A mountainous site

ts then eliminated, with an underground facility the likely siting. The

various concept option_ are Shown in Figure 3-13.

Payloads would be manufactured under the guidance of the

principal ScientiSts. They would then be Sent tO the projectile

manufacturing faciltty for Integration. The combined prOjeCtile/payload

wOuld be sent to the launch site for final checkout before launch.

he railgun projectile was sized as a scaled-down Earth-orbital projectile A 225-kg p_'ojectile was envisioned, with a square sabot

32.S cm on a side. The projectile itseif would be 3Q cm in diameter,

with a projectile/sabot length Of 1.2 m.

The rails would be from 510 tO 765 m long depending upon whether

distributed or single energy storage was used. With these lengths and

a hearty vertical launch angle to suit most mission needs, the launcher

tube would most likely be placed underground to provide structural

strength,

The coaxial projectile was also sized at 225 kg in mass, with

a payload of 160 kg. The projectile would ha.ve a diameter of 20 cm and

length of 2.8 m.

Although launch site. flexibility is highly desired for most

types of suborbital missions, 56% of the US sounding rocket launches

have been fired from two locations, with anOther 14% from a third site.

An EML suborbital launcher tube would obviously be fixed, because of

its large size and many supporting Systems.

3.1.10 S__ce-Based

The scope of this study encompassed: space applications for

both Earth-based and space-based EML systems. Two orbiting space-based

• systems were identified and are described in this Section: satellite

kick System and high-level nuclear wastedisposal system. Concept options

for both missions are presentedtn Figure 3-14.

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Required veloctty impulses are smaller from Orbit than from

the Earth'sSurface. The minimum solar system escape veiocity from Earth's

surface is 1_,7 km/s (no drag tncluded), whtle this number iS reduced

to approxi._tely 10 km/s with a lowEarth,-S orb.it launch (Section 3.3).

0rblting space-based EML systems have several inherent

disadvantages when Compared with Earth-based systems, One disadvantage

iS cOnStruction cost. FOr near-term operation, the equipment must be

built on Earth and delivered to orbit via the Space Shuttle. For long

launcher tubes, this may be expensive because of the many Shuttle fltghts -

required fOr assembly (the dimensions of the payload bay are 18.3 m long

and 4.5 m in diameter).

Another disadvantage is NewtOn's third principle: for every

action, there is an equal and opposite reaction. For the launcher, to

maintain the orbit, a stzeabie impulse is required to balance every shot

or ser_.es of shots. Drag makeup and attitude control would also be needed,

which would incUr additional propulsive maneuvers.

3.1.10.1 Sa ]lite Kick Syste.

The first Space-based EflL delivery system is a satellite kick

system which would replace the need for an 0TV or large, upper stages.

The EML satellite kick system would be based as an orbiting Space Statton.

This Space Station is envisioned to be in a 500-km Circular orbit at

28.50 inclination to facilitate delivery to and from the Station by a

Space Shuttle or derivative.

Because the launch rates would be low (one to two Iaunches

per month), the Station need not be a dedicated facility. A nuclear

power plant could be. used both to power the EML system and to supply

needed Station power requirements. Themal radiators would be used to

dissipate generated heat into space. A storage module would be required

to store the satellites after Shuttle delivery to orbit before the

des_gnatedlaunch dates.

The Space Shuttle would deliver the satellite payload to the

Space Station, where it would be stored until time of launch. At that

time, the satellite and its supporting systems would be checked out.

The payload would be loaded into the breech of the launcher and launched

to its required velocity (up to 2,5 km/s). When the proper velocity

is reached, the payload would be released from its amature/drtver and

continue towards its destination. To avoid orbital debris problems and

to maintain high reusability, a deceleration syste_ is required for the

driver mass.

The minimum velocity needed to place the satellite system into

a. geosynchronous transfer orbit with perigee altitude of 500 km is

2,36 km/S. This minimum velocity was calculated assuming a Hohmann

transfer Irom 500 km altitude. The eccentricity of the transfer orbit

is 0.72, leadl,g to a velocity of 9,97 km/s required at perigee tO sustain

iiAI'I"III. LE -- (_OI. UMilUli
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3-26

the transfer orbit. The orbtttng EHL system is traversing around the

Earth at 7.61 I_#s, the Satelitte therefore needs an additiOnal 2.36 km/s.

Lower apogee altitudes require smailer velocity increments, Therefore,

the system was stZed for ve|octttes up to 2.5 kin/s; the actual veloctty

per shot would depend upon the destred orbt¢, Thts stngle tmpuise system

would only prOvtde the AV required to put the Satellite tnto an elliptical

transfer orbit, The satellite wouid need to have an attached small

propulsion system to circularize at the apogee altitude, if this ts

required. Agatn assumtng a mtntmum energy HOhmann transfer, 1.5 km/s

would be requtrod fOr ctrCulartzatton at GE0 altttude

The payloadConststS Of the satellite and Its supporting systems,

Preliminary analysis indicated that to be competitive wtth other systems,

a useful payload mass Of 5000 kg and a minimum payload diameter of 2 m

was needed. An acceleration limit Of 1000 g'S was placed on the launcher

system, This limit was considered to be high, Considering the delicate

nature of many satellites, but was needed to 11mlt the length of the

system,

Two types of railgun systems were envisioned (Figure 3-15).

A large solid metal armature was used in both cases, The first

configuration has the satellite riding on a support structure on top

of the armature external to the rallgun. This cOncept has some obvious

structural problems to overcome. The off-center-of-gravity loads may

become severe when traveling at 2.5 km/s, and could damage.nOt only the

satellite support structure but the rails themselves, An advantage .of

thts Concept.IS that the bore need not be as large as 2 m in diameter.

The t_tellite was launched "conventionally" in the railgun

bore in the second concept. The amature acts as a pusher plate behind

the satellite. This method eliminates the structural problems associated

with the. external satellite, but increases the diameter of the bore to

at least 2.m.

Both ratlgun concepts assumed a single energy store, for ease

of assembly and maintenance. Thts increased the length of the rails

from a mintmum of 320 m to 480 m. If oniy one rail_ segment at a time

couid be transported by Shuttle, this. corresponds to at least 27 flights

to deliver the rails, alcne,.and more would required to deliver the rest

of the launch system.

The :oaxlal. accelerator concept cOnsiSts of 319 m of drive

cot1, The drive*coil inside diameter is 57 cm, which tS very small for

a satellite System. For comparison, the Delta launch vehicle shroud

ts 2.4 m (8 ft) in diameter. The Scout heatshtelds are-l.03 m _nd 0.86 m

in dlamete_.

_ce-Based

The proJecttie consists of a Small unshielded htgh*levei.waste

(cermet waste form) billet encapsuiated tn an Insulator/sabot. The btllet

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ts cylindrical in Shape, wtth a hetght of 5.868 cm and a dtameter of

S.85_cm. The mass Of the.waste btllet ts 0.95 kg.

The waste bllletS are transported by an uprated Space Shuttle

to the Orbitlng EML platfOm in two Shielded stalnless steel sphereS.

Each spheP_cal assembly contalns 3167 billets (3000 kg of wasto), for

a total of 6000_kg mass peP Shuttle launch.

To compietely dlspose Of a11 U,$. commerclal hlgh-level waste,

an avarage of 500 kg must be launched daily using the EML System (Rice,

et al, and Earhart, 1982). Therefore one Shuttle dellvery f11ght is

required every I2 days (30 Shuttle launches per year dedicated to nuclear

waste dlSposal)

Because of the very high radiation dOseS from the unshlelded

waste btllets (see Section 3,2), the EHL platform area mus_ be unmanned

or sh_eided during iaunch, etc. After the Shuttle leaves, the biilets

wouid be unloaded from the dellvery spheres and placed in the launch

sabots. The waste projectile (b_11et and sabot) would be loaded tntO

the launcher, Because the platform would be orbiting around the Earth,

a launch window of about 19 mtnutes would be available for launch on

every orbital pass. Thts corresponds to a launch every 36 seconds during

the wtndow. The enttre process from unloading the btllets to commencing

the launch procedure wouid be automated;

The platform was assumed tO be located in a 500-km circular

orbtt at 28.5 = Inclination to facilitate launches from an extsting pad

at Kennedy Space Center. A nuclear power plant, was assumed to prowide

the necessary power for ali platform operations. Large radiators would

be used to dissipate the heat generated during the 11 km/s launch. The

minimum acceleration length as a function Of launch acceie.ration ts shown

below:

............. r " " I_ ' I I I I I "_lll I I

Acceleration (g's) Length(m)

1,000 6,170.

10,000 617

25,000 247

50,000 123

75,000 82

i i .............. r" i

0bvtously, to reduce the cost of Space Shutt]e deltvery costs forassembly

and to reduce the risk of orbital debris hitS, the acceleration must

be as highas possible.

Kolm and Hongeau selected a 10,000 g acceleration lt.mit for

thetr coaxtal accelerator design, but commented that thts ltmtt could

be Increased tO 30,000 g's "without much extrapolation of present

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3-29

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technology" (Appendix D), They also modtfied the shape of the

proJectile/biliet (mass reFnalned at I kg of waste forni) to allow four

projectile coil+s,

When the preliminary rallgun system was conceptualized, it

was felt that, because the projectlle_ould be given a large enough impulse

to exit the solar system from orbit, the cOnditiOn of the waste billet

UpOn leaving the EHL was not as critical as for the Case of an Earth-based

launch whlch has greater probability of anaccident impacting the Earth.

To keep the launcher at reasonable lengths, the acceleration limit was

defined by the launcher materials, not the projectile material. Assuming

a safet_ factor of 2: and AHIIRC rails, the minimum length of the rails

Is 86 m for an acceleration of 72,000 g's and bore area of 48 cmz {3.g-Cm

radius). A distributed energy storage system would then be 86 m long.

For a single energy store, the length would beCOme 138m (at four times

the energy cost for Simplicity of system).

The projectile for the space-based EML nuclear waste disposal

System consists of commercial hlgh-level nuclear waste. The waste mix

iS defined as a highlevel- Purex {plutonium and uranium extraction) waste

with 90 percent of the cesium and strontium removed (Modified PW-4b).

The cylindriLcally-shaped waste form is the Oak Ridge National Laboratory_

Iron-based cermet.

The nuclear waste radiation analysis was performed to determine

the dosage obtained from an unShtelded cylindrical waste form billet,

with a radius (r) of 2.g26 cm and length .(I)Of 5.858 cm The analysis

was conducted at "detector points" at varlous distances from the waste

form, beginning at the billet surface to a d_stance of I km. The source

term assumed for the waste mix is the same as that used in Rice, Miller,

and Earhart {1982).

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radiation analySlSl. The dose (in rem/h,) was Calculated at a number

Of detector points at varying distances (d) from the prOjectlle. Table 3-6

summarlzesthe results of these calculations.

3,3 LaunCh k,tndow Anal is for._r S stem

llor__k 3.3,1 Previous

E payloads fromThe Battelle the Earth'sESRLsurface study toconsidered intersteliar the space impulsive(Rice,launch et al,of

1982). The launch velocity requirements xere estimated as a function !

of launch latitude, time of day, and time of year. Both vertical and i

non vertical launches were considered and atmospheric drag penalties

were inCluded. The primary conclusions were as follows: l

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FRON THE WASTEFORN PROOECTILE

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0.05

O. 10

O. 20

O. 50

1.0

2.0

5.0

10.0

20.0

50.0

100.0

200.0

500..0

1000.0

Dose (rem/hr)

5.364 x 105,

1.420 x 105.

3.553 X 104

8.984 X 103

1.433 x 103.

3,609.x 102

9.022 x 101

1,442 x 101

3.593

8.893 x 10"1

1.347 x 10-1

2.897 x 10"2

4.687 X I0-3

1.290 x i0-4

1.122 X 10-6

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3-31

e The mtntmal launch ve;ocity (approximately 16.7 km/s for

nodrag) for Solar system escape is independent of latitude

up tO a latttude of approximately 23.5 degrees.

e The penalty for launching at non-optimal times of the year

is _educed with an equatorial launch site.

e The Optimum time of day iS approximately 6 a.m. for any

latttude and laUnCh date.

e For equatorial launches, two optimal launch dates occur

each year--g0 days after the vernal equinox and the autumnal

equinox, respectively.

• For vertical equatorial launches, with drag effects included,

a launch velocity of 19 km/s could provide a da_ly launch

window ranging from 3.S to 4,6 hours. A launch velocity

of 20 km/sec could provide a daily launch window of

approximately 6 hourS,

In the current investigation, the effects of launching from

a platform in Earth orbit were examned. From a flight-mechaniCs standpoint, a launch from, orbit differs from a surface launch in the following

respects:

• The launch energy requirementsare reduced.

• Unlike a surface launch, the projectile would be injected

into its trajectory at the perigee, of the Earth.escape

hyperbola, thereby altering the time-of da_ relationship

• Since the launcher is traversing its orbit at a rate of

24 hours (local time on Earth below the vehicle) in about

gO minutes, the dally launch window is compressed.

Figure 3-17 illustrates the combined effects of the factors

just diSCuSsed for a launch from a 500 km orbit. The remaining projectile

velocity is shown at an infinite distance from the Sun for a launch on

an opttmum date.

As shown in Figure 3-17, the minimum launch velocity has been

reduced by approximately 10 km/s as compared to a surface launch; and

the best local launch time is near midnight rather than 6 a.m.

If a launch velocity of 10 km/s is assumed, the launch window

per orbit would be from about 7 p.m. to 3:30 a.m., locat time. On the

other hand, because of the orbital velocity of the launcher, this

difference in local hour-angle would be traversed in about 24 minutes.

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3-32

VL ,, 12 km/s

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Hour Angie from Midnight

LAUNCHFROM 500-1me ORBIT ON BEST DAY

VL - 12km/s

11

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Hour-Angle from Midnight

LAUNCH FROH 500-1an ORBIT ON MORSTDAY

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Figure 3-18 is a stmilar plot for a worst day of the year (at

either equinox). On these dates, the actual launch window would be reduced

to about 12 mtnutes on. each orbit, Qr approximately 3.2 hours per day,

for the lO-km/s launch velocity.

Although the launch of payloads to solar-system escape appears

reasonable from a laUnch velocity standpoint, consideration must be given

tO the fact that reactive impulse would be imparted to the launcher,

equal and oppOsite tO the. Impulse applied to the projectile. To matn_atn

the orbit of the launCheP, a sizeable corrective tmpulse would be required

after each p_ojecttle launch.

3,4 Launch Window Analjfsts for Solar _._1_stee Probe._!.

For payloads launched to destinations within the Solar system,

Only ve_ttcal launches from the Earth's surface wer___examined.

To compare soiar system probe iaunches to solar system escape

missions, it is obvious that solar system probes will require lower launch

velocities, However, the solar system probes are constrained to deilver

a payload to a particular destination. For flights to the planets, the

geometric constraints are formidable for a. fixed launcher. The analysis

of launches to .specific planets was considered to be beyond the scope

of this effort, Instead, the analysts was confined to simply establishing

the launch requirements for delivering probes to various distances from

the Sun.

For destinations outside of the Earth's orbit, vertical launches

occurring near 6 a.m. will make maximum use of the Earth's velocity. For

destinations closer to the Sun, launches near 6 p.m.. will be properly

aligned 1:o subtract most effectively from .the Earth's velocity, thereby

achieving the smallest perihelion.

Figure 3-19 shows both cases for an assJmed launch velocity

of 16 km/s (no drag penalty included). The aphelton posstble on an_optimum

day is seen to be over 20 Earth orbit radii from the Sun. As in the

case of solar system escape payloads, the optimum days are 90 days after

each equinox. On a worst day, the same launch velocity could sttl.l_produce

an aphelion beyond the orbit of Saturn.

For probes nearer the" Sun, a worst-day launch would provide

a perihelion inside the orbit of _rcury, while an optimum day would

pemit__perihelion of. less than 0.25 Earth radius f_om the Sun.

Figure 3-20 confines attention to probes to the outer solar

system, and shows aphelion as a function of launch velocity for 6 a.m..

launches. Aithough fltght ttmes to aphelion are not shown, these times

can be quite long for these minimum-energy trajectories. For the cases.

shown, the flight time to the. orbit of Juptter would ber about 2 years

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FXGURE 3-19. SOLAR sYSTER PROBE APHELION/PERIHELION RANGE

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BEST

DAY

'3 1_ _s 16 17.

Launch Velocity, km/sec

SOLAR SYSTEM PROBES**K4X][HUN DXSTASCE-FR0tl SUN

IAI"'_IRLI. E -- ¢:,OLUMSU8

16

I ,

'i

......... ,.--~ .
3-35

for the launch veioctties indicated. For destinations near the orbits

of Saturn and Uranus, the flight times would be about 4 and 11 years,

respeCtively.

3.5 Anal_sts of Requirements for LaunChing Payloads to Eart h Orbit

3.s.1 Xpom,otorI,p,lse utrweats

The use of EML concepts for placing payloads in circular

equatorial, orbits was investigated parametrically to determine the launch

velocities and apogee motor impulse requt.rements as functions of orbit

altitude and lauach angle from the horizontal.

Figure 3-21 summarizes the results. For simplicity, and in

keeping with the exploratory nature of the study, atmospheric drag was

neglected in CompUting the launch velocities.

In Figure 3-21, the solid lines are contours of constant launch

veloctty (provided by the EML) and the dashed ltnes are contours of _

constant apogee motor impulse required to circularize the orbit (provided.

by an on-board propulsion system). For vertical launches to the lower

altitudes, the apogee motor requirements aru clearly prOhibitive.

3.5.2 Range Safet_ Limitations

A further limitation appears if range safety aspects are

considered. The payload must be protected at launch by a rather massive

shield. The shield will then separate from the payload and will remain

in the transfer orbit. It will then return to Earth and impact at a

location dependent upon the apogee altitude and the launch angle.

The longitude of the shield impact point, relative to the launch

site, is shown in Figure 3-22. For vertical launches, the shield would

always appear to draft toward the. west, to an observer at the launch

site.

Figure 3_23 displays the same information, but specialized

for an equatorial launch site on the east coast of South America. For

vertical launches, the shield would fail on the South American continent

for all apogee- altitudes up to about 10,S00 km. For higher apogees,

shield impact would be in the Pacific Ocean. For this particular site,

it appears that launches could be made to any altitude (up to

eosynchronous altitude), without impacting any land mass, only for. a

aunch angle of about 55 degrees from the horizontal.

Figure 3,24 is specialized for launches from the west COaSt

of South America. This would appear to be the most promising launch

site for verttcal launches, since it would always result in a shield

impact in the Pacific Ocean.

aATTEI. I.t, -_ COt. UMBUS

)i'

i

i _ I • _

F)

°

r

.!i
I]

I!

t

s

3

I00,000

Io.ooo

1,000

100

%

3-37

Leur,ch Angle, degrees from horizontal

FIGURE 3-21. APOGEENOTORIltPULSE R£(lUIRE_NTS

FOR CZRCULAREQUATQRZALORBITS

1

(

|

IA?TII. I.| -- COI. UMmU!
4,a

,,i,,a

91:

q.

100.000

10,000

1,000

3-38 •

4O

10o i ii

8O qO 50 40 _.0

I

6O

L_unch.Ang]e, degmes from.hOrizontal

f ,

3O

FISURE 3.22. LONGITUDE OF SHIELD IMPACT RELATIVE

TO EQUATORIAl. LAUNCH "_YTE

BATTIEf,,,L| ._ GOLUNIliUI

o_.

I) (t

3

i ,t

iI

;,J

'e

i.J

i':!

It

o

E_
I00,000

3-39

SHIELD IMPACT

WEST OF LAUNCHSITE

I

|_,

I SHIELD IMPACT

EAST OF-LAUNCH SITE

i

t_

I

10,000

m

o

0

o.

0

Q.

1,000

100

90

IMPACT ON:

_SOUTH AMERICA

AFRICA

!NDONESIA •

8O 70 60 50 40 30

Launch Angle, degrees"from horizontal

FIGURE 3-23. LONGITUDEOF SHIELD IMPACT RELATIVE

TO BELEH, BRAZIL, LAUNCHSITE

IATTILL|. - {QLUMIUI

-_IP

2O

.."

i

p
1O0,000

3-40

IMPACT ON:

TH AMERICA _AFRiCA _INDONESIA

I

I

SHI ILD IMPACT I. SHIELD IMPACT

WESTOF LAUNCHSITE EAST OF LAUNCHSITE

4. J

1,000

lO0

,:.

;,:..

4O

.:,:.:.:.:..

,',-.:...

..-.:.....-....

60 50 40 30 20

Launch. Angle, degrees from horiZOntal

FIGURE 3,,Z4. LONGI/UOE OF SHIELO IlqPACT RELATIVE

TO I_JITO, EqUADOR, LAUNCHSill{.

ItAY'rMLL. i[ _ ¢OLUMIBuli

il

i-1

")

i.l
l\

l!

)

(

LF

3_4i

3.6 LU__] a !_lSStL_nchestrements for

The Impulsive velocity requirements for the lunar base resupply

mission have been computed for vertical EML launches from the equator.

Theatmospheric drag penalty has.not been considered, but could be included

as a multiplying factQr based on the projectile ballistic coefficient

as discussed in Section 3.1.4.3 of Rice, Miller, and Earhart (1982).

For this analysis, the Earth-Moon trajectory data were obtained

from direct numerical integration of the three-dimensional equations

of motion Of the projectile, under the simultaneous influence of both

the Earth and the Moon. In contrast, the retro-impulse maneuvers fOr

lunar capture or landing were computed by Considering only the

gravitational attraction Of the Moon itself.

3.6.1 F_of Launch_tttes

For verttcal equatorial launches, the projectile trajectory

wtll lies in the equatorial plane, except for small out-of-plane

perturbations in the near-vicinity of the Moon, Consequently, the arrival

of the projectile at the Moon must occur when the Moon is very close

to the intersection of the plane of the Moon's orbit and the Earth's

equatorial plane. Since the MoOn passes through the equatorial plane

twice a month, there will be two opportunities per month for a projectile

fired from a fixed launcher.

3.6.2 Launch Ve]octt_

Because of the distance to the Moon, the minimum launch velocity

for lunar capture is only slightly less than Earth-escape velocity. The

distance to the Moon varies from about 356,400 km to 406,700 km. While

this variation has only a slight effect on the minimum launch velocity

requirement, it has a pronounced effect on flight time.

Figure 3-25 Illustrates the relationships between launch

velocity, fllght time, and the impact velocity on the lunar surface for

direct-impact flights. As seen in the figure, the flight time is most

sensitive to variations in the launch veloctty when the launch velocity

is near the minimum value. The use of these lower-energy trajectories

would not be appropriate for EML launches because of the flight time

sensitivity tO launch velOcity uncertainties. The time of flight must

be known with precision to launch at the appropriate time to ensure the

simultaneous arrtval of the projectile and the Moon at the line of

Intersection of lunar plane and the equatorial plane. .

As seen in Figure 3-25, the launch velocity requirements for

lower flight times (and iower filght time sen_Itlvlty) are only s11ghtly

larger than the minimum. A greater penalty Is suffered in the impact

velocity as the fiight time is reduced. The variation in the Earth-MoOn

distance has a pronOunCed effect On the flight time, but a negltgtbie

effect On the impact velocity.

iiA'r._|.LI.| -- COI. UMilU|
m  ̧_

E

o.

S

4_

¢j0

>_

5.0-"

4.5"

v, 4.0-

i 3.5- 3,0 -

2.5-

2.0 -

e

1.se,

=J

.J

1.0"

0.5""

-70

"65

-60

-55

-50 .c

"46 _

-40 _

-3S _"

-30

_25

• |

ii .0

EI_ 3-25.

.0 a

2.0 '*"

o/

1.0-

i

.

3-42 .............

Note: VertiCal Equatorial

No Earth Atmosphere

Launches

Impact Velocity

(All DistanCes)

Earth.Moon Distance =356,400 km

iqht Time

(Earth-Moon • Distance =

406,700 km)

'"_I 'I I' I I" I

Ii.I II.2 II.3 I].4 11.5 11.6

Launch Velocity, km/s

FLIGHT TI_ AND INPACT VELOCITY VERSUS LAUNCH VELOCITY

I000 km Orbit Altitude

100 de

I000 km

II .0

FI_ 3-26.

I I l I I .... I , I ...... I .....

II .I II .2 II .3 II .4 II .5

Launch Velocity, km/s

VELOCITY iBIPULSE REQUIREDTOESTABLZSH

CIRCULAR PARKING 0RBIT rA__N

iATT|LG| _ ClOI. UMiiI.IIi,

/

/ L

.i

i

:../

k-- d' •

LI

: I

I

_J

ti

ri

)]

U
i

t']

It

[]

[j,

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@

3-43

Another parameter which was investigated is the dihedrai angle

between the orbit plane of the Moon and the plane of the Earth's equator,

This angle varies between about 18.8 and 28.59 degrees, but the effect

was found to be insignificant for the vertical equatorial launches, All

data shown are for the larger angle; but the effect Of a smaller angle

would not be discernible rin the plots.

3.5.3 Lunar Landtn Retro Re_utremnts

For direct descent to the lunar surface, the impact velocities

shown in Figure 3-25 are roughly equal to the retro velocity changes

required for soft landings. The actual requirements would be somewhat

larger, but the difference would be small for high-acceleration burns

close to the lunar surface.

An alternative scheme would invoke the use of an intermediate

parking orbit about the Moon. Although the parking orbit mode will require

a greater total tetra capability than the direct descent mode, the

intermediate Orbit may be a practical necessity to achieve a precise

landing at the Site of the lunar base,

Figure 3-26 shows the tetra impulse required tO establish a

circular orbit about the Moon. For a reasonable range of orbit altitudes,

the retro requirement is almost independent of orbit altitude, but depends

primarily on approach velocity to the Moon which, in turn, is a function

of launch velocity. The variation o+ the Earth-Moon d.lstance is not

significant.

To achieve a landing from the parking orbit,, two techniques

could be used. In the first, one tetra impulse would be used to leave

the parking orbit and the projectile would be permitted to impact the

lunar surface in a hard landing. Figure 3-27 shows the relationship

between the de-orbit impulse and the resulting velocity at the surface.

The minimum de-orbit impulse would produce a grazing, horizontal impact

while the maximum impulse would permit the projectile to fall vertically

to the lunar surface.

For a soft landing, the alternative scheme would use an

additional tetra burn just prior to impact, The magnitude of this third

retrO Impulse is approximated by the velocity a% the Surface as shown

In Figure 3-27,

The total tetra requirements for a soft landing are. summarized

in Figure 3-28 as a function o_ initial launch velocity at the Earth,

the parking orbit attitude, and the relative magnitude of the de-orbit

impulse used to depart the parking orbit. For each orbit altitude, the

uppe_ limit Of the band represents the vertical fall to the lunar surface ._

(maximum de-orbit impulse), while the lower bou,d corresponds to the

grazing approach. The accuracy of the landlng would be greatly enhanced •

by using a steep descent to the-surface..

OAT'riLLE -- _Ot. uMmuo

f

i.
3,0 i 3-4_ _ 2.5

0 1000 km Orbtt' "

_ _'___ _ 1.5

_. 1oJ ).._ _Ve,_.l I=,o_

O,0

0 1.0 2.0 3.0 4.0 5.0 6.0

De..Ov'btt impulse, km/s

FIGURE 3`27. VELOCITY AT LUNARSURFACEVERSUS DE-ORBIT INPULSE

¢J

c

5.5"

5.0"

e,5 I

4.0"

3.5 "

>_ 3.0"

4_

o 2.5-

2.0-

I;,,,,

1.sli-

¢J

_- 1.0-

0.5 '--

0

t t t. I " t.....

11.0 ll.i 11.2 1i.3 11.4 11.5

Laun,:h velocity, km/s

1000 km

O_'btt

100 km

O_'btt

FIGIJRE 3-28, TOTAL RETRO IMPULSE FOR SOFT LUNARLANDING

iiATTILL, i[ -- COI,,.tJMIBUli

[i

,(

_-_r _ b

!i

!i

1

!.

.*':,q

,,,,

,°
i

3"45

j

3.6.4 .Launch Windows

As mentioned in SectiOn 3.6.1, there will be two+ occurrences

eaCh mOnth in which the MoOn passes through the Earth's equatorial plane.

In each event, the establishment Of a satisfactory trajectory demands

ClOse control of two angles, _asured at the time of launch.

The first angle is measured tn the plane of the Moon's orbit,

and specifies the Position of the Moon, at launch time, from the ltne

of intersection wtth the Earth's-equatorial plane. The desired angle

ts directly proportional to the fltght time and depends, therefOre, on

the launch velocity.

The second control angle is measured in the Earth's equatOrial+_

plane, and corresponds to the latttude of the iaunch site, at launch

time, relative to the 1the of intersection with the lunar orbit plane.

Thts angle is related to the time uf day, whereas the loaner Corresponds

tO the day of the month.

The second angle (relative to time of day) is more critical

than the first, in that the Earth rotates about 15 degrees per hour and

the tolerance ts relatively small. In fact, the launch window Could

be as small as 4_ minutes without midcourse correction, even if it is

assumed that the launch velocity could be+ adjusted precisely to satisfy

the flight time requirement imposed by the first angle at the time of

launch.

Since the launch window appeared to be _rohtbitively small,

a brief analysts was conducted to explore the possibility of expanding

the window by the _se of a. midcourse velocity impulse.

For the computatlons, the launch velocity {without atmospheric

loss correction) was fixed at 11.4 km/sec, the minimum dlstance to the

Moon (3S6,000 km) was used, and the angle between the lunar orbit plane

and the Earth's equator was set at the maximum value of 28.59 degrees.

To further reduce the number of free parameters, the midcourse Correction

was applied S hours after launch in all Cases. The nominal flight time

for these conditions is about 28,5 hours, so the midcourse increment

was applied relatively early in the trajectory. In general, the earlier

the correction, the smaller the velocity requlrement but the greater

the required preclslon.

The results of the analysis are summarized tn Table 3-7. A

mtdcourse correction of 0.92 km/s expands the launch window to Six one-hour

opportunities per month (two opportunities of three days length). The

mtdcourse AV of 0.92 km/s is in addition to the 3.5 to 5.0 km/s lunar

landing requirement.

!

8AYTEt. LE +" c_L. UMmmu8

_i-'.: -
3-46

TAULE3-7. REPRESENTATIVE NIDCOURSE VELOCITY

INC REqmJl S

Window kltdth (hOurs) ConseCutive NO. Required

Each Day of. Days Increment (Im/sec)

0.5 1 0.37

0,5 2 0.95

0.5 3 0.76

1,0 1 0.61.

1.0 2 0.75

1.0 3 0.92

3.6.5 Conclusions

The follOwing conclusions are drawn for vertical equatorial

launches to the Hoon:

(1) gtth a midCourse correction of 0.92 km/s, st:x launch windows

of one-hour length are available each month for launching

to the Moon. Other windows of various durations are

available by changing the mtdcourse correction capability.

(2) The launch velOctty requirement, ignoring atmospheric

losses,.ranges upward from about 11.1 km/s.

(3) A launch velocity of about 11.3 km/s (no drag penalty

included) woul_ be appropriate to reduce the sensitivity

ofthe flight ttme to uncer_Lainttes _n the launch velocity.

Assuming a ballistic coefficient of 93,000 kg/m ¢, the

launch velocity accounting for atmospheric drag becomes

11.94 km/s.

(4) For practical launch velocities, and for direct impact

hard landings, tmpact veloci¢tes of 3.5 to 4 km/s would

be experienced.

..... ! L

.o

p

iI L

i'1_ '

I

r-.

i ̧

!!

(5) Fop soft lunar landings, the total retro tmpulse

requirements could be in the 4.5 to 5 km/s range tf an

tntemedtate lunar parking orbtt ts used. The parktng

orbit may be a practical necessity for accurate placement

of the payload on the lunar surface..

(6) For a lun_r-orblt destination, the total tmpulse

requirements could be in the i.5 to 2.5 km/s range.

o_

ii

iiATTEL, t..E -- ¢_OI. UMBIJli
1

3-47

3,7 Eart-h-to-GeOslmchronous Orbtt Launch ¥eloct_

This section presents the laun_ velocity requirements for

plactng a projectile iaunched from an Earth based EML Into gecs3mchronous

orbit (GEO). The basic approach used to develop these velocity

requirements is outlined _n Sectton 3,1,3.2 of the previous ESRL Fthal

RepOrt (Rice, et al, 1982). Thisapproach incorporates the following

assumptions:

e Due-east launch from the equator

• Orbit altttude of 35,800 km with 0 degrees Inclination.

Ustng the laws of conservation of energy and conservation of

angular momentum, launch velocttles and angles were calculated for several

trajectory angles at GEO altitude. These numbers were then cOrrected

for the rOtation of the Earth and for atmospheric drag. The results

are shown in Figure 3-29. Launch velocities and corresponding angles

from horizontal (9/) are drawn for several trajectory angles at GEO

a|titude (9) measured from the local horizontal. For the reco_nmended

configuration having a launch angle of 20 degrees from the horito_ta],

the resulting launch velocity ts 11,7 km/s

A propulsion system ts necessary to circularize the projectile

into the required orbtt. The law of cosines determines the necessary

veloctty Increment of the propulsion SyStem aboard the projectile. Figure

3-30 shows the velocity InCrement (AV) necessary for the circularization

as a function of projectile veloctty at GEO altitude (35,800 km). With

a launch ve]Octty of 11.7 km/s and launch angle of 20 degrees, the

projecttle velocit_ at GEO altttude ts determined by the method .deScribed

above and found to be 1.4 km/s. This corresponds to a veloctty Increment

of 1.7 km/s forwhtch the on-board propulsion system must be sized.

3.8 Projectile Concepts

To properly study certain missions, useful payloads must bo

defined. The projectile concepts for each of the four reference concepts

are defined tn this section. Also, two other revised projectile concepts

are derived tn this section. The TRU waste projectile ts derived from

the ESRL Htsston A projectile, wht:le GEO payloads are found by revising

the Earth-orbital projectile.

3.8,1 LEO Pro_ecttles

The basic requirements for the Earth-orbital projectiles were_

defined tn Rice, et al (1982) and are sumar_zed here:

• Haxtmum payload = 650 kg

•Propeiiants consist of hydraztne and

chloHne tr_f]uortde

e N2H4 = 300 kg

e CIF 3 = 850 kg

• Dry propulsion system- 450 kg

• Instruments, ACS, and astrtonicS = 105 kg.

8AT¢I_LE _ GO_UMBMS

i

_.,.:,
lO0

80

v

..,60

m

-&

20

3-48

11 12 13 14

Launc_Velocity, VL (km/s)

FIGURE 3-29. LAUNCH VELOCITY AND AN_LE REQUIREMENTS

FOR $O0--iml CIRCULAR ORI_IT

8"

v

4

0

0

FIGURE 3-30.

2 4 6 8 10

v (km/s)

DELTA-V REQUIREMENTS FOR $O0..Im CIRCULAR ORBIT

eiAT'riLL_.i -- GaLWMOWIk

! i
Ii

i

F

(

t

(

i

(

)

I!

3-49

i

This means that a minimum of 2300 kg must be launched. Remaining

projectile elements would include structure, thermal protection, nose

cone, and fins, as well as projectile coils for coaxial launches and

sabotS for railgun launches.

Figures 3_31 a,d 3-31 illustrate the Earth-orbital projectile concepts

used in this study. The LEO projectiles are more fully described in

Secttons 4.2.2.1.3 and4.3.2.1.3.

_Rocket Pro_

A useful payload of 800 kg would be available for a three-stage

solid-roCket vehicle launched at 2 km/s by an EHL. The total launched

mass fOr thts payload requirement would be approximately 15,000 kg,

Section 3.1.4 described the methods used to select the vehicle mass.

Besides the payload, other projectile elements would include

structure, three sol!d-propellant stages, nose cone/shroud, stabilization

fins, projectile coils (for coaxial system), and rear sabot/a_mature

(for railgun system). The two hybrid EHL/rocket prOjectile concepts

are illustrated in Figures 3-33 and 3-34. Sections 4.4.2.1 and 4.5.2.1

describe the projectiles In more detail.

3.8.3 Lu_ecttle

The Earth-to,orbit projectile (Section 3.8,1) was used as the

baseltne projectile to estimate payload mass for the lunar supply mission.

Available payload and projectfle mass is 2300 kg, Figure 3-35 indicates

the fraction of the available mass which is payload mass for different

values of V/I and propellant mass fraction f. _he figure was plotted

using the following equations:

-- z in -- mpl + '"pS

and f =

where

AV =.required velocity increment, in m/s

I = specific impulse, in m/s

mpl = payload mass, tn kg

mp = propellant mass, tn kg

mps propulsion system (dry) mass, in k9.

Approximately 4.5 km/s must be supplied by the projectile • to

soft-land on the Noon (3,5 km/s for lunar landing and 1.0 km/s for

mtdcourse correction). If the desired payload destination is lunar orbit,

instead of lunar impact, the propulsion system must supply about 2.75 km/s.

From Figure 3-34_ assuming an Earth-storable propulsion system with a

specific impulse of 310 s (3000 m/s), the following table Shows maximum

payload mass for different mass fractions with an available (payload

end propulsion system) mass of 2300 kg.

ItAT3"ISLI. E -- GOi..tJMBMS

f
il

i-

.l

3-50

! 0_. ,.o+,c,,.+ .o. co_"

411TIIIONiCI.

FIGURE 3-31. EARTH-ORBITAL PROJECTZLE (RAILGUN)

l flil

-.e.-

• lOS[

I

fFtm_

FIGURE 3-32. EARTH-ORBITAL PROJECTILE (COAXIAL)

BATTELLE -- COLUMBUS;

i
I.!

I.!

(:

I

(

(

(-

i

FIGURE _33.

3-51

THIRD

HYBRID P,AILGUN/ROCKET PROJECTILE

FIGURE 3-34. HYBRID COAXIAL/ROCKETPROJECTILE

BATTELLE "_ ¢OL.WMBMB

• ..,,....

!L_

I'L
AVII

1,0"

O.!

1

mpk/mT

FI_P,E 3-35. PAYLOADMASSAS A FUNCTION OF AV/I AND

PROPULSIONSYSTEMMASSFRACTION

IATTILLi.-- C 0 1. I,J M l Ill I

:I

il

Ii

!I

.,;_w

i • , .

.e

A
!

3-53

U

Mass Fraction, f Paload k

0.7 -- 325

0.75 -- 460

0.8 70 565

t!

optton.

becomes:

Lunar orbit Insertion appears to be the most feastble mtssion

If a mass fraction of 0.75 is selected, the mass breakdown then

t

 lli

(,

I

(i'

k

e Payload = 460 kg

e Propellants = 1340 kg

e Dry propulsion system = 460 kg.

The remaining projectile elements are Identical tO those deftned for

the Earth-to-o.rbit projectiles,, whtch were designed for accelerations

of 2500 g's.

3.8,4TRU Waste Projectile

The ESRL Ntsston A projectile (Rtce, et al, 1982) was used

as a baseline for determining the useful payload for a TRU waste launch.

Treated TRU wastes ace less dense than HLW in cermet form (2.03g/cm 3

vs, 6.5 g/cm3) so the ortgtnal projectile must be lengthened to have

a meaningful payload. At the same time, however, TRU wastes have less

stringent shielding requirements, so the shteld thickness may be reduced,

leaving more volume available for the TRU waste payload.

The revtsed projectile was calculated to have the same diameter

as the HLW projectile (51 cm), but is lengthened by 40 cm. The radiation

shielding is reduced by half to 6 cm. Wtth these alterzttons, the TRU

waste pay!Oad ts found to be 285 k9 per projectile.

3.8.5 GEO Projectile

A launch from Earth to geosynchronous orbit requires less

orbtt-clrcularizatton propulston than a launch from Earth to 500 km because

of smallee orbttai velocities (3.1 km/s compared with 7..6 km/s). Thts

dtrectly results tn a largerGEO payload capability for the same projectile

mass. The payload capability was calcu]ated as followS. Using the same

basic projectile that was defined for the ESRL Study (see Sectton 3.1.5.2

of Rice, et al, 1982), 2300 k9 is again avai|able for the projectile's

payload and propulsion system. From Section 3.7, the required propulsion

system AV is 1,7 km/s, The specific impulse 1s assumed to be 3000 m/s,

with the following system definitions:

e ClF3andN2H 4 propellants

e Oxidizer to-fuel ratio of 2.8

|A'r'r|l.L | -- COI. UMiiU|
P,

3-54

e Nozzle expansion ratio of 14.0

e Chamber pressure Of 100 N/cm 2 (150 psi).

where

The propellant requirement was. determined from

AV = I In ( mt

mt-mprop

mt

mps

mprop

= mps+ mprop + mpl

= mass of dry propulsion system

= mass of propellants

mpl = payload mass.

The required propellant mass.is then 995 kg. Thus, 735 k9 of C1_ 3 and 260 kg of N2H4 are necessary for orbit circularization. For a 50 second

burn time, the thrust level is defined to be 60,000 N (14,000 lbf).

The mass fraction of the. propulsion system was assumed to be

0.7, which corresponds to a total propulsion system mass of _420 kg and

a dry system mass of 425 kg. Therefore, the resulting payload mass to

-GEO is then 880 kg which is 1.35 times greater than the baseline ESRL

payload capability of 650kg .....

3.9 Nose Cone Natertal Selection

This section addresses several issues which are critical to

the selection of a material for the nose cones of the EML projectiles.

Two major concerns are discussed here: ablatton effects and economics,

3.9.1 _J)latton Considerations

If the EML .projectile is to survive its flight through the

atmosphere, the nose. cone material must not substantially ablate away.

The ablation effects are especially severe for the solar system escape

and planetary missions, where launch velocities exceed 15 kan/s.

Fi

¢

:(

b :

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The ESRL projectiles used tungsten for the nose cone material

(Rice, et al, 1982). This was determined, to be. costly, and it was

recommended that, if at all possible, steel should be used instead.

Estimates of nose cone recession due to ablation were made

by three aerodynamics experts in the previous ESRL study. These estimates

ranged from several centimeters to half the projectile diameter fo_ a

launch velocity of 18 km/s.

;]

L1

L.)

For a final honors thesis at the Ohio State University, A.

Kerslake (Kerslake, 1982) investigated this problem, studying the effects

of ablation on steel, tungsten, and graphite materials. Her results

indicated that all three materials were acceptable for the ESRL projectile

with total nose cone recession of less than 10 cmj for launch velocities

liiATTtL. LIE -- GQI. UMBU8

i
3-55

beiow 17 kmls, Above 18 km/s launch velocity, all of the materials show

Increasingly large recessions, Steel proved to be the least desirable

material and graphite the most desirable.

Early tn this study, tt was decided to use Steel for the nose

cones of the EHL projectiles, This dectston was made wtth some concern

regarding the accuracy of ablatton estimates for the high-speed projectiles

(solar system escape and planetaw). As the study progressed, the

reference concepts were chosen. None of these concepts used launch

velocities of greater than 12 km/s. Therefore, it was decided to continue

using steel as the material of choice-for the projectile nose cones.

3.9.2 Economic Constderatton_

The tungsten nose cones for the ESRL nuclear waste disposal

and Earth,orbital missions (Rice, et al, 1982) have masses of 440 kg

and 1150 kg, respectively. In 1982, tungsten metal Cost approximately

$33.00 per kilogram, and tn Hay 1983, a pure powder from which tungsten

metal would be made cost $11.14 per pound giving an tmp]tcit price of

approximately $25.00 per kilogram for tungsten metal. The nuclear waste

disposal nose cone was estimated to cost tn the range of $30K to $80K

(nominal of $50K) based upon labor factors of 2, 3.5, and 5 to manufacture

the nose cone. The Earth-orbital nose cone was costed at $40K to $250K

(nomtnal $185K) wtth similar factOrs. The low estimate was then adjusted

to be sltghtly more than the materials cost of $38K under the assumption

that additional uses would be ftund for the metal, Including pOtential

reuse.

Since the drop tn the price of tungsten is .indicative of 1;he

state of the economy rather than in new production techniques, there

is no basis for major revisions to the estimates for tungsten nose cones.

If a high strength, low alloy, structural shaped steel is

adequate for Earth orbital missions, it would have a significant materials

cost advantage ($0.239 per pound or $0.55 per kilogram), In addition,

it is more easily, worked than tungsten. A steel nose cone with a mass

of 1150 kg has a materials cost of $632.50. If it were batch-produced

to O.Ol-cm tolerances, tt quite l(kely would cost under $1,000 and almost

certainly under $2,500. If a stainless or other high-alloy steei were

required, the prtce vartance would be great. The lowest-priced stainless

. steel is currently approximately $I.25 per kilogram and T-i5 high, speed

tool steel is quoted at $21..65 to $25.57 per kilogram. Thus, T-15 steel

costs approximately the Same as tungsten.

It is dlfflcult tO detemlne costs for carbon-carbon graphltlc

materials under consideration for the nuclear waste disposal and other

Earth-escape missions in the time under consideration (beyond the year

2000). At the present time, carbon-carbon graphltics can cost thousands

of dollars per kilogram in final shapes. The cost lies not in the raw

materials (which are comparable to Iow-prlced steels), but in the

processing and quality cOntrOl labor necessary to make the final product.

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To achieve the ultt_te properties assOciated with advanced Carbon-carbon

materials, it is necessary to control the qua]try of the raw materials

from whlch the ftbers are made, and then to Control, at each step, the

spinning of fibers, weaving, shaping, pyroltztng, and final finishing.

In the time under consideration, it ts reasonable to expect that

carbon-carbon graphittcs, or other materials wtth Similar properties,

will be available at prices less than tungsten for equivalent applications.

Thus, carbon-carbon graphtttc nose cones are estimated to cost about

the same as tungsten nose cones.

3.10 Projectile StreSs AnalTsis

Calculations were made to determine the strength and stabiiity

of the ESRL Earth-orbital projectile. The compressive strength and the

buckling stability were analyzed.

The critical cross-section of the payload support structure

(PS$) for the compression and buckling analysis was selected as the

PSSwas just below the lower end of the forward sabot. The carbon-carbon

thermal protection and the oxidizer tank .wall were assumed to be

nonstructural (a conservative assumption). The applied force across

this section was calculated by summing the individual component masses

ahead of the crtttca] cross-section (instruments, astr.ionics, ACS,

propulsion system wtth propellants, nose cone, and fo_ard sabot, which

total 2950 kg) and multiplying by a 2500-g acceleration. This results

in an appiied load of 71 x IOO.N. The applied axtal compressive stress

ts calculated as:

F

where F is the applied force and A is the area over which the force is

applied. 8 The applied compressive stress was then caiculated to be 8.11 x 10. N/m 2 (117,700 psi) at the crittcal location. This stress

value is below typical yield stresses for s2atnless steels (typically

or the order of 1.035 x 10_ N/mZ); the ratio of yield-tO-applied stress

ts then 1.3.

The crttlcal buckling compressive Stress level at the selected

section was based upon JEuler's ,column formula for a column with one fixed

end:

_2 E

Ocr -- - (L/r)2

where

r x = ry = I/4 (Do 2 + 012)1/2.

E is the modulus of elasticity (for steelS, E = 2.07x 101! N/m2), r

is the radius of gyration about the bending or buckling axis, Oo is the

outside diameter of_the PSS, and Dt is the interior diameter Of the PSS.

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The effective column was treated as fixed end, because Of the iarge steel

mass in the lower section of the PSS. The crttical elastic buckltng

stress was computed to be 15.9 x 109 N/m2 (2.31 x 106 psi), aSsumtng

the Euier formula for column instability. The calculated buckling stress

was found to be 19.6 times greater than the applied compressive stress.

(1957) :

The local buckling stress was then

OCC __.

calculated from Shanley

&

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In this equation K is approximately 0.5 for this case, Et is the tangent

modulus, and t is the gwail 2thickness.5 The local buckling stress was calculated to be 3.7 x 10 N/m (5.4 x 10 psi), which is 4.5 times greater

than the applied stress. These results indicate that the structure would

be unstable at a level well above the compressive ultimate strength of

most stainless steels, so buckltn9 loads do not appearto be a problem.

It was recommended .that further detailed design should include

a structural analysis which would include .composite structural effects,

structural fastener effects.,thermal effects, and pressure loading,

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4.0 REFERENCECONCEPTS

The. four reference concepts are ldescrtbed in this section.

The reference EML concepts were selected; from those described in.the

previous sectionand are listed below:

, Ratlgun for Earth orbttal launches

, Coaxial accelerator for Earth orbit:.! launches

, Hybrid ratlgun/roCket launcher

, Hybrid coaxial accelerator/rocket launcher.

The concepts described in Section 3,1 are very preliminary and would

require _urther analysis to define the optimal detailed system designs.

The concepts are sufficiently detailed, however, to perform a comparative

assessment based upon costs, performance, technology risk, end other

factors to select themostpromtstng conceptsforfurther analysis.

This section is composed of five subsections. The first

subsection describes the evaluation process which led to the reference

concept selections. The reference ConCe_t_.Are-_e_aiied in the fol_owfng

four subsections.

4.1 Preltmtnar..T,y.Evaluat_onof EMI. Conce ts

The evaluation was based upon a combination of ten quantitative

and qualitative factors. The relatt_e importance of each of these factors

was established so that weighting factors could be assigned to each.

The EML concepts described in Section 3.1 were then rated for each of

the screening issues. EML concepts were given scores of 1 to 3 based

upon comparison with "conventional" launch methods planned for the 2020

ttmeframe. A score of 1 indicated the Concept was deficient when compared

to conventional methods. Concepts rated equal to conventional methods

for a particular screening issue were given a score of 2. A score of

3 indicated that the concept was ranked better than conventional launch

methods. A total score was computed for each concept, by multiplying

the score for each screening Criterion by its respective weighting factor

aad then summing across the board. The cOnCepts were ranked according

to their total scores, wfth selection following based upon these results.

Q.

The screening criteria and their relative weights were:

e Development cost (5)

• Operational cost (S)

• Total cost (30)

e Launcher technology rtsk (10)

, Payload technology risk (5)

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• System flexibility (5)

• Performance (25)

, Logistics (5)

• Safety (5)

• Environmental impact (5)

Ftgure 4-1 dtsplays the results for" reilgun concepts, while

coaxial accelerator results are shown in Figure 4-2. Based upon these

rankings, it was decided by Battelle and NASA/t.eRC to proceed with the

Earth-orbital mission and the hybrid EML/rocket cOncept for both the

ratlgun and coaxial accelerator concepts. These concepts were then

more fuily developed, and are described in this section of the report.

(The lunar base supply mission was added to the list late in the study,

so it was not fully develOped into .a reference concept, despite its

high score.)

4.2 Earth-to-S_pace Rat1 La_

4.2.1 Conce t l)efinttton

The Earth-to-space rail launcher concept for launching cargo

into low-Earth orbit consists of ftve major pre-launch, launch, and

post-launch activities which are discussed in this, section. The five

activitiesare:-

(1) Projectile/payload fabrication

(2) Surface transport of projecttle/payioad

(3) .Projectile/payload preparation at the launch site

(4) Launch operations

(5) Trajectory monitoring..

Definitions of the individual rail launcher system elements are presented.

in Section 4.2.2.

4_ Pa load Fabrication

The payload, projectlle,, and the orblt-circularlzatlon

propulsion System are manufactured and checked-out prior tO system

integratiOn and transport to the launch site.

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The method of surface transportation would depend upon the

selection of the launch site. If the selected site were remote, ship

or atr transportation would 11kely be used in areas where roads Or

railroads, were inaccessible, Otherwise surface transportation would

likely be accomplished using trucks or railroad caps..

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4.2,1.3 P____ecttIe Pa load Pre ration at Launch Site

Upon arrival at the launch site, the projectile/payload would

be placed in a storage fdcili)y. The projectile and payload would be

removed from storage as the scheduled launch time approached. The payload

would receive a prelaunch checkout, if .necessary. Propellants would

be loaded just prior to launch.

4.2,1.4 LaUnCh Operations

At launch time, the projectile is loaded into the preaccelerator

and all launch systems are checked out prior to launch. The launch

tube is evacuated and the homopolar generators are started.

Weather and wind conditions would be _ checked before final

countdown begins. Launch clearance would be requested from the proper

authorities and all pilots in the area warned. An alarm would be sounded

in the area of launch, so that all persons are cleared from a.designated

danger area.

When a]l systems are ready, the launch sequence and final

countdown begins° All launch systems _re computerized and fully

automated. The preaccelerator is initiated, and the projectile is

accelerated to I km/s through the preboost se)ction. As the projectile

passes through the preaccelerator into the railgun section of the launch

tube, current is automatically dumped into t!_e first rail section. A

plasma armature is formed behind the projectile. Switching of the current

into the segments is performed automatically as the projectile passes

through each section. The projectile is acce_lerated to 6.85 knVs when

it leaves the rallgun bore, Tracking Systems on the ground would be

used to verify the trajectory after the projectile has left the railgun

system.

4.2.1,5 Tra_ectoryMonitoring

The projectile is. tracked throughout its atmospheric flight

with a small radar system located near the laur,cb site. Before and

after the orblt-circularization maneuver, the projectile's three-axis

attitude control system would ensure proper projectile attitude on orbit.

The auxiliary propulsion system of the projectile would provide the

2.1 km/s necessary to. insert the payload into the proper orbit. The

payload would be taken to a Space station by the Orbital Maneuvering

Vehicle (OMV), if it. were available, or another system of this type.

4.2.2 System Eleme, t Definition

(

i

L This section presents the Earth-to-space rail launcher system elem,_nt de¢inlt!on_. The major system elements which are discussed

here are:

e Projectile/payload characteristics

e, Surface transport systems I

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4-6 i}

• Launch site support facilities

e Launcher system

• Monitoring systems

e Space destination.

4.2,2.1 Pro ecttle Pa load Characteristics

t

The payloads which are envisioned for the Earth-to-space rat auncher are bu]k-type payloads to support manned

orbiting space stations. These payloads include:

• Orbit Transfer Vehicle propellants

• Space station supply items

• Materials for space processing facilities,

Liquid hydrogen and ltquid oxygen propellants required for

Orbit Transfer Vehicles (OTVs) could be transported to space stations

in the form of water. An orbital el ectrolyzer would then be used to

transfora the water payload into the oxygen and hydrogen propellants.

The rail launcher system could also be used to supply otherpropellants

to orbit. An example of this would be a water or hydrazine payload

to support space station attitude control and drag make-up systems;

this would however be a much smaller supply requirement than are the

OTV propellants.

Space station resupply items which could be: launched by a

railgun system might include life support requirements (food, oxygen,

and nitrogen, for example), oxygen and hydrogen (in the form of water)

to supply fuel cell make-up requirements, spare parts for _tation

maintenance and emergency repairs, and other miscellaneous supply items.,

such as crew personal equipment, hygiene supplies, and ship stores,

The supply items must be able to withstand the high accelerations of

launch (1225 g's to low-Earth orbit); other delicate materials would

still require a launch by Space Shuttle.

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When materials-processing facilities are operational in space,

the raw materials necessary for product manufacture could be transported

to the orbital facilities by the rail launcher system.

The payload was determined to have a maximum mass of 650 kg

in the analysis done in the previous ESRL study. The 650 kg maximum

figure is for materials wlth the density of aluminum (2700 kgZm_) or

greater. For payloads with the density of water (IOOO kg/m j), the

correspondingmasS is 320 kg due to volume constraints.

_4.2,2.1.2 propuls!on S stm. A propulsion system is required

for the Ea_tal mission to place the payload into- orbit.. The

propulsion system assumed here is the Sah;_ one defined in the ESRL study

(Rice, Miller, and Earhart, i982). The basic characteristics of the

system are summarized here. A simple hypergolic, hlgh-propellant-density

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propulsion system was indicated; hydrazine and chlorine trlfluortde

propellants with an oxidizer-to-fuel - ratio of 2.8 were selected.

Propellant mass would be approximately 1150 kg (300 kg of N2H4 and 850

of C1F 3) with a dry propulsion system mass of 425 kg. System attitude

control and astriontcs requirements amounted to 75 kg. The specific

impulse of the auxiliary propulsion system was estimated at 310 s

(3000 m/s),,

The cold-gas attitude control system (ACS) would perfOrm the

maneuvers required to place the payload and propulsion system in the

proper position for the orbital insertiOn burn, The propulsion system

would provide the 2100 m/s necessary to circularize at500 km altitude,

after which the ACS would again be used to ensure proper attitude of

the payload for rendezvous with the Orbital Maneuvering Vehicle or other

vehicle used to transport the payload to the spacer station.

4.2.2.1.3 ProJeF_tle Ele!,nt_. The Earth-to-space

launcher projectile would consist of the following subsystems:

e Forward and aft sabots

e Nose cone

e Instrument package

e Liquid propulstoq, sy.stem, includin 9 ACS and astrionics

e Payload

e Payload support structure (P$S).

e Thermal protection system (TPS)

e Fins.

rail

The proposed projectile is illustrated in Figure 4-3. Table 4.-I presents

the projectile mass summary.

The forward and aft sabots are required to fit the round

projectile to the square bore. The aft sabot also protects the rear

of the projectile from excessive heating from the plasma armature. The

sabots are jettisoned immediately after launch so as not to detract.

from the aerodynamic characteristics of the projectile. High-strength,

non-conducting ceramic materials would be usedto construct the sabots.

The nose cone would be constructed of steel. The tip would

be slightly blunted so that the steel would evenly and smoothly melt

during atmospheric flight. As discussed in Section 3.9, the amount

of erosion during flight is not expected to be significant for the 7 km/s

launch velocity, given, the dimensiot)s of the nose cOne.

A small instrument package would be located beneath the nose

cone, The package would include a radio transmitter for trajectory

verificatiOn after leaving the rail launcher,

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HIGH.STRENGTH STEEL

SUPPORT

STRUCTURE (lIES)

GRAPHITE THERMAL

PROTECTION

Length -" 3.6 m

Projectile Diameter : 0.9 m

Sabot Diameter : 1.0 m

FIGURE 4-3. EARTH,-ORBITAL PROJECTZLE FOR

RAiL LAUNCHER SYSTEM

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TABLE 4-1., EARTH-TO-SPACE RAIL UUICHER PROJECTILE RASS SUIII_RY,

.... m_mmm .....

Project 1le Subsystem Mass .(kg).

Payload 650

Propellant 1150

ropulston • system 42550

AstrlonlCs 25

Instruments 30

PSS 2730

TPS 100 .,-

Nose cone 420

Fins 20

Forward sabot 200

Aft sabot 100

Total 5900

The payload support structure (PSS) made of high-strength

steel would provide structural support for the payload and propul.sion

system. The propulsion system would be located in the forward part

of the PSS; the nozzle would point toward the nose cone. The PSS would

also support the propellant tanks during the high-acceleration launch.

The payload would be located in the aft portion of the PSS, attached

to the propulsion system. The PSS would be jettisoned before the

orbit-insertion burn.

Another advantage of using the square bore is that the fins

can be attached more simply to the projectile, no "pop-out" mechanism.

is required. Four fins are attached to the rear of the projectile to

stabilize the projectile during the atmospheric fliqht.

4.2,2.2 Surfac_rtatton Systems

The payload, propulsion system, and projectile would be

manufactured and assembled in factilties away from the launch s_t:.

The systems would be transported to the launch site by conventio,Jal

surface transportation systems: truck, rail, aircraft, or ship. Aircraft

could be used to transport required personnel and.high?priority materials

to the launch site directly,

tnciude:

general

4.2.2.3 Launch..St te Su_ _

Launch support facilities to be located on the launch site

power plant, projectile storage and check-out facilities,

storage facilities, administration and engineering facilities,

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4-10

Industrial area, co_untty living area (if the site is remote), liquid

gas and water plant, and other miscellaneous facilities. The ;aunch

site should be selected in an area where there are no large population

areas withtn a radtus of approximately.50 Ion. Figure 4-4 is an arttst's

concept of the launch stte.

4.2,2.3,1 P r Plant. A dedicated nuclear power plant was

envlsloned'-_-'___'_rlcal power requirements of the rail

launcher system and its supporting systems. The power plant would consist

of four 50-HWe nuclear reactors similar to those used on Navy ships.

The number of power plants is higher than nominally required, because

redundancy has been considered for maintenance and unscheduled shutdowns.

4_ut Facilities. A

projectiie stor;ge ;rid c ........ t_ store

at least 70 projectiles (one week supply at a launch race of ten per

day) from time of arrival at the site until the projectiles are moved

to the storage facility located at the breech of the. rail launcher.

4..2.Z.3,3 General ....Storaqe Fa¢!lit!es. General storage

facilities are needed to support the activities Of the staff and the

operations of the launcher system. Office supplies, gasoline, and rail

l_uncher spares are am¢_lg the. items which wOuld be stored.

4.2_,2.3,r,_ A.dmtnistratton and EngiBeertng Faci!jties. Office

space would be required for the administration and engineering staff.

These. buildings would be located near the industrial area.

4,.,_2.3r5 Znd.s%Ha] A_a., Various industrial facilities

would be required to support the launch activities, including receiving

areas, repair and refurbishment shops, vehicle maintenance, and other

facilities as deemed necessary. Aircraft runways would be built to

support incoming aircraft traffic of people and supplies. The industrial

area wouldbe located in an area approprlat_ to the activity.

4.2.2,3.6 _nfty Llvtn. Are). If the launch stte were

located in a remote area with no nearby towns, a community living area

would have tO be built, at a practical distance from the rail launcher

facilitieS. The living area would likely include apartment-type housing,

schools, and some shopping and entertainment facilities to support the

launch crew and their familieS.

4_2.2.3.7 Liquid Gas la_d _al_r Plant. Liquid nitrogen ts

required to coo_ the 3600 tnductors and the preaccelerator requires

liquid hydrogen and liquid oxygen. A water supply is necessary for

the hydraulic operation of the homopolar generators and for space Station

supply payloads, as well as for iau,,;her base and personnel supplies.

A plant Is needed to support these requirements, and would likely be

located close to the power plant. Liquid nitrogen lines would be used

tO send the LN2 to the-launcher system, while LH2 and L02 would be

transported by truck. Water would be distributed throughout the launch.

site by an undergroun_ plumbing system.

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tnclude:

Other necessarysupport facilities

• Electrtc_to-hydraultc factltty

• Launch control-center

e Radar traCklng facility,

4.2.2.4 u.cher._Sys •

The Earth-to-orbit rat1 launcher system would accelerate the

projectile to a 6.85 km/s launch velocity. The launcher ts based on

a mountatnstde at an elevation angle Of 20 degrees from the horizontal.

Access for maintenance and repair has been provided, Figure 4-5 shows

a cross,sectional vtew of the rat1 launcher system, whtle Figure 4-6

presents a top view. The ratl launcherwould have a square bore, 1 m

across, The rails would be fabricated of AMZIRC (a copper-zirconium

alloy) and would be electrically Insulated and spaced with a non-asbestos

fiber-reinforced insulator material. A containment tube constructed

of Kevlarwould confine the rails and tnsulatormacertals.

The ratl launcher system would be powered with 3600 homopolar

generators with corresponding tnductors, which wOuld be evenly distributed

alon the length of the launcher. Self-activated switches would control

the current distribution from the tnduc¢ors tnto the rails.

The following subsections briefly discuss the rail launcher

systems which have been conceptualized. Launcher systems include:

e

e

e

e

e

e

e

Bore and ratls

Energy storage

Launcher support structure

Preaccelerator system

Switchtngand control

Temporary projectile storage facilities.

Service and access systems.

Launcher concept options are provided tn Figure 4-7.

4.2.2.4.1 Bo and Rails. The rat1 launcher would have a

square bore wh ch is 1 m across, he rails wou]d be 2040 m tn iength

and would be constructed Of AHZIRC, a copper alloy which ts approximately

99.9 percent copper and 0.1 percent zirconium. AHZIRC was selected

for the ratls because its strength and conductivity are better than

cooper alone.

TO calculate the mintmum distance between the rails, the

pressure applted at the base of the projectile is equal tO the pressure

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exerted on the rails. The yield strength of AMZIRC ts 4Z,100 N/cm Z

(Engineering Alloys Dtgest, 1961). Because the _orce on the proJectt.le

is 70.9 MN, the-resulting mintmum bore area is 1684 cm_,. With a square

bore, the mtntmum r_tl height ts 41 cm. This is not a preblem since

the bore must be 1 m across to accommodate the projectile. The larger

bore stze ts desirable for several reasons. For stress reasons, the

rail-launched proJectt_e should be as short as possible, whtch widens

the prOjectile diameter. Also, payloads are not as size-constrained

when the projectile is widened. Another reason for making the bore

larger ts the lower resulting pressure on the ratls which makes it easter

to. hold the ratls flat during launch. There ts a disadvantage to the

larger bore diameter, however, in.that the magnetic _teld between the

rails decreases as the distance between them increases.

The rails are separated by Insulator materials. The tube

ts wrapped in Kevlar to contain the tube stresses. Figure 4-6 illustrates

the. bore, rails, t.nsulatton, and spacers.

4,2.2.4,2 Ene _ Stora e. Several energy storage devices

were considered or use on t • ra auncher system. Homopolar generators

tn the 50 to 60-HJ range should be available in the next ftve to ten

years (telephone conversation with Dale Pryor, OIH£, 1984), A 50-HJ

HP6 shOuld have a mass of 13,000 kg (Harshall, 1984). A.single HPG

of the size required (200 GJ) would not be available in the near future

and was not considered further in this study.

The HPG/tnductor unit developed by Richard Harshall (Section

3.2 of Rice, et al 1982) was selected as the energy storage device and

is Illustrated in Figure 4-8. The deviceconststs of a 56-HaHPG coupled

with an inductor which stores 48 MJ at a current of 4HA. The aluminum

inductors should have a mass of approximately 1 to 1.5 HT each. The

inductors would be cooled to liqu!d-nitrogen temperatures because of

mass and volume efftciencles.

The energy required per unt_ length is 70.9 Ha/m. Assuming

a 72 percent energy transfer efficiency (85 percent from HPG to inductor

and 85 percent f_om inductor to ratlgun), 3585 HPG/tnductor unitS would

be required for the desired 7kra/s launch velocity. The HPG/inductor

units would be placed at 0.6 m spacing along t_e launcher tube.

4.2.24. Launher Surl;StructUre. The launcher support

structure as envsn_ gure 4-5. The launcher tube would

be partially imbedded in a cOnCrete foundation to contain a forces during

launch operations and to maintain the tube's position along its 2 km

length. The massive HPG/tnductors must be .supported as well. They

are showntn the figure supported by a half cylinder steel structure.

No structurai analysis has been conducted to support this preliminary

concept.

4_rea_tor Ssf;_z_ The preaccelerator concept

has not oeen canged from the prey ous ESRL study (Rice, Htller, and

il)

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Launcher Bore (67 x 67_a)

....... II I

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Swlcchin8 Units I duct ,

FI_ _8, __R &ENEMTOR

HYDROGEN/NITROGEN

MIXTURE

MOVABLE

.,.o. \

COMBUSTION

HIGH PRESSURE

REMOVABLE

END SUPPORT

AND INJECTOR _ AFT SABOT

h

HIGH

PRESSURE

LH 2

LO2

INJECTOR

FIGURE 4-9, PROPOSED.PREG00ST CONCEPT

iiAYTIIL'LII "' COLUMIIiJi
4-18

Earhart, 1982). The system ts requtred to prevellt eroston of the ratls

caused by the dwei1 time of the plasma amature during the intttal

acceleration. Preboost systems currently used accelerate pro_ectties

to 0,5 to 1,0 _/s with a htgh pressure (3000 5000 psi)heltum gas

injector (Hawke, et el, 1984 and O'Aoust, et al, 1984). Hellum Is

generally used because of its high speed of Sound (approximately 1 _/s

at room temperature).

The system suggested here (Figure 4-9) uses 11qutd hydrogen and liquid oxygen tn continuous high preSSUre combustion fOratng a ptS_on

against a gaseous mixture of hydrogen and nitrogen in between the piston

and sabot of the projectile. The concept is slmilar tO a gas gun, except

that tt is cOnttn,ously driven by the LH2/LO2 cOmbuStion procesS. The

preboost system would accelerate the projectile to 1 km/s before entering

the rat1 launcher.

4_,Z,2.,_.5 S_i.¢_h!,g. and..C_nt_|. Switching of the large

currents required for the rat1 launcher System is a major issue which

must be resolved before development of the system is begun. When the

HPG is fully Charged, the energy is transferred to the inductor.

Switching from the Inductor to the ralis would be activated by the

p_ojectile movement down the launcher. When discharged, the energy

stOre must be electrically disconnected from the launcher to prevent

reverse flow of energy. One possible switching mechanism using

chevron-shaped rat1 segments Insulated from each other is discussed

in Harsha11 (1984) and Rice, et el,. (i98E).

4 2.2,4.6T rat Pro ectile Star e FacilitieS'he At temporary Storage launcher.

Projectiles would be moved datly from the storage and checkout facility

tO support the day's schedu]ed launches. The facility would need to

store at least 15 projectiles.

4_2,2.4,? Service _md Access System. Crane and transport

ratls are Included in the reference concept to allow servicing of the

launcher systems, A rat1 System was chosen to facilitate movement-along

the 20-degree incltne of the system.

4.2.2,5 Monitoring System

Trajectory monitoring would be needed after the projectile -_

leaves the muzzle of the rail launcher for range safety purposes.

Telemetry wouid be included as part of the projectile package so that

the payload could be. tracked on orbit for retrieval., by vehtcles from.

the space station.

4.Z.2.G Space Oesttnatton

The destination for projectiles launched from this rail: launcher

system ts lOw-Earth orbtt, The primary destination would be a 500-km

circular orbit in which the space station system is based, but other
0

• ._ ..ww_,....*_.*,._ _- -_ocmk ql

0

I;

4-19

orbits _are possible. If the system is upgraded to provide for 12 km/s

launch velocities (see Section 3,7), launches to geosynchronous orbit

would be p0sstble. ,

4,3 E cceIerator Coacet

Oeftnt tt on

The Earth-to-space coaxial accelerator concept for launching

materfals to Earth orbit consists of five major operational activities:

(1) Projectile/payload fabrication

(2) Surface transport of projectile/payload

(3) Projectile/payload preparation at the launch site

(4) Launch operations

(5) Trajectory monitoring.

The activities are the same as described in Section 4.2.1 for the

Earth,to-space tall launcher, except for the launch opera_lons activities

which are described here, Individual Coaxial accelerator, subsystem

definitions are given in Section 4.3.2,

The projectile is removed from the storage facility just prior

to its scheduled launch time. All. launcher systems are checked out

and the projectile is loaded into the smai] preaccelerator required

to initiate projectile motion. The Brooks coil inductor is charged

and the launcher tube is evacuated.

Final launch countdown begins after weather conditions are.

checked and launch clearance is obtained from the proper authorities.

All persons are clearedfrom the immediate launch area.

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4..3.2 S_fstem Element Definition

Six major system elements have been identified and are discussed

tn this section. These system elements are:

• Projectile/payload characteristics

e Surface transport systems

e Launch site support facilities

• Launcher system

e_ Monitoring systems

e $p_ce destination.

4.3.2.1 ProjeCtile/Payload Characteristics

4._:2_1.1 Pa_loa d. The payload would be identical to that

for the Earth to-space rail launcher, described in Section 4.2.2.1._.

mATTjL. Lm -- COLUM'_Ua
f

4-20

4.3.2.1.2 Pn_ u]sfon S t4m. Propulsion system characteristics

are the same as those described in Section 4.2.2.1.2 for the

Earth-to-space rat1 launcher.

4_,2,1,3 _ ecttle. The proposed projectile is shown in

Figure 4-1_-an consi.sts of the following subsystems:

e Projectile coils , .....

e Nose cone

e Instrument package

e Liquid prOpullion system, including ACS and astrionlcs

e Payload

e Payload support structure(PSS)

e Thermal protection System (TPS)

e Fins.

i!

(

The projectile subsystems are simtlar to those described in

Sectic, n 4.2.2_Z.3 for the Earth-to-orbit rat1 launcher, except where

noted below.

Forty projectile coils (actually copper rings) would be spaced

every 16 .cm along, the projectile, The rings would be 2-cm thick with

an. outer diameter of 49 cm, Each ring would have an axtal length of

3 cm. The coils wouid be tmbedded in the carbon-carbon TPS material.

i

• i

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1

Because of the longer and narrower shape of the projectile,

the nose Cone mass is smaller than that of the railgun projectile. The

steel nose cone would, have a mass of 110 kg (versus 420 kg for the rail

launcher pr(_jectile), The projectile mass summary is shown in Table

4-2.

f.-.

,f

4,3.2,2 Surface Transportation S]stems

Surface transportation systems would be the same as those .

described in Section 4.2.2.2 for the Earth-to-orbit tall launcher concept.

Depending upon the launch site location, trucks, rail, aircraft, or

ships could be used to transport, projectiles, payloads, personnel, and

equipment.

4.3.2.3 Launch Site Support Facilities

Launch site support facilities would be similar to those

described in Section 4.2.2.3 for the corresponding rall launcher launch

site. The facilities include: power plant, projectile storage and

general storage facilities, administration and engineering offices,

industrial area, community living area for a remote launch site, liquid

gas and water plant (LN 2 tO cool Brooks coil inductor; water for supply

and launch site use), and other necessary facilities. An artist's concept

imA'rT*JL.I.f -- ¢:OL. UlVlaUS

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4-21

_ STB[L ,. NOSE

CON[ PACKAEE

EN6INE

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TANK

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TANK.

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Length = 7.0 m

Dtameter= 24.7 an

, PROJECTILE

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-LO, O_.L_ \ _!_1_/

/I n.s II_l I\ STEEL SUPPOR_T P_CTRmzLN

STRUCTURE

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FIGURE 4-10. EARTH-TO-ORBIT COAXIAL PROJECTILE

8A??ILL1 i GOLWMIUi
4-22

of the launch stte factllttes is shown in Ftgure 4-4.

TABLE 4-2, EARTH-TO-SPACE COAXIAL PROJECTILE MASS SUIIMRY

Projectile Subsystem , r,lass (kg)

,l

Payload 650 ,.-_

Propellant 1150

Dry propulsion system 425

ACS 50

AStrtontcs 25 i

w_

instruments 30 "

PSS 300 _'( TPS 200 ::

Nose cone 110

Fins Coils 30.__ o 20 i i

Total 3250

L_

4.3.2.4 Launcher System

This coaxial accelerator system would launch projectiles at

6.85 km/s to supply materials to a 500-km Earth orbit. The launcher

would be located along a 2-km stretch of mountain stde inclined at 20

degrees from the horizontal. Figure 4-11 ts a cross-sectional view

of the launcher system.

The drive coils are a continuous helical winding (2040 m long)

made of a copper-alloy material. ElectrJcal Insulation and structural

tube confinement are provided by encasing.the drive coils tn Kevlar.

i-

"t

Energy would be supplted to the coaxial launcher system from

a single large Brooks cot1 storage inductor. Energy would be distributed

to the drive cotls such that ten turns behind each projectile ring would

be acttve at all times. The turns would be switched in and out as the

projectile moves toward the launcher muzzle.

Launcher.systems discussed in this Section include:

e

Drive coils

Energy storage

Launcher support structure

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CROSS-SECTIONAL VIEH

BATTELLE -- GOt-UMEU8 t"
4-24

O

O

Preaccelerator system

Switching and control

Storage facilities

Service and access systems.

Earth-to-orbit coaxial concept opttons are presented tn Figure 4-12.

4.3.2.4.1 l)rf e Coils. The drive coils would be constructed

from one ayer 0 rectangu ar copper-alloy wlre in a continuous helical

winding. The inner radius would be 25.3 cm and the outer radius 26.3

cm. The length of the drive coil section is 2040 m. An active segment

of 8,cm length conslstlng of ten turns would follow each proJeCtlle

ring down the length Of the launcher tube.

The drive coils are encased in Kevlar for structure and for

insulation between the windings. The tube structure does not need to

be as strong as that for the corresponding tall launcher system. This

is because tube hoop stresses are lower due to the induction of the

projectile coils which adds counterbalancing stress.

_.3.2,4_2 Ene_/y _$1:orage. A single large Brooks coil inductor

would be used for the energy storage device for this concept. A Brooks

coil configuration offers the maximum inductance for a given length

of wire used. The dimensions of a Brooks coil are shown in Figure 4-13.

The kinetic energy of the projectile at the muzzle of the launcher tube

would be approximately 80 GJ. Kolm and Hongeau quote efftclencies of

98.9 percent found by dividing the kinetic energy of the projectile

at launch by the supplied energy to the launcher tube. Assuming an

85 percent transfer efficiency from the Brooks coil inductor to the

launcher itself, this corresponds to an energy storage requirement of

approximately 95 GJ. [nfomation supplied by EHL Research, Inc. lmpltes

that energy stored is a function of outside diameter (Appendix D). This

relationship is shown below:

E = 0.0207 Do3 (4-1)

where the energy E is in GJ and the outside dlamete r 0o is in meters.

Therefore a Brooks coil which would store 95 GJ would have an outside

dtameter Of dpproximately 36 m. The Brooks cotl would be made from

aluminum wire. Ltqutd nitrogenwould be usedto cool the inductor.

4.3.2.4,.,3. Launcher Support. Structure. The launcher tube

would be partially imbedded In a concrete foundation as shown in Figure

4-10. The purpose of the structure Is to help contain the hoop stresses

of launch and to add rlgldlty to the 2-km Iong tube to keep it in

alignment.

4.3,2.4.4 Preaccelerator iSystm. A preaccelerator system

was added to the reference concept to provide a small tntttai velocity

(up t t°he 100 m/s) to get the proJecttie moving through the accelerator

up 20-degree tncltne. A large system, such as that- _equtred for

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FI6URE 4-13. BROOKS COIL WITH DIMENSIOIIAL RELATIONSHIPS

the rai1 launcher system to prevent rail eroston, ts not required for

the coaxial accelerator system.

.4,3.,2,4.5 S_tC;htnq and Control. The drive current needs

to be swnchronlzed wtth the motion Of the projectile. Ten drtve cotl

turns are acl;tve behind each projectile coti and must be switched in

and out as the projectile passes through the launcher. Switching iSSues

are discussed, tn more detail tn Appendix 0 and in Section 7.1.

4.3._2.4.6 Temorary Protectt)e Stooge ..... Fa_ctllties. The

projectile storage factllty would be similar to that descr.ibed fn Section

• 4.2.2.4.6 for the Earth-to-orbit rat 1 l_uncher concept.

4.3.2.4.7 Servtce _ndAccess Sjestlm. As In the corresponding

ratl launcher concept,. _a transport rail system, would, be used to provtde

access for maintenance and servicing of the launcher system.

4.3,2.5 Honttortng System

Trajectory nmnitoring and orbital tracking systems would be identical to the Earth-orbital ,atl launcher Systems (Sectton 4.2.2.5).

4.3,2..6 Space Oesttnatton

The primary destination for projectiles launched from the

coaxtal accelerator system would be a 500-km ctrcular orbit. Other

low-Earth orbits would be possible, wtth different projectile auxiliary

propulsion systems. Should the system be upgraded to allow velocities

of 12 km/s, geosynchronous alttt.ude desttn_tt:ons would be possible.

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_rtd Ra|lu_gg__Packet Concept

4.4.1 Conce_

This sec_ton discusses the hybrid ratlgun/rocket concept for

delivering materials to low-Earth orbtt. The five major pre-launch,

launch, and post-launch activities of the referenceconcept are:

(1) Projectile payload fabrication

(2) Surface transport of projectilepayload

(3) Projectile preparation at launch st te

(4) Launch operations

(5) Trajectory monitoring.

Individual hybrid system element definitions are provided tn Section

4.4.2.

4.4.1.1 Projectile Pa_yload Fabr|catton

The solid-rocket motors and certain payloads would be

manufactured away from the launch stte. The systems would receive factory

checkout prior to delivery to the launch site. Water payloads would

be supplied at the stte.

4,4.1.2 Surface Transport of Projectile Pa]ioad

Conventional methods of surface transportation (tcuck, rail,

aircraft, and shtp) would be used to transport the pay]oads and rocket

motors. Trucks and railroad cars would be the likely candidates over

round. Ships or air transport could be used if the site were

naccesstble by other methods.

4.4.1.3 Projectile Preparation at Launch Stte

When the payloads and motors arrtve at the launch stte, they

would be placed In a storage factiity. When the scheduled launch ttme

approached, the motors would be removed from storage and stacked, The

payload would be integrated and the projectile assembied, The projectile

would then recetve a prelaunch checkout before being transported to

the launcher facility.

4,4,1,4 Launch Operations

At the scheduied launch time, the projectile Is ioaded tnto

the breech Of the launcher, All launcher systems are Checked Out and

the launch tube is evacuated prior to launch. Charging of the homopolar

generator begins,

Weather and xtnd ConditiOns would be Checked and iaunch

clearance given from the proper authorities befOrethe final Countdown

IIAT'r|LLE -- COI_iJMIIUI
4-28

begins. A launch site alarm would be sounded warning__a11 persons to

be cleared from the area.

The ftnal launch countdown begins after all systems are readied.

The launch systems are computerized and fully automated. A solid armature

is attached tO the rear of the projectile behind the protective sabot.

Current flows through .the armature providing the force required for

launch. SwitChing of the current t_to each segment is activated by

the projectile movement through the launcher tube. When the projectile

leaves the ratlgun_ it is travelling at 2 km/s. After a brief coast

period, the second stage (rocket first stage) is ignited. The stage

is jettisoned following burn-out. The rocket second stage is then ignited

and is also jettisoned after burn is cOmpleted. The rocket third stage

ts fired at orbttal altitude and places the payload in its proper orbit.

4.4.1.5 Tra ector Nonttortn

A telemetry system was assumed to be included as part of the,

projectile system. The payload could then be tracked during its

trajectory and on orbit to facilitate retrieval by the OMV for transport

to a space station.

4.4.2SystmElment Definition

Definitions of the hybrid railgun/rocket system elements are

provided in this section. The six ma_or system elements discussed here

are:

0

e

O

o

e

a

Projectile/pay!oad characteristics

Surface transport systems

Launch site support facilities

Launcher system

Monitoring systems

Space destination.

4.4.2.1_d Characteristics

_. The payloads to be delivered to low-Earth

orbit by the hybrid-ratlg_h--system would include.

• Orbit Transfer Vehicle propellants

,. Space station supply items

• Materials for space processing activities.

These payloads are described in more detail in Sections 2.1 and 4.2.2.1.1.

Despite the lower accelerations of the hybrid railgun/rocket

launch compared to the [arth_tO-Orblt rail launcher (100 g versus 1225 g),

°

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4-29

the mission model remained the same because certain supply Items and

sensitive equipment would still be transported by the Space Shuttle

with a maxtmum acceleration of 2 g.

6

The analysis documented in Section 3.1.4 indicated a maximum

payload of 800 kg for the total.projectile mass of 15,000 kg. It was

assumed, however, that 660,000 cm_of the payload volume was available.

Therefore propellant-supply payloads (i.e., water) would have a mass

_'i of 660 kg, This was the basis for the hybrid EML/rocket traffic model

LI shown in Table 3-18.

4.4.2,1.1 P,roJectile Elements. The hybrid railgun/rocket projectile would Consist of the following subsystems:

i • Rocket motor stages

e Solid propellant

. • Payload .

e Nose cone

e Instruments

e Sabot/a_ture

e Fins.

./

The proposed projectile concept is depicted in Figure 4-14. Table 4-3

presents the mass summary of the hybrid railgun/rocket Projectile.

)

..

TABLE a_.3.. HYBRID RAILGUN/ROCI(LrrPROJECTILE NASS SUI_IARY

i i r i ii

Projectile SubsyStem Mass (kg)

Payload 800

Propellant 12,700,

Propulsion system/casing 1,100

Instruments 20

Nose cone 180

Ftns 100

Sabot/armature 300

Total 15,200

i .

The rocket motor structure would have filament-wOund Cases,

which allows a lower structural mass than steel would, since

filament-wound cases can, be-"stressed, to values several times the

permissible limit of stress in an equal weight of steel" (Hill and

ilATirlI. I.II -_ ¢OI.I, IMIUII
0

4-30

't

_O_Er

" CONE/SHROUD•

PAYLOAD

VOLIJPE

t

!,l

i" q

FOURTH

STARE

THIRD

STARE

''1

l "

f

l

SECOND

IHS

Length = 12 m

Diameter - I.OIB

""t

FIGURE 4,.14. HYBRID RAiL6UN/ROCIET PROJECTILE

iATI'ELI. E -- GOLUMi.U|

., ,.¢ • ...
i 4-31

"4

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PeterSon, 1970). The cases would be loaded with solid propellant

consisting Of approximately 16 perCent aluminum, 14 percent polybutadtene,

and 70 percent ammonium perchlorate. Approximately 12,700 kg of solid

propellant (1290 sec Isp) would be requlred _:o launch an 800-kg payload

tO LEO after an initial 2 km/s boost via an EML.

The instruments and payload shroud/nP,__e cone would be strut]at

to those used on conventional expendable launch vehicles. '*Pop-out"

fins like those on mtsstle systems would be used to stabilize the

projectile durtng its atmosphere flight. The fins would be concealed

inside the structural ca._ing during launch to accommodate the round

launcher tube bore. When the projectile, exits the launcher tube, the

fins would be opened to stabilize the vehicle. The lO0*g acceieratton

environment should not provide any real problems, SinCe pop-Out fins

. are used for gun-launched missiles which also experience high

accelerationS.

A sabot/armature system would be attached to the back of the

projectile. A so.lid armature was assumed because of the 2 km/s velocity

limit. The sabot would tnsulate the projectile from the conducting

armature and would be made of a non-conducting material.

4.4.2,2 Surface Transportation Systems

The payloads and solid-rocket motors would be manufactured

in facilities away from the launch site. Conventional methods of surface

transportation would be used to transport the payloads and motors to

the launch site. A combtnation of truck, rat1, aircraft, and ship systems

would be used, depending upon the locatton of the launch site. Aircraft

could be used to transport personnel and high-priority materials directly.

4.4.2.3 Launch Site Sup.port Facilities

Facilities which would support launch site activities include:

power plant, projecti:le storage and check-out facilities, general storage

facilities, administration and engineering facilities, industrial area,

community living area, iiqutdgas and water plant, and other mtscelianeous

facilities. Except for the power plant and projectile facilities, the

areas would be stmilar tO those described in Section 4.2.2.3 for the

Earth-to-orbitraii launcher and are notdiscussed here.

4.4.2,3.1 Power Plant. Because of the lower power

requirements, a dedicated nuc ear power plant is not required for the

hybrid EML/rocket launches. Utility power would be adequate.

Transmission ltnes would be required and a substation located on the

launch site may be required aS weli.

4.4.2,3.2 ProjectileStoraeandCheckout Facilities. A

faciitty _o store, process, lnte rgate, and checkout the projectiles

would be required. Storage o_ at least 35 projectiles (105 rocket motors)

and payloads would be required, corresponding to One week's supply at

k

iATTELL E.- CO_UMIU|
4-32

a launch rate. of five per day. The sOlid .rocket motors would be stored

from time o.f arrtval Until just before launch, when the three-stage

rocket would be stacked and the. payload would be integrated. At that

time, the assembled.projectile would be moved to the temporary storage

facility at the base of the lauNCher..

4.4.2.4 _.

The hybrid railgun/rOcket system would launch supply payloads

to Earth orbit, The rail gun would: accelerate the projectile at 2 kin/s,

after whl.ch the. three-stage rocket would supply the remainder of the

required AV to reach the desired orbit, The launcher system would be

based on a mountaln Side inclined .at an angle of 20*degrees from the

horizontal.

Figure 4-15 and 4,16 provide cross-sectional and. side views

of the railgun launcher system. The rail launcher would have a circular

bore, I m in diameter. AMZIRC (a copper-zirconium a11oy) would be used

to construct the rails., The rails would be electrically insulated and

spaced by a non-asbestos fiber-reinforced insulator material. A Kevlar

force containment tube would confine the rails and _insulator materials.

Distributed along the launcher tube, 750 homopolar generators

and inductors would provide energy to the launcher. Switching of the

current would be activated by the .projectile movement through the rail

Segments.

Launcher systems which

subsection_ are listed below:

are discussed

• Bore and ratls

e Energy storage

• Launchersupport structure

• Switching and control

Temporary projectile storage facilities

Service and access.systems.

in the followin9

The launcher system concept options are shown in Figure 4-17.

4.4.2.4.1 Bore and Ratls. The, rail launcher would have a

round bor__er, The minimum bore diameter was

calculated for AMZIRC. rails. With a force on the projectile of 14.7 MN,

• 2

the resulting minimum bore,area is 350 cm. This corresponds to a minimum

bore diameter of 21 cm (for a round bore). Since the projectile was

conceptuaiized at I m diameter, the bore is much larger than the minimum

required for Pall structural integrity.

AMZIRC was chosen for the rail material because of its good

strength and conductivity propertieS. The rails would be separated

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by non-asbestos fiber-reinforced insulator materials. To contain tube

stresses durtng launch, the launcher tube iS wrapped in Kevlar. The

bore Sectton is seen in Figure 4-15.

4.4.2.4.2 Erie Store . The homopolar generators (HPGS)

and Jnductors discussed in SeCtion 4.2,2.4.2 would be used as the energy

storage devtces for the hybrid launcher system. Assuming the same

transfer efftcienctes from HPG to inductor to ratl, 750 HPG/tnductor

units (56-HJ HPG) would be required to launch the 15,000 kg projectile

at 2 km/s. The storage devices would line the length of the launcher

tube at 2.7 m center-to-center spacing, as shown in Figure 4-16.

..2.4.3Launcher Su rt Structure. The proposed launcher

support st_lstrated in Figure 4-15, A concrete foundation

would support the launcher tube maintaining its alignment over its 2-km

length, The concrete bed would also help to prevent ,rail separation

and damage caused by launch forces. The homOpolar generators and

inductors would rest alon 9 one side of the foundation, eliminating the

need for an additional HPG support structure.

........ As indicated in Figure 4-17,

a preboost system y required because a solid armature

is used (no erosion of the rails due to plashy, dwell time). However,

it was felt that a small preaccelerator may be advantageous to initiate

the projectile morton up the 35-degree elevation angle. Velocities

of up to 100 m/s were feltto be sufflcient.

4_Stcht_ and Control. Switching mechanisms would

be stmilar_or_er (Section 4.2.2.4.5).

.4_4_2_ar Proect_e Stora • F ctltttes. Each

day, proJ___f_m the arge storage ;_d check-out

facility in preparation for launch. Tempora_ storage of projectiles

would be accomplished in a facility located at the base of the rail

launcher system. The facility would be required to store one day's

supply of projectiles or at least 10 projectiles (2050 high r_del).

4 4.2.4.7 Se. f e .and Ace ss s . . Access to the launcher

system wou d e r_ovidedbyr crane and transport rails shown in

Figure 4-15. This access would _11ow routine maintenance or emergency

repair to occur. The rail system was chosen over others because of

the 35-degree l_uncher elevation angle.

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for range safety reasons and in fltght to insure the rocket motors are

perfoming as expected. Telemetry would be included in the rocket

projectiie tO communicate the required data. The projectile would also

need to be monitored on orbit to facilitate retrieva_ for delivery to

the space station.

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4.4.2,6 S._ace Destination

Low-Earth orbit is the primary destination of the hybrid rail

launcher. Specifically, the system was designed for launch to a 500-km

circular orbit. Other orbits would be possible, at the expense (or

gain) of payload mass,

4_rtd Coaxial AcceleratorlRocketConcept

4.5.1 ConceptDeftnttton

The hybrid coaxtal accelerator/rocket concept for material

delivery to low-Earth orbit consists of five major operational activities:

(1) Projectile fabrication,

(2) Surface transport of projectile

(3) Projectile preparation at launch site

(4) Launch operations

(5) Trajectorymonltoring.

These activities are the same as those described in Section 4.4.1 for

the hybrid rallgun/rocket concept, except for the launch operations

activities whlch are discussed here. Definltions of the individual

hybrid system elementsaYe presented in Section 4.5.2.

The projectile is removed from the temporary storage facility

and placed in the launcher tube. The Brooks coil energy storage inductor

is charged. All launcher systems are checked out and the launch tube

is evacuated before final launch countdown begins.

Countdown begins after weather•and wind conditions are checked.

The proper authorities would be notified to obtain clearance to launch.

All persons would be cleared from the immediate launch vicinity.

When all systems are readied, the computerized launch sequence

begins. Projectile motion is initiated and switching of current into

the active segments behind the projectile cotls occurs, The muzzle

launch velocity is 2 km/s.. Rocket procedures arethe same as described

in Section 4.4.!.4.

4.5.2 System Element Definition

The six major coaxial accelerator system elements described

in this section are:

e

e

e

Projectilepayload characteristics

Surface transport systems

Launch site support facllitles

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described

S-.._._._,,,__. The proposed projectile for the hybrid

coaxiai/rocket-- Concept s, shown

following subsystems:

e Launcher system

• HOnttortng. systems

• Space destination.

4. S. 2,1 Projectile/Payload Character1 sttcs

'1 The pay]oad would be Identical to that for the hybrid rail gun/rocket system.

4.!

tn ,Figure 4-18 and consists of the

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• Rocket motor stages

• Solld propellant

e Payload

• Nose cone

• Projectile coils

• Fins.

The three solid-rOcket motor stages are similar to those assumed

fop the hybrid railgun/rocket (Section 4.4.2.1.2). During fabrication,

however, 80 copper rtngs (projectile coils) would be tmbedded in the

structura] castng. Each copper rtng would be 2-cm thick with an outer

dtameter of 0.8 m. The axial length of the rings would be 6 cm.

The rest of the projectile is stmllar to the hybrid

ratlgun/rocket projectile. A, mass summary of the proposed projectile

is given in Table 4-4.

TABLE 4-4. HYBRID COAXIAL/ROCKET PROJECTILE NASS SIJ_RY

Projectile ,Subsystem Mass(kg)

Payload 800

Propellant 12,700

Propulsion System/Casing 900

Instruments 20

Nose Cone 180

Fins 100

Projectile Rings 650

Total 15,350

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4,5.2.2 Surface Transportati_

SurfaCe transportation wou]d be accomplished by systems

identica| to those- used for the hybrtd rai]gun/rocket concept. The

methods would depend upon the location of the launch stte.

4.5.2.3 Launch Site Su_

Launch site support facilities would be simtlar to the

facilities described in Section 4.4.2.3 for the hybrtd ratlgun/rocket

concept. The facilities would include: uttltty power plant substation

and/or transmission 11nes, projectile storage and general storage

facilities, administration and engineering offices, Industrial area,

community living area for a remote site, ltquld gas (LN2 required to

c0ol Brooks cotl energy storage inductor) and water plant, and other

necessary facilities.

4.5.2.4 Launcher

The hybrid coaxial/rocket system would launch a rocket

projectile to a veloctty of 2 km/s using a coaxial accelerator. The

remaining velocity required to reach low-Earth orbit would be supplied

by the three-stage rocket. The launcher would be located on a mountain

side with an elevation angle of 35 degrees. Figure 4-19 provides a

cross-sec_tonal view of the launcher system.

The drive cotls are 2040 m long and made of a copper-alloy

material. The drtve cotls are encased in Kevlar to provide electrical

insulation and tube structural containment.

A single large Brooks coii energy storage inductor would supply

energy to the drive coils. Switching would be coordinated so that ten

turns behind each of the 80 projectile coils would be active at a11

times as.the projectile Js accelerated through the launcher tube.

This section describes the following launcher systems:

• Ortve cotls

• Energy storage

• Launcher support structure

• Preaccelerator system

• Switching and control

• Storage facilities

e Servtce and access systems.

The hybrid coaxtal acceleratorrocket concept options are provided in

Ftgure 4-20.

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4._ol Drive Cot]s, The drive coils would be fabricated

from a st_of_ectangular copper-alloy wire into a continuouS

helical winding. Dimensions of the drtve coils are: inner radius of

40,2 Cm, Outer radtus of 41.2 Cm, length (of heltx) of 20.40 m. Ten

tutus would be active behind each projectile co11 with an active axial

length of 10 Cm.

A Kevlar tube structure would encase the dr_ve colts to provide

electrical insulation between turns and to cOntain hoop stresses during

launch. The Kev]ar structure does not need to be as thick as for the

hybrid railgun/rocket because the hoop stresses are lower due to induction

effects from the projectile cotls.

4.5.2.4.2 Ene Stora . The 1energy storage device for thts concept--would be a sing le Brooks col inductor, Muzzle kinetic

energy of the projectile would be 30 GJ. Again assuming an 85 percent

transfer efficiency (see Section 4.3.2.4.2), the energy storage

requirement is 35.3 GJ. The Brooks COt1 would be made of aluminum wire

and would be cooled to liquid nitrogen temperatures.

_Launcher Su rt Str cture. Figure 4-19 shows

the launcher tube part n a concrete foundation. The

support structure ts used to matntatn launcher rigidity over its 2-km

length and to prevent drive coil damage due to hoop stresSes during

launch.

4.5.._4 Pre ccele _r S . A small

system w_nitiate projectile motion.

no mere .than 100 m/s were seen to be necessary.

preaccelerator

Velocities o.f

4.5.2,4,5 SM htn an Cont 1. Switching and Control of

currents _yh,tS+e_,;ou]dbe stmilar to that discussed in Section 4.3.2.4.5 for the Earth t(>-orbit coaxial accelerator system.

4_eStora Facilities. The

projectile storage aunch_r would

be similar tO that described tn Section. 4._.2.4.6 for the hybrid

ratlgun/rocket launcher.

4,52_4,7 _rilceandAccess S teNs. To provide access

to the launc repair purposes, a ratl

transport system similar to the hybrid rai|gun/rocket concept would

be used.

4,5,2,5 Non| to_

Trajectory monitoring and orbital tracking systems would be

identical to the corres_ rat| launcher system (Section 4.4.2.5).

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4.5.2.6 Sce Destination

The primary payload destination would be a 500-km circular

tn which a space station was presumed to be orbiting. OCher

orbits would be possible with varytng payload masses.

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5.0 SAFETY AND ENVIRONIqENTAL IMPACT ASSESSMENT

This section documents the preliminary safety and environmental

impact assessments of the reference conceptS. No significant differences

were seen between, the r_Ltlgun and coaxial accelerator Concepts; therefore,

they are considered to be of similar envit,onmental risk, except where

specifically noted, Section 5.1 presents the preliminary safety

assessment. The environmental impact analysts is summarized in Section *

5.2. The 1982 ESRL study should be referred, to for additional .data

and infomatton regarding safety and environmental impact.

A preliminary safety assessment identified the foliowingmaJOr

accident events:

o Reentry of hardware (deliberate and accidental)

e Projectile break-up

e Liquid propellant sptlis (Earth-to-orbit concepts)

e Propellant fires at the launch site.

These events were identified as those which could impose a hazard to

the human population Or the biosphere.

The largest safety risk was seen to be reentry of the nose

cone and payload support structure for Earth-to-orbit launches and of

the nose cone and spent stages for hybrid systems, as well as the

possibility of projectile break-up in the atmosphere with its resultant

release of matertal and the uncontrolied reentry of projectile pieces.

Should these events occur over populated areas, the damage could be

significant. The launch site location must be sited "so aS to avoid

any overflight of populated areas.

Propellantorelated riSks would be comparable to those

experienced b_ existing NASA space programs. Adequate safety procedurqs,

similar tO those required for other programs, must be used to prevent

the possibility Of toxic propellant spills and "on-pad" f4_es. Every

effort must be made to avoid toxic exposures to workers a,. the local

uncontrolled human populatton.

5.2En_ vt_nuental Impact Assessment

An in-depth environmental impact, assessment of the reference

concepts was not possible in: this study. However, we did review the

reference concepts to assess certain environmental impact areas to determine if any critical issues exist which might prevent EML development

for space missions; none were found.

The environmental impact areas fo_ the reference concepts

were. assessed in two major categories: (I) faclllty development and

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conStrUction and (2) testing and operations. These areas are discussed

in the next bye sections..

5.2.1 Facilities 0eve10 ent and Construction

The environmental impact for facilities development and

construction ts highly dependent upon. the location at which the site

is constructed. NO Specific location for the launch site was selected

tn this study. If the launch site were placed in a remote area, the

environmental impacts to the area could be significant; however, the

overall _mpact tO the quality of the human environment from a. remote

site would likely be lower than the impact from a non-remote site. Major

impacts for launch site development might involve the following:

relocation of inhabitants, destructlon of vegetation and wildlife

habitats, extinction of local animal specles, and disturbance of

archeologlcal Sites. Site Selection criteria Should take account of

these posslblelmpacts and could be used to minimize these effects.

The types of facilities that are anticipated for .the concepts

are listed below:

e Launcher system

• Power plant and utility lines

• Roads

e Airfield and rail lines

•. Buildings.

The types of environmental impacts caused by the construction of these

facilities should be typical of any construction activity in an

undeveloped area. The construction of the launcher system and airfield

runways could pose significant impact to the area. The construction

of power plant or utility lines, roads, buildings, and rail lines is

not expected to. poSe Significant environmental effects.

Materials usage must also be addressed. Although large amounts

of aluminum, copper, and steel are necessary for construction of the

launcher systems, the impact to U.S. and world annual consumption estimated for the year 2000 is expected to be minimal.. The materials requirements for construction of each .of the reference concepts are presented

in Table 5_I.

5.2.2 Iestinl and _rations

This section discusses the expected environmental ImpaCtS r

from. normal testing and operation of the reference concept systems.

The impacts were assessed assuming the fifty-year average of flights

per year from the traffic models shown in Tables 2-5 and 2-18. Ten

flights per day were assumed for the Earth-tO-orblt systems and! five

flights per day were assumed for the hybrid EML/rocket systems. Areas

of concern relating to this impact, area include:

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e. Sonic.booms.

e Power plant emissions,

• Chemical effluents.

• , Solid waste dtsposal

e Hzteriais usage

• .Orbital debris (hybrid EML/rocket concepts).

The following subsections discuss each of these concerns.

5.2.2.1 Sonic Boo_

The impact of sonic boom generation from the EML launches

was investigated. The magnitude of the sonic booms produced by the

reference concepts was calculated following the procedure described

in Section. 5.3.2.1 of Rice, el al (1982). The equations used in the

assessment were derived from Sedov (195g). The derived pressure-rise

equation was:

Ap = O'082pV2Cdd2

x2

where Ap = preSSure rise

P = atmospheric density

V =.projectile velocity

Cd = drag coefficient

d = projectile diameter

X .= radial distance from the disturbance.

Table 5-2 lists the distances calculated for various sonic boom overpressures. The critica! distances were calculated assuming a drag

coefficient of 1.0 which represented the sabot Jettison for the

Earth-orbital railgun Concept and the stabilization fin extension for

the other concepts. Large differences in critical distance are evident

between the Earth-orbital concepts and are explained by the factor-of-two

difference in projectile diameter (1.0 m for the railgun projectile

and 0.5 m for the coaxial projectile). The hybrid projectile .diameters

are similar (1.0 m for railgun and 0.8 m for coaxial).

The overpressure limits were provided in CPIA (1972), At

an overpressure of 20.7 N/cm 2, the lethal threshold is reached. Persons.

within the critical distance for lethality (within 15 m for the worst

case--_arth-orbital railgun concept) would likely be killed. An average

human being would experience, ruptured eardrums at 3..45 N/cm 2 within

40 m of the launcher muzzle for the worst case, Window breakage of

typic_1 glass would occur at 0,345 N/cm2, while "uncontrolled areas"

would typically experience 0,138 N/cm z. Sonic booms on the order of

supersonic aircraft at high altltudes would occur between 165 and 700 m

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depending upon which reference cOnCept ts assumed. At distances of

the order of hundreds of meters away from the launcher muzzle, most

antmals would leave and seek other habitats. People living or working

within several kilometers would 1,tkely be annoyed by the boom. Ba:;ed

upon this preliminary assessment, it is conciuded that localized damage

to the biosphere would be evident in the region near the muzzle of the

launcher and that antmal species in the vtcinity of the muzzle would

probably migrate to other locations. Effects at larger distances are

not belteved ,to be significant.

5.2.2.2 Power Plant Emissions

A nuclear power plant was. assumed for the Earth-orbital

reference concepts_ Normal emissions from nuclear reactors are not

expected to pose-significant hazard to the environment. The radiation

dose to plant workers and the risk of accidents are not expected to

be different from other, nuclear facilities currently operating under

federal guidelines, The hybrid EML/rocket concepts assume the use of

utility power. The environmental impact should be similar to existing

power plants. Emissions from the power plant are not expected to be

Of significant envirOnmental impact..

5.2.2.3 Chemical Effluents

The major chemical effluent resulting from testing and

operations are those related to the use of solid propellant in the hybrid

systems. HCf emitted from the solid propellant motors at high altitude

could cause a reduction in the ozone concentration in the stratosphere

and cause ionospheric disturbances. Hybrid solid systems, however,

would be of less consequence than an all solid launch vehicle that burnS,

propellant in the lower toposphere. Also, it is expected that various

types of cleaning solvents and various propellant contaminants could

be released into the biosphere (air and water). Tnese activities are

not expected to be of major significance and are expected to be comparable

to those of current .Space .Shuttle launch activities.

5.2.2.4 Solid Hastes

it is expected that solid wastes, including the production

of waste propellant in solid motor production for the hybrid system,

would be similar to those from current Space Shuttle operatiOns or from

typical industrial operations. No significant environmental impact

is expected from the generation of solid waste produced from the testing

and operation of the reference concept systems.

5.2.2.5 Naterlals Usage

Table 5-3 presents the major material requirements for the

Earth-to-orbit reference concept projectiles at the rate of ten launches

per day and compares these• requirements with the estimated annual U.S.-

and world consumption in the year 2000° The major materials requirement

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for the hybrid EML/rocket concepts is solid propellant, At five launches

per day, 23,180 MT Of solid propellant is required per year. At first

glance, this seems to be a large amount; however, one Space Shuttle

launch uses I010 MT of solld propellant. The annual hybrid EML/rocket

requirement is then approximately equal to 23 STS launches which are

reasonably expected by 2000. The solid propellant ingredients are

aluminum {roughly 16percent), polybutadlene (about 14 percent), and

ammonium perchlo_ate (approximately 70 percent) which are common

matertals. _

Although significant upgrade may be necessary to produce certain

aterials, namely C1F3, N2H4, and solid propellant, the use of these

tems is not critical to the consUmpttonof the Earth's resources.

5.2.2.6 Reentry of Itmrdware

This issue was previously addressed in the safety assessment

(Section 5.13 and in the mission requirements analyseS (Section 3.5.2).

During testing and operational activities, the nose cone and steel payload

support Structure would be jettisoned and would reenter for every launch

of t_e Earth-orbital concepts (ten par day average). Spentstages and

nose cones would reenter for hybrid EML/rocket launches {average of

five per day). Minimal risk to the population and the potential for

recovery of materials (if economical) would occur if the launch site

and resulting landing areas were properly selected (see Section 3.5.2),

5.2.2.7 Orbital Oebris

Orbital debris is not expected to be a problem for the Earthto-orblt concepts. An Orbital Maneuvering Vehicle (OMV) would rendezvous

with the payload on orbit and transport it tO the Space Station. The

propulsion systems would be returned to Earth via the Shuttle.

Solid rocket exhaust particulate is Currently receiving attention in the orbital debris area. The micron-size particulate may cause

pitting of spacecraft and damage to solar arrays. The final stage burn

would be an orbital insertion maneuver into the Space Station orbit,

Most of the. particulate would deorbit; however it is possible that. some

particulate would, remain in orbit and impact the Space Station systems.

Further analysis is required, including detailed trajectory analysis

and an investigation Of alternatives such as a liquid-propellant upper

stage. This debris issue should not prevent hybrid EML/rocket

development.

S.3 Conclusions

Based upon the preliminary safety and environmental impact

assessments for the reference concepts, no significant safety or environmental impact problems have been found. Population overflight must

be considered when selection of the launch site is made. The initial

construction of facilities is expected to create some environmental

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mpact to the local area; however, this is not expected to be significant.

ontc booms would create localized problems for antmals SurrOunding

the launcher system, but few effects are expected on the human population.

No major issues have been found thus far tn the safety and environmental

!mpact evaluation to prevent EML development for space missiOns.

Economics appears to be the most important non-technical issue.

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6,0 _ COST ESTZItATES

Th]S section documents costlest]mates for the coax]a1 and ratlaun

electromagnetic launchers and examines the cost competit]veness of the

EHL concepts as applted to near-term (2000-2010) bulk supply miss]ons,

The EHL concepts are compared to conventional chemical launch systems,

such as the Space Shuttle and a so1]d-prop911ant rocket.

Costs for the EML reference concepts (Earth-to-orbit EHL and

hybHd EHL/rocket concepts) ustng ra]lgun and coaxial accelerators have

been estimated accord]n9 to the Work Breakdown Structure (HBS) ]n Tables

6-1 and 6-2. The resulting investment costs estimates are shov._ ]n Table

5-3. The goa| of these estimates is to provide a comparison of the ra]lgun

and coaxial technologies ]n the area of costs, and to examine the cost

competitiveness of these concepts versus fully chemically powered launch

systems such as the Space Shuttle and an all-solid-rocket launch vehicle.

The cost information developed here relies strongly on our previous

investigation of ratlguns (Rice, Hiller and Earhart, 1982).

Both the ESRL report and this report use 1981 dollars. To

adjust total cost estimates given in 1981 dollars tO 1984 dollars, multiply

by i.16 to reflect Consumer Price Index trtflatJon of about 5 percent

per year. Hater]a1 costs, however, have not exhtbJ.ted uniform inflation,

and tn some cases are lower than in 1981.

The Hork Breakdown Structure does not inclute research,

technology development, and design efforts pr]or to formal ]nit]at]on

of an EHL development. These are not included because the research is

applicable to many other activities and some of these costs may be paid

by those activities, and because advanced .reSearch and technology costs

are highly uncertain.

The research and design costs used for the Earth-tO-orbit EML

designs are the same as used tn the ESRL report, about 10 percent of

the initial investment for that hig h-capability system. The expected

cost for research and design is $466 H, and the low and high estimates

are $320 H and $633 H. The low and high costs are believed to represent

the 90 percent confidence ]nterval for research and design expenditures,

while the expected value approximates the mean of the cost distr]but]on.

This interpretation ts given to a11 use of the terms lOw, expected and

h]gh cost estimates. For the hybrid EHL/rocket, the research and design

COstS are expected to be somewhat lower, refiecting the lower level of

capability required. The costs are assumed to be two-thirdS of the

estimates for the Earth-orbital EMLS. This. procedure gives a range frOm

$200 H to $400 H with the expected value $300 H. Project resources do

not permit the detailed investigation of the resource requirements for

research and design for.either cOncept. The compaHson of the associated

costs o_ the ra_lgun and coax]a1 EHL concepts ts conducted at the System

1eve1 with every attempt to keep the systems mission-equivalent. At

the present level of understanding of these concepts, tt appears that

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6-2

Ell DEVELOPRENT AND INVESTlf.NI MORK BREAKDO_ STRUCTURE (MBS)

'4

1.0 Facil

2.0

tttes and Supporting Systems

1.1 Land

1.2 Power Plant or Substation

1.3 Personnel Support Facilities (housing, roads, sanitation, school)

1.4 Shipping Docks, Storage, and TransportationFacillties

1.5 Airfteld and Hanger

1.6 Industrial Area (Equipment Refurbishment)

1.7 Administration/Engineering Buildings

Launcher Systems

2.1 Hountainside Structures

2.1.1 Launcher Tube Housing

2.2 Launcher Tubes

2.3

2.4

2.5

2.6

2.7

2.8

2.9

2.10

2.11

2.2.1 COpper Alloy Conductors (rails Or coils)

2.2.2 Spacers-Insulation

2.2.3 KevlarContainment

2.2.4 Vacuum Container and Exterior insulation

Launcher Energy Storage (includes hydraulic motors and hydrau]ic

distribution) and Supporting Structures

Inductors and Switches (includes LN2. distribution system)

Preboost System

Power Conversion Facilities

Water Distillation Plant_

Gas Handling Facilities ..

2,8,1 Liquid Nitrogen Plant and Storage

2,8,2 Evacuation System for Launcher Tube

2,8,3 Water Electroiysls Plant

E]evator Systems and .Projectile Handling Devices

Control Center, Controls, and Monitoring Systems

Tracking Systems

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ENL OEVELOPIfdTrTEST PRO6RANAND OPERATIONS

WORKBRBK_ STRUCTURE(_BS)

3.0 Projectiles and Hisston Peculiar Equipment

4.0 Operations

4.1 Management and Support

4.1.1 Management

4.1.2 Engineering

4.1.3 Facility SuppoV't

4.2 Power Plant Operations (Supplies and Crew)

4.3 Technical Personnel and Supplies

4.3.1

4.3.2

4.3.3

4.3.4

4.3.5

4.3.6

4.3.7

Control Center Crew

Launcher Equipment Support Crew

Equipment Refurbishment Crew

Power Conversion Facility Crew

LN2 Plant/Vacuum System Crew

ProjectilePayload Operations Support Crew

Faciltty Utilities Crew

5.0 Development Test Program

5.1

5.2

5.3

5.4

Test of Launcher Segment(s)

Development of Projectiles

Transient Housing at Launch Site

Launcher Operations Costs During Tests

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the coaxtal accelerators provtde lower levels of .stress (for normal.

operations) on both the lau,ncher and the. proJectt.le, This lower level

of stress leads to lower design requirements and less mass tn the launcher

and p_oJecttle, and therefore a lower level Of costs. It also appears

that- -the non kinettc electrical energy Storage devices (Brooks cotls)

for the coaxial concepts cost less than the homopolar generators (HPGs)

used for the rat]gun concepts, Railgun technology, however, has been

demonstrated at htgher accelerations and velocities than has coaxtal

technology and the rat lgun technology may yet prove to have other

advantages whtch may not be apparent at the present time, especially

for htgh energy miSsions. The system Cost estimates here tnclude: (1)

system development and construction, (2) an inttfal f11ght test program,

and (3) thtrty years of operations. A cost summary sectton presents

an overvtew of cOsts developed and the cost per unit mass of payload.

_nt and 1nvestle_

Thts sectton discusses the derivation of development and

investment costs for factltJ;tes and supporting systems and for the launcher

systems. Low, expected and htgh estimates are presented tn Table 6-3.

The low and htgh estimates Can be considered an estimate of the 90 percent

confidence Interval for expected Costs. To matntatn visibility of the

cost differences between railgun and coaxtal systems, facilities costs

are the same for both technologies uniess there tS an obvious reason

to change them.

6.1.1 Facilities and Supportfng System

Stx bastc categories are constdeted and costed (1) land, (2)

electrical power, (3) personnel SuppOrt, (4) transport facilities, (5)

industrial area, and (6) administrative/engineering buildings and access

roads.

6.1.1.1 Land

For all concepts a mountainous island or a mainland site near

the equator iS assumed. Slopes of 20 degrees (Earth-to-orbit EML) and

35 degrees (hybrid EHL/rocket) wo_ld be requtred for distances of at

least two kilometers. For the hybrid EHL, a stte at 28.5 degrees latttude

may be acceptable to facilitate launches, to orbits with this inclination,

but the number of launch opportunities for a single.space station would

be limited to one per day. For an equatorial site, there would be 16

launch windows per day to a single equatorial space station. AnequatOrJal

site would also be desirable for launches to higher orbits such as

geosynch¢onous equatorial orbit, but these missions are outside the primary

scope of this investigation.

While free use of land is a possibility, a cost of _2470 per

hecture ($1000 per acre) is assumed. Because of Its powe_ plant

requirement, an Earth-to-orbit !auncher facillty ts expected to occupy

from 24 to 40 km2 (_5 to 25 sq mi) while the hybrid EHL/rocket launcher

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facility is expected to occupy from 12. to 25 k_2 (5 to 10 sq mi),. The

estimated costs for land for the Earth-orbital facility would.then range

f_Om $g.6 M to $16.0: H. with an expected cost of $i2,0 M. For the hybridEML factltty the land cost estimates range from $3.2 M to $6.4 H wlth

an expected cost of $5 M.

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For the hybrid EML/rocket launcher, the power requirements

are such that an independent base-load plant would not be efficiently

used. Use of commercial power or saie of excess power fro_l a jotnt venture

power plan would be. desired for thts concept. For the Earth-to-orbit

EHL, 100 HWe capacity is considered necessary for all purposes and this

should be dtvtded into two 50 HWe units tO provide a maintenance reserve.

(The energy requirentent for 8 launch t_ LEO is about 42,4 HWh.) Civil

reactor availability currently runs abOut (_6 percent due to both scheduled

(about 20 percent) and unscheduled (abOut 14 percent) maint, enante. For

economic reasons, ctvil power reactors are larger than 200 MNe, but naval

and other small nuclear reactOrS are believed to be available in the

appP.opriate size range. Given the use of two independent reactors,

avallabiltt_( of power from at least one reactor should be about 90 per_nt

[1-(1-0.66) z - 0,89] and should improve in the future. For further

discussion of the factors influencing this choice, see Rice, et ai, 1982, _

pages 6-7 and 6-8.

For the hybrid EML/rocket concept, 15 to 30 MWe of capacity

would be required fop both the EHL itself and for personnel support.

Clearly, use of comercial power would be desired because baseload plants

below 100 MWe are not considered to be economical in this country. Based

on available documentation of plant costs (Frtedlander, Electrical World,

October, 1981), a 200MWe baseload nuclear plan is pvoJe_m

$2200 to $2800 per kwh capacity with $2500 per kwh expected.

A IO0-MWe baseload coal plant is expected to cost from $1400

to $1800 per kwh, with an expected cost of $1600 per kwh of capacity

(Friedlander, 1981). Extrapolating the curves given by Friedlander for

coal*powered units to 30 MWe, however, leads to estimates of $2000 to

$2600 per kwh of capacity, with an expected value of $2300 per kwh.

Accordingly, the untfoPm cost per kwh of capacity used ranges

from $2200 I;o $2800 with an expected value of $2500 per kwh. This leads

to estimates of $220 ;4 to $280 H with $250 H expected for -the

Earth-to-orbit EML facility wtth a lO0-MNe nuclear power plant. For

the hybrid EML/rocket facility, a 30-MWe non-nuclear plant would be

expected to cost from $66 M to $84 M, wtth an expected cost of $75 H.

If, as expected, comercial power ts used, these estimates would be high,

and an estimate of $10 M to $25 M, wtth an expected va]ue of $15 M might

be expected to cover transmission lines, a substation, and a small amount

of emrgency outage power-generation capability. This estimate is based

upon twenty percent of the estimates for providing power at the site

and is in turn based upon Friedlander, 1981, and other articles in

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E_ectrical _orld, The estimates are strongly dependent upon the launch

_t_e belng located within a reasonable servtcedtstance from the commercial

plant or major transmission ltne_ To preserve comparability between

concepts, the higher estimates are used.

6,1.1.3 Per;sonnel Support, Facilities

PersOnnel support facilities include houstng, roads, sanitation,

and school buildings. It is assumed that there would be a permanent

workers' community with reasonable a_entties, and that most workers would

work in some sort of butldfng. The cost Of these facilities is estimated

at $100,000 per pe_n_nently stationed employee. If the factitty could

be located in the U.S., some of this cost would be borne by the workers

and local government, but most missions under consideration appear_to

requfre a lacation outside the Continental U.S.

Ratlgun Concepts, because, they use homopolar generators, would

require a maintenance crew larger than foreseen for coaxial concepts,

but this impact on facilities costs is expected tO be small. AccOrdingly

personnel facilities are costed on the basis of personnel estimates for

launchers as given in Section 6.2. From 300 to 900 personnel are

estimated, with an expected number of 550. This leads to personnel

facility estinlates of from $30 M to $90 M, with an expected estimate

of $53 M,

G.1,1.4 Transport Facilities

Transport facilities capable of handling large objects in

relztively high volume over short periods of time would b_ required.

These include air cargo, land and shipping facilities appropriate to

the location of the launcher. To the extent that existing facilities

can be used, additional: costs could be avoided; Sea, air, and land

facilities arediscussedbelow are estinmted to co_t from $60 M t_ $200 M

with an expected cost of $130 M.

5.1.i,_1 Shipping Docks, S,torage...,pnd Transport Facilities.

The.cOst-of surface transport and storage facilities required would 'depend

on the spectftc features of the site, such as, terrain, how much

deve|opment already exists, and whether there is a natural harbor. While

the initial site development could probably be supported by small ships,

it Is expected that a protected pier would be required for the construction

phase. Accordingly, these facilities are estimated in the range from

nothing to $40 M with a $20 M expected cost. For example, if the site

were located in northern Mexico, a port would probably not be needed.

For equatorial sites, it is very ltkely that, at a minimum, existing

port facilities would.need to be improved.

6.1.1,4.2 Land Translpmrt FaCilities. Two-lane heavy-traffic

roads cost from $1M to $9 M 'per kllomete'F _pendtng on the nature of

the terrain (Ohio Department of Highways)° Heavy-duty rail lines cost

from $4 M to $7 M on terrain-sultabl_ tn rail use. Because a specific

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site is not proposed here, detatled land transport cost estimates cannot

be made. An allowance of 20 miles of roads and/or railroads is made

with cost esti_ted in the range of $40 H to $80 M with an expected cost

of $_GOM.

G.I.I.G Alr'f.teld and _nqar

Since high-value ok sensitive cargos would not be u_ed, an

airport capable of accommodating frequent flights of the largest _ircraft

would not be necessary, lhe local geography would be the major determinant

nf the airfteid cost, and this cost could vary by as much as a factor

of ten depending upon the site chosen.

The ESRL report provided for an airport which could handle

the largest standard cargo aircraft and had two GO00 m runways, taxiways,

100 m x 100 m hangar, and a fuei depot; the airport was estimated to

cost from $26.3 M to $100 M with an expected estimate of $5G M. Because_

the facility is not envisioned, for transporting hazardous and/or sensitive

payloads, it is probable that one runway would suffice _nd that the cost

could be constrained within estimates of $20 M to $80 M _ith an expected

estimate of $50 M. Cost considerations, given in the ESRL report, include.

$3000 to $5000 per lineal meter of runway with minimal soil preparatien,

factors of I tO 3 aRplied to these estimates, for grading, $450/m z for

hangar construction, $I M for the fuel depot, and ten percent of runway

costs-for taxiways.

6.1.1,:6 industl"t.al _A_a

Since the. EML concept would employ a considerable amount of.

moving machinery (such as homopolars, gas liquefaction compressors, etc.),

n_merous maintenance and repair activities are anticipated. Thus, a

facility which could, repair and refurbish the equipment would be needed.

There would also be a .need to store replacement hardware components in

a warehouse Because of the uncertainty of the requirements for this

facility, it-is arbitrarily estimated at $40 M to $80 M with an expected

cost of $G0 M, including bo_hbUildings and industrial equipment.

6.1.1.7 AdsJnistration/Engtneertng Butidtngs

The .administration and engineering functions are expected to

reach a peak during development and initial operations, and .then drop

toa lower level as initial operational problems are resolved. Activities

would rise to higher levels only if additional demand, justifying new

or replacement launchers, is achieved. The englneerlng development staff

wo_ld most likely be accommodated in inexpensive buildings which could

be used later to accommodate transient personnel during the operaticnal

phase. Since the initial motivation for construction of buildings would

likely be the development test program, an estimate of $3 M to $5 M is

charged to the.development .test program,

For the permanent staff, however, there are expected to be

from 100 to.400 people who will need permanent office or other working

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space. This is expected to cost about $20,000 per worker, resulting

in an administration/engineering buildings Cost of $2 M, $5 M, and $8 H

for low, expected, and high estimates of i00, 250, and 400 workers needing

these facilitieS. _

6.1.2

All launcher syste;as are considered to he constructed along

mountain sides. This would avoid the considerable cost of constricting

tunnels over 2 kmlong aS well as extensive undergrouno eccess and working

areas. This cost*reduction potential, however, restricts the number

of potential sites, because slopes of approximately 20 degrees will be

required for the Earth Orbital launcher concepts and slopes of

approximately 35 degrees will be required for the hybrid EML/rOcket

conceptS, Despite their appearance, mOuntain slopes of 20 to 35 degrees

ovee a length of 2 km or more are relatively rare, and other geographical

factors such as slope orientation and down-range safety zOnes further

restrict the number of acceptable sites. An artificial mountain with

a 2 km-long Slope at 35 degrees would be approximately 1 km high. The

earth-moving and stabilization problems associated with constructing

an artificial mountain 1 km high would result in higher costs than

tunneling_, _

6.1.2.1 Launcher Tube Housinq

FOr al_ concepts, the launcher tube and associated equipment

would need to be fimly anchored to the side of the mountain and have

a substantial cover from which overhead cranes can hang tO move launcher

equipment. Such a structure is assumed to be at least 2.1 km long to

aCcOmmodate equipment at the breech and muzzle. This housing is expected

to have the complexity of a four-lane Superhighway structure carrying

elevated traffic. Four-lane superhighways currently cost from $2.5 M

to $18 M per-km. ($4 M to $30 M: per mile) (Ohio Department _of Highways)

with the lower figure representing construction on level farmland and

the higher figure representing elevated structures. The launcher housing

would be built in a remote, mountainous area and would require substantial

amount of site preparation. For a 2.1 km length, this leads to estimates

of two to four times the maximum superhighway cost ($38 M to $76 M, with

$57 M the expected cost). (If a subterranian complex were to be selected

the facility construction cost is estimated to cost from $250 M to $540 H

with an expected cost o_ $300 M--see discusSiOn on pages 6-11 to 6-13

of Rice, et al, 1982-.)

6.1.2.2 Launcher Tubes

A11 concepts would use launcher tubes with copper a11oy

conductors, insulation to hold the conductors in place, a Kevlar wrapping

to contain normai and some accidental launch forces, and a vacuum

container. The coaxial designs would operate at a lower current and

higher voltage than railgun _launchers, Since launch, stresses are

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6-10

proportiOnal to the square of the Current, and the coaxial iauncher's

lower current stresses are reduced by induction effects of the projectile

coils, this results in a much lower 1eve1 of radtal forces on the launcher

tube. The coaxtal launcher tube, however, must still be protected against

the effects of accidental currents and accidental ignition of propellants.

The additional wrapping to contain accidental overloads reduces the

potential savings in tube construction costs that could be possible if

overloads were nOt considered. Launch stresses for railguns are much

higher than for coaxtal launchers. AcCordingly, the stress Confinement

wrappings allocated for nomai stress confinement for the ratlgun designs

are Considered to be adequate for accident protectiOn,

Calculations of the launcher tube costs are presented in Table

6-4. The tubes would have common components--conductors, insulation

force confinement wrapping and an exterior vacuum container which would

also provide a mechanical connection tO adjacent tube segments, It is

expected that the tubes would be manufactured in segments of convenient

length to pemtt replacement of worn or damaged segments, The launchers

would have circular bore cross Sections with the exception of the railgun

for Earth-orbital missions, which would have a square-bore cross-section.

The square-bore design is carried over from the ESRL report. Launcher

tube configurations and their cost implications are discussed in the

following four sections.

6_11 Launc r Conductors. The launcher

conductors wou ses which should not

provide so much heat that active cooling is required. The material

selected is ANZIRC which is approximately 99.85 percent copper and 0.15

percent zirconium (Engineering Alloys Dtgest, Inc., 1961), Based on

a typical price for copper foms of $1.76 per kg and a price for zirconium

of $16,50 per kg, the conductors would cost $1 .78 per .kg. The density

for AMZIRC is taken as 8.96 g/cc,

It is expected that the conductOrs would be formed, heat treated,.

surface machined, and later assembled into a complete tube segment amenable

to handling and transport. Because these are traditional manufacturing

practices and no advanced technology appears to be involved, the

appropriate labor factor for fabrication and installation in quantities

this large is tn the range of 2 to 3; a labor factor of 2.5 is the midpoint

andis used to form the expected cost.

6_ri ] Insulation, The rails or drive coils

would requtre mechanical support and e ectrical insulation. For the

coaxial launchers, where the voltages are high and the forces relatively

low, tt is expected that a synthetic rubber or plastic should suffice.

Railgun launchers will _ need insulattve materials with high compressive

strength. Many potential candidates currently contain asbestos, which

is considered to have unacceptabie, human health concerns in most

applications, it is expected that a substitute can be found and would

cost leSs than $1.00 per kg and would have a density in the range of

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3 to 5 g/cc. For the coaxial accelerators, the synthetic rubber or piastic

is expected to cost less than $2,20 per kg and have a density of 1.2

g/cc, Labor f_ctors of 1 (materials cost tncludes labor), 1,5, and 2

are used to estimate the low, expected, and high estimates of tube

Insulator costs.

6_1.2.2r3. Radial_o_ce Contatnl_nt. To hold the launcher

tube tn place against the pressures developed during launching, the tube

would have to be wrapped for support. Kevlar fiber wrapping ts belteved

to be the best material available havtng the required strength at a

reasonable cost. The Kevlar thickness required is currently estimated

to be between 5 and 10 cm (6.5 cm for the coaxial hybrid EML/rocket

launcher). Kevlar is made from two components, yarn and epoxy resin.

The yarn iS currently befng Sold in quantity at $26.40 to $44.00 per

kg depending upon quality control. The epoxy resin is currently selling.

at $4.40 per kg. The degree of epoxy impregnation is a design variable,

and a typical mix is 60 percent fibers/40 percent epoxy. This combination

has a denstty of 1.38 g/cc as contrasted to Kevlar fibers with a density

of 1,44 g/Cc (Kevlar-49 Data Manual, 1976, 1982). The calculated cost

_2_. kg for a combination, using aerospace-grade yarn (at $44/kg) is then 20/kg of composite. Direct costs of labor to fabricate are .given

by DuPont personnel as being approximately equal to material costs. Since

the winding will have to be penetrated by conductors, it is most likely

that a COmplex buildup pattern would be selected and a machine would

be used to. make the winding. For this type of operation, a direct labor

factor of 2 times the material cost is .used with.the low material cost

estimate from the low force containment estimates. For the expected and

high cost estimates, labor factors of 2.5 and 3 are used with the high

materials cost estimates.

6.1.2.2.4 V_cum Co, tainlent. An exterior container would

be required to provide a vacuum seai and a mechanical connection for

the segments, This is assumed to be 1 cm thick on the coaxial launchers

and one or three cm thick on the railgun launchers. An.inexpensive plastic

coating would provide electrical protection for -normal Operations,

Aluminum is assumed to be used, but steel could be used if additional

strength iS desired. Aluminum in simple forms costs $i.68 per kg and

has a density of 2.7 g/cc. The cost of the insuiative coating is contained

in labor factors of 1.5, 2, and 3, used to form the low, expected and.

high cost estimates for the segment containers.

6,1.E.3 Electrical Power Storage

For the coaxial launchers, a single liquid, nitrogen-cooled

Brooks coil is envisioned to store electrical energy needed during launch.

For the railgun launchers, man_ homopolar generators (HPGs) with

liquld-nitrOgencooied inductors are. selected. The coaxial launchers .

use relatively high voltages and relatively low current to deiiver power

to the projectile while the railgunS use high current at relatively low

voltages to deliver their power to the projectiles, Additionai equipment

Including large transformers would be required to use the Brooks cotl

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with a rallg,n; other design Changes Would be required to use HPGs with

a coaxial aCuelerator. Accordingly, potential appliCatiOns Of mixtures

of these technologies are not considered. Technologically advanced

concepts such as superconducting storage ring unde_ investigation by

the University of Wisconsin are not considered because the costs are

htghly uncertain. The costs for the Brooks coils and hon_polar generators

are derived tn Table 6-5 usin_ information provided by P. Mongeau (Brooks

toil) and R. Marshall (HPGs). The electrical energy storage for all

designs is increased over the kinetic energy requirement of the.projectiles

to cover energy losses and impedance mismatches.

Brooks Coils. A Brooks cotl has a "life-saver"

shape (see Figure 4-13 and uses a stngle strand of conductor; alumtnum

was selected over copper because its lower density (2.7 vs 8.9 g/cc)

would result in a lower cost for cotls of the same size. Cooltng the

coil with liquid nitrogen would result in a decrease tn resistance (by

factor of about 10) and _reatly increase electrical energy storage

efficiency. The major design problems for the cooled Brooks coil are

considered tO be stress confinement and fnsulattng the coil loops from

each Other. The stress confinement requirement is assumed to be managed

by a reinforced concrete structure emplaced at the site and backed by

the rock of the site. The loop insulation requirement is assumed to

be settsfied, by Insulating blocks which are not fastened to the conductor

and have a thermal expansion coefficient which is very close to that

for aluminum, such as Bakelite. Such large amounts of conductor are

required to make the cotls needed for the launchers that the labor to

bufld the Coil is relatively small ";n comparison to the cost of the metal.

For this reason, labor factors of 1.5, 2, and 3 are used to provide the

low, expected, and high cost estimates. These estimates include the

confinement structure, internal LN2 plumbing, and insulation, but do

not include switching and control costs. For the Earth-orbital launcher,

the cost of the Brooks coils is estimated to cost from $237.5 M to

$475.1 M, with an expected cost of $316.7 M. For the hybrid EML, the

cost estimates range from $93.7 M to $187.3 M, with $124.8 M expected.

Only one Brooks coil is costed, and this could severely affect system.

availability if the coil were severely damaged.

6.1,2,3_2 HomRoljr Generat@rs(HPGs). The homopolar generators

(HPGS) are considered to be the largest source of uncertainty in the

mechanical design of the rat3guns. While very capable machines have

been bu!_t tn laborat_ries, the HPG experience Sttll represents a

relativel_ immature technology. _lere is also little experience tn their

manufacture or their use in operational systems, and there is no experience

to indicate a reasonable number of spares.

To avoid heat build-up in the launcher facility from conversion

of electrical into mechanical energy in the HPGs, it is likely that some

form of conversion external to the launcher area would be required.

Hydraulic conversion is selected for the HPGs because the hydraulic fluid

could carry away excess heat; ReconversiOn from mechanical to electrical

energy wou]d also generate heat; provision must be made to provide

circulating fluids or air to preclude heat build-up.

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6-14

TABLE 65. ENERGY STORAGE REQUIREHENTS

Velocity Energy

ProjeGti,le Energy Mass, kg (m/sec) (JOules,)

Coaxlal--Hybrid EML (I/2)

--Earth-Orbl tal

Rai/gun--Hybrid EML

--Earth-Orbi tal.

m

lS,4oo x (zooo).2 = 3.o x iolO

3,250 x (7000)_. = .....8.0 x 1010

15,200 x _(2000) z = 3.0 x 1010

5,900 x (7000) z = 1.5 x 1011

Brooks Coil Size, Mass, Cost (Scaled from Kolm and Mongeau, 1983).

Mass (AI, MT) = EJ_ Diameter (m) = '

2._x Z_

Energy increased by 15 percent for impedance .losses:

Coaxial Hybrid SHL Coil Energy = 3.0 x 1010 J x 1.15 = 3.45 x 1010

Coaxial Earth-Orbital Coil Energy = 8.0 x 1010 J x 1.15 = 9.15 x 1010

Brooks Coil _ _. Diameter (m)

Hybrid EML 3.45 x 1010 37,166 25.5

Earth Orbital 9.15 x 1010 98,646 35.3

Aluminum Cost, $1680/MT; Labor Factors: 1.5, 2, 3

Single BrOoks Coll

Cost Estimates $, M, 1981

Low E_._Dected Hi_/__

Hybrid EML 93.7 124.8 189.3

Earth-Orbital 248.5 331.4 497.2

56 Hegajoule Homopolar.Generato_s; Number and Cost

Energy increased by 38 percent for impedance losses in transfer from

HPG to inductor (85 percent efficiency) and inductor to rails (85 percent

efficiency):

Railgun--Hybrid EML HPG Energy 3.0 x 1010 x 1,38 = 4.15 x 1010 Joules

Railgun--Earth-Orbital HPG Energy = 1.5 x 1011 x i.38 = 2.0 x 1011 Joules

--56 MJ per HPG--

Number of Hybrid EML HPGs = 4.15 x 1010/56 x 106 = 742

Number of Earth-Orbital HPGs = 2.0 x i011/56 x 106 = 3573

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Adjust for spares:

Number of Hybrid ENL HPGs = 800

Number of Earth-Orbital HPGs - 3800

Costof HPGs:

For production runs of 10,000, $1,000 to $1,500 per MJ (R. Marshall,

UT) or $56,000 to $84,000 per HPG wtthan expected cost of $70,000

per HPG.

For production runs of about 1000, multiply by 3.

For production runs of abOut 4000, mu]tiply by 2

cost [stt 

Homopolar Gene_ Lo.__ww "__

Hybrid EML (800)

Earth-Orbital EML (3800)

94.0 168.0 201.6

425.6 532.0 638.4

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6-I6

The preliminary est4mates of the size of a 56-RJ homopolar

generator believed approprtiLte for these designs are 1.8 m in diameter

and 1.B m |ong, wtth a mass of about 10 MT. The size of these devices

tS such that they could be placed on one stde of the hybrid EHL/rOcket

launcher tube, but would have to be arranged tn a crescent about the

7 km/sec Earth-to-Orbit launcher tube.

Information on existing HPGs and possible production runs on

the Order of 10,000 indicates that HPGs should cost between $56,000 and

$84,000 each, with an expected cost of $70,000. FOr a device wetghtng

10 MT, the expected cost is $7.00 pe.r k.g, about that for an automobile,

Since the production runs for the Earth orbital launcher and the Hybrid

EHL/rocket launcher concepts will be about 3600 and 800, respectively,

the effects of learning on production cOSts are not expected to be as

great as for a production run of about 10,000. Accordingly, the costs

for the hybrid ERL r atlgun rocket launcher concept which would require

742 installed HPGs (750 for redundancy) are increased by a factor of

3. The Costs for the Earth-orbital launcher concept, which would require

3585 installed HPGs (3600 for redundancy), are increased by a factor

of 2. Since an appropriate level of spare HPGs cannot be forecast at

the present _lme, a level of about 5 percent is used. Thus 800 HPGs

for the hybrtd EHL/rocket launcher concept would cost between $94.0 H

and $201.6 H with an expected cost of $168.0 H. The 3800 HPGs for the

Earth-Orbital launcher would cost between $425.6 H and $638.4 M, with

an expected cost of $$32.0 M.

6.1.2.4 Swttchtn and ERL Inductors

Reusable Switches capable of handling the power and switching

rates contemplated for both coaxial and railgun launchers represent a

major area of technological uncertainty. The costs for the Switches

land their development) accordingly are very uncertain.

Sto_age tnductors would be. required for the railgun Systems,

because the HPGs would riot be able to convert their mechanical energy

into electrical energy quickly enough to feed the launcher efficiently,

Since the coaxial launchers use a Brooks cot1 to store electrical energy,

it is possible that secondary storage tnductors may not. be needed aiong

the launcher; however, the self-induCtance Of the long conductors needed

to transmit the current from the Brooks cotl to the. launcher-tube coils

and the short action times of these coils do not permit this assumption.

Accordingly, secondary inductors in equal numbers and ratings, are costed

for both coaxtal and railgun launches.

For the rallgun launchers, preliminary calculations by R,

Marshall (Rice, et al, 1982) indicate the secondary Inductors must store

approximately 48 HJ of energy at a current of 4 HA tO achievethe assumed

efficiency of 85 percent. To prevent resistive energy losses, the inductor

must also have a resistance of less than 2,7 x I0"oohms. For inductors

of coaxial or torOldai configurations, mass is sensitive to the number

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of turns and the conductivity Of the material. Since nO_mal conducting

metais drop in reSiStance by approximateiy a factor of 10 when their

temperature drops from room tempet'ature to that o_ liquid nttrogen,(LN2),

it is presently considered desirable to use LN2 cooled tnductors. This

results in a calculated significant reduction in inductor mass. Harshall's

preliminary calculations tndicate that a four-turn inductor Of this stze

would have a diameter of 1.5 m and a length Of 1.8 m. The indUctOr can

alsO reasonably be expected to contain the LN2 used to cool it. Foamed

insulation currentiy has problems _th cracking and separation upon

repeated cryogenic cycles; research ts being conducted in this area for

application 1:o reusable space vehicles. Thus, it ts reasonable to expect

that foamed insuiatton would be appropriate at the time of Implementation.

Contained foam (prefomed) insulation would always be available as a

back-up technology. Accordingly, the tnductors are costed wtth a labor

factor of 10 times the raw material prt_e to. reflect the uncertainty

of the switching technology. The current price for aluminum ingots ts

$1.68/kg (AI_, 1984). The requirement for low conductivity translates

into a requirement for controlled purity and thus may bring the price

up to $2.00/kg. Thus, the materials cost for an inductor of 1 to 1.5

tons iS $2000 to $3000. Other materials and labor, at a factor of 10

times the primary materials prtce, raise the cost per inductor to $20,000

to $30,000 for each unit. The low, expected, and high estimates for

the inductor subsystem are then formed by unit prices of $20,000, $25,000, __

and $30,000, and the same level of .spares, as for the HP.Gs.

Switching for the coaxial accelerators is expected to have

a slightly different set of requirements in that individual loops or

grOups of loops in the launcher coil must be controlled.to produce several

current pulses in very rapid succession. Because a detailed design is

not available, a COst analysis is not possible at this time. It is

expected that the costs of distributing and switching coaxial launcher

tube coils would be about the same as for the railgun accelerators. The

cost estimates, accordingly, are the same as for the ratlgun accelerators.

For the hybrid railgun EML, 800 inductors/switches (including

spares) would be required with unit p'ices in: the range of $20,000 to

$30,000; this leads to an estimate of $16.0 M to $24.0 M, with an expected

value of $2.0.0 M. For the Earth-orbital railgun, 3800 tnductors/switches

would be required; these are estimated to cost in the range of 76.0 M

to $114.0 M with an expected cost of $g5.0 M. These costs are also used

for the equivalent coaxial accelerators.

_ecttle 1n_ection Systems

A preboost system is believed to be desirable for optimum

launcher operation and would be required for plasma,amature raJlguns

to reduce rail erosion at low velocities, A design for the preboost

systems has not been developed, only preliminary concepts are suggested.

These all use gas to overcome the standing inertia of the projectile.

In ail caseS, marginal Operating costs are expected to be low ($50-$200

per launch), Except for the Earth orbital EML, the capttal Costs are

alsO relatively small,

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For both hybrid EML (coaxial and railgun) concepts, stored

compressed air is belteved tO be sufficient tO drive the projectile along

a 100 to 150 m tube with sufficiently high acceleration that initial

launcher acceleration would be Smooth. If the acceleration achieved

by thts system is as low as 1 g, the velocity upon entering the launcher

would still be about 50 m/s. The Cost Of a compressed air system to

tnClude compressors, storage tanks, piping and valves, is believed to k

cost in the range of $1 M to $3 M with an expected cost of $2 M. Because

the coaxial Earth-orbital EML does not have arc-erosion problems, the

compreSSed gas p_boost system is also considered appropriate, and is

charged to thts launcher. The Earth-to-orbit railgun launcher, however,

is required to have higher preboost velocities to operate properly. For

this reason a hydrogen/oxygen driven piston system is proposed. This

pre-accelerator, alSo about 100 to 150m long, would be used to accelerate

the projectiles. A piston would drive a mixture of nitrogen and hydrogen

gas which would in turn accelerate the pro:iectile. For the rallgun the

sabot necessary to distribute electromagnetic launch stresses would also

suffice tO prevent leakage. The cost of the propellants, given the

availability of electrical power and a gas liqulfication plant, w_ll

be in the range of $100-$200 per launch. The cost to design, manufacture,

and install the launching barrel segments together with the breech Section

are expected to far outweigh the cost of the steel used. It iS estimated

that the design, manufacture and installation of the barrel could be

accomplished for $80 to $120 M with an expected cost of $100 M.

The high cost of the conceptually stmple system is due to the

need to inject a large quantity of gas at high pressure in a ShOrt time

(0.4 Sec). High-pressure rocket engine pumps would be required to tnject

the liquids. These would have an operating time of only a few seconds

per shot, SO that service life should be very long (10 to 20 years),

based On modest extrapolation from current Space Shuttle experience.

Whtle rocket engine technology would be used, there are incentives to

permit large increases in mass of components and housings to provide

safety. Thts is One area where growth in allowable mass could reduce

costs. While an SSME currently costs on the order of $20 to $30 M, many

Components, such as nozzles, engine mounts, etc., would not_be needed.

Thus, it is reasonable t° expeCtdcoul that hardware components adapted to this task, Including spares, be purchased at about half the cost

for one SSME at the present time. The design effort, however, would

be significant and accounts fop most of the costs estimated.

6.1.2.6 Power Converston Plant (P_silgun Launchers)

Because motoring the homopolar generators would significantly

reduce brush life, and to reduce heat build up in the launcher facility,

railgun HPGs are envisioned to use hydraulic power from a conversion

facility neap the launcher. It is expected that hydraulic motors would

both save space in the launcher facility and be somewhat lo_er in cost

than electric motors, While it would be possible to transmit power in

the form of steam from the power plant, the tranSmiSSion flexibility

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of electrical power suggests that it wouid be better to accept the

converSion inefficiencies of reconverttng the electrical power back into

mechanical power at a station near the launcher. Such conversion would

also permit an ambient-temperature hydraulic fluid to carry away the

heat from driving the HPG rotor. This form of thermal control is

Considered necessary for the Earth-orbital railgun and desirable for

the hybrid railgun/rocket system, even though the presence of the hydraulic

fluid (assumed to be water-based) may increase the risk of an electrical

accident.

Since electric to high-pressure hydraulic cOnversions of this

size do not appear to have been undertaken previously, no good analogy

is available to draw upon. The conversiOn power cost was established

aS CoSting one-fourth that of the power plant on .a per kilowatt of capacity Or $550 to $700 per kwh, with $625 expected Conversion capacity of

15,000 kwh would be required for the hybrid EML, with costs in the range

of $8.3 M to $10,5 M ($9.4 M expected). For the Earth-orbital launcher

with a launch velocity of 7 km/s, 50,000 kwh of capacity would be required

with a cost range of $27.5 M to $35.0 M and an expected CoSt of $31.3 M.

The conversion power level selected would permit recharging the homopolar

generators in one hour, rather than the. one and one-half hours.expected

between iaunches.

6.1.2.7 Water Distilling Plant

The launch site cannot be assumed to have sufficient fresh

water either to support the launcher operations (power plant, hydraulic.

conversion, LN2 plant cooling, etc.) or the operating personnel and their

families. Whiie much of the water used in launcher operations would

be recycled, the same cannot be said for water for human consumption

and household use. Accordingiy, a water distilling plant may be needed.

The plant is sized at 1,000,000 liters per day, representing 400 liters

per person per day for 2500 people. This is expected to have reserve

Capacity for the crew as well as families and transients. The distillation

plant would use the heat rejected from the nuclear plants in their bottom

cycle and would_therefore represent a predominantly capital cost. This

type of facility is expected to cost about $2.5 per liter-day of capacity,

Or $2.5 M. Because of the effective integration of this system into

the power plant, the uncertainty in cost is very high--it may cOst much

less or somewhat more depending upon the Specific designs selected. Solar

evaporation and condensation is'also available in this price range, but

would have higher operating costs, Solar evaporation _would be used if

the site does not require a nuclear plant.

6.1,2.8 Gas Ha_ Facilities

Three types of gas handling facilities are expected: (I) a

liquid nitrogen plant and storage area; (2) an evacuation system for

the launcher tubes; and (3) hydrogen and oxygen .storage for the preboost

System needed for the Earth*orbital railgun, These are discussed in

the following three sections.

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6,1.2,8_1 L|qu|d ,ilttro_en Plant .and Storage. To provide

acceptable inductor masses, the resistance of their conductive material

must be dropped by approximately an Orde_ of magnitude from that available

at room temperatures. Liquid nitrogen cooling of the tnductors was

selected over superconduction because the state of superconducting

technology does not permit reasonable cost estimates in the foreseeable

future. The major uncertain!ties in selecting :N 2 cooling are the

requirements of LN2 due to tnsulattve losses in the tnductorS and their

plumbing and to the efficiency of transmitting electrical power through

the inductors tO the launcher. These would involve both thermal and

electrical losses, placing a heat-sink requirement on the LN2 and requiring

insulated plUmbing/duCtworks. The major uncertainty, however, iS believed

to be the insulation requirements and the costs needed to meet the_,,.

BaSed on the heat of vaporization for LN2, 47..6 kcal/kg (0,05534 klqh/kg),

and an assumed 15 percent of input energy as a combined thermal, and

electrical inefficiency causing LN2 boil*off, the requirements for LN2

are calculated for the maximum launch rate of 16 iaunches per day. From

information prOvided by J. Cost, Air Products Company, a plant providing

325 MT of LN2 per day wOuld cost $4 M and would scale upward by a 0.6.

power law on cost. Mr. Cost believes the 0.6 power law is slightly

optimistic, so a 0.7 power law is used to Calculate the expected costs,

The. costs of multiple units of 325 MT per day of LN2 capactty are used

to calculate the high cost estimate. A Summary of the calculations is

presented in Table 6-6, where the dissipated energy is 15 percent of

115 percent of the i/2mv 2 energy, or 17.5 percent of the required

projectile energy for each launcher. The cost of the LN2 plant is small

in relation to other costs, as shown in Table 6-6.

6.1,2_.8,2 Evacuation S_'Srtem for Launcher . Tube. The evacuation

Of the small-diameter launcher tubes to approximal:ely 1/100 (7.6 mm Hg)

atmosphere would require the removal of less than 1300 kg of air. For

the larger tubes, removal of u_ to 3000 kg would be required. This could

be accomplished with rotating impeller pumps, able to achieve high volume

thrOughput. The removal of 99 percent of the air would leave 13 to 30 kg

of air in the bore. At least three pumps are estimated to be required.

Each pump would be able to handle the evacuation unassisted. The

installation is estimated at $1 to $1.5 M for ductwork, shelters, pump_,

and motors.

6,1,2.8,3 Water Electrolysis Plant. To provide hydrogen and

oxygen for the Earth-orbital ratlgun S preboost system, a water

electrolysis plant would be needed. Since hydrogen has much lower

viscosity than air, it has been used in large electrical generators to

reduce the atmospheric friction between rotors and stators. This hydrogen

has usually been produced by electrolysis of water with the electricity

produced by the generators. Accordingly, the cost of the electrolysis

plant is contained within the estimate for the power plant. The facilities

to liquify the gases are provided within the estimate for the liquid

nitrogen plant. In addition to these elef_ents, there would also be

additional costs for storage and handling facilities. These are estimated

at $0.2 M, $0.3 M, and $0.4 M for the low, expected, and high costs of

these facilitieS,

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TABLE 6-6. LIOUIO NITROGEN REQUI_NTS

Detemtninq Ass.umption_ E_ergy losses resulttn_ in LN2 boil off are

17.5 percent of projectile energy. _

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LaUncher .

Coaxlal--Hybrld EML

--Earth Orbital

Ratl gun--Hybrtd EML

--Earth Orbital

Projectile E_ergy,

aouies

3.0 x lo!o

8.0 x 10!o

3.0 _ 1010. "

1.5 x lO11

LN2 Energy

16 Shots/Day @

17.5percent_ Joules

8.4 x 1010

2.2 x 1011

8.4 x 1010

4.2 x 1011

LN2 heat of vaporization - 47.6 kcal/kg -- 199,254 aoules/kg

Launcher Plaht Capacit_N2 Plant Output/Day, MT

Hybrid EML

Coaxlal--Earth .Orbital

Railgun--Earth Orbital

421.5

1104.1

2107.9

Capacity

Plant Cost Scaling: C ($, M, 1981) = ($4 M) 325 MT "

x " 0.6, 0.7, 1.0

Multiply by 1.5 for storage, plumbing, etc.

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Hybrid EML 7.0 7.2

Coaxial--EO 12.6 14.3

Ratlgun--EO 18.5 22.2

Plant Cost+ $, M, 1981

Expected

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6-22

6.1.2.9 itandltng I)evices/_Systems

Handltng devices, principally cranes and small ratlcars, would

be needed to manipulate the projectiles and tube segments, as well as

hOmopolar generators and/or tnductOrs. The handling devices are envisioned

to travel a|ong suspended beams attached to the floor, roof, or walls

of the launcher tube housing. Light elevators can cost as low as $500/m

and heavy .freight elevators can cost $1250/m (personal conversation,

M,, Minelt, Otis F,levator Co., March, 1982). It is expected that there

would be multtpl( cars or cranes per set of tracks and that they would

be of very heav; duty construction. Accordingly, unit costs for the

handling systems a,_e believed to fall in the range from $1000/m to $5000/m

with an expected value of $3000/m. For a iauncher with a nominal length.

of 2000 m and a loading-preboost length of 100 to 200 m, the estimated

costs for the handling devices and associated equipment are estimated

to cost: in range of $2._Mto$11 M with anexpected value of $6.6 M.

6.1.2.10 Control Center, Controls and Nonitortnq S_sCems

A preiiminary system design, as well as specification of the

control requirements, is needed before accurate estimates of the control

costs can be made. It is assumed that inductors and coils or homopolar

generators and inductors could be monitored and their switches controlled

from a master control center for a relatively low cost per unit. A

t_ntative e_timate of $1000 to $10,000 per HPG/tnductor set is used for

the railgun launchers, and $5000 is used as the expected value. Because.

the switching and control systems are not well described for the coaxial

launchers, the equivalent costs are assumed for the coaxial EMLs as for

the railguns, An additional $5.0 M is added for the central, control

system and other unidentified costs, For the hybrid EML/rocket launcher

concepts, the costs are estimated to range from $6.0 M to $15 M for the

800 HPG/inductor sets (or equivalent; inductors/sw.ttches for the coaxial

launcher), with an expected value of $10 M. FOr the Earth-orbital EML,

concepts the costs range from $8,0 M to $39 M ($21 M expected) for 3800

HPG/inductor sets or their equivalents for. the coaxial launcher.

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A tra.cking, system would be required to monitor the trajectory

of launched projectiles. Since _hese launchers are not _intended to handle

high value or 'tazardous payloads, the necessity to have accurate knowledge

of the trajectory under abnormal conditions can be relaxed. For this

reason, adoption of a military tactical radar station is selected to

cover the near-launch-site trajectory, with long-range monitoring and

cor_trol in space conducted elsewhere. Accordingly, low, expected, and

high values for the radar station are established as $5 M, $10 M, and

$15 M.

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6-23

6.1.2.12 Acctdent Recover $ tams

Because hazardous m_¢erials are not anticipated to be transported

by the launchers, no formal accident recovery system is planned, and

none tS costed, It is expected that misfired projectiies which could

be recovered eastly would be, but this would, be a very minor part of

operations activity.

_ons Cost Esttmtes

The costs to operate a launcher facility have, two components:

the recurring costs, associated with each projectile, and the annual costs

for personnel and supplte_; projectile costs are summarized in Table

6-7; personnel and supply costs are summarized in Table6-8.

e

While the projectile is part of the launch system, payloads

are usually considered to be part of some Other mission or activity and

are not costed here. The cost of the bulk payloads considered for the

supply missions would be very low, in any case. The costs for their

development .are addressed in. the Cost estimates for -an Operational Test

Program.

The annual cost estimates cover operation of the. facility for

use as a launch site only, While provisions are made for people and

consumables to load the bulk payloads, this would be a simple prOcedure.

Only tasks such as loading of liquid propellants and other fluids,

initiation of guidance systems, and verification of status would normally

be undertaken at the facility. Use of the facility for research or other

programs would involve additional costs..

6.2.1 Pmjecttles and Ntsston Peculiar Equi_nt

The costs for Earth-orbital projectiles are highly uncertain..

due to technological advances needed, to achieve and demonstrate the.

capabilities required. In addition, the annual quantities required

(hundreds or thousands) may not be large enough to assure that major

savings through mass production (in the manner of automobiles) can be

achieved, The ability to achieve the low estimates depends upon keeping

labor costs low.

For the hybrid EML/rocket launcher concept, the projectiles would

be derivatives of existing solid-propellant stages with relatively low

technology risks. Those stages must, however, be produced at rates much

higher than achieved to date. The next three major sections discuss

the costs for the four projectiles and their mission peculiar equipment,

Costs are summarized in Table 6;-'7.

_rld EXL._ocket Pro ectile Coaxial and Rail un

6.2,1.1,1 Payload, Payload costs are not considered in our

analyses. _Oadspay are expected to be bulk materta]s and should

not be affected by the accelerations of the electromagnetic booster.

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6 .1.1.2 Nose a r Paload S h Ud. FOr the hybrid EML

concepts, the nose COne or pay e an adaptto, of existing

designs. This shrOud is intended for a 2-3 km/sec aerodynamic regime

which is similar to those currently experienced by expendable launch

vehicleS. Production rates of the adapted design are expected to reduce

the costs significantly to yield a range of $500 tO $3000, with an expected

value of $1000.

6.2.1.1.3 S rutture. The primary structure for the hybrid

EML projectile wo_of the solid rocket motor cases and their

interstages or Connectors. The provision of the inductive rings on or

near the surface of the coaxial projectile is expected to have only a

small effect on the design costs for the coaxial projectile, and result

in no appreciable increase in the manufacturing cost Over a case wtthout

provision for rings. The cost of the rings isexpected 1:o be approximately

equal to the cost of the sabot needed for the ratlgun projectile and

ts discussed in Section 6.2.1.1.6. The railgun projectile would require

a stronger case, especially in the rear part of the projectile where

provision muSt be made for thrust transfer from the sabot. However,

because both cases must be designed to withstand lO0-g, accelerations,

the cost differential between the rallgun coaxial projectiles due to

structure strengthening ts expected to be about 15 percent. Production

rate effects associated with the high volume of projectiles are expected

to keep the costs of the structures low in relation to other costs.

Because the projectile case is an integral part of the solid rocket motors,

specific cost estimates are made in the propuiston section (6.2.1.1.8).

_.1.4 ema] Proctton S Side . For the

hybrid EML concepts, the r)rojectl not require thq protection.

_. These projectiles would require stabilization

(not steering) fins. Their operating regime would not be severe and

their cost is expected tO be tn the range of $100 to $500 with $300 being

the expected cost.

6. 1..6 ix)t or R n s. Because of relatively low launch

stresses, the sabots for the rai gun hybrid: EML projectile would not

require exotic materials, but rather an inexpensive nOn-conductor tO i_

protect the projectile together with a conductive plate, probably made t_

with copper or alumtnum alloy. The sabot cost shouid be little more

than the price of simple constructs from a manufacturer in the range )(

from $1.30 to $2.00 per kg; labor factors of 1.1 to 1.3 are appropriate. _}

The coils or rings used on the coaxial accelerator projectile are expected

to be made of a copper alloy and have no unusual requirements. For both I_

the coaxial and ratlgun launchers, cost estimates are in the range of (J

$1000 to $1700, with the expected value of $1400 based on a mass of 650 kg

for the coaxial projectile rings. U

6_1. nt. and Supoltes, The n

projectile gatiOn and .......

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control system would require ftnal checkout and preparations at the launch

site, This checkout procedure is expected to require a few man days.

Low, expected, and high cost estimates are $1000, $2000 and $5000 for

time and materials.

6.2.1.1.8 inst tit Packa . The instrument packages for

all proje_gned for high production and low unit costs.

It ts assumed that assemblies with moving parts (gyros, etc.) would be

avoided or designed to be cost effective, and complex functions would

be handled by software rather than hardware. To keep costs down,

multi-year procurements of components/assemblies would be used, even

if final assembly is regulated by demand. Based on these conditions

and annual demand for thousands of instrument packages, it is likely

that the cost of the instrument package (20 kg for hybrid EML projectiles)

would lte in the range of $1000 to $3000. The expected value is selected

as the mid-range or .$2000. _

6"2"l"l"gPr°ulst°n S. tea. The projectiles for the-hybrid

EML concepts would consist Of three Stages with total, masses (Including

payload) of 15,400 k9 for the coaxial projectile and 15,200 kg for the_

railgun projectile.. In contrast, the current Scout expendable launch

vehicles has a mass Of about 21,500 kg .(excluding payload) and the three

Scout upper stages have a .total mass of.6687 kg, or about 44 percent.

of the proposed projectiles.

For the latest production run of 15 Scout vehicles, production

costs were about $4 M per vehicle and the total launch cost is in the

rang e of $8 M to $10 M at a launch rate of about three per year (EDD,

1976). In the late 1960s, Scout Launch vehicles were being manufactured

for $I M each, with launch services on a basis equivalent to current

charges being about another $I M. At that time, the launch rate was

about I0 per year. Thus, in addition to inflation, the launch rate

decrease has resulted in a significant cost increase because overhead

costs cannot be spread across many launches.

The projectiles must be mass-produced at _,uch lower costs.

Because the stages would essentially be cannisters filled with relatively

stmple and inexpensive chemicals, it is expected that they could be

prOduced in quantity at about the same cost per kilogram as an automobile

or $6.00 per kilogram. Material costs for propellants (or $3.00 per

kg), however, are expected to keep the cost from going much below this

level. The structure of the projectile is expected to be formed from

fiber matrix composites (e.g., Kevlar) which currentiy costs about $40

per kg, but contribute only 5 to i0 percent of the projectile's mass.

It iS also reasonable to expect the cost of these materials .to drop in

the future. Accordingly, the cost per kilogram f, • the propulsion system

(which includes the structure) is estimated to be in the range from $6

to $12 with $9 expected. This leads to estimates of the coaxial projectile

from $78,000 to $156,000, with an expected value of $117,000 for propellant

and Structural casing mass of about i3,000 kg, This estimate tncludeS

a small auxiliary or attitude control propuisiOn SyStem. For the Pailgun

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prOjeCtile, the estimates are increased by 15 percent to consider the

stronger motor cases and other structural components. This yields an

estimated range of $90,000 to $180,000 and an expected estimate of

$135,000,

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6.2.1.1.10 Transrtatton, Transportation charges for all

projecttlesare established at $500, reflecting surface transportation

in economical lots. This is consistant with the projectile's major use

as a method of bulk transport.

6.2.1.2 Earth Orbttal Pro.lecttle (Coax|a1 Launcher)

6.2.1.2.1 Pa loa.d, Payload cOsts are not considered in our

analyses. _a_'epa_ expected to be relatively inexpensive bulk

materials which would not be affected by the high accelerations of the

launcher.

,,_,Nose Cone, The nose cones for the Earth-orbital

launchers wOuld undergo a much more severe thermal environment than that

for the hybrid EML projectile, and an ablative nose cone was assumed.

Considerable technical verification isneeded to develOp experience and

confidence in metallic ablators, but inexpensive metals such as steel

prObably can be used and their COSTS can be quite low. Thus, materials

cost per kg can range from less than $1.00 for modestly priced alloy

steels to $10 for some specialty stainless steels, Labor factors for

high production rates are expected to be in the range from 2 to 5. Thus

for a nose cone mass of 110 kg, the costs can range from $220 to $5500.

Accordingly, a low estimate of $220 reflects use of modestly priced steel

and a labor factor of 2, an expected estimate of $1000 reflects use of

higher priced steel and a labor factor of 5, while rthe high estimate

of $5500 assumes use of expensive steel ($10 per kg) and a labor faci:_r

of 5.

6,2.1,2 3 Stru ture. An inexpensive steel structure with

a mass of 300 kg wou d be needed to provide strength. The materials

cost for this steel would be in the range of $0.55/kg to $1.00/kg. Low

labor factors (in the range of 2 to 5) would yield estimates in the range

from $165 to $1500. The low end of this range is not cOnsidered feasible,

so the estimates are increased to $1000 to $3000 with $2000 being the

expected value.

" _ Protection S $idebod , A thermal

protectton hsySt/_m.#woUm_t_b_ ?-_qutred and ts envisioned tO be made Of carbon-carbon (C C) material (or other advanced composites) and have

a mass Of 200 kg. At .the present time C-C materials are expensive to

fabricate because they are labor intensive in all manufacturing stages

and in quality control. Given historical progress in materials

development, it is expected that C-C materials can be prOduced in the

time frame of the launchers' at an installed price equivalent to that

of tungsten metal or $33 per kg, A likely range in this estimate is

from $15 to $70 per kg. Thus for 200 kg of sidebody structure and thermal

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protection, the costs would range from $3000 to $14,000 wtth an expected

value of $6,(;00.

_. The use of ftns for aerod_mamtc stability

would be requ!red, and their mass ts estimated at ZO kg, At thts low

mass, exottc materta]S such as carbon-carbon composites or tungsten do

not produce htgh Costs and are associated wtth materials costs tn the

$30 to $40 per kg range and wtth ]abor factors of 2 to 5. The expected

]abor factor ts 3. Thts yte]ds cost estimates in the range from $1300

to $3300 per proJectt]e, wtth an expected estimate of $2000.

6.2.1.2.6 Coils. The Inductive cotls for the projectile would

have a mass of 300 kg and wou]d be made from copper a]]oy. Materials

costs are less than SZ.OOper kg and labor factors of 2, 3, and 4 yte!ds

estimates tn the range of $1200 to $2400 wtth $1800 the expected value.

_u Eu_ nt a dll S 1 S. The launcher site

checkout is expected to requ ew man aays, and is estimated to cost

from $1000 to $5000, with an expected cost of $2000 (see

Section 6.2.1.1.7).

6...2.8 !nstrume t Pa . The instrument packages for

a11 projectiles are expected to cost about the same. Through design

or htgh production rates, this cost ts expected to be in the range from

1000 to $3000, with an expected cost of $2000 for a 20-kg package.

6.2.1.2,9 Pro uist, S s . From the definition of the

projectile, the dry mass of the propu sion system would be 425 kg, and

the mass of the propellants wou]d be to achteve a 500-km circular orbtt

would be 1150 kg. The available cost Information on propel|ants exhibits

wide range--from $0.18/kg for ch]ortne to $30/kg for prOpe]lant-grade

hydraztnes. For the currently expensive propellants, new production

facilities would be required. With economical production facilities,

the propellant costs are expected to be about $6/kg; for the i150 kg

propellants tn this projectile, the propellants wouid cos_ $6900.

Because the projectile for the coaxial accelerator would have

a]tght support structure/thermal protection system ttts expected to

require tanks to contain the fuel (N2H4) and Oxidizer (ClF3). It ts

conceptually possible that the support structure could serve as tankage,

but safety considerations wtll probably not pemtt such a single-point

failure mode. The tankage requirements, however, are not expected to

be a major cost problem on the basts that tanks wt]l be produced at rates

well tnto the thousands per year.

To achieve a reasonably attractive cost for propulsion units,

ttts necessary to hypothesize advances in several production techno]ogtes

{e.g., cOmputerized machining, we]dtng, and testing), tn addttton to

htgh production rates. At the present time, the major cost problems

are believed to be tn the area .of engine mechanical con%to] rathe_ than

tn the engine or tankage, it ts in these areas that improvements tn

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production and testing technologies are belteved to have the btggest

payoff in reduced costs. High production rates for a 425-kg (dry weight)

propulsion system and 75-kg ACS/astrtontcS package are believed tO be

able to reduCe unit costs to the level of an automobile, or about $10,000,

but this ts considered optimistic, espectally for the early part of the

program. The low cost ts therefore established at $25,000 with an expected

cost of $100,000, and a high cost estimate of $200,000, all of which

tnclude the cost of propellants. The esttmate of $100,000 ts about

one-tenth that of a cOmparably sized low-<j untt produced today at a

production rate of 5 to 10 units per year.

6.2.1.2,10 Transrtatton. Transportation charges for all

projecttoes are establ shed at 500. See Section 6.2.1.1.10 fop additional

Information.

6.2.1.3 Earth-Orbital Pro_uncher)

_.3.1 _. Payload costs are not considered in this

analysis, Host payloads are expected to be bulk materials with relatively

low unit costs, and resistant to damage from launch forces.

6.2.. ,2 Nose Co e, The nose cone for the Earth-orbital

launcher Is expected to undergo a more severe thermal environment than

that of the hybrid EHL. An ablative nose cone is considered to be

required. The discussfon of Section 6.2.1.2,2 indicates that the cost

of materials can have a wtde range.

Inexpensive or moderately-priced steel is expected to be

appropriate for thts application. Tile least expensive steel costs about

$0.55/kg and some heat resistant alloys cost about $1.00/kg.

Hoderately-priced stainless steels cost about $3.00/kg. Some very high

temperature steel alloys can cost in the range of $10/kg, Appropriate

labor factors for high production rates are in the range from 2 to 5.

Thus, for a nose cone mass of 420 kg, the cost per projectile can

potentially range from $420 to $42,000. The expected cost estimate is

$1260 reflecting steel at $1.00/kg and a labor factor of 3. The range

selected for this assessment ts $840 to $2100 reflecting low and

moderately-priced steels and labor factors of 2 and 5.

3311Z_j_J]l_g_. The structure Of the railgun projectile

would be required to sustain substantial launch stresses and must be

themally protected against atmospheric heating. The support Structure

is envisiOned to be made of inexpensive steel wtth a mass Of 2730 kg,

and the thermal protection system (Section 6,2.i.3.4) would be made of

reiattvely expensive composites with a mass of 100 kg. The structural

steel would: have a cost of $0.55 to $1.00 per kg, and labor factors are

expected to be tn the range of 2.0 to 4.0. ACcordingly, the estimated

cost for the steel structure is from $3000 to $11,000 with an expected

cost at the midrange of $7000.

6.2.1.3.4 TherBai\_¢t_ S stem. The TPS ma_Hal would

be made o_ an advanced comr)osl¢e wnlcn l:s expected to have an installed

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Cost in the future about that of tungsten metal or $33/kg. The range

on thts cost ts expected to be plus or minus a factor of two or about

$15 to $70. Accordingly, the TPS for the ratlgun projectile is estimated

to have a cost range of $1500 to $7000 wtth an expected co_t of $3300.

6.2.1 3.5 Fins, Use of small fins for aerodynamic stability

would be required, and the4r mass 4s estimated at 20 kg. As In the case

of the coaxtal £arth-orb4tal projectile, the cost range _s $1300 to $3300

per project41e, vtth an expected estimate of $2000.

6o2.1.3._6Sabl)tS. Both a forward and aft sabot wou]d be needed

to support_correct]y in the tube during acce]eration. The

aft sabot would also serve to conduct the railgun current across the

back of the projectile. The sabot masses are esttmted at 200 kg for

the forward sabot and 100 kg for the aft sabot. The aft sabot ts expected

to be made of an Inexpensive non-conductor to protect the projectile

and a conductive plate to provide a current path. The forward sabot

would be made Of a non-conductive material, possibly a plastic. Under

the assumption that the strength associated with exotic materials is

not required for the ratigun accelerations, tt is ]ike|y that the copper

alloy will cost approximately $1.78 per kg whi]e the plastic will cost

$2.00 per kg. The labor factors for the sabots should be low (1.5 to

2.0), so the cost range is estimated at $900 to $1200 with an expected

value of $1000.

6_c out £ utnt and Su lies. AThes launch-site checkout is expected to requ y a few man days. for all other

projectiles, this is estimated to cost from $1000 to $5000 with an expected

cost of $2000.

6. 1.3. Str t Parka e, The instrument packages for

a]l project1 _ected to cost about the same. Through design

for high production rates, this cost is expected to be In the range from

$1000 to.$3000, with an expected va|ue of $2000.

6.2 1.3. Pro u] on S st . The propulsion system and

ACS/astrtonlcs for t e rai gun proe--_ile is very simtla_ in mass to

that of the COaXial projectile, and the same costs are used. These are

1o_ and high costs of $25,000 and $200,000, with an expected cost of

$100,000.......

6.2.1.3.10_ Trans rtat!on. Transportation charges for al1

projectiles are estab]lShed et $500. See Section 6.2.i.1.9 for additional

4nfOmatton.

5.2.2

Operating perSonnei and operations support would be located

both tn the continental U.S., as well as at a remote mountainous stte.

Stnce most mtSstons have been Identified as gOtng Into.either a 28.S-degree

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orbit or a O-degree (equatorial) orbit, the. mountainous s_te wOuld not

be located in the continental U.S.. For multtple daily launches to a

single space station, the launch site would be located within a few degrees

of the equator and would launch to an equatorial space station.

In addition to the technicians required for the launcher and

projectile support, an ongoing engineering effort would be necessary

tO maintain and improve the launchers and their Operational effectiveness

during their lifetime. This level of engineering effort cannot be forecast

precisely at this time. It is expected, however, that most Of the design,

assembly, and repair/rebuilding efforts would be accomplished by the

operations staff envisioned in the cost estimates. Manufacturing efforts,

however, would be contracted. Because of the difficulty in transporting

large, very heavy equipment, it is expected that much of the final

assembly, repair, and rebutld effort would be conducted on site.

Accordingly, for the personnel estimates, it is expected that about half

of the managers and englneer_ would be located in the continental U.S.

and the others would be located on site. It is expected that projectiles.

would be butlt and checked out before being transported to the site.

Any final efforts would consist of: loading propellants and flutds,

setting initial conditions for guidance hardware, and making a ftnal

test of satisfactory payload conditions using a preprogrammed computer.

The personnel requirements are estimated to range from about

305 to 885 people, with expected estimates tn the range of 500 to 540

people, depending on the specific system being considered (see Table

6-8). The cost estimates are dependent upon the assumption that the

equipment is inherently reliable. The cost estimates also provide for

sufficient spares that rebuilding of equipment can proceed on a schedule

with little disruption fOr emergency repairs. For example, it is assumed

that the brushes on the homopolar generators would have a normal wear

life sufficient for thousands of launches, which would permit at least

two years between brush replacement. The a,nuai purchase of supplies

tO support the launcher facility is estimated to be about twice that

needed to support the power plant. There is insufficient information

to justify a specific level for supplies for the installation; these

levels were selected as being a reasonable assumption. For the hybrid

EML/rocket launchers, the power plant costs represent a small, non-nuclear

plant; commercial power would be preferred if available, if commercial

ower were available, these supply costs would be considerably reduced

one-half to two-thirds), butthe personnel would be needed to malnta_n

the local d4strlbution system.

The personnel estimates, as well as supply estimates, are given

tn Table 6-8. A value of $50,000 per man-year ts used in estimating the cost of the staffing, Thts tncludes an allomnce for launch site

personnel overhead, e.g., transportation for vacation purposes, and is

a direct cost estimate. No estimate of support for programs which use

the launcher for scientific or technical purposes is included in the

estimates of annual expenses which range from $16.0 M to $64.0 M with

expected estimates tn the range from $32.0 M to $36.5 M.

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Oetatls of a development test program are difficult to project

at this time because of uncertainty in the technology. At this time,

the development tests are envisioned as having two major aspects. The

first aspect would be a test of one or two launcher tube segments to

verify the perfomance Of launcher segments and other subsystemS. An

inert test projeCtile _ould be accelerated at full thrust to give

confidence in the design before any major construction at the launch

site were undertaken.

The second aspect of the development test would be part of

the transitiOn from construction to operation. At this time, it is

expected that most of the investment in the launcher would be complete.

The deveiopment test wOuld concentrate on asSuring that the controls

operate correctly and that the terminal velocities can -be achieved

reliably. In addition, there would be a need to verify adequacy of

projectile designs. While some of this verification can be done by

subscale electromagnetic launchers, there would still be a need to verify

fullscale designs. If the fullscale tests are successful, the development

test would be expected to last about one year. if they would be

unsuccessful, rework of either the launcher or the projectiles would

be required and non-prOductive costs would mount. The first Component

of the development test, that of full*sized segments(s) of the launcher,

is expected to preclude the need for any major investment period rework

of the launcher design.

In addition to the construction crew costs, which are included

in the development/inveStment cost estimate, there are operating crew

training costs, which would start during the construction phase and

continue through the development test phase. These costs are approximated

by using two years Of operations crew costs.

These considerations are taken into account in the development

test programwhich includes:

Testing of an all-up launcher segment (in the continental

U.S.)

Development of projectiles, estimated at 400, 600, and

800 man-years of effort at $100,000 per man-year

e Transient housing fOP 500 to 1000 people at $10,000 per

person and convertible to pemanent transient housing

of 250 units

• Two years of launcher operations cost.

The development test program estimates are given tn Table 6-g. A summary

of investment and operations coSt estimates is given in Table 6_i0.

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TABLE 6-9. DEVELOPMENT.TEST PROGRAN COST

ESTINATES ($, N, 1981)

rim rl i m f fT r • i i mmi __ •

Low Hi gh

ii i i i

Launcher Segment Tests. 75 100 150

Development of Projectiles 40 GO 80

Mtscellaneous Facf ltttes 3 4 5

Two Years of Operations 40 70 13._00

Totais 158 234 365

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Expected

6.4 Economic Evaluation of ENL Concepts

The EMLs considered in this report are one alternative for

the Earth-to-space transport of bulk materials. This section provides

Cost information for some near-term alternatives to the EMLs and then

provides a comparison of transport costs for two mission applications

using EMLs or their alternatives. Section 2.0 describes and develops

transport projections for both LEO space station and manned lunar base

supply missions. This section evaluates the.economic usefulness of EMLs

to these missions.

G.4.1 Launch Systems Alte_atfves

In addition to the Space Shuttle and its potential .derivatives,

other launch systems are also considered which might be available in

the year 2000 and might be competitive with EMLs. The ;,ybrid EML/rocket

concepts with EML velocity capabilities of 2 km/s are viewed as being

in this category in that they are within reach of a moderately ambitious

program whi.le the Earth-orbital EMLs have slightly higher costs, as well

as higher technological uncertainty. Accordingly, within the context

of supply missions starting, in the year 2000, the hybrid EML/rocket

concepts are used. for cost comparison. For the more demanding lunar

base. supply mission, where launch velocities of 12 km/s or more would

be required, an uprated Earth-orbital EML would be needed. Such a facility

would, require several launch tubes and its costs would be at. least, those

of the two-launcher system discussed in the ESRL report.

are:

The alterna.tive launch systems considered and described b+low

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Hybrid EML/rocket

LEO and lunar suppiy EML

Four-stage solld propellant rocket

Single-stage gas gun

Rocket-powered sled

aceShuttle and Shuttle-derived Unmanned Launch Vehicles

ULVS).

These alternatives are nomaltzed to provide an equivalent level of

transport: 800-kg payload to the LEO Space Station at 500-km altitude

or 500-kg payload to a lunar space station. For both the current and

derived STS vehicles, payloads might be flown on a space-avallable basis,

but in any case would be flown in larger quantities; they are considered

to bear the same cost per kilogram as general cargo.

_. The expected total capital investment

of the 2-km/s hybrid coax a auncher is estimated at $1,134 M with annual

operatiOnS costs of $32 M. Projectiles for this concept have an expected

unit cost of $124 K at a production rate Of ten per day. At lower

production rates, the unit costs are expected to be abe.lt 50 percent

higher because production overhead would not be. amortized as thoroughly.

The expected costs for the railgun hybrid EML are slightly higher and

areshOwn in later tables.

LEO and L nat Su 1 ENL, This concept has not been

investigated In detail tn this report. It would require several launcher

tubes to accomodate the expected launch rate. The factltty requirements

would approximate those of the ESRL system (Rice, et al, 1982). The

investment accordingly is estimated to be $5 B, and annual support costs

are estimated at $68 M. The projectiles, however, would cost about the

same as the Earth-orbital projectiles discussed in Sections 6.2.1.2 and

6.2.1.3. The high investment cost is due to the ltmtted launch window

which will require multiple launches within a few hours on a few days

each month. Launches to the LEO space station, however, would have daily

launch windows.

Four-Sta Rocket. The four-stage rocket envisioned for

use t n th_'-_,e, sion would be a. modified version of the

three-stage projectile used for the ratlgun hybrid EHL together with.

a first stage which would replace the boost provided by the EHL. The

entire vehicle would be reoptimized to reflect the lower levels of

acceleration over longer periods of time which would be required while

flying in the lower atmosphere. These iower levels of acceleration would

pemit the entire vehtcle to be designed to s!Ightly lower mechanical

strength conditions than the EML projectiles, but no major cost saving_

is expected. After the intttal design and test phase, the costs for

the four stage rocket would be strongly dependent upon materials usage

rathe_ than manpower as is the case with current vehicles. The production

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6-37

philosophy, moreover, must be the same as for the EML projectiles: capital

(machines) should be used in preference to labor.

The four-stage rocket would have a total mass of 48,300 kg

with the first stage having a mass of 33,300 kg and the three upper stag_s

a total _ss of 15,000 kg. The three-stage rat lgun projectile is estimated

tO cost $142 K with a range from $94 K to $194 K at a production rate

of ten per day (see Table 6-73. By scaling the first stage cost from

the 0.6 power law on the ratio of the mass of the three-stage projectile

to the first stage, the first stage ts expected to cost $229 K, wtth

a range from $150 K to $313 K. The entire four-stage vehtcle, at a

production and launch rate of ten per day would then. cost from $244 K

to $507 K with an expected cost of$371K.

To check the valtdtty of these estimates, J. L. Van Cleave,

Deputy Scout Project Manager, NASA/LaRC, was contacted. Based on a brief

look at his information, Van Cleave estimated that the Scout could be

produced for about $500 K at a rate of ten per day, but that the production

philosophy would have to be greatly modified. For example, the Scout

vehicle is assembled and checked out at the factory and again at the

launch st te. To meet the quoted price, labor intensive processes such

as the double checkout would have to be eliminated. Accordingly, the

expected Cost of the larger four-stage supply vehicle is considered to•

be reasonable given that it would be a new design and that production

facilities and equipme, L would be scaled to produce the desired number

of vehicles without labor intensive procedures.

The cost increase per unit due to lower • production rates

of one or two per day is not expected to be much higher (on the order

of SO percent higher). This increase would be principally due to poor

amortization of overhead costs which would not be spread as well as at

higher production rates. The expected cost per unit at a rate of 365

per year is then $557.K, wtth a range from $336 K to $760 K. Launch

services are tncluded in these estimates, but ongoing engineering and.

support would cost about $10 M per year.

The original Scout development cost was about $16 million

In Ig60 dollars or about $56 mllilon in current dollars (_,

Langley Research Center, 1976). The development of the four-stage rocket

is expected to cost about $I00 M because of its larger size and higher

performance, especially in terminal guidance accuracy for retrieval by

Space station personnel. Production facilities to achieve the rates

considered {365 to 3650 per year) are expected to cost on the order of

$50 M. Final assembly and checkout at the launch site, payload mating,

and launch and post-launch control facilities are included in the

development and production estimates which total about $150 M.

_. A single-stage gas gun was briefly examined and

found to b_cally feasible for boosting a 15,000 kg projectile

to 2 km/s within a 2 km-long tube with a cro.ss-sectlonal area of I meter.

Using the ideal gas equation {pV-- nRT - _v,Zl)with temperatures of 1273 K

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and 1773 K, 51,003 kg and 36,619 kg of steam (from H2/O 2 propellants)

wOuld be required to drive the projectile at 100 g's wtth 148 atmospheres

of pressure. The temperatures were selected as the range which could

be telerated by steel alloys after the steam expands adiabatically within

a refractory combustion chamber. To achieve hlgher launch velocities,

however, higher temperatures would be required and these could lead to

materials problems. The major technical challenge for the 2-km/s gas-gun,

however, ltes tn the tnl;roductton of 36 to 50 HT of cryogenic propell_nts

into the combustion chamber tn two seconds.

The gas gun envisioned for the LEO Space Station supply mission

is expected to cost less than the hybrid EHL/rocket launcher because

of its inherent simplicity and relatively low technological uncertainty.

SinCe many of the same facilities would be required, and the propellant

manufacture and storage requirements would be significant, the

development/Investment costs of a gas-gun facility would be greater than

thOSe for the four-stage rocket. Also, the potential for growth of a

single-stage gas gun beyond 2 _/s is limtted by the achievable propellant

temperatures and by the abtlity tO provide materials which can withstand

those temperatures, Accordingly, higher velocities would require

additional Stages and thus would increase the complexity and costs of

the launcher system.

The brief investigation indicates that the gas gun performance

potentia| should be equivalent to the hybrid EML concepts and if a

fixed-azimuth launcher in this velocity regton iS needed, the concept

should be kept in consideration.

S1_!. A liquid-propellant rocket sled was considered

as an alternative tO the hybrid EML concepts. The rocket sled would

be placed on the side of a mountain slope with rails of over 2-km length.

The rocket sled would accelerate a 15,000 kg prOjectile for 2 kJn, release

the projectile, decelePate, and return for reuse. Inttial calculations

assumed the use of H2/O 2 propellants and modified Space Shuttle Main

Engines (SSMEs) with a mass of 6000 kg per engine. More than 10 SSHEs

would be needed to produce the required thrust. A solid-propellant rocket

sled was also considered. Preliminary calculations indicated that a

very rapidly burning solid rocket moto_ (two-second burn time) with a

diameter at least 100 times greater than the length might also be capable

of providing the desired performance. Previous solid motor design

experience has been with length-to-diameter ratios from about 1:1 extending

to long, thin motors, This so]td _ocket motor would have problems

maintaining adequate chamber pressure while permitting a yew high mass •

flow. The fourr-stage solid rocket discussed above was felt to be superior

to the rocket sled because the motors would be more conventional, the

acceleration could_ be lower than 100 g, and a vertical launch would be

preferrable.

_pace Shuttle, The current Space Shuttle has a nomlnal payload

capabtitty of_ 29.51_ ('65,000 pounds) to a 296-km (160 n.mt.) orbit at

28.5 degrees Inclination. Projections of the recurring Shuttle cost

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per launch after Its Inltlal learnlng problems are resolved and higher

launch rates are achieved, fall in the range from $41.7 M to $58 M in

1981 dollars. The low estimate includes the effects of probable

modifications to the Shuttle to reduce operating costs and/or increase

performance. The low estimate was originally provided by Frank Williams

of Martin Marietta Corporation and Mike Van Hook of NASA/MSFC as part

of the work for, but not reported tn, the Phase ii Final Review,.

"Technology Requtremants. Study, Shuttle Derived Vehicles", April 26,

1982. The high estimate tS made-by Battelle; it reflects inflation since

the Shuttle pricing base date is 1975 and does not include additional

investment in modifications to the Shuttle.

The Space Shuttle wtll have a performance penalty to achieve

the higher Space Station orbit (500 km versus 300 km nominal STS orbit),

It would also have another performance penalty, or a new launch site

may be required, to support a space station at an orbttal inclination

other than 28.5 degrees. To account for both performance penalties

associated with flights to a 500-km orbit and ongoing operational

improvement costs, the $58 M recurring cost is used for the Space Shuttle,

and the lift capability of 29.5 KT is reduced 1;o 25 MT. However, to

reflect the likelihood that much of the bulk cargo may be flown on a

space-available basiS, nO additional penalty is assumed for container

weight. This leads to a calculated cost of $2320 per kilogram.

Advallt_ &1hand an n V htcle . There are several

concepts for unmanne Vs) or Unmanned Launch

Vehtcles (ULVs), in addition to several approaches to uprattng the manned

Shuttle fOr transport to LEO. There wtll also be a need for Orbit Transfer

Vehicles (OTVs) with the capabflity of transpOrting a manned module beyond

LEO. The OTV (and very ltkely the ULV) would be required to transport

large equipment and/or personnel for the lunar base program. If EMLs

were deveioped to transport bulk materials (and especially propellants)

a substantial reduction in the number Of ULV and OTV flights wOuld be

possible, but the development of advanced chemical propulsion vehicles

cannot be avoided. Deve:lopment estimates for these vehicles lie in the

$1,0 B to $2.4 B range• and are dependant upon modifying current technology

and/or .hardware (e.g. Shuttle or IUS).. Recurring cost estimates forEarth-to LEO vehicles range from $36 M (Martin Marietta Uprated Shuttle)

to $48 M (Martin Marietta- Shuttle Derived .Vehicle). Since the ULV would

be largely expendable:, this estimate is considered low in cOmparison

with other large expendable vehicles. A reusable 0TV is expected to

have somewhat lower recurring costs ($5 M to $10 M per round trip) but

only when the cost Of transporting its propellants to the space station

is not included. For transport of bulk materials between LEO and a lunar

base or a lunar space station, bOth an Earth-to-LE0 vehicle and an 0TV

will be required. The ULV is believed to have a relatively low development

cost ($1.5 B), but a relatively high recurring_cost ($70M) reflecting

the relatively low reusability of this vehicle. The 68 MT payload yields

a transport cost of $1029 per kilogram.

The 0TV is estimated to have a deveiopment cost of $1.0 B,

and a recur'ring cost of $5 M per round-trip ts considered reasonable

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if the cost of transporting propellants or water (and processing cost)

is not Included, The cost for these propellants largely Consists of

their transport to Orbit by the ULVs, which are costed separately. When

the Cost of propellant transport by ULV (and by OTV to the lunar space

station) ts included, the cost per ktlogram for OTV cargo would be about

$13,200 ($6000/1b).

6.4,.2 rison of Alternatives

Coalparison of the EMLs with the alternatives indicates that

the EMLs could perform these missions with considerable cost advantage.

This advantage also holds under the more stringent criterion of discounting

cost streams at ten percent. Thts comparison technique takes into

consideration the value of monetary tnves_ents over time; itsmajor

effect is to penalize large early investments which, as for the EMLs,

must be made before benefits of using the investment can be realized.

The discounted and undtscounted cost comparisons are shown in Tables

6-11 and 6-12 for the LEO space station supply mission and the lunar

base supply mission, respectively.

The results of this cost assessment of two missions with

their multiple launches per day tend, however, to conceal the fact that

high launch rates are essential to the economic use of the EMLs. If

the launch rate were to fall to one launch per day (365 per year), the

four-stage rocket would produce lower discounted total program costs

than the EML. This is illustrated tn Figure 6-1 wherediscounted program

COSTS are plotted as a function of equivalent 800-kg payloads per day.

This figure illustrates the total program costs for the LEO Space Station

supply mission alternatives over the period 2000 to 2029 and presents

the four-Stage rocket and EML programs with development costs included.

Space Shuttle, ULV, and OTV are shown in terms of total discounted program

cost per kilogram of cargo without development costs. A $500/kg line

is included to show its close approximation to the four-stage rocket

program. A series of possible EML programs is also presented. These

are hypothetical programs with development costs ranging from $1 B to

$5 B. The same projectile unit costs are used for all five programs.

The hybrid EML reference concept falls between the two lower EML program

curves and has lower discounted costs than the four-stage rocket as long

as multiple launches per day are achieved.

The iunar base bulk supply mission has a similar analysis

and results. The major changes for this mission are the use of a $5 B

launcher Complex, use of projectiles with 500-kg payloads (versus 800 kg

for the LEO mission), and the fact that the four-stage rocket cannot

be used to supply the lunar space station. If the four-stage rocket r

were used to supply the LEO portion of this application, the lunar space

station portion of the EML launches would be less economical because

of lower utilization of the investment in EMLs. The cost of transporting

propellants or other bulk cargo to the lunar space station by chemtcel

propulsion techniques (e.g, OTV) is nevertheless sufficiently high (about

$13,EOO/kg), that the EML wouid still be justified for this mission alone.

Other uses would enhance this advantage.

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TABLE 6-12. COST COMPARISON OF ALTERNATIVES FOR 2010-2030 LUNAR

BASE SUPPLY IqlSSION ($, N, 1981, EXPECTED)

All

Chem|cal Propulsion

Category (STS, ULV, OTV)

(Investment Period)

Total Investment

(STS-Patd;/ ULV-II.SB;

OTV-$1.0B; EMI.-$6B)

(Flights: 2010-2030)

Cost Per F11ght

Cost of E119hts

30-Year Total Cost

30-Year Total Cost

Discounted at 10 PePcent

2000-2004

$2500M,

STS ULV OTV

126 546 651

$58N $70M $5M

$48,738M

$51,238M

Chemtcal

Propulsion and EML

(ST$, ULV, OTV, EML)

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$7500M.

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0 126 17,584

-- $5M $0.L70"

$12,145N

$19,645R

$2,635M $1,891M

* Plus $ 58M annual operations.

It should aiso be noted that the two mtsstons discussed are

tndependant and therefore have addtttve launch rates. Thts would help

tn achieving economical use of an EML factltty for LEO and lunar

operations It would be difficult to justtfy an expensive (more than

$2 B) EML fac_11ty tn the face Of competition from the four-stage rocket,

tf fewer than two launches per day were expected.

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7.0 TECHNOLOGYASSESSMENT

The objective of thts task was to assess the status of current

EML technology as it applies to electromagnetic launcher Concepts for

space applications, and to provide supporting research and technology

(SR&T) recommendations to NASA. The tnfomatton developed over the course

of this study has been used to assess the technology. This section of

the ftnal report ts broken down into two major subsections. First, the

technology evaluation subsection, wtil discuss some of the technological

areas that need further work. The second subsection provides a suggested

plan for NASA to follow, tf tt destres to develop EML technology for propulsion applications tn the space:program.

7.1 Technoloo_yEvaluatton

Technology development areas were Identified based upon review

of the open literature, discussions wtth EML experts, and attendance at

several EML conferences to assess the state-of-the-art technology. As

a result of the technology evaluation activity, recommendations were made

regarding certain areas whtch requtre NASA supporting research and

technology-SR&T (see Secti.on 7.2). The SR&T recommendations were made

tn those areas that NASA must resolve before an EML Earth-orbit supply

launcher development program couid be implemented. It is expected that

the DoD will, over the next several years, develop many areas of technol.ogy

needed for EHL Earth-orbit supply systems.

The prtmary areas requiring technology development ,.,llch were

identified in this study are listed below:

e Scale-up of existing EML systems

e Energy distribution and switching

• Projectiles

• Energy storage systems

e Projectile brushes and amatures

• HPG brush materiai

e Launcher tube bore

• Preboost systems

• Launcher structural support.

These areas are discussed in this section.

7.1,1 Scale-uE

e

Scale-up of extsttng SyStems is _equlred for Space missions.

Existing EML systems are under 10 m in length (the Los AlamOs HYVAX Pailgun

will be 13-m long to accelerate projectiles at 15 km/s); to be effective

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space launchers, 2000-m long EML systems must be operational, Projectiles

must also be scaled. The reference Concept projectiles range fl'om 3000

to 15,000 kg and must be launched between 2.and 12 km/s, Ratlguns have

reached 10 km/s velocity, but for gram, size prOjectiles. Coaxtal

accelerators have launched 5 kg projectiles, but at launch velocities

on the order of 0.1 km/s. Scalerup of current systems must be demonstrated

before beginning.development of EML systems.

7,1.2 Enero.y_Otstrtbutton and Switching

Demonstrattonsof distributed energy storage (DES) ratlgun systems

have been progressing over the last few years. A DES demonstration program

was begun at the University of Texas at Austin Center for Electromechanics

(UT-CEM) tn 1980. In July 1983, UT-CEM fired a ten-stage 4-m railgun

accelerating the projectile to 3 _/s. The investigation goal of UT-CEM

iS to accelerate I to 3-g projectiles to I0 km/s wlth this railgun (Holland,

1984). Vought, in connection with UT-CEM, is building a 5 stage 3,65-m

railgun with. a design capability of propelling a BOg projectile to

3-4 km/s.

By nature, coaxial accelerators are distributed systems. The

dM/dz parameter in the thrust equation is not constant, so the current.

must be synchronized with the projectile motion (see Section 3.1.2 and

Appendix D).

For a11 distributed energy store systems, switching of current

at the right time is crucial. For railgun systems, current must be switched

when the projectile reaches a new segment. The coaxial EML reference

concepts continually switch drive coil turns in and out .so that ten turns

remain active behind each projectile coil.

Switching is considered a critical technology area, until

switching issues are resolved EML system development will not take place,

Critical issues are switch performance and life times for an EML system

which launches an average of ten times per day over a 30-year system life.

The switches must perform over many launch cycles without being replaced,

Frequent replacement would be prohibitively costly in terms of time and

economlcS.

7.1.3 Pro_[ecttie Developmnt

Critical issues to the design of an EML-launched projectile

are: high acceleration constraints, aerodynamic considerations, and

materials selection. All of the above are critical to EML system

development and all require in-depth study,

The high accelerations of launch (100-3600 g's) produce unique

requirements for NASA programs. Acceleration effects on propulsion systems (sensitive liquid propellant lines and tankage, and solid propellant grain)

need to be.investlgated. A finite-element analysls is required to establish.

confidence in conceptuai deslgns, particularly in projectiie structure.

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Instrumentation and stabilization pop-out mechanisms (fins) must also

withstand the acceleration loads. Existing DoD expertise in the area

of high-acceleration tube-:aunChed artillery and misstles should be able

to contribute significantly to this investigation,

Aerodynamic drag considerations are important from the standpoint

of the velocity required to launch a projectile thrOugh the atmosphere

to obtain the desired end condition. The higher the drag, the greater

is the loss of velocity and energy along the flight trajectory. The higher

the drag, the greater will be the sonic boom generated by the projectile

as it transverses the atmosphere. Daniel and Mtlton (1980) indicated

that low-drag bodies are possible; experimental research is required tO

verify this.

Aerodynamic stability as the projectile leaves the muzzle of

the launcher tube is another crttical issue. Preliminary assessment

conducted during this study indicates that stability of the vehicle is

critical to the performance of the System, Initial concepts for

establishing flight stability include the use of fins at the rear of the

projectile and the center of mass being no_e forward. (If the pitching

rates can be made fairly low, the vehicle will not have a chance to pitch

very far during the few seconds that it flies through the atmosphere.

Pitching moments of the order of perhaps 10 degrees per second would still

allow the vehicle to fly out of the atmosphere without any problem.) A

very important consideration in the launch of the Earth-to-orbit railgun

projectile is the jettison of the sabot, as the projectile leaves the

muzzle Of the rail launcher 'ube. The sabot would have to break free

in a very timely way so that a pitching moment is not imparted to the

vehicle. Pop-out fins are used in the round-bore systems of the other

references concepts. If stability should prove to be a problem, rOund

bores could be used for all concepts and the projectiles spun-up prior

to launch. Theoretical and experimental work is required in this area.

Aerodynamic heating is also a very critical aspect of the EML

concepts, especially the Earth-to-orbit concepts. Initial assessment

indicates that because the projectile flies rapidly through the atmosphere

(7 km/s), there is little time for aerodynamic heating tO melt the nose

cone to any significant degree, The heating rates are very high, and

it is expected that a fairly s_gnificant area at the stagnation point

would be melted away, depending upOn the latent heat of fusion and the

melting temperature of the nose tip material. Steel was used for the

nose cone in the Earth-to-orbit reference concepts. Experimentation is

needed to determine appropriate materials for use.

7.1.4 Ene Stora • S tems

Energy storage technology is critical especially_for near-term

concepts. Homopolar generators (HPGs) are commercially available . at

6.25 MJ; UT-CEM has a 10 MJ HPG. In the next five to ten years, an

order-of-magnitude more storage capacity for HPGs should be expected

(telephone conversation with Dale Pryor, OIME), Othe_ railgun energy

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storage devices _nlch may merit further investigation are batteries and

the inverse rallgue (Marshall, 1984b).

Various storage inductors are available for the coaxial

accelerators; a Brooks coll was selected. The slngle Inductor is simple,

but does represent a possible single-point failure area• The multiple

HPGs on the railgun systems are more complex and massive, but offer some

redundancy.

When a detailed system design is conducted, energy storage devices

will need to be considered• carefully before selection is made. This is

not a critical area of system development, however.

7.1.5 projectile Brushes and Armatures

Projectile brush and armature selection are dependent upon launch

velocity• Several coaxial accelerator concepts which were considered

used projectile brushes to pick up excitation currents from the drive

coils. These concepts were not selected for reference concepts, because

of uncertainties of brush survtval at velocities above 1-2 km/s. Westinghouse is conducting advanced brush work• At 1 km/s, brushes survived

with no rail erosion. William Snow (EML Research) expects no problems,

even at 7 km/s. If. experimental evidence shows that brush deterioration

is not a problem during launch, these launcher concepts may be revisited

for consideration t n a system design.

Railgun armature selection depends upon velocity of launch. Below

2-3 km/s, a solid metal armature may be used (similar argument as for

projectile brushes). Copper and aluminum are the most commonly used

armature materials. Aluminum armatures have been used at velocities of

up to 3. km/s. Above these velocities, plasma armatures are needed; but

at the initial low velocities, erosion of the railS may occur unless

preboosting of the projectile is done. Solid/plasma armatures have been

suggested as the optimum solution. Extensive experimental research is

needed before this "armature is selected for system design.

7.1.6 HP6 Brush Matert als

The ability for HPGs to store energy depends directly upon the

speed at which the homopolars can operate. Critical to HPG reuse economics.

are the brush materials that transfer; the charge. Discussions with William

Weldon at the University of Texas Center for Electromechanics indicate.

that HPG speeds are currently limited to approximately 220 m/s with long

life at the brush interface, The use of advanced brush materials would

allow increased speeds which would in turn allow more energy to be stored.

The University of Texas Center for Electromechanics is investigating HPG

brush technology and has improved the brush materials to the point where

they can obtain speeds of 300 m/s using advanced materials, but still

experience a great deal of erosion.

A major issue in the EML systems would be the required replacement

rate of brush material. It is desirable to operate the HPGs at high speeds

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and with mtnimal maintenance over long pertods of time. The ease of brush

replacement tS also a major issue. Currently, the commercially-available

OiME HPGmust be disassembled to replace the brushes (teiephone conversation

with Dale Pryor, OIME). When 750 to 3600 HPGs are used, disassembl_ would

reqUire an excessive amount of time and would ltkely limtt the launch

rate severely. These issues are critical to the operational cost of the

system. Therefore, there Is a need to advance the state of technology

tn the area of brush materials for HPGs,

7.1.7 Launcher Tube Bore

The bore deSign and especially bore/projectile interfaces are

key factors in reusability of EML tubes for bnth ratlguns and coaxial

accelerations. The selection of bore shape, materials, and support

structures is important for launcher tube longevity.

Round-bore railguns have been experimented with more since the

previous ESRL study. Lawrence Liven.ore National Laboratory and Log Alamos

Scientific Laboratory are the primary investigators. A round bore offers

several advantages over a square bore. A round bore allows the capability

for spin-stabilizing of the projectile. Pop-out fins can be used, reducing

the bore diameter, and thus the launcher system mass. A round,bore railgun

alSo allows for the poSsible remachintng of the bore which would permit

a long operationa! life for the system.

The proper selection of materials for tube structural support

is crtttcal to maintaining launcher alignment and reducing bore damage

due to launch bursting forces.

Another critical technology area that needs to be investigated

prior to development of an EML system, is the evaluation of: (1)

projectile/bore friction, (2) sabot/projectile-friction for rallgun systems

during the launch phase, and (3) the tolerances that are required to

avoid projectile destruction during the launch phase. Aspects of friction

Should be evaluated for both square and round bore-launchers. Analytical

and experimental work should be conducted tO establish the significance

of friction and critical dimensions of the bore and the projectile at

the time of launch. Experiments could be conducted in existing facilities.

The problem of launcher tube movement as a result of continued firings

is an important aspect related to the bore tolerances.. This aspect

determines the reusability of the launcher bore after numerous firings.

7.1.8 Preboost

ReuSability of the rails is a major issue for plasma-armature

ratlguns. The long dwell time of the plasma during initial accelerations

tends to erode the railS. For a large-scale railgun system to be developed,.

rail damage must be prevented. Pre-accelerating the projectile .to 1 km/s

or so greatly reduces the damage.

Preboost systems were used in ali four reference concepts in

this study, even though only one was a plasz_a-armature ratlgun. Preboost

iAl'Tml, t.m -- COLUMNUg
7-6

systems were assumed to provtde-a small velocity (50 m/s or so) to initiate

projectile motion to overcome the downward forces due to the tncline of

the lO-degree and 30-degree launcher elevation angles.

Helium gas injectors are commonly _uSed to preaccelera_e

projectiles before entering the ratlgun section. Typically the projectile

is accelerated to velocities of 0.5 to 1.0 km/s using high-pressure helium

(3000 to 5000 psi). Various preboost concepts would need to be tested

before Selecting an appropriate method. Preboost systems are not considered

to be critical to the development of an, EML system in that the basic

technology is avai]able.

7.1.9 EIILS__VStem Structural So_

Large support structures are required for the EHL systems for

space missions _or several reasons. The launcher tubes are very long

(,the reference concepts would be 2040 m long) and correct alignment must

be maintained, The bursting force on the rails or drive coils during

launch must be contained. The structure for the coaxial accelerators

to contain tube stresses cou]d be somewhat lower in mass than, for railguns,

because the projectile coil presence .decreases the forces.. Launcher

structural support technology is not expected to be critical to the

deveiopment of EML space-mission, concepts.

7.2 Supporttnq Research and _Technolog), (SR&T) Recommendations

Based upon the results of this study, supporting research and

technology (SR&T) efforts have been prioritized and funding estimates

have been made, Table 7-1 provides our estimates.of 5-year funding requirements for. our.recomendations in 1984dollars.

Four major- areas of activity have, been categorized: (I) EML

experimental research; (2) EML mission requirements studies; (3) EML

systems studies; and (4) special studies. The phllosophy in developing

the schedule for EML SR&T was based" upon the fact that DOD agencies are

funding EML technology development at an ever increasing level and that

NASA should only complement this effort for the space transport mission.

Subsections below discuss the suggested areas of NASA-funded SR&T.

7.2.1 _ Experiment#,.] Research

To keep abreast of the latest technological developments in

the EML area, Battelle recommends that NASA/LeRC continue a low-level

experimental work profile with the equipment in hand and concentrate the

effort on hypervelocity impact testing of Space Station power system

components (space debris and meteoroids) and continue to investigate

switching technology for distributed energy railgun systems and coaxial

systems. Funding levels are provided in Table 7-I. Funding for Item A

increases in the latter years because of testing with Space Station power

system prototypes,

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7"7

ESTINATED FUNDINS REQUIRENENTS(1984, KS) FOR,

RE_NDED _ SR&T ACTIVITIES

ActiVity FY-85 FY-86, FY-87 FY-88, FY-89 Total

EML Ex erimental Research

A. EML Demonstration and

Hypervelocity Impact

Testing 150 150 200 250

B. Energy Distribution and

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C. Lunar Base Supply

D. Space Station Interfaces

EML Systems Studies

E.

F.

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P_ojectile/Sabot Design

Propulsion Systems and

instrumentation Design

Total Launcher Systems

Design Study

Special Studies

I. Environmental Impact

Assessment

(200) (200) (200) (250) (300) (zzso)

200 100....... 300

75 75 ...... 150"

(275) (175) (--) (--) (--) (450)

-- - 75 ...... 75

.... 150 t50 -- 300

.... 150 150 -- 300

...... 100 500 600

(--) (75) (300) (400) (500) (1275)

150 150

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Total SR&T

50 50

...L_.o).

525 500

50 50 50 250

(SO). (200) (400)

550 .700 L000 3275

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7.2.2 E]4LXtsston Requirelents Studtes

Two study areas need further development to provide better insight

and benefit quantification for EML missions. Our study showed that the

'far-term application of EMLs_stems for large bulk-mass transport showed

promise for two NASA programs: the Space Station and thelunar+base.. These

studies (discussed below) should be done soon to provide guidance to the

rest o_ the recommended SR&T activity.

A consideration of EML support of Space Station has not been

part of the +initial Space Station pre-program and program planning (and

should not be), EML support may introduce new and/o_ addltional requirements to a growth Space Station facility. Trades such as projectile

retrieval using on-board automated systems or OMV-type systems should

be considered. How many OMVs are needed and how are storageand handling

of both the payloads and projectiles accomplished) For example, the Shuttle

could return spent projectiles to the surface. The study of these and+

other related issues provide a better understanding of costs, burdens,

and benefits of using an EML system,

Key issues that may affect the benefits of EML Supply of a lunar

base include the lunar base Characterization, transportation alternatives,

and on-orbit support capability. Some important considerations of the

lunar base characterization include the lunar-produced oxygen benefits,

crew slze and rotation, cryogenic storage at the lunar poles, and lunar

resource/product return (possibly using a lunar-based EML). Comparison

of the EHL supply to alternative, more-conventional vehicles is key to

deriving the benefits of EML. Transportation trade studies would also

include various types, sizes, and performances of Earth-based, space-based,

and lunar-based transport vehicles used to supplement the EML with manned

transportation. For example, the _ize of an OTV-type vehicle and the

types of propellants are critical considerations. The studies of on-orbit

support alternatives would address effects of storage, retrieval, and

refurbishment capabilities for projectiles, vehicles, propellants, and

payloads. The range of effects of any issue may have'a large impact on

cost benefit_.

7.2.3 EML S]st_s Studies

After the refinement of missions requirements (see Section 7.2.2)

for promising EML systems applications, various EML studies need to be

accomplished. Because of the need for a preboost system on railgun EMLs

and the potential of this system being a replacement for the EML hybrid.

launcher, the first study would be to investigate and analyze preboost

systems, in particular, large light-gas gun systems. Second, projectile,

sabot, propulsion systems,, and instrument designs need to be designed

in detail based upon Items C.and D (see Table 7-I). Also, trade studies

need to be accomplished on solid or liquid propulsion systemS. These

studies would also investigate projectile aerodynamics, materials, and

structural problemS. The goal would be to arrive at designs for both

a railgun and coaxiai launcher system for Earth orbit and lunar missions.

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Total fundtng for these is estimated at $600 K tn FY-87 and -88 (see

Table 7-1). Once these studtes have been completed, a total launcher

systems design study would be needed to detemtne the benefits of both

ratlgun and coaxtal systems versus conventtonil methods of space transport,

In FY-88, a preliminary study would be conducted; a larger study $500 K

would be acCompliShed tn FY-89.

7.2.4 S_pecial Studies

Special studies are defined to include, unassigned studies and.

an environmental impact assessment tO be accomplished in paralle1 effort

with Item H.

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In addition to ral_guns, which were tnve3tigated as an

" t s

alternative to chemica]1y-propelled launcher systems in Battelle

Earth-to-space roll launcher (ESRL) study (Rice, M111er, and Earhart,

1982), a number of othe_ electromagnetic and electrothermal propulsion

systems have been proposed. These c_ncepts are reviewed briefly in this

section and their suitability fOr the space missions considered in this

report are assessed. Other' concepts such as magnetoplasmadynamic and

free-radical thrusters were not reviewed as they are considered to be

primarily low-thrust techniques, at least for the foreseeable future.

Those concepts that • are reviewed include coaxial accelerators,

electrothermal ramjets., electromagnetic rocket g_n, and electromagnetic

there gun. Soviet work in electromagnetic acceleratlon i.salso .reviewed.

8.1 C_uaxial Accele_atorConcepts

Two additional coaxial accelerator concepts were reviewed and

were not selected for inclusion in the reference concepts. These were

the Frequency Controlled Coil Driver., designed by O. K. Mawardi of

Collaborative Planners, Inc., and the Solenoid Earth Launcher, designed

by William Snow of the U.S. Army Armament Research_and Development Center•

and currently with EML Research, Inc. Summaries of these, concepts are

given in this sectio,.

8.1.1 Frequenc¥Con&ro|led Coil Driver

NASA/LeRC contracted with Collaborative Planners, Inc., and

EML Research, Inc., to supply coaxial accelerator concepts as input to

this study. Appendix E Contains the report of Collaborative Planners'

effort which .is brlefly summarized here.. Twn. launcher • concepts were

designed by O. K. Mawardi, a hybrid EML/rocket concept and an

Earth-to-orbit EML concept. Because of the limited payload mass and

the technical uncertainty of superconducting the projectile coils,_ these.

concepts were not selected as reference, concepts for further Study,

Figure 8-I presents a cut-away view o_ Mawardi's launcher

concept. The launcher lenath was calculated at 2.0 km; this corresponds

to 8888 drive coils for the hybrid mission. The projectile coils would

be made of aluminum and would have a mass of 406 kg. From the results

shown in Appendix E for the Earth-tO-orbit cOncept, mass summaries were

derived and are shown in Table 8-I. These numbers indicate that because.

of the large size of the projectile coils, the available mass is limited.

The available mass must include the payload, orbit,circularizatlon

propulsion system, and structural mass. The velocity for the projectiles

is. also insufficient to reach orbital altitudes.

For the hybrid EML/rocket concept, ten projectile coils were

assumed, To orbit an 800-kg payload (payload •requirement, see Table 3-I),

approximately 12,700 kgOf solid prOpellant_is required. After subtracting

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the projectile cot.1 mass (4060 kg), the remaining projectile mass would

be 10,940 kg. To use this concept, either an optimization study must

be done to increase the size of the three-stage rocket or the lower payload

mass must be. accepted.,

TABLE 8,-1. Kq3S AND VELOCITY SUiUIIARY FOR KqMARDI'S

EARTI_TG-ORBIT PROJECTILE

ProjeCtile Projectile Available

_ss (kg) Cotl Mass (kg) Mass (kg)

Final

Veloctty (km/s)

650 406 244 4.8

1000 816 184 3.3

6500 40.60 2440 2.9

The concepts use brushes and guide rails to feed excitation current to

the projectile coils. Because "it is essential to maintain current density

in the brushes at a low safe value" to guarantee mechanical stability,

the concepts are velocity ltmited. To launch at velocities above 5 km/s,

Mawardi suggested that the projectile coils be cooled to superconducting

temperatures with the current induced in the cotls before launch; this

eliminates the need for brushes entirely. However, superconducting

projectile coils, present new problems. For example, if the superconducting

coils Should become normal befOre leavtng the launcher tube, damage to

the projectile and to the accelerator could occur..Superconduction of

projectile coils requires further analysis.

8.1,2 Solenoid Earth Launcher

Another coaxial accelerator concept which was brtefly reviewed.

is the Solenoid Earth Launcher, designed by Wtlltam Snow of the U.S. Amy

Armament Research and Deveiopment Cente¢, now with EML Research, Inc.

The review was based upon notes and viewgraphs (Snow, 1984) supplied

late in the study effort and telephone conversations with Snow in April

1984_ The purpose of this launche_ would be to launch projectiles to

Earth-orbital altitudes at velocities of 6 to 7 km/s. The launcher would

consist of a stngle helix drive coil built in segments and a projectile

with a single pr.ojectile coil and current pick-up brushes.

The proposed accelerator would consist of 1000 2-m drive cotl

segments, with each segment containing 60 cotls. Approximately 2600 HT

of copper would be required to butld the launcher tube with a tube

thickness of 5 cm. Two homOpolar generators connected in series would

suppiy energy to each segment at 70 MJ pep segment (35 MJ peP HPG);

therefore, 2000 HPGs would be required. Since these wOuid be self-excited

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HPGs, Snow claims that there would be no need for intermediate storage

Solenoids. (telephone conversation with Wtlliam Snow, Aprtl 1984).

Triggered vacuum gap switches would be timed to open and close to activate

the drive coils as a function of projectile position. A maximum of 60,000

switches would be required (one for each drive coil). Figure 8-2 presents

the Solenoid Earth Launcher concept.

Snow investigated three projectiie masses: 2000 kg; 3000 kg,

and 4000 kg. Each projectile consisted of a payload, propulsion system

for Orbit insertion, structure, and one projectile cotl. The projectile

cotl for each of the three projectles studied was assumed to be made

of aluminum wtth a mass of 273 kg. Current would be fed to the projectile

by means of brushes which would pick up the current from the drive coil.

The 3000-kg and 4000-kg projectlles were rejected by Snow because

of heating and stress during launch. The projectiles were assumed to

be precooled to 150 K, while the solenoid would be water-Cooled during

launch. Snow's simulatlons indicated that the 2000-kg projectile (with

a sln_le 273-kg projectile co11) would experience a 180 K temperature

rise (to 330 K) and a 100,O00-psi stress of which 40,000 psi would be

solenoid compensating so that the back of the projectile ceil would see

60,000 psi. The 4000-kg projectile with the same coii mass would

experience a 500 Ktemperature rise and a stress of 200,000 psi.

This concept does not use one inherent advantage that the

reference concept coaxial accelerators have over the railgun. The use

of multiple projectile coils provides lower projectile launch stresses

by distributing the stress throughout the body, rather than having the

stress applied at the back of the projectile as in the case of a railgun

launch. While one projectile col1 in the center of the projectile (like

Snow's concept) will reduce stresses somewhat over a similarly-sized

railgun projectile; to get a useful payload for space delivery, the

structural mass needS to be at a minimum. Multiple projectile coils

h_ip to increase the payload-to-projectile mass ratio by lowering

projectile launch stresses which decreases the amount Of structural mass

required; however, this occurs at the expense of added system Complexity.

Table 8-2 lists the payload versus structural mass for Snow's 2000-kg

projecti]e assuming a similar propulsion system to that used in the

reference Concepts (payload mass to propulslon mass ratiO of 0.28, see

Section 3.8). For comparison, the Earth-tO-Orbit railgun projectile

concept contained 2730 kg of structure and the Earth.to-orbit coaxial

projectile (with, 40 projectile coils) contained 300 kg of structure.

Another concern is the use o£ projectile brushes at the required

7 km/s launch velocity. Sliding brushes have been used up to I km/s

with no damage (see Section 7,1.5). Snow believes that the brushes could

take velocities of 10 km/s or more with out damage. Since the brushes

would be attached to the projectile, if they could Survive long enough.

to feed current through the launch phase, it would be sufficient as reuse

would not be required.

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TABLE 8-2. PAYLOAO MASSES AS A FUNCTION OF STRUCTURAL MASS

FOR 2000-1{6 PROJECTILE

Projectile Propulsion

Total Coil System Structural Payload

Mass, kg Mass, kg Mass, kg Mass, kg Mass, kg

2000 273 1238 0 489

2000 273 1095 200 432

2000 273 951 400 376

2000 273 808 600 319

2000 -- 273 665 800 262

2000 273 521 1000 206

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If a detailed system design for an Earth-to-orbit EML system

is initiated, the Solenoid Earth Launcher should be considered along

with those of the reference concepts (Section 4o0), especially if further

research and experimentation yields promisin_ results for use of brushes

at high launch velocities.

8.2 Electrotl_rmal Thrusters

In additiOn to the electromagnetic acceleration concepts covered

later in this report, electrothermal means of producing accelerations

were also investigated. The work summarized in this section was done

by Wilbur, Mitchell, and Shaw of Colorado State University under NASA

Grant NGR-06-O02-112 (Wilbur, 1983).

8.2.1 Electrothermmal Rocket

Figure 8-3 shows a simple schematic of an electrothermal rocket.

concept. In this concept, an on-board propellant with a high Isp (such

as hydrogen) is heated electrically, then exhausted through a nozzle

in order to produce thrust, Heating of the propellant could be

accomplished by several different processes. In one concept, the

reststojet, propellant would be heated by being passed over

current-carrying resistive elements. Exhaust velocities of up to 8,500 m/s

have been reported using this method, which is limited by material

properties of the heating elements. In a second concept, heating of

the propellant is accomplished by passing electrical current directly

through the propellant in the heating chamber. In this concept, labeled.

an arcjet thruster, exhaust velocities of up to 15 km/s have been claimed.

In both the reslstojet and the. arcjet rockets, propellant is carried

aboard the vehicle and is metered into the heating chamber.

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A COncept in which no propeliant is carried on the vehicle --

is shown schematically in Figure _4. This concept, an electrothermal

ramjet, requires that the prope11_nt be distributed within a sealed launch

tube. The body of the ramjet could be fabricated from the payload material

reinforced, as necessary, by other structural materials. The ramjet

engtne, moving through the dispersed propellant, would gather the

oropellant in through the inlet diffuser, heat it electrically, and provide

thrust by expelling tt through the nozzle. Energy could be conducted

into the heating chamber either electromagnettcally or through pulsed

electHcal discharges occurring sequentially throughout the length of

the launcher tube wall. A frangible cover would be needed to contain

propellant withtn the tube and pemit the ramjet to escape the tube with

minimum loss of energy. Other aspects of ramjet operation are discussed

following presentation: of an annular flow ramjet configuration that might

present fewer problems in fabrication and operation than the configuration

shown in Figure 8-4.

8.2.3 Annula_lrlo_ Electrothermal Ran et

If the Concept depicted in Figure 8-4 is modified so that the

heating chamber is a space between a solid payload package and the launch

tube wall, a configuration similar to that shown in Figure 8-5 results.

This concept is similar to that of Figure 8-4 in that the launch tube

is sealed and is filled with propellant. However, in the annular flow

concept, the heat-addition regiOn iS not inside the payload; it is designed

to be located between the moving payload and the stationary launch tube

wall. One advantage this configuration presents is that there is no

need for a "window" in the payload that woulcl be required .to permit .passage

of energy into the heating chamber as is required in the. case of the

concept Shown in Figure 8-4. However, either of the concepts of Figure 8-4

and Figure 8-5 would need a first stage to accelerate a payload to the

velocity needed for commencement of efficient ramjet operation. This

acceleration could be prOduced by an electromagnetic driver (railgun

or coaxial accelerator), a booster rocket, or a light-gas gun. This

need for a hybrid launch mechanism using two technologies creates the

potential for more problems than would use of a single technology for

the iauncher. Some of these possible problems are discussed later in

this section.

8.2,4 Launcher Re utrments

Wilbur (1983) presented the results for analysis of an

Earth-escape mission for iO-kg payloads. Requirements were to provide

sufficient thrust to constantly accelerate a payload Of i0 kg at 30,000 g

to a final velocity of 15,000 m/s. The acceleration and final velocity

requirements yield a launch tube length of about 400 meters and a launch

time of approximately 50 msec. The payload was assumed to have a shape

similar to the payload shown in Figure 8-5 with maximum diameter being

16 cm at both diffuser and nozzle throats and minimum diameter at the

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INTAKE

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INTAKE __.

DIFFUSER_ PROPELLANT

PAYLOAD _

PROPELLANT

[LECTROTHERMAL_

HEATING

REGION

NOZZLE --_

LAUNCH I1_

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center of the payload being 8 cm to provide a heating chamber between,

the center portion of the payload and the launcher tube wall. The diffuser

cone half-angle was chosen to be about 22.5 degrees to permit exhaust

velocities to be as high as possible. The chosen propellant was hydrogen

and was assumed to be in the launch tube initially at 300 degrees K and

at a pressure of 30 atmospheres. At the end of the launch sequence,

the exhaust temperature was calculated to be approxie_tely 5,000 degrees K

and the exhaust pressure, about 400 atmospheres.

The la,nch tube iS required to be tapered as a function of

the payload speed in order to. permit efficient functioning of the ramjet

along the length of the tube. The tapered profile of the launch tube

needed, for the hypothetical application is shown in Figure 8-6. This

taper is needed so that a Math number of H=l at the nozzle throat can

be maintained throughout the launch; this, in turn, leads to most efficient

operation of the ramjet portion of the launcher. This requirement is

based upon both thermodynamic and aerodynamic considerations.

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In addition to its tapered profile, the launch tube would have

to be designed (1) to withstand the temperatures andpressures associated

with the launch, (2) to have a htgh duty cycle with a correspondingly

ShOrt turn-around time, and (3) to provide for transferring, external

energ_ to the internal, moving electrothermal heating region.

8.2.5 Erie TPan.sfe._._._

Figure 8-7 displays an instantaneous power proftle of the themal

energy that must be added to, the heat addition _egton of the electrothen, al

ramjet in order to Sustain the destgfi acceleration of 30,000 g until

the speed of 15 km/S is reached at the end of the launcher tube, The

total thermal, energy input, found by integrating under the power curve,

is about 2.4 GJ. Using informatlon, provided in Tables 3-3 and 3-4 of

Rice, et a1., lg82, and scaling to the same final veloctty and p_yload

size, gives about 1.6 GJ as the energy needed to perform an equivalent

mission with a hypothetical railgun launcher.

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A large amount of the difference between the .energy requirements

for the two concepts is the relattve inefficiency of the electrothemal

ramjet tn converting themal energy in the heat chamber to kinetic energy

of the payload projectile. An instantaneous efficiency curve for the

electrothemal ramjet is presented as Figure 8-8.. Energy conversion

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effictenctes of up to 85 percent have been postulated for railguns (Rtce,

et al,. 1982). The electrothermal ramjet w0uld, requtre additional energy

during the pre-boost phase of launch, maktng it even leSs attractive

than a ratlgun alternative.

The preferred method, of Wilbur (1983) for transferring energy

into, the thermal heating region is by means of lasers arranged outside

the launcher tube wall projecting energy through windows in the wall

in order to provlde heating of the hydrogen propellant. When both quantum

and thermodynamic efflciencies of the lasers are taken Into consideration _

it is obvious that iarge penaltleS In terms of non-productlve energy

use wou]d accrue to the iaser-drlven, electrothermal ramjet concept. In

addition, electrical energy generation and storage capabilities similar

to those described in Rice, et al, 1982, for a ral]gun launcher would

be needed to drive the lasers.

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An alternative method of supplying heating energy to the moving

thermat heating region would be by means of a sequentially-fired series

of electrical discharge mechanisms built into the Inside of the launcher

tube. This concept would have the advantage of transferring energy

directly to the heating chamber without the intermediate (and inefficient)

conversion to optical energy in the lasers. Wilbur (1983; did not present

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any design or efficiency data on 1his proposed Concept, but it would

appear to have both inittal and conttr_uing COSt advantages over the last

concept unless it contains a fatal flaw that would disqualify it from

consideration. This concept still has the major disadvantage that massive

amounts of electrical energy would have to be stored, then transferred

to the electrothermal heating region extremely rapidly and in a precision,

rapid sequence. These requirements result in very high capttal costs

that have been discussed at some length for rail launcher concepts in

Rice, et al (1982). However, there appears to be another concept, not

mentioned in Wilbur (1983), that could be more attractive than either

scheme discussed to this point.

Consider again the general annular ramjet configuration shown

in Figure 8-5, However, instead of a pure hydrogen environment, assume.

that the tube is. filled with a mixture of 90 to 94 percent, hydrogen and

6 tO 10 percent oxygen,, a mixture that would sustain combustion in the

presence of- an igniter, but One-that would not be explosive. Low*energy

spark generators could be Implanted°In the inner walls .of the. launch

tube and fired in sequence tO maintain combustion in the moving heating

region. The major advantage to this concept, which more closely resembles

a pure ramjet, is that the high capital costs for electrical energy

generation, storage, and conversion could be bypassed. Feasibility of

this concept was not pursued because of project time and monetary

constraints. Even if theoretical feasibility could be established, this

concept shares a major problem with any other ramjet scheme; that is,

the ramjet must be accelerated b_ other means to efficient operating

velocities..

8,2.6 Sumar:f

A ramjet starts to become an efficient mechanism only at.

velocities considered to be extremely high under most ordinary conditions,

For example, in the electrothermal ramjet configuration shown in Figure 8-5

and discussed in Wilbur (1983), the authors assumed the need of a pre-boost

phase to a speed of 5,000 m/s before ramjet action would be initiated,

Even if this value is quite conservative, an initial stage is needed

to produce some high value of velocity before efficient ramjet action

can occur. This, in effect, leads to a hybr.id launcher.. The first stage

could be a railgun or. a light-gas gun. In either case, addtional

complexity is introduced and stage interface pro.blems would be encountered.

Additional areas that would have to be investigated in order

to determine .both feasibility and competitiveness of any ramjet concept

include:

• Payload size limitations

• Repetition rate of launch

• Life of launcher tube

• Payload ablation both in the tube and in the atmosphere

:4W_

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8-14 I(

• Tube recondtttoni.g requirements between launches

• Structural and thermal compatibility between tube structural

matertal and, window material in those concepts including

laser-drive heating.

8.3 Electrom_gnetlc Rocket Gun

Dr. F. Winterberg has rocently proposed and theoretically

analyzed a concept for a high-velocity accelerator which he calls an

electromagnetic rocket gun {Wlnterberg, 1983). This concept Is presented

schematically in Figure 8-9. The launch tube walls are made up of field

coils that are sequenced to make up a traveling magnetic wave. The

projectile consists of two partS; a payload section in front, and a

conducting, cylindrical, metallic section in back. The hollow portion

of the cylinder is filled with propellant (in Winterberg, 1983, the

preferred propellant is solid hydrogen).

The magnetic field is controlled so that it has a variable

magnitude with respect to the projectile as Shown in Figure 8-I0. The

leading edge of the field includes currents in the thin walls of the

back portion of the projectile. The heat generated by these currents

vaporizes the propellant which then encounters a region in which the

magnetic field is rapidly rising. The now gaseous propellant is heated,

ionized, and compressed by the rising force of the magnetic field and

transformed into a plasma. After reaching the point of maximum field

strength, the hydrogen plasma iS ejected from the magnetic field as a

high-speed exhaust jet, imparting an accelerating force to the projectile.

In effect, the magnetic field establishes a "nozzle" through which the

ionized propellant moves _o produce thrust.

Advantages claimed for the electromagnetic rocket gun are as

follows:

• The exhaust jet can be many times as long as the projectile,

permitting switching of the current necessary to establish

and move the magnetic field to be done at a lower rate than

is necessary in alternative electromagnetic launchers such

as synchronous accelerators.

• The magnetic field can be utilized to center the projectile

in the launch tube. This eliminates contact between the

projectile and the side wall which has been a source of

problems in other launchers such as railguns at higher

velocities.

• Superconducting or ferromagnetic materia]s are not needed..

A large amount-of effort would be required to determine whether

the electromagnetic rocket could be effectively used for the applications

considered in this study. AS far as is known, no experimental work has

been performed to validate the concept. Since there are several unique

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FIGURE 8-9. ELECTRORA6NETICROCKET.GUN

FIELD STRENGTH

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features included in this concept (solid hydrogen aS a propellant,

electromagnetic nozzle) and since the author is COncerned primarily with

small size, ultra-high veloctty (100 km/s) applications, consideration

of the electromagnetic rocket gun was not pursued further.

8.4 E|ect,romqnettc Theta Sun

The electromagnetic theta gun is a concept developed at Sandta

National Laboratories. primarily for military applications (Burgess.

1980). The basic idea of this launcher is shown in Figure 8-11. The

projectile is tubular in shape and is made of a hlghly conducting material,

Driver coils are embedded in the cylindrical wall of the launcher and

are energized sequentially as the projectile proceeds along, the length

of the launcher.. Linear motion along the launcher tube is produced in

a similar manner as rotational motion is produced in an,induction motor.

The magnetic field produced by the drive coils induces a current in the

closed, Conducting path of the projectile. This current, in turn, creates

its own magneticfield which interacts mutually with the driver coil

fields to produce a force in the desired direction. The projectile is

guided along the launcher by a nonconducting mandrel extending the length

of the launcher. If, as has been suggested (Burgess, 1980), the radial

forces would tend to Swage the projectile, onto the mandrel, a slightly

different configuration from that. shown in Figure 8-11 could be designed.

A second set of driver coils could be set into the mandrel., thus providing

counteracting radial forces as well as adding to the axial forces producing

acceleration along the launcher.

PROJECTILE DRIVER COll.

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FIGURE 8-11 SCHEMATIC OF. ELECTROI_ENETIC THETA GUN

The theta gun has several characteriStics that render it

unsuitable for.the applications considered under this program:

e The cantilevered mandrel would present formidable design,

manufacturing, and construction problems, since much larger

launcher lengths are required than in the original theta-gun

cOncept.

e The thin-wall-cylinder projectile configuration would be

very inefficient for the .types of payloads at which this

study is aimed.

iTIATTIELL t -- {:OI. UMIUli

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8-17

e The necessity for the projectile to be highly conductive

represents a major problem, as well.

Thus, although the theta-gun ConCept may be promising for certain military

applications, it does_ not warrant additional study for the space EML

missions considered in this report.

8.5 Soviet Electrom nettc ACcelerators

8,5.1 Induction Acceleration

A major thrust in electromagnetic acceleration work tn the

USSR is induction acceleration. Bondaletov has been the prime investigator in this area. The Soviet work consists of a one-coil inductor

through which a current is pulsed. A conducting projectile (coil or

ring) is accelerated by the magnetic field pulse from the adjacent

inductor. Beryllium, copper, and aluminum are the preferred metals for

projectile materials. The Soviets have reached velocities of 200 m/s

for large projectile masses (200g). Emphasis on the acceleration of

larger projectiles has been on industrial applications, including magnetic

forming of metals and high-speed current switching devices (Golovln,

19 ).

Work in the acceleration of smaller projectileshas emphasized

performance. In 1976 velocities of 5 km/s were reached for projectile

masses on the order of i gram. The Soviets have since reached 10.5 km/s..

8.5.2 Rall_uns

There is no evidence in the open literature .of Soviet. work

in the railgun acceieratton of solid projectiles, Golovin (1982)

speculated that this is because the Soviets are waiting untii other

countries, the U.S. in particular, have published sufficient information

before performing a comprehensive study of the technology, and setting

their own direction in the field.

iiAI'TiI. L| -- ¢OI. uMmuI

,._'.
bj_

k

A-1

American Metal Markets, March 3, 1982.

Battelle Columbus Laboratories, 1983.

June.

Outside Users Paxload Model,

Benford, J., et al, 1984. "High Energy Ratlgun Designs", IEEE

Transactions one, Volume HAG-20, Number 2, Mach.

Boeing Aerospace Company, 1983. "Space Station Needs, Attributes, and

Architectural Options Study", Volume 6, Final Report, D180-27477-6,

April 5.

Boeing Aerospace Company, 1982. "Analysts of Space Systems for the

Space Disposal of Nuclear Haste Follow-On Study", Volume 2, Technical

Report, NASACR 161992.

Boeing Aerospace Company! 1980. "Orbit Transfer Vehicle Concept

Definition Study", Final Briefing, Executive Summary, July.

Burgess, T. J., et al, 1982. "The Electromagnetic Theta Gun and Tubular Projectiles" IEEE Transactions on Ha netics, Volume HAG 18, Number 1,

January.

Burgess, T. J., et al, 1980. "The E1ectromagnetlc Theta Gun and Tubular

Projectiles", Sandia National Laboratories, SAND_-J_801988, December.

Burke, J. D., 1977.

October..

"Where Do We Locate the Moon Base?", $Spacefliqht,

Chemical Propulsion Information Agency, 1972. Hazards of Chemical Rockets

__ts _ndbook, Volume I, "General

Criterla , C iA 194,

Carlisle, R. F., and J. M. Romero, 1982. "Space Station Technology

Readiness", ASME Winter Annual Meeting, November 17.

Choo, Y, et al, 1983. "Presentation: Impact of Orbiter/ETPropellant

Scavenging on Power Generation for Space Station", NASA/LeRC PIR Number

18, February.

D'Aoust, J. O., et al, 1984. "Development of a Modular,, Electromagnetic

Launcher System for the Air Force Armament Laboratory., IEEE Transactions

_, Volume HAG-20, Number 2, March.

Daniel, D. C., and J. E. Milton, 1980. "A Drag and Stability Analysis

of Hypersonic. Spin Stabilized Projectties", AIAA 80-1587.

iiAlrTll, l.i -- ¢_OLUMIUl

)
A-2

q

Davis, E. E., 1982. "Future Orbital Transfer Vehicle TechnolOgy Study",

Volumes 1 and 2, NASA CR 3535 and NASA CR 3536, Boeing Aerospace Company,

May,

Dossey, J. R., and Trottt, G. L., 1975. "Counterpoint: A Lunar colony",

_, July.

Engineering Alloys Digest, inc., 1961. Data on World Wid_ Metals and

AItoys, Aprll.

Faith, W. L,, D. B. Keyes, and R. L. Clark, 1965. Industrial Chemicals,

Third Edition, John Wiley &Sons, Inc., New York, NY,

Friedlander, 1981. Electrical World, October.

General Oynamlcs Convair Division, 1983. "A Study of Space Station

Needs, Attributes, and Architectural Options", Final Briefing, April

5.

General Dynamics, 1980. "Orbit Transfer Vehicle Concept Definition

Study", Final Study Review, July 8.

Gilmore, W.,. 1984.

Oeveloplng", Paper

Conference, June.

"Scavenging of STS Propellants A Resource Worth

Presented at AIAA/SAE/ASME 20th Joint Propu]sion

Goiovtn, M. N., 1982. "A Survey of Soviet Research on Electric Guns

and High-Power Batteries", TACTEC Report R-6133, .Battelle Columbus

Laboratories, September.

Grumman Aerospace Corporation, 1983. "Space Station Needs, Attributes.,

and Architectural Options", Final Briefing, SA-SSP-RP 009, April S.

Guy, W. W., 1983. "Partially Closed System Will Save Big Money",

Astronautics and Aeronautics, Volume 21, Number 3, March.

Hamilton Standard, 1983..

e its r and Ph sics, i975-1976. 56th Edition, CRC Press,

Hannah, E. C., 1977. "Radiation Protection for Space Colonies", Journal

Of theBritish Inter lanetar_, Volume 30.

Hawke, R. S., et al, 1984. "Rail Accelerator Development for Ultra-High

Pressure Research", IEEE Tra_netics, Volume MAG-20, Number

2, March.

Heald, O. A., et al, 1978. "Orbital Propellant. Handling and Storage

Systems for Large Space ProgramS"j_Volume_.A II, Finai Report, General O_amics/Convair Division, CASD-ASP-IB Ou_, pril 14..

• #

[11

it

IIA'iITiI. Llil -" ¢OI. MMBMIi

........ _ ..... :Li:i: .....
!

L,

I !:

I

Ii

I

A-3

Htll, P. 6., and C. R. Peterson,: 1970, _chan_Cs and Thermodynamics

of Propulsion, Addison-Wesley Publishing Company, Inc.

Hoberecht, M. A., et el, 1984. "Design Considerations for a lO-kW

Zntegrated Hydrogen-Oxygen Regenerative Fuel Cell .System",' NASA TM

83664.

Holland, L. D., 1984. "Distributed-Current-Feed and Distributed-EnergyStore Ratlguns", IEEE Transactions on Maqnetics, Volume NAG-Z0, Number

2, March

Johnson, R. H., 1969.

Number 5.

"The Lunar Colony", Science Journal, Volume 5,.

Kerslake, A. P., 1983. "Aerod_mamtc-Thermodynamtc Considerations of

Electromagnetic Railgun Launched Hyperveloctty Vehicle", Final Honors

Thesis, The Ohio State University, May 25.

Kevla_ 4_ Da_a Manual, 1976,1982. E. I. du Pont de Nemours & Company,

Wi.lmington,. D. E.

Koelle, H. H., 1982. "Manufacturing on the Moon?"_ Spaceflight, Volume

24, Number 4, April.

Lockheed Htssiles and Space Company, 1983. "Space Station Needs,

Attributes, and Archttectura_ OptionS", Volume 1,. Executive Summary,

D889718, April 5.

Marshall, R. A., 1984a. "The Use of Nested Chevron Ratls in a DiStributed

Energy Store Ratlgun", IEEE Transactions on Magnetics, Volume RAG-2_,

Number 2, March.

Marshal1, R. A., 1984b. "A Reusable Inverse Railgun Magnetic Flux

Compression Generator to Suit the Earth-to-Space Rail Launcher", .IEE..._EE

Transactions on Magnetics, Volume RAG-ZO,.Number Z, March.

MarSin Marietta Corporation, 1984, "STS Propellant Scavenging Systems

Study", Performance Review Report, March.

Martin Marietta Corporation, 1983.

and Architectural Options Study",

SOC-SE-02-02, April.

"Space Station Needs, Attributes,

Final Review/Executive Summary,

McDonnell Douglas Astronautics Company, 1980.

Restraints Manu_l", DAC 61687D, July,

"Delta_ Spacecraft Design

McDonnell Douglas Astronautics Company, 1983. "Space Station Needs,

Attributes, and Architectural Options", MissiOn Requirements, HDC H0533,

April.

Mongeau, P. P., 1981. "Coaxial Air CoPe Electromagnetic Accelerators",

PhD Thesis, MassaChusetts Institute of Technology, OCtober.

i

!

i .

mATirmL._.m - =oL.uMmum
A-*4 •

Morrtson, D., and Hinners, N, W., 1983. "A Program for Planetary

Exploration", Sctenc_.._e, Vo;ume 220, Number 4597, May 6.

NASA, 1984. "Space. Station AdvanCed Development Program", July2.

NASA, 1973a. "Environmental Statement for the National Aeronautics

and Space Administrationi Office of Space Science Launch Vehicle and

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NASA, 1973b. "Launch Vehicle Es_Imatlng Factors for Advanced Mission

Planning", NHB 7100.5B.

NASA/JSC, 1984a. Lunar Base. Workshop, Working Papers and Miscellaneous

Notes, Draft, April23-27.

NASA/JSC, 1984b. Report of Lunar Base Workshp, Draft, April 23-27..

NASA/JSC, 1984c. Lunar Surface Return, •V_ewgraphs, March..

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November 29.

NASA/MSFC, 1983. "Teleoperator Maneuvering System Preliminary Definition

Study", June.

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_, Volume 19.

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Report on P_eliminary Feasibility Assessment for Earth-to-space

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NAS CR-167886, June 30.

Rice, E. E., R. S. Denning, and A. L. Friedlander, 1982. "Preliminary

Risk AssesSment of Nuclear Waste Disposal in Space", NASA CR-162028

and NASA CR-16202g, Ba_telle Columbus Laboratories.

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Quarterly Review, March 27-28.

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Attribute_, and Architectural Options Study", Volume II, Program Options,

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January 1.

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Academic Press, New York, N. Y.

BA'i''r EI.I.| -- I_GiLUMImU§

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Shanley, 1957. ___terials, McGrawHi11 Book Company, New

York.

Sheppard, O. J., 197_. "An Alternative Technology for the Lunar Colony",

_, Volume 19, July.

Snow, W. R., 1984. "DNA Solenoid Accelerator Performance Analysts",

Vtewgraphs, U.S. Army Armament Research and Development Center.

_t, 1983, "Homegrown is Best", Volume 25, Number 12, December

TRW, 1983, "Space Station Reeds, Attributes, and Architectural Options

Study't, Executive Summary Briefing, Final Review, April 5.

Teeter, R.,and B. Reynolds, 1983. "Final Report on the NASA Suborbital

Pr'ogram: A Status Review", Battelle Columbus Laboratories, January.

Teeter, R. R., and W. M. Jamieson, 1980. "Preliminary Materi_l Assessment

for the Satellite Power System (SPS)", OOE/ER-O038, Battelle Columbus

Laborato¢ies, January.

U. S. Department of Energy, 1982. "Spent Fuel Storage Requirements,

An Update of DOE/SR-O07", DOE/RL-82-1, June.

U. S. Department of Energy, 19xx.."Technology for Commercial Radioactive

Waste Management", DOE/ET-O028.

U. S. Department of Energy, 197S. U. S. Energy Research and Development

Administration--ERDA Manual, Chapter 0524, "Standards for Radiation

Protection".

i

University of Wisconsin, 1983. "Wisconsin Superconductive Energy Stowage

Project", Volume IV, DOE/ET/26602-35, July.

Usaba, S, K. Kondo, and A. Sawaoka, 1984. "Development of Railgun

Accelerator Combined with Two-Stage Light Gas Gun", IEEE Transactions

_, Volume MAG-2O, Number 2, March.

Weldon, J.M', and D. H. Pryor, 1984. "Producibility improvements and

Manufacture of a Prototype Commercial Compact Pulsed Homopolar Generetor",

IEEE Transactions on Ma_, Volume MAG-20, Number 2, March.

Wilbur, P. J., 1983. "Advanced Ion Thruster and Electrothermal Launcher

Research", Colorado State University, NASA CR-168083, January.

Wilbur, P. J., C. E. Mitchell,. and B. D. Shaw, 1982.. "The Electrothermal

Ramjet", AIAA-82-1216, June.

Winterber9, F., 1983. "The Electromagnetic Rocket Gun--A Means to Reach

,!

Ultrahlgh Velocities , Atomkernenerqie-Kerntechnik, Volume 43, Number 2,

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ACRONYMSANDABBREVIATIONS

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cubiCcentimeter

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Cryogenic Fluid Management Facility

centimeter

Contracting O_ficer's Representative

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distributed energy storage

Department of DefenSe

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environmental control/life support system

electromagnetic launcher

Earth-to-space railgun launcher

External Tank

extra-vehicular activity

gram

Generic Environmental ImpaCt Statement

geosynchronous orbit

gigawatt electric

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high-level waste

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low-Earth orbit

lunar .liquidoxygen

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Lunar Space Transportation SyStem

Lewis Research Center

meter

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metric ton heavy metal

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National Aeronautics and Space Administration

nautical mile

oxidizer-to-fuel ratio

Orbital Maneuvering Vehicle

O_bit Transfer Vehicle

payload Support structure

roentgen equivalent to man

regneratlve fuel cell

Surface Construction and Transport Vehicle.

Surface.Crawler Observer Unpressurized Transportation

supporting research and technology,

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Solar System Exploration Committee

Single.stage-to-orbit vehicle

Space Transportation System

Transatmospheric Vehtcle

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APPENDIX C

RETRiCIENGLISH UNIT CONVEIISIONFACTORS

To convert

atmospheres (arm) ....

atmospheres (arm) ....

calories (cal) ......

Calories per gram

(c.al/g) ..........

centimeters (cm) .....

centimeters (cm) .....

centimeters (cm) ....

cubic centimeters (cm3).

cubic meters (m3). _ . .

degrees Centigrade (%).

degrees Kelvin (°K)...

grams (g) .........

ki 1ograms (kg) .....

ktlometers (kin) ......

ki.lometers (kin) .....

kilometers (km) ......

kilowatts (kW) .....

meters (m) .......

meters (m) ........

meters (m) .......

meters per second .(,n/s).

metric tons (MT) .....

metric tons (_) ......

micrometers (_) .....

Newtons (N) .......

Newtons per cm2 (N/cm2).

into

m

pounds per square inch (psi)...

pOunds per square feet (psf). , .

British themal units (Stu) ....

Brttish themal units per

pound (Btu/lb) ...........

inches (in) .............

feet (ft) ..............

yards (yd)............

cubic inches (in 3) .........

gallonS (gal) ...........

degrees Fahrenheit (°F) .....

degrees Ranktne (°R) .......

pounds (lb) ...........

pounds (lb) ............

statute miles (mi) ........

nautical miles (n.mi.) .......

feet (ft) ..............

Btu per hour (Btu/hr) . ......

inches (tn) ............

feet (ft) ...............

yards (yd) .............

feet per second (ft/s) .......

pounds (lb) ..............

tons (T) ..............

meters (m) ............

pounds force (lbf) .........

pounds per square Inch (psi)...

multiply by

14.70

2116.8

3.9685 X 10-3`

1.80

O. 3937

3.281 X 10''2

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0.0610

264.2

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1.3

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2. 205

0.6214

O. 540

3281

3413

39.37

3. 281

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