SPACE TUG AUTOMATIC DOCKING CONTROL STUDYthe User's Manual are shown in Figs. 1-1 through 1-4. in the brackets of Fig. 1-4, Vehicle Coordinate System, are the APS engine thrust numbers

SPACE TUG AUTOMATIC DOCKING CONTROL STUDY

(NASA-CR-12G578) SPACE ‘IUG AUTOMATIC DCCKING C C N T E C L STUCY F i n a l Report ( L o c k h e e d tlissiles and Space Co.) 125 p HC $5.25 CSCL 228

N75-14820

Uoclas G3/18 08C61

FINAL REPORT

CONTRACT NAS 8-29747 LMSC-D

LOCKHEED MISSILES & SPACE COMPANY- INC. A S U B S I D I A R Y O f t O C K H C E D A I R C R A F T C O R P O R A T I O N

https://ntrs.nasa.gov/search.jsp?R=19750006748 2020-03-11T17:51:39+00:00Z

SPACE TUG AuMMATfC DOCKING CONTROL SrmM

FINAL REPORT

Repared for Nationsl Aeronautics and Spece AdministratSon

Mcrrshall Space Flight Center Huntsville, Alabama

Contract No. VAS8-29747

Author: J. W o h l

l h i s study inmstigbted the docking sensor reqy- s, the influence of the docking mechanism, and the implic8tiom and effects of a docking abort. hrr- Lng the study a digital simulation, which included the primary aspects of the docking manewer, was developed.

LOCKHEED MISSILES & SPACE COMPANY

Foreword

This final report of the Space Tug Automtic Docking Control

Study was prepared f o r the k t i o n a l Aeronautics and Space Administration George C. Marshall Space F l ight Center by Inckhetd Missiles & Space C a n p a n y , Inc. In accordance with Contract WS8-29747

!be study effort herein was conducted under the d i rec t ion of National Aeronautics and Space Administration Study Manager, &. -10 lL-BinfUrth; Mr. Eaner C. Pack, alternate. report ma prep- by the Lockheed Missiles & space campany, Inc . , Sunnyvale, by Mr. Jack Wohl, IXSC Study Manager. study results were developed during the period From August

1973, throue July 1974.

!he -

The

!!here are two parts t o this report:

(1) Final Technical Report (2) LOCDOK User's Manual

Requests f o r addi t ional information should be addressed to:

I&. Mario H. I&einfimth Chief, Dpiamics and Trajectory Analysis Branch, Control Systems Division Systems Dynamics Laboratory . . - Ea 1.5

Marshall Space Fllgbt Center krshall Space plight Center, A l a . 35812 Telephone ( 2 0 5 ) 453-2470

iii


CONTENTS

Section

1

2

3

4

5

6

7

8

9

10

11

FOREWORD

INTRODUCTION

STANDARD CHARACTERISTICS FOR ANALYSIS

LITERATURE SURVEY

DOCKING CONTROL STRATEGIES

DOCKING SENSOR REQUIREMnVTS

DOCKING MECHAhiSM DESIGN

ABORT CONSIDERATIONS

LOCDOK SIMULATION AND USER'S MANUAL

CONCLUSIONS

SUGGESTED FURTHER STUDIES

REFERENCES

PRIXEDING. PAGE B M NOT

Page

ii

1-1

2-1

3-1

4-1

5 - 1

6-1

7-1

8-1

9-1.

10-1

11-1

V

LOCKHEED MISSILES 81 SPACE COMPANY

L I S T OF ILLUSTRATIONS

Figure

1-1

1-2

1- 3 1-4 1-5

4-1 4- 2

5-1

5-2

6-1 6-2 6-3 6-4 6-5 6-6

7-1 7-2

7-3 7-4

7-5 7 -6 7-7 7-8

7-9 7-10

INERTIAL COORDINATE SYSTEM

EARTH-ORIENTED COORDINATE SYSTEM

ORBITAL COORDINATE SYSTEM

VEHICLE COORDINATE SYSTEM

BASELINE TUG CONFIGURATION

AVIONICS CONFIGURATION FOR RENDEZVOUS AND DOCKING

AUTONOMOUS DOCKING PHASES

SENSOR F'IELD O F VIEW REQUIREMENTS

ENCOUNTER " J I G T R Y RELATIONSHIPS

GEMINI DOCKING SYSTEM

APOLLO DOCKING SYSTEM MENASCO DOCKING SYSTEM

MTE3NATIONAL DOCKING SYSTEM

SQUARE FRAME DOCKING SYSTEM

ABORT BURN IMPULSE V S DOCKING MECHANISM ALLOWANCE

FORWARD THRUSTER IMPINGEDENT

FORWARD THRUSTER IMPINGEMENT o CM SEPARATION

FORWARD THRUSTEZ3 IMPINGEMENT 254 CM SEPARATION

FORWARD THRUSTER IMPINGEMENT 508 CM SEPARATION

FORWARD THRUSTER IMPINGEMENT 762 CM SEPARATION

FORWARD THRUSTER IMPINGEMENT 1 0 1 6 ~ ~ SEPARATION

FORWARD THRUSTER IMPINGEMENT 1270CM SEPARATION

FORWARD THRUSTER IMPINGEMENT 1 5 2 4 C M SEPARATION

NORMAL THRUSTER IMPINGEMENT

NORMAL THRUSTER IMPINGEMENT 0 CM SEPARATION

Page

1-2

1-2

1-3 1-4 1-6

4-16 4-17

5-6

5-7

6-2 6-2 6-3

6-3 6-4

6-11

7-9 7-10 7-11 7 -12

7-13 7-14

7-15 7-16

7-17 7-18

PRECFDING PAGE BLANK NOT F'ILMHY vii


L I S T OF ILLUSTRATIONS ( C o n t ' d )

Figure Page

7-11 7-12

7-13 7-14 7-15 7-16 7-17 7-18 7-19 7-20

7-21

7-22

7-23 7-24

8-1 8-2 8- 3 8-4 8- 5

8-6 8-10

NORMAL THRUSTER IMPINGEMENT 254CM SEPARATION



NORMAL THRUSTER IMPINGEMENT 1 0 1 6 ~ ~ SEPARATION

NORMAL THRUSTER IMPINGEMENT 1 2 7 0 C M SEPARATION

NORMAL THRUSTER IMPINGEMENT 1 5 2 4 C M SEPARATION

PAYLOAD TORQUE AND FORCE V S X-DISTANCE FORWARD THRUSTER

PAYLOAD TORQUE AND FORCE VS X-DISTANCE NORMAL THRUSTER

PAYLOAD TORQUE AND FORCE V S X-DISTANCE NORMAL THRUSTERS

ABORT CALCULATION COORDINATES

ERROR NORMAL TO DOCKING A X I S V S ABORT RANGE & NORMAL VELOCITY

PAYLOAD CLEARANCE BELOW DOCKING A X I S VS ABORT RANGE DOCKING VEL. AND ABORT BURN TIME

MINIMUM RANGE BS ABORT BURN TIME

X-DISTANCE AT CLOSEST APPROACH VS ABORT RANGE, DOCKING Vn. AND ABORT BURN TIME V S Y-TORQUE

7-19 7-20

7-21 7-22

7-23 7-24

7-25 7-26

7-27 7-28 7-29 7-30

7-31 7-32

COMPUTER PLOT - ATTITUDE CONTROL TOTAL IMPULSE 8-3 COMPUTER PLOT - REORIENTATION OF TUG TO ACQUIRE PAYLOAD 8-3 COMPUTER PLOT - TUG ATTITUDE 8-4 COMPUTER PLOT - A P S TOTAL IMmTLSE AND ENGINE NO. 8-4 COMPUTER OUTPUT - S I U N I T S , PERFECT'ATTITUDE CONTROL, INPUT VARIABLES

8-5

COMPUTER OUTPUT - S I U N I T S , PERFECT ATTITUDE CONTROL - DATA 8-7 COMPUTER OUTPUT - ENGLISH UNITS, DETAILED ATTITUDE 8-20 CONTROL - DATA

viii

& SPACE COMPANY

T a b l e

L I S T OF TABLES

Page

4-1 4-2

4-3 4-4

6-1 6-2

10-1

10-2

AUTONOMOUS DOCKING PHASES 4-15 PARTIALS OF P O S I T I O N AND VELOCITY A T A LATER TIME WITH RESPECT TO P O S I T I O N AND VELOCITY AT AN EARLIER TIME

TRANSITION MATRIX 4-10 MATRIX OF PARTIAL MEASUREMENTS WITH RESPECT TO THE 4-11

4-9

O3BITAL FRAME

EVALUATION OF DOCKING SYSTEMS DOCKING ACCURACY STRUCTURAL S P E C I F I C A T I O N S

NAVIGATION SYSTEM MECHANIZATION

TYPICAL KALMAN FILTER EQUATIONS

6-9

10-4 10-6


Section 1

INTRODUCTION

1.1 BACKGROUND

An important mission of the Space Tug i s the recovery of s a t e l l i t e s a t o r below synchronous o r b i t a l a l t i tudes for re turn t o the Space Shuttle.

The docking operation i s t o be automatic w i t h the poss ib i l i t y of TV

remote control available as a backup. The Tug must be able t o automatically dock with high probabili ty on the f irst attempt.

i s intended t o provide a bas i s f o r designing such a system.

This study

1 .2 STUDY OBJECTIVES

( a ) Develop terminal docking control s t ra teg ies and determine the sensor requirements.

Assess the influence of the docking mechanism design on the type and accuracy of sensor data required and the probabi l i ty of successful docking on the f i r s t attempt.

(b )

( c ) Assess the effects of a missed docking attempt on the Tug propellant consumption and on the payload a t t i t ude control system.

( d ) Provide documentation of the resul tant computer program. This i s

t o include a user ' s manual, decks and/or tapes, and flow charts

suf f ic ien t for: running the program. A l s o included w i l l be test cases with the description of inputs and outputs.

_. 1.3 AXES CONVENTIONS

The axes conventions used i n t h i s report , the LOCDOK Simulation, and t h e User's Manual are shown in Figs. 1-1 through 1-4. i n the brackets of Fig. 1-4, Vehicle Coordinate System, are the APS engine

thrus t numbers. In the LOCDOK printout the axial engines ( fore-af t )

The numbers shown


t GHA DEFINED FOR ZERO SECONDS MJD

yI

'FIRST POINT OF ARIES

FIG. 1 1 INE1sTIAL COORDINATE SYSTEM

zE

FIG. 1.2 EARTH-CENTERED CO0;U)INATE SYSTEM

1- 2

ASCENDING

/ 0' LONG

%ATELUTE

E ECCENTRIC ANOMALY P = TRUE ANOMALY

FIG. 1-3 ORBITAL COORDINATE SYSTEM

1-3

s a t. +

I- S c3 - A U U 0

7 7 + -- $a I

I

L n-

vi 0 * E L -

L I +

1-14

LOCKHEED MISSILES & SPACE COMPANY. INC.

d c 4 m k 0 0 t’

-+ - I

rl e ac ic

have a value of' 1 fo r forward thrust . Or! the p lo ts the ax ia l thrust has the value 9. Which engines are thrust ing may be derived from the following algebraic equations.

LOCDOK Printout

Axial engine No. = 1, + th rus t - 1, - th rus t

r Lateral engines No. = 1 , + z L r u s t ] + [ zll; -y th rus t ]

-1,- z th rus t +y th rus t

On the 4020 p lo ts

3 GUID Eng No. = [+/-, t h m s t ] + 3[-/+, y thrust] + 9 [+/-, x thrus t

1.4 VEHICLE CONFIGURATION

The Baseline Tug Configuration used i n t h i s study is shown i n Fig. 1-5.

1-5


8

"B

p !3

8 ri

1-6

LOCKHEED MlSSlLES & SPACE COMPANY

1.5 conv5?sIom To SI

Inches (in.) Feet (ft)

centinetera (em) meters (8)

Nautical Wler (uml) k l l a m e t e ~ ~ (k.) Pounds, farce (lbf) newtons (n) hss (slugs) kilo€?== (ke) Torque (ft-lbf) meter-ncnrtoas (m-n) Moment of Inertia (slug-rt 2 ) Pounds, k8s (la,) kilo- (ks)

2 kilogram-meters ( kg-m2)

constants

d e g 0.01745329

In.* 2.54 tt+ 0.3048 a* 1.852 lb* 4.44822

14.5939 f t - lbw 1.35582

2 slug-ft * 1.35582 lbm* 0.4532267

* C O ~ ~ I O I I mtants fraar HASA SP-7012, Rtf. 14.

1-7


Cnn+:m.n 3 U F L L # . L U A A C

STANDARD CFaCTERISTICS FOR ANALYSIS

2.1 INTRODUCTION

The Standard Characteristics for Analysis is a compilation and specification of the many vehicle and system parameters necessary to simulate and analyze Automatic Docking of the Space Tug.

The characteristics should be considered a living document that will be updated, modified, and added to as the vehicle and subsystem parameters are more definitized.

These Characteristics could be used as the specifications for Space Tug requirements.

The values shown in Section 2.2 are used for the preset data in the LOCDOK Simulation. If these characteristics are changed the preset data should be changed also.

2-1


2.2.1 Wei@t

kg-m2 ' (slug4t2) se,sar.7 43,448 9 8 5g4* 43# 7,157.4 5,279

541.0 cm 1 23.3 (in.)

Dry Weiet Max Retrieval Wt for SaE Tug Ryloed Isterfsce start DDckilagwt

Burnout Weight

Tots1 Rupellants & m e a

Ignition Weieplt

kga2 (alug-rt2) 24,071.2 1 ~ 7 % 23,751.3 17,518 6,495.7 4,791

612.1 cm 241 (in.)

14-m2 P i t ch ( I n ) 93,742.8 Yaw ( IZZ) 93,437.7 R o l l (I=) 7,81900 Tug C.G. 518.2 ca

2,341.4 1,164.8

14,810.5 669.0

2,815.9 23,385 -2

24,326.5

( rlug-rt2)

69,141 68,914 5,767 204 (in.)

2.2.3 Attitude Control

*Bustline !hag DefiDitioa Docum& Rev. A, 26 June lW2 tu amended by drta for we OLI Spsce Tug Autoratic Docking Control, 12 Decedmr 1973, Bopcr Pack.

2-2


Attitude CaatrOl hIM

IT (Tail off) (nominal)

n4ec

14,234.3

22.2

22.2

n-icc

1,123.43

2.22

2.22

S t a r t Ikcklna

2-3


AZianrth

Elevation Gimbal Rster Q Acceleration

Angular Mte (Acq. Aver.)

Ebx Angular Rete (Tracki-) Angular Acceleration Acquisition Range

Acquisition Scan Pattern (H probbil l ty Of 8 C @ S i t i O l l )

Beaswidth Bsndvidth

Rsnge Rssolutiar

0.003713 &/em; 0.2128 (dedeec)

0.001745 1.0 (deg/sec)

W A 143.72 km; (77.6) m n ~ H

2-4


~~~~

Poritioa Velocity

2.2.8 Docking wecharrism RcQlti ( 3 4 p g )

Docking AXIS Miss Mstance HS6

Loag. Velocity

Lateral Velocity

Aneular v-its

2.2.9 e Harinstfaa update (3 -sigum)

poeitlar, Each Axis 5-89 (3.163 -1

o to 0.3048 m (1.0 fi) - 005235 rad (2 3 deg) 0.03b8 m/sec (.1 ft/sec) to 0.3048 m / s e c (1.0 f t / d e c )

o to 0.09144 m / s e c (0.3 ft/sec) 0 to 0.0017k5 rad/sac (1.0 dcg/sec)

+

*Must be divided by 217,945.9 m (715,045.6 f't) for input to U)CDOK.


A li-ture rurve7 vas mde at the beginning of the contract. -€JW24-- nicakl Inforprrtioa Center intefiagated the Dcc and NASA data barer,- c&Ssified 8s w e l l a4 unclrsrified. In addition, UW's Dialog data bsst vas surveyed.

The following Descriptors BIB- and in cumbination were used:

--- - _ _ - - _ _

3p8ccclaft

3.2 SURVEY .. _ _

3-1


LOCKHEED MISSILES & SPACE C O M P A N Y

mch. ~tport miwo-2, M igo Bodley, C. S., aad A. C. Park Final *port Contract W-23280, -in CW.


AUTOMATIC RQmEzvWS Ill SPACE

Forelm Technology Mv Wright-Pattereon AFB ab10 (14l6W) Author: frgostaev, V. P., Iiarrrhenbakh, B. V.

6105E1 FLD: 22A, 22C (USGRDFi6913) 5 Dec 68 31

Edited trans. of mono. Congress of the International Astronautical Federation (19th) Hew Yor, 1968.

Report l l ~ . FPD-BP-23-13b6-68

Report n.p., 1968 pl-24

D. Ka01-k.

~ I U M af AtrmMATIc CoIcPRoL IR SPACE (2nd) (SELEmED Kn3SIorJS) Forelga Technology Dlv U r i ~ t - p s t t e F s o n AFB Qlo (14l600)

14 J\m 68 65p

Edited machine trans. of Synpo8lum on Autoaatlc Control in Space (2nd) Vienna, 4-8 Sep 67 pl-43

5732B3FLD: 22B us=-

Report M o o PTD-m-24-127-68

3-4


SPACE SHUTTLE dufJMCB, RAVICATIOII ASD CONTROL DESIQf Sau(R0aS VQLUME 3 ORBITAL 0pERATIolls

Naticmal Ibrolrautlcs aad Space Admlnlstration cfsnnad Spacecraft Center, Hourrton, 'kx. A4995K3 FLD: 22A, &A sspAIIlOl6

1-n 5% RcpoFt NO. HASA-SX-68368, m-042l7-RW-B Mlsc 0 rer ia

3-5


A P O I U SPACECRAPT SY- AHALYSIS PROGRAM. PEARL FL3m TESr DATA

AMLYSIS OF RERDEZVUJS RAMR

TRW &vatamfJ Group, Reddo Beach, ea. Author: Dobbby, S. D., M y , M. G. 5855lA PZD: 22B Sl?AR07q 4Rov68 46p

Contrsct: HAsg-8166 Report Ro. W-CR-92493, TW-lU76-W-EEO-00


c ~

TBE PRO- OF D0C-G UIm A PASSIVE OFBITING OBJECT WBICH PaSsESSES ANW- LARI%xmmuM

71x10 921q Issue 4 Page 1'77 Category 30 NAsA-cR-l22853

nGR-394B-s2 71/09/25 25 P ~ @ S ( D e v e l ~ n t snd Analysis of IhChnlque8 for Retrieving Uncooperative s p i a n i ~ Objects in Space Environment)

Pennrrylvunla S t a t e Unlverslty, Uaimrsity Park.

Presented at the 22nd Congress of the Internstional Astronsutlcal Fed,

' A/Kcplsn, M. E. (Dept. of Aerospscc

Engb=-e)

h ~ S t l 6 , 20-25 s p t . igi

3-7


tlMWWED ~ Z v O u S APPLICATIONS FOR SPACE RESCUE 73Al 1156 Issue 1 Page 105 Category 30 71/00/00 5 pages A/Hateley, J. C. (m, Sunnyvale, Ca)

International symposium on space technology and science, 9th

Proceedings (A73-lll01 01-31) Tokyo, AQIE Publishing, Inc. TO-, ~apsn, ~ a y 17-22 imi

1971, P* 557-561

CONT€UBU!PIOIUS To TBE STUDY OF A EXIROPEWl IfipERoRBfiAL TUG 7 1 ~ 3 70309 Issue 19 p a p 3150 category 3 71/00/00 12 -gee

In Italian (European Unaranned In t e ro rb i t a l Tug, Investi-ting Configurations Structure, Hookup System, Docking and Propellant Supply)

A/Pom, M. (AA/Fiat S.P.A., Mvisian Aviazione, Turin, Italy Rome, Rasseepra Internazionale Elet t ronics Uncleare TelersdiocinuPstografica, convention sponsored by the Ministero Deai Affari Ester1 and the! Associazione Induatr ie Aerspaziali

IIOSllTUTE OF IVAVIGATION, HATIOHAL SPACE m G ON SF'ACE SBIRPI-SPACE STATION-IWCIZAR SEUTTLE NAVIGATIOH

7lA3 5051 Issue 17 Page 2772 Catego- 21 71/00/00 540 pages HASA Mu-shll Space Flight Center, Huntsville, Ala., Feb 23-25, 1971 Proceedings (Space &ut tu , Space station and r j u c l e ~ r a u t t l e Mavi&ion-Coaference, Huntsville, Ala. , Feb 1971)

AIRYmATIC COrpm#IL I10 SPACE 3, IlFpERlQATIOHAL FEIlERATIOlV OF A-C CONTROL

71~1 9526 Issue 7 mp ll54 category 21 70/00/00 815 pages International Conference, 3rd %u,lowe, Ranee, &wch 2-6, 1970

UHWIWED €?EN'ISZVOUS, STATION KEEPIJSG, ABD D O C m G FOREXTRAVEEICULAR SPACE ACPIVPPIES

7u47 5159 68/00/00 22 pages A / M , #. 1. B/hmbert, A. I. General Electr ic Co., Philadelphia, PS ( M s s i l e & Space Div)

C/Gido, J. F.

FIFPB AEROSPACE lBCHARIsKs SYMPOSIUM 72141 3391 Issue 4 Page 485 Category15 HASA-SP-282 71/00/00 (Conference of Structurs l Design Principles and Mechanical Engineering Methods for Aerospace Mechanisms Used in Orbital snd Spsce Flights)

3-8


Washington Rmceedings of Conf. Held in Greenbelt, Md. , 15-16 June 1970; Sponsored by rsASA. Goddad Space Flight Center, Santa C l a r a University,

and Wieed Missiies (k Space C a p m y

RERDEzvouS M SPACB. HOW TBB AUTOMA!CIC SATELUTES F'OUHD EACH OTBER I N ORBIT 6gm 5331 Issue 5 page 8gg category 31 AD-678411

(Docking of C o s n m - 1 8 6 and Cosmos=l87) A / b r t i n i n , Y. Air Force Systems Connnand, Wrl@t-Pstterson AFB, Ohio (Foreign "ethnology Mv.)

-ET-23-160547 67/12/06

SW-SATEUITE SZRVICING VEHICLES FIWG mRT, AUG 1968 6gNl 0561 Issue 1 Page 176 Category 31 NASA-CR-97483 NSR-44~5-059 68/09/00 355 pages (Relimlnarg System Design for S a t e l l i t e Servicing Vehicles - Manned and ummnned Models) A/Colwell, R. G.; B/Mckerson, S. L.:

H a s t a n University, Texas C/F'aul, A. N.

APPLICATIONS OF RADAR To S P A C E C W ARD SPACEFLIm

69N3 6488 Issue 23 page 4020 Category 7 NMA-!lM-X-60473

(Spacecraft Radar Present and Future Applications in Rendezvous and

Soft Tanding) A/Broderick, R. F.; B/Chick, R. : C/Fenner, R. G. Hationdl Aeronautics and Space Mminiatration, L. B. Johnson Space Center, Houston, Texas Presented a t the Southwestern I n s t of Elec and Electron Engr, 1 April 1966

~ / ~ / o o 6mms

3-9


Section 4 DOCKING CONTROL STRATEGIES

4.1 INTRODUCTION

Docking Control s t ra teg ies must be formulated f o r the seven autonmous docking phases shown i n Table 4-1. cmpass the events shown i n the table, but a l so select ion of a data f i l t e r .

The s t ra teg ies not only have t o en-

To cmple t e ly define a l l the s t ra tegies would require knowledge of the Tug's

Avionics Configuration, mission definit ion, and operational constraints. One possible Avionics configuration i s shown i n Fig. 4-1. configuration has the equipment needed f o r autonomous navigation.

Note tha t this

For the following discussion, refer t o Table 4-1.

4.2

The f irst phase f o r docking begins a f t e r the rendezvous in jec t ion burn, which should place the Tug a t t h e nominal aim-point. aim point i s discussed i n Section 5. The posi t ion of the aim-point must consider t he aspect of the sun, moon, o r ear th with respect t o the f ie ld-of-

view of the docking sensor. If a docking sensor i s selected t h a t i s i n the

v isua l o r infrared spectrum an additional constraint would be t o have the payload sun-illuminated.

Phase 1 - Rendezvous Injection Burn

Calculation of the nominal

It would be advantageous t o have the Tug perform an autonomous navigation

update a t t h i s time t o reduce the uncertainty i n i t s posi t ion and reduce t h e sensor FOV requirements, see Section 5. If the Tug missions include mult iple

payload servicing o r deployment, then the navigation update would be required

if the Tug's autonamy leve l i s I or 11.


4.3 Phase 2 - Reorientation

The Tug's a t t i tude a f t e r the in jec t ion burn w i l l normally require reorienta- t ion of t h e T u g t o point t h e center of the sensor 's search pat tern a t the

center of the uncertainty volume of t he payload.

The primary decision f o r this phase would be the time a l lo t t ed f o r the maneuver. The longer the reorientat ion time the less A P S propellant would be used.

4.4 Phase 3 - Acquisition of the Payload

The primary strategy t o formulate during t h i s phase would be i f the payload was not acquired. Several a l ternat ives are shown i n Table 4-1.

If the payload i s acquired but there i s a poss ib i l i t y t h a t t he Tug might impact the payload i n a short time, o r the Tug might move out of the acquisi- t i on range of t he sensor, an immediate evaluation must be made. performs a rapid data taking and evaluation a f t e r lock-on t o e i the r stop the motion of t h e Tug o r reverse i t s veloci ty i f it i s moving away from the

payload.

LOCDOK

4.5

Given the docking axis i n the payload o r b i t a l coordinate system and knowing the Tug's s t a t e vector r e l a t ive t o the payload from the sensor, the Tu@; can

now compute a gross t ransfer t o the docking axis. tha t t he Tug's s t a t e vector i s now known very accurately as the payload posi- t i o n i s knm t o f 1.852 (1 lhni) th ree sips.

Phase 4 - Gross Transfer t o the Docking Axis

.

It i s in te res t ing t o note

The control s t ra tegy i n LOCDOK fo r this phase checks t o see that t h e Tug's

t r a j ec to ry t o the docking a x i s does not v io la te the required miss d i s t m c e threshold (see Fig. 4-2). The t r a j ec to ry then i s calculated so as t o termi- nate the gross transfer beyond the minimum gross t ransfer distance specified. The t ransfer distance i s selected so +,hat the Tug can nul l a l l posi t ions and veloci ty errors normal t o the docking axis before it reaches the stand-off range.

Tug will reach t h e a x i s i n a maximum specified t h e . The average veloci ty toward the docking axis i s ccanputed so tha t the

4-2


During the Tug's t ransfer t o the axis the sensor i s always pointed toward the payload (1) t o keep the payload within the FOV of t h e body mounted sensor ac6 (2 ) t~ aut~~atizally a e q i ~ i r c the dockhg a i d on tine payioad a f t e r t h e

f i n a l gross t r ans fe r burn.

Mid-course corrections are periodically made t o correct t ra jec tory errors.

The f i n a l gross t ransfer burn i s computed, allowing f o r the long thrust ing period, so tha t t he docking axis i s not crossed.

docking axis should be t h a t specified by mission requirements.

The veloci ty along t h e

The t ransfer t o the docking axis is considered complete i f the docking aid i s within the M)V of the docking sensor. If it i s not the Tug would m a k e an addi t ional f a s t t ransfer t o the axis maneuver. I n the event t h a t the docking

aid i s not acquired, ( the a t t i t ude of the payload has d r i f t e d the docking aid out of t he FOV of the sensor). A means f o r acquiring t h e docking aid must be

implemented. navigate the payload un t i l t he a i d is acquired. have t h i s capabili ty. However , a circumnavigation simulation has been developed by LMSC and could be integrated in to LOCDOK a t a l a t e r date. Capabili ty

f o r t h i s addition are provided i n LOCDOK.

One method of accomplishing t h i s would be t o have the Tug circum- A t present LOCDOK does not

It should be understood t h a t i f the payload i s ro ta t ing rapidly, during any

phase o f docking, the docking attempt must be cumpletely aborted.

4.6

The bas ic guidance s t ra tegy fo r Phase 5 i s t o n u l l the posit ion and veloci ty

e r rors normal t o the docking a x i s while maintaining the veloci ty along the

ads .

propellant usage. who can se lec t the exponent, G17 i n the input dictionary.

Phase 5 - Transfer down t h e Docking AXib

This portion of the guidance uses an exponential logic t o minimize The rap id i ty of convergence i s controlled by the operator

4-3


The f i n a l b u r n down t h e docking axis normally reduces the Tug's velocity t o t ha t permitted by the docking mechanism and the poten t ia l abort maneuver.

Note tha t the r e t r o burn has t o be made f a r enough from the Tug so tha t th rus te r impingement does not disturb the payload. From the r e t r o burn

point on, a l l forward thrust ing engines must be disabled because of impingement.

4.7

A t t he stand-off point the Tug t o payload posi t ion, velocity, and a t t i t ude

i s evaluated. If t h e tolerances dictated by the docking mechanism are ex-

ceeded the Tug should abort the attempt. The stand-off point select ion i s

detai led i n Section 7.

Phase 6 - Evaluation a t t he Stand-off Point

There should be some provision made t o inspect t h e payload docking mechanism t o see tha t it i s not obstructed o r damaged.

4.8

During t h i s phase a l l th rus te rs m u s t be disabled except for an emergency abort capabili ty. It should be understood tha t an abort a t t h i s t h e w i l l

severely d i s t u r b the payload and may make fu ture docking attempt impossible.

Phase 7 - Coast t o Latch-Up

4-4


~~ ~ ~~ ~

4.9 DATA PILTERIWG

4.9.1 Introduction

As a l l sensor measurements are noisy, some method of data f i l t e r i n g must be

employed to determine t h e best estimate of the payload s t a t e vector v i t h respect to the Tug.

This section summarizes known resu l t s i n l i nea r i zed and l inear estimation theory, The classical least squares m a x i m u m likelihood version of the mn- l i n e a r estimation pmblem is outlined and the sequential version of the op- t i m u m f i l t e r ing solution (as derived by Kalman). The Kalman equations have the aaVantage that dyumdc noise I n the model is eas i ly handled, b u t both

the least squares version and the Kalman equations can be used with any de termini s ti c mode 1 . -

A six-by-six sequential Kalman was selected for the Tug data f i l t e r i n g as being t h e best campmnlse for on-board pmcessing, The Tug has knowledge of I ts accelerations from on-board instrwnentstion and it is assumed tha t

the payload would not maneuver.

This data filter has been incorporated i n the LI>CDOK simulation i n subroutine

HEST. For additional details see Reference 12 and 13.

4.9.2 Notation

I n general, lover case letters in the equation notations (i.e., u, v, x, and

z) denote column vectors while upper case letters (i.e., A, B, F, C, H, and E) denote matrices, The components of a matrix A and vector u are desig-

nated by subscripts as A and u The letter I represents the iden t i ty 13 3'

4-5


matrix. The superscript prime, as i n A' o r u', will denote the transpose of

the matrix o r the transpose of a column vector (which becomes a r o w vector). The product of a column vector and a m w vector, such 86 XI' is a matrix. The symbol E represents the expected value so that E(A) represents the expected value of the quantity A. The subscript -1 as in A'l means the in-

verse of the matrix A.

the matrix exists, although, qui te often, the inverse can be replaced by a pseudo-inverse o r generalized inverse without changing the r e s u l t s .

I n all cases it will be assumed that the inverse of

A set of equations used to model nonlinear estimation for a deterministic system can be w r i t t e n as shown below where A 56 the number of measurement6, z is the 1 X 1 vector representing the actus1 measurements, wk is the 1 X 1

v e c t o r representing the uncorrelated noise on tk measurements, x is the mXl vector representing the state of the system, %(x) is an 1 x 1 vector representing perfect measurements, and is the I X I covariance matrix of the

noise

k

5 = %<XI + k

cw Wk = F$ for k = 1,2,...,IV

It w i l l be assumed the noise has zem mean and it is uncorrelated from one measurement to the next.

E [w~] = 0

cov w w - 0 j f k [ j kl

The best estimate of the s t a t e 5: can be written as shown below, where ?$ is the 1 X m matrix of partial derivatives and xo is the i n i t i a l nominal value of the state

A


~~

The inverse of M is the covarLance of the error i n the estimate. If the model is linear, the mnlinear function \($&) is replaced by its linear equivalent fi, xo and the expressions involving eo cancel out as shown below. A

If there is prior information that the state x has a value z0 with covariance Po, it can be included i n the above analysis by extending the sum so it in- eludes k = 0 and defining zo = Xo, Eo = identity, and Ro = Po.

-

For the discrete version of the linear estimation problem, the system to be estimsted can be described by the following set of matrix difference equa-

tions.

4-7


!he l inea r measurements obtained from the system are given by another set

of matrix equations.

% = % "k 'k for k = 1 , 2 , , . , , N

The matrices @

'hble 4-4 represent known quant i t ies which can change f r o m one measurement

to the next, while the vector z represents kmwn measurements.

not known exactly, but are zero mean independent random variables with known covariance. noise) while the variable w represents random changes i n the measurements (measurement mise) . a t the t i m e of the kth measurement. zero f o r a l l t i m e , the system is said to be "detenninistic." ThC covariances of the zem mean dynamic noise and measurement noise are shown b e l o w :

( t r ans i t i on matrix) Tables 4-2 and 4-3 and H (output matrix)

The vector x represents the estimated state of the system The vectors u and w are

The variable u represents randm changes i n the state (dynamic

The subscr ipt k represents the value of the quant i t ies If the dynamic noise u is iden t i ca l ly

E(ujui) = Qk if j = k and zero otherwise

E(wjwL) F$ if j = k and zero otherwise

Most physical systems w i l l involve nonlinear equations, but it is assumed the

above set of l inear equations can be obtained by l inear iz ing about some nominal values for the State and the measurements. cal system is governed by a set of l i n e a r (or l inear ized) d i f f e r e n t i a l equa- t ions although the measurements w i l l take place at discrete t i m e s . case, the o r ig ina l system d i f f e r e n t i a l equations must be integrated to obtain the required difference equation r e l a t i n g the change i n state from one measurement to the next, Conversely, under ce r t a in conditions, in the l i m i t - ing case as time between wasurements goes to zero, the d i sc re t e system w i l l

approach a continuous system.

It may be that the physi-

I n that

4 4

LOCKHEED MtSSILES & SPACE COMPANY

~

TABLE 4-2 PARTIAIS OF POSITION AND VEIDCITY AT A LATER TIME WITH RESPECT To

POSITION AND VELDCITY AT AN EARLIER TIME. 0 (TRANSITION MATRIX)

4-9


r

1.0 -6 Sin Wt+6 W t

0.0 1.0

0.0 0.0

0.0 -6W Cos Wt+6w

0.0 3 W Sin W t

0.0 0.0 -

TABU 4-3 TRANSITION MATRIX

0.0

0.0

1.0

0.0

0.0

0.0

4 Sin w t - 3 i7

2 cos Wt-2 w’ i j

0.0

1.0

-2 Sin W t

0.0

-2 cos wt+2 - W

1 S i n W t v 0.0

2 Sin W t

1.0

0.0


0.0

0.0

t

0.0

0.0

1

TABLE 4-4 MATRIX OF PARTIAL MEAS- WITH RESPECT TO THE ORBITAL FRAME

IH1 =

a s / a c r 1

[HI = 1 cos El/S

lo

s = r a n g c

Rad = M a l position difference

IT = In-track position difference

El = Elevation angle

Az = Azimuth angle

4-11


The optimum estimate w i l l be the l i nea r estimate which minimizes the mean square e m r . Calculating the estimate requires knowing the mean and covariance of al l the random variables of interest , but no higher moments. If a l l the random variables have a normal probabi l i ty dis t r ibut ion, the

estimate w i l l be t he conditional mean of the state given the measurements. Sometimes the estimate is also cal led the Maximum Likelihood estimate be-

cause it maximizes the conditional probabi l i ty distribution.

A Let x denote the optimum estimate of the state x given a l l the measurements up to If j is greater than or equal to k, it is called f i l t e r i n g and prediction. If j is less than k, it is called smoothing. The error i n the optimum estimate is the difference between the actual value of the state and the estimate.

3 /k 3

The covariance matrix of the error, P k/J, is defined:

The sequential version of Kalman, can be w r i t t e n as filter.

h A

xk/k %/k-1

the optimum f i l t e r i n g solution, as derived by shown below where \ is the gain on the Kalmau

A

%l/k = k+l %/k

The covariance matrix P can also be calculated sequentially:

The init ial

'k/k

'k+l/k = @ k+l 'k/k

(I - Bk %) 'k/k-l

i+l + 'IC

conditions for the f i l t e r i n g m l u t i o n are based on the a priori - information, which I s that the state variable x1 ha8 a kmwn mean, Go, and covariance Po.

0

4-12


- For computational reasons, it is necessary that the matrix Po be non-singulm. I f the actual a p r i o r i inforatation is not su f f i c i en t to m a k e h, non-singular, usually it can be modified empirically, by trial and error, to make it nee-

s ingular vi thout having a substantial effect on later calculations.

An a l t e rna t ive sequential version of the optimum f i l t e r i n g solution sakes

use of two relations:

.-

-1 A -1 *k/k %/k 'k/k-l %/k-l + % 5' 'k

These relations arise natural ly when using the classical maximum likelihood derivation, The first r e l a t ion can be proved by multiplying P$ by Pklk t,?

g e t the ident i ty; the second, by showing that

Bk = 'k/k % 5;'

4-13


(4-13)

4-14


(3 z_ a a 0

f

rl I .*

LOCKHEED MISSILES & SPACE C O M P A N Y

4-16

Section 5 DOCKING SENSOR R E Q W

5.1 INTRODUCTION -.

The Rxking Sensor is the key piere of hardware for Autonomous Docking. The

following requimments can be modified by tmde-offs with other parameters both in t e rna l and external to the sensor. be a judicious mupmmise of a l l t h e requirements in order to optimize the

The final requirements should

total system,

5.2 ACQUISITION FUUVGE

The sensor sha l l have a 0.99 probability of acquiring a passive cooperative payload a t a minimum range of (7'7.6 ma) or 143.72 he

This range is based on the 3-sigma guidance accmcy and paylaad uncertainty as specified i n Section 2. ing within the acquis i t ion range of the sensor and the boresight of the sensor l o pointed a t the center of the search volume.

It assumes t h a t the tug reorients p r ior to enter-

The nominal a i m po in t for t h e rendezvous burn is computed by:

where: PU =

GAPl =

GAVl =

SFr =

m = DA =

3 Payload posi t ion uncertainty, (rrm)km

3 Guidance posi t ion accuracy, (w)km

3 Search Fmme Time, see

Data Taking In t ewa l , see Deacceleration %me, sec

Guidance veloci ty accuracy of CAP, ( f t / sec) ; 6076 km/sec

5-1


This Aim Point w i l l insure that there cannot be an impact w i t h the payload

no matter what the guidance dispersions are perpendicular to GAPl or with

3 - s i p dispersions i n p 8 Y l O a d position, GAP o r GAV.

motion has been neglected as i ts e f f e c t is second order.

The re l a t ive o r b i t a l

The t i m e needed to cancel the guidance veloci ty error to aver t impact is:

Deecceleration T i m e (DA, SEC) = (CAV)~ M n o where :

M = mass of vehicle (slugs), kg

T = Thrus t ( lbf) , n

The acquisition range (ACQ) then is:

Thus for 100 sensor measurements, a search frame t i m e of 1.41 sec., a vehicle mass of 14810.5 kg (1014.84 slugs) and a retro th rus t of 44.8 n (100 1.b) along an axis which has the maximum GAP, the acquisit ion range is 143.72 km

(77.6 m).

5.3 FIELD OF VIEW

Fig. 5-1 is a graph of the sensor's field of v i e w requirement, with and with-

ou t a guidance update after the rendezvous burn, rendezvous is 250.2 m (821 f t , ) or .e Ian (-135 ma). This distance and the

payload uncertainty i n the r e U t i v e posi t ion of the tug are the drivers for FOV requirements. If it i s desired to guarantee that the payload is within the Fotr then the FOV required i s 21-57 x t1.57 rad (two x +,90°). guidance update the FOV requirements decrease slowly as the I n i t i a l rsw increases. rapidly.

+.96 - x t.96 rad (t55' x t 55O) (assuming zero dispersions perpendicular to

The closest range after

.-____ -

With no

The FDV requirements with a guidance update decreases much more A 99$ probabi l i ty that the payload w i l l be within the M>v requires

5-2


the payload - A i m Point axis). -_ -

With horizon sensors or s t a r trerckers to provide the i n i t i a l altitude reference and a requirement of 0.14 torquing accuracy f o r t h e maneuver, a l t i t u d e errors w i l l not appreciably increase the FW requirements.

- _ - -- -_

5.4 SEARCH FRAME TIME (SFr) -

The prinary requirement for SFT is to achieve lock-on before t h e payload can

d r i f t ou t of the field of view. pendicular to the l i n e of sight again neglecting o r b i t a l dynamics:

If w e RSS the guidance velocity error per-

Ll = 6.04 m/sec (19.8 f t / sec)

Requiring the addition to the F'QV a t closest range be no more than lo$ due to SFJ so as to be negligible when RSS With the sensor then:

-- -

I-

SEP = initial range x tan (,I FOV) ..

sfi = (821) x (.1%38) = 8.06 eec 19.8

- -. .. 5.5 RANGE, ELMATION AWD AZIMUTH ACCURACY

-~ - P r e l i m i n a r y simulations show that the followlng 3 - s i p sccuracics would allow successful latch-up. -

- .

Range: .I$ of range

Angle: OsooOs725 rad (.OS0)

5-3


5.6 BIAS AND RESOLUTION

B i a s is the most d i f f i c u l t sensor error to accoaanodate.

can be measured optically and by other methods. should be compensated for. resolution and 3 - s i p biases:

Many sensor biases The known systematic error

Preliminary simulations a l l o w the following

Resolution: b n g e .Og m, (0.3 f t ) Angle ,0436 m md (0.0025 deg. )

Bias : Range ,0046 m (0.015 f t ) Angle ,0034 rad (0.0002 deg. )

5.7 ACQUISITION AND TRACKING RATES

The following minimum rates are suggested:

Acquisition: 0.0279 rad/sec (1. @/see) Tracking: 0.0506 rad/sec (2.9"/sec. )

Fig. 5-2 I s the encounter relatf&Ips. 5.8 W S S OF IDCK-ON

If for any reason t h e sensor loses lock-on for three consecutive measurements, the semor should start an expanding squares or spiral search about the last known position. after the loss of the payload then a loss of payload signal should be provided to abort the docking during a cr i t ical phase and the sensor should then i n i t i a t e the f u l l M)v raster scan.

If the payload is not reacquired within one second

5 . 9 DISSRIMIRATION

The sensor should be able b discriminate against objects other than the payload i n the FOV. Space debris

could be eliminated to some extent by range gating and relative veloci ty

These would be primarily the star background.

5-4


discrimination. There should be an indicator i f the* is nore than one ob-

j e c t i n the N V after discrimination.

S The sensor should be able to operate if the sun, moon, o r earth l tmL,s are more than 0.08725 rad (5 ' ) f r o m the FOV. the sensor i f the sun, moon, o r earth appear i n the Mw and the sersor should recover normal operation v l th in 10 sec.

There cannot be any Wage to

5.10 DATA F R E Q ~ C Y RATE

The data frequency rate is usually driven by other sensor requirements s m h

as the pulse repe t i t ion frequency, data processing method, acquisiticr, sr3

tracking rate requirenents, e tc , Preliminary simulation shows that a m i t ? ? -

mum of 16 range, azimuth, and +levation aeasurements per second is adeuuatc.

5-5


(1(w

1 .n1 (90)

1 ..*A . .--- .

'

t

CLOSEST INITIAL APPROACH .25 km (.135 N-MI) - \ \

n Y d h v n Y

\ \ \ \ \ \

VN0 UPDATE

.047 (60) -

.. WITH GUlD UPDATE 5.858 km (i3.163 N-MI), 3 0

0 ' i I I I I I I

INITIAL RANGE (N-MI) kn


0 Qo . c d s d 0

9 cy 0 d

c

cy

9 * 9

4 4

.-


Section 6 DOCKING MECHANISM DESIGN

"0 assess the influence of the docking mechanism design on the type and

accuracy of the data required, t he description and mode of operation of ex is t ing and projected docking mechanisms were retrieved from the l i t e r a - ture surrey genexmted ear l ie r . i n Ref. 17, only the Gemini (Fig. 6-1), the Apollo (Fig. 6-2) and the

Menasco (Fig. 6-3) systems were retained f o r fu r the r evaluation. vere added the androgynous internat ional docking system (Fig. 6-4) develcped

f o r the Apollo-Soyuz docking experiment and the "Square Frane" csncept (Fig. 6-5) pro3ected for the !?@&e Tug.

conaepts can be found i n Ref. 20.

O f the l ist of docking systems described

To these

Description of these more recent

The requirements for an autoamtic docking system are formulated i r , 3ef. 23 also. They can be expressed as follows:

The docking system shall be:

1. Comprised of iaechanically mated, automatically operated pa r t s which self-al ign and self-actuate on contact to provide a load carrying mechanical connection between chaser and target vehicles.

2. Simple.

3. Reliable.

4. Itow i n weight.

5. Capable of independent release on command using _?over furnished by the chaser vehicle.


Mg. 6-1 Gemlnl Docking System

Fig. 6-2 Apollo Docking System

6-2


.. .

- -- -- -

-SPACE STATION \ /

ACTIVE DOCKIHC SYSTEM

Fig. 6-3 Menasco Dacking System

PASSIVE DOCKING SYSTEH -

.-

ATTENUATORS-

Fig. 6-4 International Docking System

6-3


6-4


6.

7.

8.

9.

10.

11.

Restored to a ready condition, on both the target and the chaser

vehicles, p r io r to undocking o r separation.

F i t ted with components so as to glve a c lear f ield of view to optical , radar, o r laser sensors on the chaser vehicle and re- flectors on the target vehicle during rendezvous and i n i t i a l capture.

Equipped, of possible, with three latching points for the docking system design.

than three is not s t ruc tura l ly stable o r e f f ic ien t . )

Equipped with latch and contac t points located near the vehicle mold l i n e to minimize loads due to bending.

Equipped wi th automatic latching devices designed to carry loads during boost fmm earth to earth orb i t , as w e l l as loads during lnter-orbi t t ransfer operations.

Designed with st ructure to absorb impact loads without shock absorbers o r load attenuators, i f possible.

(More t h a n three is unnecessarily redundant - less

It is believed that the following requirements should be added:

12, Design system to perform r o l l indexing and establish a hard l i n e e l e c t r i c a l connection.

Such a capabi l i ty w i l l enable the chaser vehicle to ac t iva te devices aboard a disabled target vehicle. as antennas and solar panes1 vi11 have to be retracted or jet tfsoned before any retrieval mission can be acccnnplished.

r e q u i r e a precise roll alignment of the spacecraft,

It is envisioned that pmtruding components such

A service mission would a l so

I n order to make a first attempt at the evaluation of the f ive docking con-

cepts mentioned, values ranging from 0 to 3 were at t r ibuted to each of the requirements .

c- 5


LOCKHEED

Not enough information was gathered to evaluate the docking systems on re-

quirements 4 and 11. capture a t first attempt harder to achieve and w i l l bring a weight penalty to any concept designed w i t h o u t impact attenuators.

It is believed t h a t requirement 11 w i l l make posi t ive

Table 6-1 s h o w s the r e su l t s of the preliminary evaluation of t h e f ive dock-

ing systems selected.

Table 6-1

Requirements Gemini Apollo Menasco Internat ional Square Frarne

1 3 3 3 3 3 2 3 3 1 2 3 3 3 2 1 2 1

5 0 3 3 3 3 6 3 3 3 3 3 7 3 1 3 3 3 8 3 3 3 3 2

9 2 1 3 3 3 10 3 3 3 3 2

12 3 0 0 2 2

TOTAL 26 22 23 27 25

Rating based on evaluation of docking 3 = ycs

system capabi l i ty to meet requirements 2 = probably

listed. 1 = doubtful o r n o

6.2 F'IVE DOCKING CONCEPTS S'NDIED

The following camnents on each docking system mryr help i n a fur ther appraisal of the five concepts.

6-6

MISSILES &

~~

SPACE

~

COMPANY

(1) M i n i Conceptf I n r e l a t ion to t h e other systems, t h i s concept is losing p o i n t s on requirements 5 and 9.

The Gemini Agena Target Vehicle was supplylng a l l the power required f o r the actuat ion of the docking mechanism. performed by the chaser vehicle bu t a l l the docking ac t ive la tching sad mooring operations w e r e perforwd by the target vehicle. If this concept is considered f o r automatic docking, the ac t ive mechanism, ( i n t h i s cue, the in t e rna l docking cone) should be in s t a l l ed on the chase vehicle and the

passive external cone be par t of t h e target vehicle. vould have the o ther advantage to provide more space inside the i n t e r n a l docking cone to i n s t a l l the optical , radar and laser sensors required for the automatic rendezvous and docking operations.

The docking approach manewers were

This new configuration

If such a modification of the Gemini docking concept proves feasible, then

the r a t i n g to requirement 5, i n a b l e 6-1 should be changed to 3, and the

btal bectnues 29 instead of 26.

---_I^_ ~- _ _ _ -_

A s far as requirement 9 is concerned, the docking Latch receptacles are in- stalled i n the external docking cone which has a diameter of approximstely 81.28 cm (32 in.). If the t a r g e t vehicles are i n a diemeter range of 1.52 lm 2.13 m ( 5 to 7 f t ) and the center of gravi ty is located within the pennissible space to prevent jackknifing of the spacecraft on impact, the ex is t ing

Gemini hardware could probably be used. and where their center of gravl ty are outside the pennissible l i m i t s , it I s

conceivable that a larger diameter Gemini docking system could be designed to m e e t the conditions of requirement 9.

For vehicles having larger diameters

(2) Apollo Concept: The reason this concept was s l i g h t l y derated on require-

ment 3, was that th rus t has to be applied to the chaser vehicle following impact to achieve a successful capture. (Ref. 19)

The fact t h a t the probe head and drogue capture receptacle are on the center l i n e of the r e s 6 c t i v e spacecraf't, i n s t a l l a t i o n of the rendezvous and docking


sensors is made more d i f f i c u l t and the i r f ie ld of view is restricted by the

extended probe mechanism of this concept. ra t ing on requirement 7.

These are the reasons f o r a l o w

A l o w ra t ing on requirement 9, is due to the fact that the f irst impact load

is reacted by the probe head which is on the center l i n e of the chaser ve-

hicle, and thus has a tendency to cause the vehicles to jackknife. docking mechanism must resist a greater bending moment to a l ign the vehicles after impact and capture.

This

No means to correct an angular misalignment i n the roll axis during the

docking o r mooring operations resul ted in a l o w r a t ing for requirement 12.

(3) Menasco Concept: This concept, with its latch hooks running up and dmm the radial movable anns, is not simple. conducted successfully on a ful l -scale prototype mechanism (Ref . 17), it seems complicated and not as reliable as the other selected systems. It does not show any provision to al ign the two vehicles i n the lpll

axis.

A l t h o u g h tests have been

The above remarks are the reasons for l o w crstings on conditions 2, 3, and 12.

(4) International Concept: It is believed tha t t h i s system has not been

f l i gh t tested yet and this is the reason it was s l i g h t l y derated under the requirerPents 2 and 3. i n the roll axis, it does not seem to be as accurate a8 in the Gemini

Although it provides an angular alignment

SyStem.

(5) Square Frame Concept: the l i t e r a tu re about this concept, an attempt was made a t r a t ing it against the other four better known systems. known models of this concept have been b u i l t or tested, a l o w r a t ing

on reliabil i ty was given.

Although not much information has been found i n

Due to the f a c t that no

LOCKH EE D MISSILES & SPACE COMPANY

b u t it may not accommodate as large an angular misalignment i n the roll axis. design. components was the reason for a lower ra t ing on requirement 10.

It is not i n the scope of this study to invest igate such a change i n Iack of infomation about the s t ruc tu ra l i n t eg r i ty of the latching

-

A lower ra t ing on requirement 12 has been given because it is believed t h a t

t h i s concept does not provide a rol l indexing as accurate as i n the Gemini system.

Our evaluation study would not be complete without a comparison of the capabi l i t ies of each docking system. G e m i n i , Apollo and the projected Space Tug could be found. Frame" concept is beliwed to be designed to the Space Tug Oocking specif i - cations. Table 6-2.

To date only the specifications of the

The "Square

(Ref. 22). A comparison of these specif icat ions is shown i n

Table 6-2

Baseline Gemini A p o l l o Space Tug

Uni ts Ref. 18 Ref. 7 Ref. 22 Centerline Miss ( f t ) (0 to 1.C) ( 0 to 1.0) ( 0 to 1.0) Distance m o t o . 3 0 ~ 0 to .3048 0 tc .3048 Kiss Angle (0 to 10.0) (0 to 10.0) ( 0 tQ 5.0)

rad 0 to .17b5 0 to .1745 0 t c 090725 Iongitudinal ( f t/sec ) (1.5 max) (0.1 to 1.0) (0.1 tc 1.0)

lateral Velocity ( f t/sec ) ( 0 to 0.5) (0 to 0.5) (0 t~ 0.30) Velocity m sec

m/sec Ang?Jlar Velocity i n (deg/sec

Pitch, rad/sec (0 to 0.75) 0 to .0130g

0 to .01309

0 to -1745

Yaw, (0 to 0.75)

Roll, (0 to 10.0)

Combined 0 to -1745 (c to 1.0) (0 to 0.50) 0 t4 00175 0 to .008725

6-9


6.2 I N F L ~ C E OF DOCKING MECHANISM DESIGN

Analysis of the specif icat ions i n Table 6-2 shows t h a t any of t h e concepts would be sat isfactory fo r t he Space Tug, although Table 6-1 r a t e s the In te r - national system the bes t .

The specification tha t influences the Space Tug the grea tes t i s the center

l i n e m i s s distance. Figure 6-6 shows the t o t a l impulse required, fo r an abort , with a 15.24 cm (6 i n ) and 30.48 cm (12 i n ) specif icat ion f o r the center l ine miss distance. (2200 LBf-sec) f o r 15.24 cm t o 4092 N-sec (920 LBf-sec) f o r 30.48 cm allowance.

"he t o t a l impulse required drops from 9786 N-sec

It can be concluded tha t t o minimize the influence on the Tug the center l ine m i s s distance should be made as large as pract icable consistant with the docking mechanism optimization.

6-10

LOCKHEED MISSILES f3t SPACE COMPANY

~ ~~

ABORT BURN IMPULSE VS. DOCK!NrJ MECHANISM ERROR AiiaWKNCE

d

I-

NORMAL ERROR (IN .) CM

6-11


7.1 II!ITRODUCTION

This study has shown tha t a severe constraint on the abort process is imposed by thruster impingement on the payload. control of the posit ion and velocity e r rors normal t o the docking axis are t h e

other major drivers.

Docking mechanism requirements and

7.2 AUXILIARY PFtO~SIoIJ SYEEEM ( A B ) IMPINGEDEN" EFFECTS ON DOCKING

A n impingement study of t h e APS thrusters was made applicable t o any payload shape. of two thrusters firing, which will be the usual case, can be eas i ly derived as the impingement i s symmetrical.

Figs. 7-1 through 7-8 are for one forward f i r i n g thruster. The e f f ec t

Fig. 7-9 through 7-16 are f o r thruster f i r i n g normal t o the Tug X-axis. The impingement forces and torques a re a function of the distance the payload i s

.from Tug s ta t ions 457 and the payload are exposed t o the plumes. - ~ -.

- - - - The forces and torques f o r any payload shape may be estimated by. using-the - -

overlay included with t h i s report over a su i tab le scaled payload - - outboard - - -

pro f i l e . t o f ind the force on the payload.

t he segment yielding the torque. .Note t h a t the pressures are symmetricd -

about the v e r t i c a l axis and t h a t the force and torque from the complementary thruster must be included.

- -- The average segment surface pressure may be graphically integrated

-

The center of pressure i s a t centroid of -

It is obvious from Fig. 7-8 tha t even 1524 cm (600 i n ) from the payload the X-force for two th rus te rs tending t o push a 1270 cm (500 i n ) radius payload

7-1 PREZEDING PAGE BLANK NOT FILMELl

LOCKHEED MlSSlLES & S P A C E COMPANY

away from the Tug is 171.3 newton (38.5 l b f ) which i s highly undesirable. Fig. 7-17 shows the t o t a l force and Torque on the payload as a function of the x-displacement f o r a 1270 cm (500 i n ) radius and a 508 cm (200 i n ) radius

payload. the payload while the torque peaks a t 609.6 cm (240 i n ) . radius payload is subjected t o much l e s s force, the force i s s t i l l 6% of the 1270 cm payload a t zero displacement. load were perfectly symmetrical, t he thrusts equal, and the Tug posit ion exactly on the axis of the payload, an unlikely s i tuat ion.

Note tha t the x-force peaks when the tug is 215.9 cm (85 i n ) from While the 508 cm

The & torque would cancel i f the pay-

For the same conditions as above, but f i r i n g thrusters normal t o the X-axis

the X-force i s only 0.417 netwons (0.094 l b f ) a factor of 410 times l e s s than f o r forward f i r i ng thrusters while the torque i s 203 times less.

Figs. 7-18 and 7-19 show the torques and forces for a 1270 cm and 508 cm radius payloads. ment beyond 762 cm (300 i n ) .

Note from Fig. 7-19 tha t there are no e f fec ts of impinge-

This suggests tha t the abort maneuver should be made by f i r i n g the thrusters normal t o Tug X-axis.

The APS impingement e f fec ts on the Tug i t s e l f were not studies, however it can be concluded that the fore-aft th rus te r should be canted upward t o reduce the impingement torques on the Tug itself and reduce the high stagnation tempera-

tures on the vehicle.

7.2 ABORT

7.2.1 Introduction

Abort s t ra tegy and implementation a re very important. A n abort, which can occur f o r many reasons, has a severe impact on the mission because of time and t o t a l impulse requirements. The fundamental tenets for abort would be:

7-2


(I) Tug and payload safety must not be compromised.

(2) m-e payload att i tude mrt; not; be dest&?il izcd.

(3)

(4)

The additional t o t a l impulse required must be minimized.

The time required t o redock must also 'be minimized.

The following analysis assumes:

-Z axis t h m t e r s f o r the abort burn, tha t is, the tug clears the underside of the payload.

The payload docking mechanism is i n the center of the payload and requires 508 cm clearance below the docking axis.

The payload cannot control more than 13.6 cm-n (1.2 in-lbf) torque.

The Tug's posit ion and velocity errors normal t o the docking axis are independent and normal, assuming 10.8 cm and .004 m/sec

(3-Sigma)

The !hg i s as defined in Section 2.

The coordinates f o r the abort problems are shown i n Fig. 7-20.

The docking axis miss distance .a52 m (.5 f't) and docking veloci ty = .152 m/sec (.5 f t /sec) .

Mininnun allowable range t o payload 1.829 m (6 f t ) .

I n general, the smaller the t i m e a l lo t ted f o r a kneuver o r phase the greater

the t o t a l impulse consumed will be.

7.2.2 Selection of the Stand O f f Point (SOP) and Docking Velocity ( D O C K )

The f irst s tep is t o determine the minimum abort range required. docking mechanism normal ( t o the docking axis) e r ror and the normal veloci ty

Given the

' 7-3


(+sigma) allocated a t the SOP, Fig. 7-21 can be used t o determine the minimum abort range.

the minimum abort range is 4.176 m (13.7 f t ) . As an example, using the parameters i n Section 2 and Para. 7.2.1,

The m i n i m abort distance should be selected t o minimize the docking veloci ty while not violat ing other constraints. o r (.5 f t / s e c ) has been selected.

Here the docking veloci ty .152 m/sec

The next s tep i s t o determine the abort burn time from Fig'. 7-22.

t r a t ion uses 508 cm as the required clearance below the docking mechanism axis. Thus, the abort burn time and t o t a l impulse can be extrapolated as 10.8 sec or 4804 N-sec (1080 lbf-sec) .

The i l l u s -

The minimum range can be checked with Fig. 7-23. (6 f t ) i s safe by .21 m (.7 f t ) .

In t h i s case, the 1.829 m

The l a s t check, payload torque, i s made using Fig. 7-24. using the abort burn time. velocity-minimum abort range curve a l i n e i s drawn horizontally t o in te rsec t the Y torque curve and the magnitude of the Y-torque i s read from the abscissa. For the example it is 11.98 cm-n (1.06 in-Lbf); less than the specif icat ion of

13.56 cm-n.

The graph is entered A t the point t h i s time in te rsec ts the docking

For t h i s case the SOP chosen would be 4.176m and the docking veloci ty .152m/sec.

7.2.3 Abort Impulse Calculation

The t o t a l impulse required f o r abort t h a t is implemented i n LOCDOK is:

8 (Yc + 7.33) a ITA = 50 [ e - \i

7-4


~ ~~

l -

= ('2 min + 23.3 + Pw) 'Retro

safety - Yc + Y V - tT

RY

(7.3)

(7.4)

where :

. = Total impulse required f o r abort and redocking It

Ra = Standoff point distance

YC = Clearance required below docking mechanism axis

= Clearance below Yc required fo r antenna, e tc . 'safety

a = Acceleration o f Tug thrust/vehicle mass

= Docking veloci ty 'dock

= Distance frompayload on the docking axis t o start f i n a l docking approach R2 min

= time allocated t o re turn t o R2 min from below payload after abort

= Payload width

V = Velocity along the Y-axis RY

= Total impulse required t o reach the SOP from R2 m i n ITOD

= I n i t i a l abort burn t o t a l impulse ITA

7-5


= Velocity along the docking axis required t o reach R 2 min 'Retro a f t e r abort

Equation 7.1 accounts fo r a l l the t o t a l impulse needed t o abort , re turn t o a

point on the docking axis f o r another f i n a l approach, and re turn t o t h e SOP.

It does not include the t o t a l impulse needed by the a t t i t ude control system.

To recognize the fac tors t ha t contribute the grea tes t amount t o the abort impulse required, the following typica l values i n addition t o those specified previously a re assumed:

'OD = .152 m/sec (.5 f t / s ec )

= 66723.3 N-sec (15,000 lbf-sec) from LOCDOK simulation ITOD

pw = 10.16 m (33.34 f t )

= 304.8 (1000 ft) R2 min

'safety = 22.63 m (74.25 f t )

= 3600 sec

MASS = 14593.9 kg (1000 slugs)

From Equation 7.1

It = 2160 + 15000 + 1000 (2 (.0253 + .5) + .2935)

= 2160 + 15000 + 1399.5 .It

7-6


~~

It immediateiy becomes obvious that the most impulse i s used by the redocking

run down the docking axis. quired t o regain R2 ~~n the l ea s t . impulse i s inversely proportional to the t i m e a l l o t t ed i n arriving a t R

and poten t ia l ly it could become very large if the time a l lo t t ed i s short .

The abort impulse is next, with the impulse re- Note, however, f r o m Equation 7.3 tha t t h i s

2 min

7.2.4 Abort Time Calculation

The t i m e required tA t o redock the tug is:

= tC + tCB + tI + SOD + tOD tA

tC = Ra/VDOCK

'Retro2 +

'DOCK2 2a Pw + 23.3 + - -

vDOCK

%OD = 'OD/"

- /VOD tOD - R2min

where :

= Velocity down the docking axis &om R2 min t o Ra 'OD

From Equation 7.5:

= 27.4 + 116.6 + 3600 + 5 + 2000 tA

= 5749 sec, 1.6 hrs tA

7-7

(7.7)


More than half the t i m e required for redocking i s used i n re-establishing the

Tug's posi t ion for a run down the docking axis.

While t h i s time can be shortened, it would increase the t o t a l impulse needed

fo r the abort.

The second largest increment i s used for the run down the docking exis. The

veloci ty VOD selected was .152 m/sec (.5 f t / sec) t o save the t o t a l impulse required t o accelerate t o VOD and deaccelerate t o M O C K .

7-8


~~ ~~ ~

(IlaVd 11x3) 3NWd 11x3 3lZZON W O U 33NVlSlO 1VIW

7-Y


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7-10


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7-12


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7-13

LOCKHEED MISSILES & S P A C E C O M P A N Y

7-14

LOCKH E E D MISSILES & SPACE

~~ ~ ~

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7-16


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7-18

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7-20


7-21

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7-22


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7-23


A

z Y

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7-24

(28) 124.55 z

(24) 106.76

(4500) 50843.3

(4400) 49713.4

(4300) 48503.6

z Y

(14) 62.28

(3900) 61064.2

(3800) 42934.3

(10) 44.48

(3700) 41804.5

(8) 118.72

61 26.69

(3600) 40674.6

(4) 17.79

(3500) 39544.8

X -DISPLACEMENT (IN.) CM

FIG. 7-17 P A W TOR- AND mRCE VS X-DISTAEE FORWARD -8ZER

7-25

LOCKHEED MISSILES 8c SPACE COMPANY

(30) 339.0

(20) 226.0

(10) 113.0

0

PAYLOAD TORQUE AND FORCE VS X-DISTANCE

254 5108 762 1016 1270 1524

X-DISPLACEMENT (IN .) CM

7-26


PAYLOAD TORQUE AND FORCE W X-DISTANCE

X-DISPLACEMENT (IN .) CM

7-21


7-sd

SPACE

_____

COMPANY.

~

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ERROR NORMAL TO DOCKING AXIS VS ABORT RANGE AND NORMAL VEL.

Qc

ec Qc w

0

(.013 FTBEC) VDXK .003% W E C

3 VN = (.0065FT/SEC)

4 VN = (.013FT/SEC) -00396 MhEC

VDZK = (.7FT/SEC) (2.5) .00198 M E C .213 W E C .762

= (.5FT/SEC) "OCK .152 MhEC

0 (5) (10) (15) (20) (25) (30) (35) (40) (45) (50) (55) (60) (65) (70) 1.52 3.05 4.57 6.10 7.62 9.14 10.67 12.19 13.72 15.24 16.76 18.29 19.81 21-34

ABORT RANGE (FT.) M

PfG. 7-2l ERROR NO= TO DocIczRlc m S VS ABm RANGE AnD mRMAL YgLOcITY

7-29

LOCKHEED MISSILES 8c SPACE COMPANY

0 2 4 6 8 10 12 14 16 ABORT BURN TIME SEC

I I I I I I I I I 0 200 400 600 800 loo0 1200 1400 1600

TOTAL IMPULSE (LBCSEC) 889.6 1779.32668.9 3558.64448.25337.96227.5 7117.2

N-SEC


MIN RANGE VS ABORT BURN T I M (32)

z t h

Y

Z 3

4.75

(24) 7.32

(4) 1.22

0

(200 IN.) 508 CM CLEARANCI r7 7 -

F V DOCK- = (.7 FT/SEC) .213

0 1 2 4 6 8 10 12 14 16

ABORT BURN TIME SEC

FIG. 7-23 MINIMUM BURN TIME VS ABOKP BURH TIME

7-31


X-DISTANCE AT CLOSEST APPROACH VS. ABORT RANGE DOCKING VELOCITY AND ABORT BURN TIME VS Y-TORQUE

(360) 914.4

(W 836.6

(320) 813.8

(300)

(280)

(260) 660.4

(240) 3 609.6

2 (220) ;3 558.8 6 2 (200) I 508.0

k (180) 9 457.2

762.0

711.2

h

c

a

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2 355.6

f 304.8

254.0

(80) 203.2

(60) 152.4

(40) 101.6

(20) 50.8

8 &?i 5 (140)

3 (120)

5 (loo) X

'0 2 4 6 8 10 12 14 16 ABORTBURNTIMESEC.

0 (.2) (.4) (.6) (.8) (1.0) (1.2) (1.4) (1.6) (1.8) (2.0) (2.2) (2.4) 2.3 4.5 6.8 9.0 11.3 13.6 15.8 18.1 20.3 22.6 24.9 27.1

Y-TORQUE (IN-LBICM

7-32

Section 8 LOCDOK SIMULATIm AND USER'S MANUAL

8.1 ~ O D U C T I O N

LOCWK is a d i g i t a l simulation of a space vehicle docking with a payload. This simulation was abstracted and modified f r o m a much la rger simulation, AVION. I n constructing the program ce r t a in features , not required by the study were not included but the interfaces t o permit incorporation i n the future were retained. of the payload, and maintenance of fixed posit ion.

Some of these features are rendezvous, circumnavigation

8.2 LOCDOK SIMULATION

The LOCDOK simulation i s wri t ten in Fortran IV fo r t h e Univac 1108 computer.

segmented. The program is modularized, containing 81 subroutines. The simulation may be run i n English or i n the Internat ional System of Units.

It requires 62,782 decimal words of core storage and i s not

The program. input data is preset t o tha t only variables t h a t require change

f o r a pa r t i cu la r run need be inputed. multiple cases may be run without resubmission on the job.

The input i s designed so that

The program has a ve r sa t i l e SC4020 plot capabi l i ty which has the following

features: allow rep lo t t ing at a later date without re-running the e n t i r e simulation.

In addition, there is the capabili ty t o perform mathematical manipulation

of the variables.

multiplied; and of course one of the var iables maybe a constant. program automatically scales a l l variables so as t o f i l l the e n t i r e plot . If there a re several dependent variables, the program automatically scales them so t h a t adequate resolution i s obtained, and then annotates each

variable with i t s t i t l e and the scale factor , see Figs. 8-1 t o 8-4.

A l l plo t t ing data can be stored on computer magnetic tape t o

These variables m a y be added, subtracted, divided, or

The


Fig. 8-5 through 8-U are sample pages of the output. labeled i n Eng l i sh together with the important Fortran names. of the output variables are self-evident as can be seen from the output.

Figures 8-5 and 8-6 are sample page outputs i n metric un i t s f o r perfect

a t t i t ude control. Figures 8-7 and 8-6 are similar except i n English uni ts . Figures 8-9 and 8-10 i l l u s t r a t e t h e difference i n output from the runs w i t h

detailed a t t i t ude ccntrol. Additional LOCDOK d e t a i l s may be found i n the User's Manual, LMSC/D424229, which i s submitted as a separate volume with t h i s report .

A l l variables are

Definitions

8.2 USER'S MANUAL

The User's Manual, MSC/D424229, i s published separate from t h i s volume fo r ease of use. information necessary t o run the program can be found i n the User's

Manual, although occasional reference t o the Final Report may be required a t f irst .

Once an operator becomes familiar with LOCDOK, a l l

All the input variables required t o run the program are i n the Input Dictionary Section of the User's Manual. names of the input var iables , def ini t ion of the variables, together w i t h

l imitat ions and notes, are a lso i n the Dictionary.

Preset data values, Fortran

8-2

M lSSl LES

~

S P A C E C O M P A N Y

~~

: 2 5 5

TIME, SK

FIG. 8-2 CCMRJPBR PIDT - REORWITATION OF TUG TO AC- PAYtQAD

8-3


C - %

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Section 9

CONCLUS IONS

9.1 DOCKING CONTROL STRATEGIES

A posi t ion update (autonomous navigation) could be advantageous after the

rendezvous in jec t ion burn if the autonomy leve l selected f o r the Tug i s I or 11.

Strategies are required i n the event the Tug does not acquire the payload during the acquisit ion phase, o r the dockirig a i d f o r the f i n a l docking phase.

Impingement could be a very severe problem during docking o r abort. forward thrus te r should be disabled far frm the Tug. The a t t i t ude control system logic should be mechanized so t ha t no forward th rus te r will be f i r e d i n the payload vicini ty . The affect on a t t i t ude accuracy and r a t e and t o t a l impulse should be studied t o assess the impact of t h i s mechanization includ- ing the cross-products of i n e r t i a and center of gravi ty of fse t s and t rave l . During the coast from the standoff point t o latch-up a l l thrusters should be

disabled except i n the case of an emergency abort.

only thruster normal t o the x-axis should be used. t o inspect t he payload docking mechanism t o see that it i s not obstructed o r damaged.

The APS

If an abort i s required

A provision should be made -

Low-G propellant slosh could be avery serious problem f o r the Tug.

it could have an even grea te r e f fec t on the vehicle then impingement.

of t h i s problem i s strongly recommended.

Potent ia l ly Analysis

A s i x by six K a h n sequential f i l t e r f o r the docking sensor data i s recommended.

9-1


9.2 DOCKING SENSOR REQUIREMENTS

A posit ion update capabi l i ty would reduce the f i e l d of view requirements. For the specified rendezvous inject ion accuracies a minimum 143.72 KM

( 7 7 . 6 ~ ~ ) acquisition range i s recmended.

9.3 DOCKING MECHANISM DESIGN

The androgynous international docking mechanism seem sui table fo r the Space Tug. possible.

The center-line m i s s distance specification should be made as large as

9.4 ABORT

If an abort i s required only thrus te r perpendicular t o the vehicle x-axis should b e f i red . Additional impingement studies are necessary.

For a given payload configuration the driving parameters on the t o t a l impulse used f o r the abort are:

The docking mechanism posit ion e r ror allowance normal t o

t h e docking axis and the maximum latch-up velocity.

The t ranslat ional control capabi l i ty t o reduce t h e velocity normal t o t h e docking axis t o a small value.

The time allowed f o r the redocking attempt.

The time available t o c lear the payload a f te r t h e i n i t i a - t ion o f the abort burn. This time i s simply the distance from the payload a t the abort time divided by the docking velocity (VDOCK) .

The run down the docking axis t o redock consumes the most APS t o t a l impulse

f o r an abort.

9-2


Section 10

SUGGESTED FuRllIER SPVDIES

Ihe suggested studies in this section w a u l d provide needed analysis t o fbr-

ther define the requirements and configuraficm for the Space Tug and to provide support t o the Aero-Astrionlcs Laboratory.

(1)

LOCDOK is a fairly complex simulation. PrevZous experience has shown that a minimum of two week8 of instruction and customer usage, a t MSFt, is required t o proficiently u t i l i ze a cauplex simulation.

Training at MSFC i n the U s e of UXDOK

The training w i l l comprise lectures, nel, and aid i n debueging problapps.

(2) - Modif lcstlon

The documentation provided under the

to permit experienced progmmers to period of use, desirable changes and

supemlsion of runs by customer person-

present contract is not detailed enough modif'y fx)(=DoIc. Invariably, after a additions become a d & . The necessary

programing support can be provided t o the Aero-Astricmics Laboratory.

(3) I n c o r p O ~ t i ~ Of LOU-G Prope l l an t Slosh Into the Space Tug Automatic Dock- Simlat ion

Lm-G propellant slouh could have a very large impact 011 the Space Tug Mis- sion. Propellant slosh could affect the misrion capability spd require redesign of the Tug subsystans. Lockheed Missiles t !~ Space Company has been involved wlth low-G propellant slosh and prapellant

many years. I n addition t o extensive analytical efforts, Lockheed has written a technical brief, Ref. 15, that d e t a i l s the suggested effort .

systems for

10-1


(4) Control System Modification

The re su l t s of the impingement study have sham that it is undesirable t o f ire the forward thrusters i n the v ic in i ty of the payload.

a t t i t ude control simulation should be modified t o simulate and analyze the ef fec t of this mechanization on the Tug's a t t i t ude control t o t a l impulse, pointing accuracy, and limit cycle rate.

LOCDOK's detailed

(5) Autonomous Navigation

For higher autonomy levels, the Space Tug will probably require an autoncrnous n a v i s t i o n capability. requirements for the docking sensor.

A posit ion update capabi l i ty will reduce the RJV

Table 10-1 shows potent ia l concepts f o r autoncmcm navigation. It is assumed that stellar i n e r t i a l reference, computing, and t i m e refvence capabi l i ty are available fo r on-board navi&ion. Specifically, space vehicle navigation i s

perfonned with all posit ions and veloci ty conpatations done on board the vehicle. each of these approaches for purposes of measuring accelerations due t o tank venting and other unscheduled vehicle perturbations.

It is further assumed that a low-g acceleraneter is considered i n

Certain of these sensor types (e.g., horizon sensor) can be considered and f o r backup. A block diagram for a typ ica l shown i n Fig. 10-1. Since three posit ions and must be corrected w i t h perhaps only one o r two is just one messurement), the select ion of how

f o r use during i n i t i a l i z a t i o n orbit navigational system is three veloci ty coordinates measurements (range, f o r example, much correction t o make t o each

state must follow an orderly process i f the solut ion is t o converge. man filter has the algorithms for the orderly "sequential f i l t e r i n g " of the measurements. Table 10-2 contains excerpts of Kalme,n filter, l i n e a r system, and noise equations that would hsve t o be studied and huplemented f o r autono- mour navimtian.

The Kal-

Anather suggested study would determine the optimum Kaban filter f o r Space

h;lg. as a data f i l t e r for the docking sensor.

!be study would include KalraRn filter compatlbillty and poesible use

10-2


I ( 6 ) Imp ingement Studies

A aigW’lcsat study would be the effects of impingement on sane typical payloads. Payloads with solar panels, antema, and other asymmetrical shapes should be examined. which give r i s e t o torques end hi& temperatures on the skin should be analyzed.

In addition, the impingement effects on the Tbg itself

lhis study would optimize, wlthln the constraints specified by the customer, the cant angle of the APS thn;lsters parallel to the vehicle

I x-alds

Section 2 allocates 411.9 Kg (906.8 lbm) of APS propellants a t the start of docking. ‘Ibis allocation is marginal for a n o d docking and would not be sufficient if an abort or non-nominal trajectory occurred. It is suggested that a study to optimize the tats1 APS impulse be made and the allocation of propellant be ma8sessed.

10.3


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10-5


104


Section 11 REFERENCE

1. LMSC, Guidance Equations f o r Non-Agile S a t e l l i t e Interceptors, by W. H. Barling, Jr., and D. I. Gildenberg, LMSC TM 55-31-OM-15, L M S C - D ~ O ~ ~ ~ ~ , Sunnyvale, California, 9 April 1968.

. . .. _. -

-- 2. LMSC, Relative Motion of Close S a t e l l i t e s i n Highly Eccentric Orbits,

by D. I. Gildenberg and W. H. Barling, Jr., LMSC TM 55-31-0M-029, Sunnyvale, California, 14 September 1967.

3. LMSC, Camments on Explicit Guidance Problems Based on t h e Gauss T i m e - _ - -

Equation, by R. F Burt, LMSC IDC 55-33-2146, Sunnyvale, California, 8 March 1968.

4. I l l i n o i s I n s t i t u t e of Technology, Research Ins t i t u t e , Digi ta l Simulation of Low Relative Velocity Intercept Problems, by K. Jakstas e t a l . , Technical Report AFAL-TR-68-177.

5. LMSC, Orbi ta l Intercept/Rendezvous Guidance, Interim Report ( U ) , LMSC-B247907Ay Sunnyvale, California, 4 January 1971 (S) .

6 . LMSC , Orbital Interceptor/Rendezvous Guidance Development ( U ) , LMSC -B244086, Sunnyvale , California, 15 January 1970 ( S ) .

7. LMSC, SKY, A Real i s t ic Radar Data Generator for Sa te l l i t e - to-Sa te l l i t e Trackin&, by W. H. Barling, Jr., D. I. Gildenberg, and W. G. Uplinger, LMSC TM 55-31-0M-27, LMSC-DOO3410, Sunnyvale, California, 6 March 1968.

8. LMSC , Equations f o r Sa te l l i t e - to-Sa te l l i t e Tracking fo r E l l i p t i c a l Orbits of Arbitrary Incl inat ion and Relative Incl inat ion, by D. I. Gildenberg and W. H. Barling, Jr. , LMSC TM 55-31-OM-029, LMSC-Do31992, S m v a l e , Cal i fornia , 29 August 1967.

9. LMSC, Orbit Determination Using (?.well Orbit Integration, by R. J. Yo0 and Y. L. Bradfield, LMSC Tracking Note 55, LMSC-577906, Sunnyvale, California, 3 November 1964.

10. LMSC, SC-4020 Manual, edited by J. N. Dyer, Organization 19-32, LiJISC, Sunnyvale , California, 1 December 1965.

11. National Aeronautics and Space Administration, LOCDOK/Space Tug Docking; Simulation Users Manual, J. Wohl, LMSC-Db24229, 24 July 1974.

12. LMSC, A Plethora o f Paradigims and Algorithms f o r Trajectory Estimation, H. E. Rauch, LMSC-D267398, Palo Alto, California, 1 May 1972.

11-1


13

14.

15

16.

17.

18.

19

20.

21.

22.

23

LMSC, Documentation of Estimation and Smoothing Computer Program, H. E. Rauch, Unpublished, 5 October 1972.

National Aeronautics and Space Administration, The Internat ional Syster?. of Units, E. A. Mechtly, NASA SP-7012, Marshall Space Flight Center, Alabama, October 1964.

LMSC, Incorporation of Low-G Propellant Slosh Into the Space Tug Pluto- - .- matic Docking Simulation, LMSC-D387917, Sunnyvale, California, 29 March 1974.

National Aeronautics and Space Administration, Development of Generation No. 3 Scanning Laser Radar, NAS8-20833, Phase 111, January 1974.

American I n s t i t u t e of Aeronautics and Astronautics, Survey of Docking Mechanisms Applicable t o Logistics Spacecraft Systems, T. J. Nishizaka, AIM Paper No. 67-908, October 1967.

LMSC, Gemini Agena Target Vehicle Familiarization HDBK, LMSC-A602521, April 1964.

Martin Marietta Corp., Response of Flexible Space Vehicles t o Docking Impact, Final Report, Contract NAS8-21280, Martin Corp. Tech. MCR-70-2, March 1970.

North American Rockwell Corp., Guidance, Navigation and Control f o r Auto- matic Rendezvous, Docking, and Separation of s-11 Derivative Vehicles, Contract NAS7-200, North American Rockwell, Tech. Report SD73-SA-0009, February 1973.

North American Rockwell Corp., Space Tug Point Design Study, Final Report. Vol. I11 Tech. Report SD72-SA-0032, February 1972.

National Aeronautics and Astronautics, Baseline Tug Definition Document, Rev. A. , NASA-MSFC, June 1972.

National Aeronautics and Astronautics, Neuter Docking Mechanism Study, James C . Jones, NASA-MSC, 6th Aerospace Mechanisms Symposium, Moffet Field. California, September 1972, NAS TM X-2557.

11-2


Documents

SPACE TUG AUTOMATIC DOCKING CONTROL STUDYthe User's Manual are shown in Figs. 1-1 through 1-4. in the brackets of Fig. 1-4, Vehicle Coordinate System, are the APS engine thrust numbers